Patent Publication Number: US-2005138914-A1

Title: Turbo rocket with real carnot cycle

Description:
REFERENCES TO PRIOR APPLICATIONS  
      This application relies on the priority of the following provisional applications:  
      U.S. Provisional Application Ser. No. 60/466,270, filed Apr. 28, 2003, entitled, “Turbo Rocket with Real Carnot Cycle;” 
      U.S. Provisional Application Ser. No. 60/470,706, filed May 15, 2003, entitled, “Turbo Rocket with Real Carnot Cycle and Transpiration Blades;” 
      U.S. Provisional Application Ser. No. 60/486,637, filed Jul. 11, 2003, entitled, “Turbo Rocket with Real Carnot Cycle Continued;” 
      U.S. Provisional Application Ser. No. 60/507,400, filed Sep. 30, 2003, entitled, “Turbo Rocket with Real Carnot Cycle Continued Second.” 
    
    
     BACKGROUND OF THE INVENTION  
      This invention relates primarily to a combined turbine and rocket engine that implements the Carnot cycle. The combined turbine and rocket engine, also termed a turbine-rocket or turbo rocket, is designed to provide a highly fuel-efficient propulsion system for high altitude flight where available oxygen diminishes to levels where rocket propulsion with supplemental oxygen is required. To accommodate different levels of oxygen availability from atmospheric, where oxygen is freely available, to space, where oxygen is absent, new engine designs are required.  
      The Carnot cycle has been considered an ideal thermodynamic cycle maximum theoretical efficiency. However, the real Carnot cycle has not heretofore been implemental in a physical embodiment that effectively follows the four phases of the cycle. As indicated in the cycle diagrams included in this specification, the real thermodynic Carnot cycle includes the following four basic phases in a T-S (temperature-enthalpy) diagram: 
           1 - 2  isothermal compression;      2 - 3  polytropic (adiabatic) compression;      3 - 4  isothermal-stoichiometric combustion-expansion;      4 - 1 . a  polytropic (adiabatic) final expansion.        

      The rocket engines of this invention incorporate the Carnot cycle and are represented in several different embodiments. Many of the component elements of the specific embodiments of the turbo rocket engine are derived from prior turbine engine designs and turbojet engine designs of this inventor.  
     SUMMARY OF THE INVENTION  
      The turbo rocket engine of this invention is designed to incorporate the real Carnot cycle into physical embodiments, primarily for high altitude propulsion. In addition, the embodiments of the turbo rocket engine define propulsion systems for aircraft that are operable in atmospheric and stratospheric conditions at maximum efficiency. Certain embodiments of the engines omit the turbine component and other embodiments are adapted for power generation. The engine cycle is by definition a universal Carnot cycle that is advantageous for air and near space propulsion.  
      High efficiencies are achieved by use of a ram air intake to enhance compression and drive associated air turbines that operate a counter rotating axial compressor for ultra high compression of combustion air. Where altitudes cause a diminishing supply of air, the process is supplemented by liquid oxygen supplied in progressive proportions.  
    
    
     BRIEF DESCRIPTION OF THE DRAWINGS  
       FIG. 1  is a cycle diagram, illustrating a Carnot cycle for certain turbo rocket engines of this invention.  
       FIG. 2  is a schematic illustration of a first embodiment of the turbo rocket engine.  
       FIG. 3  is a schematic illustration of the engine of  FIG. 2 , showing gas flow streams.  
       FIG. 4  is an enlarged schematic illustration of a part of the engine of  FIG. 3 .  
       FIG. 5  is a schematic illustration of a second embodiment of the turbo rocket engine.  
       FIG. 6  is a schematic illustration of a third embodiment of the turbo rocket engine.  
       FIG. 6A  is a cycle diagram, illustrating a Carnot cycle for the turbo rocket engine of  FIG. 6   
       FIG. 7  is a schematic illustration of a fourth embodiment of the turbo rocket engine.  
       FIG. 7A  is a cycle diagram, illustrating a Carnot cycle for the turbo rocket engine of  FIG. 7 .  
       FIG. 8  is a perspective view, partially in cross section, of a turbine blade with internal cooling.  
       FIG. 9  is a perspective view, partially in cross section, of the turbine blade of  FIG. 8  in a turbine rotor.  
       FIG. 10  is a cross sectional view of the turbine blade of  FIG. 8  and an adjacent stator blade.  
