Abstract:
A new dual-mode ramjet combustor used for operation over a wide flight Mach number range is described. Subsonic combustion mode is usable to lower flight Mach numbers than current dual-mode scramjets. High speed mode is characterized by supersonic combustion in a free-jet that traverses the subsonic combustion chamber to a variable nozzle throat. Although a variable combustor exit aperture is required, the need for fuel staging to accommodate the combustion process is eliminated. Local heating from shock-boundary-layer interactions on combustor walls is also eliminated.

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
     This patent application claims priority to U.S. Provisional Patent Application Ser. No. 61/262,756 filed Nov. 19, 2009 which is incorporated herein by reference hereto in its entirety. 
    
    
     ORIGIN OF THE INVENTION 
     The invention described herein was made by employees of the United States Government and may be manufactured and used by or for the Government for Government purposes without the payment of any royalties thereon or therefor. 
    
    
     FIELD OF THE INVENTION 
     The invention is in the field of dual-mode combustors for use as a ramjet and a scramjet. 
     BACKGROUND OF THE INVENTION 
     Combined-cycle propulsion is considered when the high efficiency of air-breathing propulsion is desired over a broad Mach number range. Air-breathing access to space is one such application of current interest to NASA. The dual-mode scramjet is central to most combined-cycle schemes. Turbine-based combined-cycle (TBCC) systems use a turbine engine for low speed acceleration, and operate to a maximum flight Mach number in scramjet mode dictated by system considerations. TBCC systems are normally assumed to take-off horizontally, and use a second, rocket-powered stage to achieve orbit. Rocket-based combined-cycle (RBCC) schemes use chemical rocket propulsion for low speed acceleration. The high thrust-to-weight ratio of the rocket component allows for its integration within the air-breathing duct. RBCC systems are normally assumed to be launched vertically, and can operate from lift-off to orbit. Turbine-engines reach temperature and thrust limitations as Mach number increases. Rocket thrusters provide a high ratio of thrust-to-weight at any speed, but are very inefficient from the standpoint of specific impulse. In either case, it is advantageous to extend dual-mode scramjet operation to as low a Mach number as possible. 
     Supersonic combustion has long been recognized as a solution to problems associated with the severe stagnation conditions within a ramjet engine at high flight Mach number. Diffuser momentum loss, dissociation, non-equilibrium expansion losses, and structural loading are all relieved by transition to a supersonic combustion process. In general, the cross-sectional area of the supersonic combustor increases in the downstream direction to avoid thermal choking and excessive pressure gradients. The subsequent nozzle expansion process requires a more dramatic increase in cross-sectional area and is usually integrated with the vehicle aft end to provide the maximum possible area ratio. 
     In order to extend the operable flight Mach number range of the scramjet engine downward, toward the upper limit for turbojets or to limit rocket operation to as low a ΔV (speed range) as possible, “dual-mode” operation was introduced by Curran, et al. in U.S. Pat. No. 3,667,233. U.S. Pat. No. 3,667,233 is incorporated herein by reference hereto. 
       FIG. 1  is a prior art drawing from Curran et al., U.S. Pat. No. 3,667,233, and, in particular, is a schematic diagram partially in block form of a dual mode combustion chamber according to the invention. 
       FIG. 2  is a prior art drawing from Curran et al., U.S. Pat. No. 3,667,233, and, in particular, is a schematic cross section of the device of  FIG. 1  showing one possible arrangement of the fuel injectors. 
       FIG. 3  is a prior art drawing from Curran et al., U.S. Pat. No. 3,667,233, and, in particular, is a schematic diagram partially in block form showing an annular configuration for the combustion chamber of  FIG. 1 . 
       FIG. 4  is a prior art drawing from Curran et al., U.S. Pat. No. 3,667,233, and, in particular, is a schematic end view of the device of  FIG. 3  from the exhaust end. 
       FIG. 5  is a prior art drawing from Curran et al., U.S. Pat. No. 3,667,233, and, in particular, is a schematic diagram partially in block form of a modified fuel supply system for the device of  FIG. 1 . 
     Conceptually, a thermally-choked combustion process is established in the aft regions of the scramjet flowpath where the cross-sectional areas are greatest. As depicted in  FIG. 7 , the diverging scramjet duct acts as a subsonic diffuser, and the thermal throat is positioned so as to produce the required back-pressure. 
       FIGS. 6 and 7  are another illustration of the structure and process of the prior art Curran et al., U.S. Pat. No. 3,667,233. 
     Curran et al., U.S. Pat. No. 3,667,233, states at col. 1, Ins. 29 et seq. that: 
     “A combustor with a fixed geometry has one parallel combustion section with a substantially uniform cross section along its length. Fuel is injected into this section and the flame is stabilized on recessed flameholders. As the fuel burns it causes choked flow in this section which sends a shock wave upstream to convert the normal supersonic flow through the combustor to subsonic flow. For transition from subsonic mode to the supersonic mode, fuel is injected into a diverging section upstream of the parallel section which causes the shock to move downstream until it is ejected from the engine. In the final transition to supersonic mode, fuel is supplied only to the upstream injectors.” 
     Further, Curran et al., U.S. Pat. No. 3,667,233, states at col. 2, Ins. 4 et seq. that: 
     “At these speeds fuel is supplied to nozzles  36 . Burning of the fuel in the uniform cross section combustion chamber  24  causes choked flow which sends a shock wave upstream of the flow to convert the supersonic flow to subsonic flow within the combustion chamber. As the speed of the aircraft increases to a speed between Mach 4 and Mach 5, fuel control  30  starts a flow of fuel to nozzles  32  as the fuel control  34  gradually decreases the fuel flow to nozzles  36 . This causes the shock wave to gradually recede as fuel to nozzles  32  is increased and fuel flow is decreased to nozzles  36 . At a speed of about Mach 8 fuel to nozzles  36  is further reduced and supersonic combustion now occurs throughout the divergent and parallel ducts. The expansion of the heated gases in expansion section  22  permits higher Mach speeds to be attained.” 
     The cross-sectional area of the thermal throat must increase as flight Mach number decreases, unless fuel-to-air ratio is reduced. For a given duct, this effect determines the minimum flight Mach number for dual-mode operation. At Mach 2.5, the required thermal throat area approaches that of the inlet capture area. The primary technical challenges in practical application of the dual-mode scramjet scheme of Curran et al., U.S. Pat. No. 3,667,233, are modulation of the thermal throat location, modulation of fuel distribution, ignition, and flame-holding in the large cross-section. Any in-stream devices must be retractable or expendable so as not to inhibit supersonic combustion operation. 
     Curran et al., U.S. Pat. No. 3,667,233, controls fuel flow to modulate the position of the thermal throat at low flight Mach numbers and then, subsequently, to transition to supersonic ramjet operation. If Curran doesn&#39;t modulate the position of the choked flow correctly, the shock wave moves further upstream into the inlet passage 21 of Curran and un-start of the engine may occur. 
       FIG. 6  shows a cross-sectional view  600  of a prior art (Curran et al.) scramjet engine operating in the scramjet mode. Processes that govern scramjet efficiency are inlet momentum losses, Rayleigh losses due to heat addition, heat loss to the combustor walls, skin friction, and non-equilibrium expansion. Other factors that must be considered include separation of boundary layers due to adverse pressure gradients, intense local heating at re-attachment points and shock impingements, and fuel staging or variable geometry to accommodate the variation of combustion area ratio with free stream stagnation enthalpy. 
     Referring to  FIG. 6 , fuel injection nozzle  601 , inlet contraction section  602 , diverging supersonic combustion section  603 , and exit nozzle  604  are illustrated. As stated above, in the scramjet mode, this engine is fed with fuel injector  601 . Reference numeral  608  illustrates and internal wall of the engine. Reference numeral  606  signifies incoming air being compressed. Reference numeral  605  represents the fuel-air mixture being combusted. And, reference numeral  607  signifies expanded gas/combustion products being expelled from the engine. 
       FIG. 7  is the cross-sectional view  700  of  FIG. 6  (Curran et al. prior art engine) in the ramjet mode illustrating choked flow  702  and a shock waves  701 . Fuel injectors  703 ,  704  are illustrated and are operable in the ramjet mode. Curran et al. must delicately control the insertion of fuel. First, fuel is inserted with injectors  703 ,  704  and then fuel is inserted using injector  601  to prevent the shock wave from being expelled leftwardly into the inlet contraction section  602  which may result in un-start of the engine. Reference numeral  606 A indicates incoming compressed air and reference numeral  607 A represents combustion products expelled from the engine. 