       FIG. 11  is a schematic illustration of a fifth embodiment of the turbo rocket engine.  
       FIG. 11A  is a cycle diagram, illustrating the Carnot cycle for the turbo rocket engine of  FIG. 11 .  
       FIG. 12  is a schematic illustration of a sixth embodiment of the turbo rocket engine.  
       FIG. 13  is a cycle diagram, illustrating the Carnot cycle for the turbo rocket engine of  FIG. 12 .  
       FIG. 14A  is a schematic illustration of the side of a ram-air rocket engine.  
       FIG. 14B  is a schematic illustration of the top of the engine of  FIG. 14   
       FIG. 14C  is a cycle diagram, illustrating the Carnot cycle for the engine of  FIGS. 14A and 14B .  
       FIG. 15  is a schematic illustration of a cryogenic rocket engine.  
       FIG. 16  is a schematic illustration of a seventh embodiment of the turbo rocket engine.  
       FIG. 17  is a schematic illustration of an enlarged portion of the engine of  FIG. 16 .  
       FIG. 18  is a cycle diagram, illustrating the Carnot cycle for the engine of  FIG. 16 .  
    
    
     DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS  
      Referring to  FIGS. 1 and 2 , the turbo rocket engine shown schematically in  FIG. 2 , and identified generally by the reference numeral  10 , follows the real Carnot cycle diagrammatically depicted in  FIG. 1 . In the T-S diagram of  FIG. 1 , the first embodiment  20  of the turbo rocket engine  10  undergoes isothermic compression from points  1  to  2 , polytropic (adiabatic) compression from points  2  to  3 , and a staged progressive isothermal stoichiometric combustion and expansion from  3 . 1  to  3 . 2  to  3 . 3  to  3 . 4  to  3 . 5  to  3 . 6  to  4 . With a final polytropic (adiabatic) expansion to  1 . a , the cycle is essentially complete.  
      In the turbo rocket engine  10  of the figures, there is typically a central core  12  with an outer housing  14  with an intake opening  16  and a discharge orifice  18 , with a complex of passages from intake opening  16  to the discharge orifice  18  that generates the unique cycle of operation.  
      In the engine embodiment  20  of  FIG. 2 , there is included a centrifugal, heat exchanging air-turbine rotor  21  that initially compresses air that is then delivered through heat exchanging struts  22  to a staged axial compressor  19 . Outwardly directed, axial compressor blades  21 . a - 21 . f  on an extended air turbine rotor hub  21 . 1  are driven in a counter rotating direction from a complimentary set of inwardly directed compressor blades  24 . a - 24 . f  on the core  24  of an axial air turbine  23  with outer fan-like blades  23 . 1 .  
      Centrifugal compression of air in the hollow core  21 . 2  of the air turbine blades  21 . 3  is cooled by the cold by-pass air around the blades which usually has a 5-20 by-pass ratio over air funneled through the air-turbine rotor  21  and hollow struts  22 , which function as an intercooler between the centrifugal and axial compressors to further cool the centrifugally compressed air. The intensive cooling of the centrifugally compressed air results in an isothermal compression indicated in  FIG. 1  by the cycle phase  1 - 2 .  
      The staged axial compressor  19  is provided with alternating counter rotating, outwardly directed blades  21 . a - 21 . f , and inwardly directed blades  24 . a - 24 . f  to produce a polytropic (adiabatic) compression indicated in  FIG. 1  by phase  2 - 3  of the cycle.  
      Isothermic-stoichiometric combustion and expansion represented by phase  3 - 4  in  FIG. 1  for the embodiment  20  of  FIG. 2  is shown with air and gas flows in the schematic illustration of  FIGS. 3 and 4 .  
      The majority of the intake air flows through the engine  10 , by-passing the compressor apparatus except to drive the hollow blades  21 . 3  of the air turbine rotor in one direction and the fan blades  23 . 1  of the fan turbine  23  in the opposite direction.  
      Air that is highly compressed in the centrifugal and axial compressors from the phase  1 - 2  and  2 - 3  points, is divided into two flows  3 . a  and  3 . b  as shown in the enlarged schematic of  FIG. 4 . An air injection chamber  26  and primary combustion chamber  25  are divided by a conical expansion nozzle  32  provided with a series of peripheral windows  3 . 1 - 3 . 6  and  4 . A complimentary series of injector nozzles  31  injects fuel into the compressed air streaming through the nozzle windows  3 . 1 - 3 . 6  to mix and expand with the axial central compressed air flow  3 . 1  that picks up fuel from central slinger injectors  30  when entering the combustion chamber  25  at phase point  4 . The final polytropic (adiabatic) expansion in the phase from  4 - 1 . a  closes the real Carnot cycle for the internal systems of the turbo rocket engine  10 .  