     SUMMARY OF THE INVENTION 
     The supersonic free jet mode of a new combined-cycle combustor is disclosed herein at various scramjet flight Mach numbers including 5, 8, and 12. The dual-mode combustor has the ability to operate in ramjet mode to lower flight Mach numbers than current dual-mode scramjets, thereby bridging the gap between turbine or rocket-based low speed propulsion and scramjets. 
     One important feature of the invention is the use of an unconfined free jet for supersonic combustion operation at high flight mach numbers. The free-jet traverses a larger combustion chamber that is used for subsonic combustion operation at lower flight Mach numbers. The free-jet expands at constant pressure due to combustion and rejoins the nozzle throat contour. Recirculating flow in the combustion chamber equilibrates to a pressure slightly lower than that of the free-jet causing under-expansion features to appear in the free-jet. The free-jet joins the nozzle throat contour with little interaction and expands through the nozzle expansion section. 
     At scramjet flight Mach numbers from 5 to 12, the supersonic free jet traverses the combustion chamber and rejoins the nozzle contour at the combustor exit. Periodic wave structure occurs in the free-jet and is initiated by an entry interaction caused by pressure mismatch and rapid mixing and combustion at the combustion chamber entrance and upstream in the inlet section. The periodic nature of the free-jet also led to an exit interaction determined by the phase of the wave structure with respect to the throat location. The effect of reducing nozzle throat area was to increase the combustion chamber pressure, and reduce the period of the wave structure, but not its amplitude. A viscous loss due to momentum transfer to the recirculation zone is also apparent in each case. 
     Calculated heat loads were commensurate with previous estimates for air breathing systems. Peak heat flux occurred upstream of the throat at an impingement point separating the free-jet from recirculation zone. For a given wall temperature, heat load depends on the recirculation zone temperature and volume, the severity of the exit interaction, and the fuel injection scheme. 
     The new combustor is disclosed for use over a wide range of flight Mach numbers, operating in both subsonic and supersonic combustion modes. It operates as a conventional ramjet at low speed, eliminating the aforementioned issues with dual-mode operation. Transition to supersonic combustion in a free-jet mode occurs at the appropriate flight condition upon the rapid opening of the nozzle throat. 
     A supersonic combustion ramjet engine is disclosed and claimed. The terms supersonic combustion ramjet engine, supersonic combustion ramjet and dual-mode combustor are used interchangeably herein. The supersonic combustion ramjet engine is operable in a ramjet mode and a scramjet mode. The ramjet mode extends from about flight Mach number 2.5 up to about flight Mach number 6. The scramjet mode extends from about flight Mach number 5 up to about flight Mach number 12. An inlet passageway receives compressed combustion air from a supersonic diffuser. The inlet passageway includes a fuel injector. A subsonic diffuser and a combustion chamber follow the inlet passageway. The subsonic diffuser (sometimes referred to herein as the diffusion section) includes an inner periphery. A radial step is interposed between and links the inlet passageway and the diffusion section. 
     The inlet passageway is in communication with the diffusion section and the diffusion section is in communication with the combustion chamber. A ramjet-mode flame holder array is located between the subsonic diffuser and the combustion chamber. The flame holder array includes a central circular aperture therethrough. The flame holders are affixed to the inner periphery of the combustion chamber. 
     The engine also includes a contraction section, a variable nozzle throat and an expansion section. The combustion chamber is in communication with the nozzle contraction section and the nozzle contraction section is in communication with the variable nozzle throat. And, the variable nozzle throat is in communication with the expansion section. 
     A nozzle positioner drives and moves the arc section forming the variable nozzle throat to a desired diametrical opening according to an algorithm which is a function of flight Mach number and combustor mode. The algorithm has a discontinuity at a given flight mach number transitioning from the ramjet mode to the scramjet mode forming a free jet from the inlet section, through the subsonic combustion chamber and reattaching at the variable nozzle throat. The ramjet mode includes subsonic operation from about flight Mach number 2.5 up to about flight Mach number 5.0 to 6.0. The scramjet mode includes supersonic operation from about flight Mach number 5.0 to 6.0 up to about flight Mach number 12.0 and greater. The nozzle positioner divergingly adjusts the nozzle throat to a relatively larger diameter between about flight Mach number 5.0 to 6.0 transitioning from the ramjet mode to the scramjet mode forming a free jet extending from the inlet section at the location of the radial step to the variable nozzle throat. The free-jet does not engage the subsonic diffuser. Nor does the free jet engage the combustion chamber. The free-jet rejoins the variable nozzle throat. 