      The variable volume intake control  27  regulates the high volume of the by-pass air flow  28 - 29  and the compressor air flow, and regulates the angularity of the prewhirl air at the entrance to the engine  10 . The ram compressed, by-pass air mixes with the combustion gases for a final heated expansion through the discharge orifice  18 .  
      The combustion process can be stoichiometric, since the products of combustion do not drive a turbine for compression but expand in a combustion-staged nozzle, producing a maximum power density.  
      For high altitude flight, ram air can be supplemented by liquid oxygen injected by one or more nozzles, similar to the fuel injection nozzles  30  and  31 . Oxygen is added to the air stream in progressive proportions as available high altitude air diminishes.  
      In  FIG. 5 , a second embodiment  50  of the turbo rocket engine is shown. The turbo rocket engine embodiment  50  has a variable multi-stage Carnot cycle with components similar to those of  FIGS. 2-4 . In addition, a modified isothermal centrifugal compressor  51  receives ram air at phase point  1 , and delivers compressed air through the inter-cooling struts  22  at the temperature and pressure of phase point  2 , before compression in the axial compressor  19 . The hollow bladed centrifugal compressor  50  rotates in an opposite direction to the free wheeling air turbine  23 , with its outwardly directed fan-like blades  23 . 1  and its inwardly directed compressor blades  24 . 1  to  24 . 6 .  
      In the engine embodiment  50  of  FIG. 5 , the modified isothermal centrifugal compressor  51  includes a planetary gear assembly  52 , connecting the centrifugal compressor  51  to a turbo-compressor shaft with outwardly directed axial compressor blades  53 . 1 - 53 . 5 , which co-act with counter-rotating, inwardly directed axial compressor blades  24 . 1 - 24 . 6  of the associated air turbine  23 . Additionally, the turbo-compressor shaft has gas turbine blades  54 . 1 - 54 . 3 , which are supplied compressed air at the pressure and temperature at phase point  3 , and fuel from injector  30  for expansion of combusted gases from the combustor  3 . 1  through a gas turbine  54  with staged additions of compressed air from chamber  26  and injectors  56 . 1 - 56 . 5 . The staged air and fuel supply produces an isothermal, stochiometric combustion and expansion typical of the Carnot cycle until cycle phase point  4 .  
      The final adiabatic expansion  4 - 5  closes the total real Carnot cycle. In parallel with the internal Carnot cycle, the ram compressed by-pass air proportionately provides greater thrust and is enhanced at higher elevations by liquid oxygen injectors  58  at the intake  16 , and liquid hydrogen or liquid natural gas injectors  59  at the final mixing nozzle  57  before the discharge orifice  18 , extending a combined cycle phase from  4 - 5 - 6  with a final adiabatic pressure-temperature expansion from phase points  6 - 7 .  
      At very high speed, the energy of the ram air drives the centrifugal compressor  51  and counter-rotating air turbine  23 , and the power required for the gas turbine  54  diminishes. To prevent over rotation, the peripheral combustion chamber  26  has a variable opening valve  55 , which regulates flow through the gas turbine  54  by diverting the gas flow directly into the by-pass flow  29  for mixing and discharge through the discharge orifice  18 .  
      Referring to  FIG. 6 , the turbo rocket engine  10  has an embodiment  60  with a front end and back end that are the same as the embodiment  50  of  FIG. 5 . In a middle section, the free wheeling air turbine  19  is replaced by first centrifugal compressor  61  and a second counter rotating centrifugal compressor  62  radially staged from the first centrifugal compressor  61 . Final heat exchanging struts  63  provide an intercooling before entry of the compressed air into the combustion chambers and gas turbine  54 .  
      Using the multiple compressor stages, including the axial compressor  19  with counter rotating blades driven by centrifugal compressors  51  and  62 , pressure ratios over 100 can be created and the efficiencies of the modified Carnot cycle maximized as shown in  FIG. 6A . The equivalent maximum temperature for the Brayton cycle  1 - 2   a - 3   a - 4   a - 1 - 3   a  at the temperature  3 . 1 , generates insignificant power when compared to the Carnot cycle  1 - 2 - 3 . 1 - 4 . 1 - 1 , with the addition of liquid oxygen is extended to  4 . 1 - 6  and final expansion  6 - 7 - 1 .  