     A supersonic diffuser is used to compress combustion air into a combustion air passageway. Fuel is injected from the combustion air passageway into the combustion air in the combustion air passageway creating a stoichiometric fuel-air mixture. In the scramjet mode, the stoichiometric fuel-air mixture is fed from the combustion air passageway into a free jet that traverses the subsonic diffuser. Operation in the scramjet mode is premised on previous operation and ignition in the ramjet mode using flame holders in the subsonic diffuser. 
     In ramjet mode, fuel is in injected from the combustion air passageway. In the ramjet mode, the fuel-air mixture is combusted in the combustion chamber. The combusted fuel-air mixture is evacuated from the combustion chamber and into the variable area nozzle throat. The variable nozzle throat is modulated and positioned according to an algorithm creating and controlling the position of a terminal shock in the subsonic diffuser. The algorithm is a function of flight Mach number. 
     The step of discontinuing operation of the igniters, and the step of modulating and positioning a variable nozzle throat according to an algorithm, includes transitioning, using the algorithm, the dual-mode combustor from a ramjet mode to a scramjet mode by rapidly opening the variable nozzle throat at a specified flight Mach number. The algorithm includes a discontinuity where there are two values for a specified flight mach number and it is this discontinuity, and the action based upon it, which shifts the dual-mode combustor from the ramjet mode to the scramjet mode. Shifting from the scramjet mode to the ramjet mode is also possible. 
     The algorithm includes the variable nozzle position as a ratio A/Ac of the actual nozzle throat area, A, to the inlet capture area, Ac, of the supersonic diffuser. The nozzle position varies from a ratio of about 0.8=A/Ac at about flight Mach number 2.5 in the ramjet mode to a ratio of about 0.18=A/Ac at about flight Mach number 5.0 in the ramjet mode. The nozzle position varies rapidly from a ratio of about 0.18=A/Ac at about flight Mach number 5.0 in the ramjet mode to a ratio of about 0.41=A/Ac at about flight Mach number 5.0 transitioning to the scramjet mode. Thereafter, the nozzle position varies from about 0.41=A/Ac at flight Mach number 5.0 in the scramjet mode to a ratio of about 0.15=A/Ac at about flight Mach number 12 in the scramjet mode. 
     In the scramjet mode, the fuel-air mixture and the combustion products are separated into a free-jet beginning at the exit of the combustion air passageway/radial step and extends to the variable nozzle throat. The free jet does not engage the subsonic diffuser, the combustion chamber or the contraction section. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  is a prior art drawing from Curran U.S. Pat. No. 3,667,233 and, in particular, is a schematic diagram partially in block form of a dual mode combustion chamber according to the invention. 
         FIG. 2  is a prior art drawing from Curran U.S. Pat. No. 3,667,233 and, in particular, is a schematic cross section of the device of  FIG. 1  showing one possible arrangement of the fuel inlet jets. 
         FIG. 3  is a prior art drawing from Curran U.S. Pat. No. 3,667,233 and, in particular, is a schematic diagram partially in block form showing an annular configuration for the combustion chamber of  FIG. 1 . 
         FIG. 4  is a prior art drawing from Curran U.S. Pat. No. 3,667,233 and, in particular, is a schematic end view of the device of  FIG. 3  from the exhaust end. 
         FIG. 5  is a prior art drawing from Curran U.S. Pat. No. 3,667,233 and, in particular, is a schematic diagram partially in block form of a modified fuel supply system for the device of  FIG. 1 . 
         FIG. 6  is another cross-sectional view of the prior art Curran device operating in scramjet mode. 
         FIG. 7  is the cross-sectional view of  FIG. 6  in the ramjet mode illustrating choked flow and a shock wave. 
         FIG. 8  is a perspective view of the dual-mode combustor in the ramjet mode. 
         FIG. 8A  is a cross-sectional schematic view of the dual-mode combustor of  FIG. 8  in the ramjet mode. 
         FIG. 8B  is a quarter-sectional schematic view of the dual-mode combustor of  FIG. 8  in the ramjet mode. 
         FIG. 8C  is an enlargement of a portion of  FIG. 8B  illustrating, diagrammatically, the step between the inlet cylinder and the subsonic diffuser. 
         FIG. 9  is a perspective view of the dual-mode combustor in the scramjet mode. 
         FIG. 9A  is a cross-sectional schematic view of the dual-mode combustor of  FIG. 9  in the scramjet mode illustrating the free-jet extending from the inlet cylinder to the variable nozzle throat. 
         FIG. 9B  is a quarter-sectional schematic view of the dual-mode combustor of  FIG. 9  in the scramjet mode. 
         FIG. 9C  is a sectioned view of the dual-mode combustor of  FIG. 9  illustrating the flame-holder having a central aperture therein for the passage of the free-jet. 
         FIG. 10  is another example of the dual-mode combustor employing different geometry. 
         FIG. 11  is a quarter-sectional diagrammatic view of the dual mode combustor in the scramjet mode for flight mach number 8 illustrating a step, a hinged diffuser section, a fixed combustion chamber section, a hinged contraction section, a hinged nozzle throat section (circular arc section) and a hinged expansion section. 
         FIG. 11A  is a table of values relating to  FIG. 11 . 
         FIG. 11B  is a schematic view of an example of a receiving joint forming the nozzle throat 
         FIG. 12  is a plot of the prior art thermal throat of Curran, the geometric/nozzle throat of the dual-mode combustor in ramjet mode and in scramjet mode, and the inlet diameter of the dual-mode combustor as a function of the inlet capture area. 
         FIG. 12A  is a table of inlet contraction ratios as a function of the inlet capture area for a range of flight Mach numbers. 
         FIG. 12  B is a control system for positioning the variable (geometric) nozzle throat. 
         FIG. 13A  is a generalized quarter-sectional diagrammatic view of the flight Mach number 2.5 ramjet. 
         FIG. 13B  is a generalized quarter-sectional diagrammatic view of the flight Mach number 3.0 ramjet. 
         FIG. 13C  is a generalized quarter-sectional diagrammatic view of the flight Mach number 4.0 ramjet. 
         FIG. 14A  is a generalized quarter-sectional diagrammatic view of the flight Mach number 5.0 ramjet. 
         FIG. 14B  is a generalized quarter-sectional diagrammatic view of the flight Mach number 5.0 scramjet. 
         FIG. 15A  is a generalized quarter-sectional diagrammatic view of the flight Mach number 6.0 ramjet. 
         FIG. 15B  is a generalized quarter-sectional diagrammatic view of the flight Mach number 6.0 scramjet. 
         FIG. 16A  is a generalized quarter-sectional diagrammatic view of the flight Mach number 8.0 scramjet. 
         FIG. 16B  is a generalized quarter-sectional diagrammatic view of the flight Mach number 10.0 scramjet. 
         FIG. 16C  is a generalized quarter-sectional diagrammatic view of the flight Mach number 12.0 scramjet. 
         FIG. 17A  is an illustration of the pressure contours within the engine for the flight Mach number 5.0 scramjet. 
         FIG. 17B  is an illustration of the pressure contours within the engine for the flight Mach number 8.0 scramjet. 
         FIG. 17C  is an illustration of the pressure contours within the engine for the flight Mach number 12.0 scramjet. 
         FIG. 18A  is an illustration of the Mach number contours within the engine for the flight Mach number 5.0 scramjet. 
         FIG. 18B  is an illustration of the Mach number contours within the engine for the flight. Mach number 8.0 scramjet. 
         FIG. 18C  is an illustration of the Mach number contours within the engine for the flight Mach number 12.0 scramjet. 
         FIG. 19A  is an illustration of the static temperature contours within the engine for the flight Mach number 5.0 scramjet. 
         FIG. 19B  is an illustration of the static temperature contours within the engine for the flight Mach number 8.0 scramjet. 
         FIG. 19C  is an illustration of the static temperature contours within the engine for the flight Mach number 12.0 scramjet. 
         FIG. 20  illustrates ideal net thrust per unit airflow against flight Mach numbers for a conventional ramjet, thermally-choked ramjet and a scramjet. 
         FIG. 21  illustrates the mass-averaged static pressure distributions with the pressure at the nozzle throat station (supersonic combustor exit) denoted by symbols. 
         FIG. 21A  illustrates the mass-averaged axial velocity distributions for various flight conditions. 
         FIG. 21B  illustrates the mass-averaged temperature distributions for flight Mach numbers 5, 8 and 12. 
         FIG. 22A  illustrates the static pressure ratio for scramjet mode flight Mach number 8 with the variable nozzle throat positioned at 110% of the design operating point. 