      At high speed, the ram air will raise the pressure ratios for precombustion at  3 . 2 , followed by isothermal combustion  3 . 2 - 4 , constant pressure combustion  4 - 5  and maximum stoichiometric combustion  5 - 6  and final expansion  6 - 7 .  
      Referring to  FIGS. 7 and 7 A, a further embodiment  70  of the turbo rocket engine  10  is shown, with its extended Carnot cycle diagram. In  FIG. 7 , the engine embodiment  70  has a front end and mid section similar to the embodiment  60  of  FIG. 6 . The engine embodiment  70  includes the internal isothermal centrifugal compressor  51  and intercooler  22 , the axial compressor  19  and the radially staged first and second centrifugal compressors  61  and  62  with the final intercooler struts  63 .  
      The high pressure compressed air is supplied to a combustion chamber  71  with central fuel injectors  72  and staged fuel and air injectors  73  along the venturi section  75  with a variable geometry nozzle control  74  to regulate expansion. The liquid oxygen injectors  58  and liquid hydrogen or liquid natural gas injectors  59  provide supplemental oxygen as needed and added thrust for the final expansion in the mixing nozzle  57  before ejection from the discharge orifice  18 .  
      Again, with the additional three stages of combustion, the Carnot cycle is extended from  3 . 2 - 4 - 5 - 6 - 7  for maximized power and efficiency.  
      Although embodiments of the turbo rocket engine that omit a gas turbine can easily operate at the stoichiometric level of fuel combustion, those embodiments including a gas turbine operate most efficiently with cooling of the rotor and stator blades to achieve stoichiometric levels.  
      Referring the FIGS.  8  to  10 , a design for internal and external cooling of the blades of the gas turbine  54  of  FIGS. 5 and 6  are shown. The concepts are similar to those described in my U.S. Pat. No. 5,177,954, but modified by including bleed holes  81  and  82  that communicate with the hollow core  83  of the modified rotor blade  80  of  FIG. 8  and core  84  of the stator blade  85 , shown in  FIG. 10 . As illustrated, fuel and air are forced into the rotor blade  80  and stator blade  85  and out the bleed holes to coat the blades and contribute to the combusted gas flow while cooling the blades.  
      Referring to  FIGS. 11 and 11 A, a multi-stage isothermal stoichiometric gas turbine engine  110  is illustrated, which incorporates three phases of the Carnot cycle. The cycle phases include adiabatic, polytropic compression  1 - 2 ′; isothermal, stoichiometric combustion and expansion  2 ′- 3 , polytropic adiabatic expansion  3 - 4 , and a rejection of residual heat  4 - 1 , which closes the cycle as depicted in  FIG. 11A .  
      In the gas turbine engine  110  of  FIG. 11 , an axial compressor  120  has an isothermal compression stage  120 . 1  using a spray of water through injectors  134  to maintain a constant temperature and an adiabatic compression stage  120 . 2  for final compression to phase point  2 ″. The compressed air is supplied to peripheral chamber  121  surrounding central annular combustion chamber  122 . A staged supply of fuel through injectors  135  and  138  is in part carried through windows  139  into the isothermal gas turbine  123  of the type described with reference to  FIG. 5 . The isothermal gas turbine  123  is provided with variable geometry gas turbine nozzles  124  and  125  to control the volume and maintain the pressure of motive gases delivered to the multi-stage axial power turbine  126  at variable demand levels. The power turbine  126  drives the power generator  127  by shaft  128  with spent motive gases released through pipe  130 . The isothermal gas turbine  123  drives the multi-stage compressor  120  by the shaft  138  and is operated at stoichiometric levels by spraying the conical wall  131  of the combustion chamber  122  with fuel by injectors  135  and by the staged spray of fuel through the windows  139  by injectors  138  in the peripheral chamber  121 .  
      Additional cooling is provided by water or fuel injection into and through the blades of the turbines as described with reference to  FIGS. 8-10 .  
      The double process of evaporative cooling of the combustion chamber walls and the air and super fine spray of vaporizing fuel into the multiple stages of the gas turbine  123  permits a complete process of combustion at stoichiometric levels. The controlled maintenance of the maximum pressure at part loads maintains the cycle efficiency and the lowered pressure and consequently lower efficiency at part loads, typical of the Brayton cycle.  
      Referring to  FIG. 12  and to the T-S diagram of  FIG. 13 , a universal thermodynamic gas turbine  140  is depicted. The universal gas turbine  140  is comprised of three main functional assemblies.  
      The first assembly includes a centrifugal two-stage compressor  236  with a first isothermal stage central rotor  141  surrounded by a second adiabatic stage, counter rotating peripheral rotor  142  driven by the electric motors  143  and  144 , respectively.  
      A series of water injectors  145  spray cooling water into the central compressor rotor  141  proportional to the compression level to generate a polytropic-isothermic effect for cooling the first stage compression. This phase is indicated on the T-S diagram of  FIG. 13  by the evolution from phase point  1  to  2 . The second stage compression by the counter rotating peripheral rotor  142  produces a polytropic adiabatic compression indicated by phase points  2 - 3  in the cycle diagram of  FIG. 13 .  
      A second assembly is formed by a gas turbine  146  with isothermal combustion. The gas turbine  146  drives the electric generator  237  by shaft  238 . The gas turbine  146  is similar in construction to the gas turbines in  FIGS. 5 and 11 , and evolves the cycle from phase points  3 - 4 . 1  and  4 . 1 - 4 . 2 .  
      The third assembly is formed by an axial, polytropic, adiabatic power turbine  147  which drives the electric generator  148  by the shaft  149 . Controlled injection of water into the turbine blades  234  and stators  239  allows the gas turbines  146  and  147  to maintain temperatures that are consistent with the temperature limits of the materials of the turbines.  
      The exhaust of the gases and final expansion completes the cycle from phase points  4 . 2 - 5  and  5 - 1 , closing the cycle.  
      The universal thermodynamic gas turbine cycle, as depicted in  FIG. 13 , includes a Carnot cycle  1 - 2 - 3 - 4 . 1 - 1  for the highest pressure and highest temperature at part loads, thereby maximizing the thermodynamic efficiency.  
      Also, at full load including the isothermal-stoichiometric phases of the cycle  1 - 4 . 1 ,  4 . 2 - 5 . 1 - 1 , a maximum power can be generated. For comparison, the Brayton cycle  1 . a - 2   a - 4 . 2 - 5 . 1 - 1  and diesel cycle  1   a - 2 . a - 4 . a - 5 . a - 1   a  are included in the diagram of  FIG. 13 .  
      Referring to  FIGS. 14A and 14B , a ram-jet rocket engine, designated by the reference numeral  150 , is shown. The engine  150  is designed for high speed, supersonic operation in atmospheric, stratospheric and space conditions and is operable with, or preferably independent of, a turbo component. For example, the engine may be incorporated into an aircraft that is launched from a ground accelerator or from an airborne carrier where sufficient speed is attained to sustain an ignition and independent acceleration on fuel combustion. Alternately, the engine  150  may be included on an aircraft having a conventional engine or an engine  10  as described herein for lower speed atmospheric operation and independent operation at stratospheric and space operation.  
      The ram-jet rocket engine  150  has an outer body  151  having a variable geometry intake port  152  regulated by an intake control valve  153 . A central primary combustion chamber  155  is followed by an expanding multi-stage isothermal combustion chamber  156  of the type described with reference to  FIGS. 5, 11  and  12 . The expanding isothermal combustion chamber  156  is followed by an adiabatic, multi-stage combustion and expansion nozzle  157 . A peripheral air plenum  154  supplies compressed ram air that is supplemented or replaced by oxygen from liquid oxygen nozzles  158  as available air diminishes at high altitudes and space flight. Fuel is injected through fuel injector nozzles  159  in the primary combustion chamber  155  and along the windows  156 . 1  of the multi-stage, isothermal combustion chamber  156  to maintain isothermal conditions. The additional air and/or oxygen entering the windows  157 . 1  of the multi-stage combustion and expansion nozzle  157  cools the surface of the expanding nozzle structure. The injected fuel and injected liquid oxygen has a combined cooling effect over the internal and external surfaces of the multi-stage isothermal combustion chamber  156  and the multi-stage adiabatic combustion and expansion nozzle  157 . The staged fuel vaporization and super mixing provides for perfect stoichiometric combustion and high core temperatures to the ejecting gas stream.  
      The configured engine  150  of  FIGS. 14A and 14B  operates as a ram-jet and scram-jet for super high speed where air is available, and as a hybrid scram-jet rocket as air is supplemented by liquid oxygen. In space, where oxygen is lacking, the variable geometry intake valve  153  is closed and the engine  150  operates as a pure rocket with the full capacity of the oxygen injectors  158 .  
      As illustrated in  FIG. 14C , the cycle proceed from the stage points in  FIG. 14A  from stages  1 - 2 - 3 - 4 .- 1 . At progressively higher speeds where the compression pressure increases by ram air, the cycle improves its efficiency and follows the stage points in  FIG. 14C  from  1 - 2   i - 3   i - 4 - 1 .  
      Referring to  FIG. 15 , a staged rocket engine  190  is shown. The staged rocket engine  190  has a housing  191  with a cryogenic oxygen compartment  192  that forms a plenum  193  around the core nozzle  194 . Liquid oxygen is injected into the plenum  193  through one or more injectors  195  and forms a cryogenic gaseous oxygen. The core nozzle  194  is equipped with a lead venturi nozzle  195  and fuel injector  197 .  
      A series of multiple conical venturi nozzles  198  of increasing size with accompanying staged fuel injectors  199  form an injection cascade of fuel and cryogenic oxygen through the nested windows  200 . The isothermal combustion and expansion continues to the final ejection nozzle  201  where the adiabatic expansion within the cooled walls  202  provides final propulsion in a nozzle structure that is sufficiently cooled to allow for stoichiometric combustion. In this manner, by definition, the motive gas flow has a maximum density providing a super powerful reactive mass flow for propulsion.  
      The successive addition of new heat energy to the central adiabatic flow produces an isothermal Carnot cycle stage to maintain the outer nozzle structure within thermal limits until the final adiabatic expansion in the discharge nozzle  201 .  
      Referring to  FIG. 16 , an embodiment  160  turbo generator engine  10  operating under a real Carnot cycle with a supplemental super pressure cycle is shown. The engine embodiment  160  is suitable for power generation where high efficiencies and low fuel consumption is desired.  
      In the embodiments of  FIGS. 16 and 17 , an axial compressor  161  and a centrifugal compressor  161 . 1  are coupled and driven at least in part by a motor generator  162 , having a common shaft  164 , which also connects to an axial gas turbine  163 . A high pressure chamber  178  contains a group of high pressure compressors and turbines shown in the enlarged view of  FIG. 17 . A high pressure centrifugal compressor  165  is driven by an electric motor  166  by interconnecting shaft  167 . A final ultra high pressure centrifugal compressor  168  rotates counter to the high pressure centrifugal compressor  165  and is driven by a motor generator  169  and/or a gas turbine  170  through the common shaft  177 .  
      The gas turbine  170  is configured with the annular combustion chamber  171 , into which compressed air from the staged high pressure compressors is delivered with a toroidal swirl for complete mixing and combustion. By measured injection of water, the combustion chamber and expansion in the turbine is isothermal with staged entry of the motive gases by the window features as previously described with reference to  FIGS. 11 and 12 .  
      The torroidal rotation of the air and fuel around the gas turbine  170  produces the maximum mixing and complete combustion even for heavy, inferior fuels.  
      From the high pressure chamber  178 , an exhaust pipe  172  conveys the medium pressure motive gases to a medium pressure combustion chamber  173 .  
      In the medium pressure combustion chamber  173 , the motive gases are introduced with a torroidal swirl where fuel may be added through staged injectors  179  for the isothermal gas turbine  163  before final expansion through an adiabatic power turbine  74  and exit through a discharge nozzle or conduit  175 .  
      In a preferred configuration, the power turbine  174  drives an electric generator  176 .  
      As diagramatically illustrated in  FIG. 18 , the super pressure Carnot cycle shows isothermal compression from phase points  1 - 2 , which is produced in the axial and centrifugal compressors  161  and  161 . 1 . High pressure adiabatic compression is indicated by phase points  2 - 3 , produced by counter rotating centrifugal compressors  165  and  168  in the pressurized chamber  178 .  
      Isothermal ultra high combustion and expansion indicated by phase points  4 - 5  is produced in the isothermal combustion chamber  171  and isothermal gas turbine  170 .  
      Isothermal medium combustion and expansion with torroidal swirl is indicated by phase points  5 - 6  and produced in the combustion chamber  173  and gas turbine  163 .  
      Final adiabatic expansion indicated by phase points  6 - 7  is produced in power turbine  174 .  
      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.