         FIG. 22B  illustrates the static pressure ratio for scramjet mode flight Mach number 8 with the variable nozzle throat positioned at 100% of the design operating point. 
         FIG. 22C  illustrates the static pressure ratio for scramjet mode flight Mach number 8 with the variable nozzle throat positioned at 90% of the design operating point. 
         FIG. 22D  illustrates the static pressure ratio for scramjet mode flight Mach number 8 with the variable nozzle throat positioned at 80% of the design operating point. 
         FIG. 23  illustrates the effect of nozzle throat area variation for scramjet mode flight Mach number 8 on the rate of ethylene fuel-depletion. 
         FIG. 23A  illustrates the ethylene mass fraction for flight Mach numbers 5, 8 and 12 versus axial position. 
         FIG. 24  illustrates the effect of nozzle throat area variation on mass-averaged static pressure distribution for scramjet mode flight Mach number 8. 
         FIG. 25  illustrates the ideal net thrust per unit of airflow plotted against combustor exit pressure ratio and nozzle throat area variation for scramjet mode flight Mach number 8. 
     
    
    
     DESCRIPTION OF THE INVENTION 
     In the design of the dual-mode combustor all processes were assumed adiabatic. Air capture, inlet contraction ratio, and total pressure recovery were specified as a function of flight Mach number illustrated in  FIG. 12A . These characteristics are representative of a single-cone, axi-symmetric inlet design with forebody pre-compression. Air was assumed to be a mixture of nitrogen and oxygen at 78.85% and 21.15% by volume, respectively. 
     In the analysis of all ramjet cases, ethylene fuel entered at sonic velocity, normal to the propulsion axis at 5180 R. The energy required to raise the ethylene fuel to this condition was ignored. Constant-area combustion in a cross-sectional area equal to 83.3% of the inlet capture area was assumed. This area was chosen to allow operation at a minimum flight Mach number of 2.5 without thermal choking. For comparison, calculations were also done assuming a thermally-choked combustion process. For these cases, the diffuser exit Mach number was set to result in a combustion area ratio of 1.5. 
     The AIAA (American Institute of Aeronautics and Astronautics), paper entitled Supersonic Free-Jet Combustion in a Ramjet Burner, by Charles J. Trefny and Vance F. Dippold III, NASA Glenn Research Center, Cleveland, Ohio, 44135 was published and presented on Jul. 26, 2010 is incorporated herein by reference hereto. 
     The dual-mode combustor is illustrated in  FIGS. 9 ,  9 A and  9 B in scramjet mode wherein supersonic combustion in an unconfined free-jet  943  traverses a larger subsonic combustion chamber  805 , a contraction section  806 , and a variable nozzle throat  807 .  FIG. 9  is a perspective view  900  of the dual-mode combustor in the scramjet mode.  FIG. 9A  is a cross-sectional schematic view  900 A of the dual-mode combustor  899  of  FIG. 9  in the scramjet mode illustrating the free-jet  943  extending from the inlet cylinder  802  to the variable nozzle throat  807  which yields a nozzle throat diameter D 1 . Nozzle throat diameter D 1  as illustrated in  FIG. 9A  is larger than nozzle throat diameter D illustrated in  FIG. 8A . The nozzle throat area is dictated by the curves illustrated in  FIG. 12  for both the ramjet and the scramjet. The examples of nozzle position  807  given here for the ramjet and the scramjet are the portions of the curves  1202 ,  1205  where  FIG. 8A  ramjet mode uses a smaller nozzle throat area than  FIG. 9A  (scramjet mode). 
       FIG. 9B  is a quarter-sectional schematic view  900 B of the dual-mode combustor of  FIG. 9  in the scramjet mode. In the scramjet mode, reference numeral  845 A signifies supersonic combustion and reference numeral  847 A represents expansion.  FIG. 9C  is a sectioned view  900 C of the dual-mode combustor of  FIG. 9C  illustrating the flame-holder  810  having a central aperture  850  therein for the passage of the free jet  943  there through. 
     During scramjet mode, of operation, the propulsive stream  943  is not in contact with the combustor walls  805 , and equilibrates  943 A to the combustion chamber pressure  944 . Boundary  943 A represents the interface of the free-jet/propulsive stream  943  with the recirculation zone/combustion chamber pressure  944 . Thermodynamic efficiency is similar to that of a traditional scramjet, under the assumption of constant-pressure combustion. Qualitatively, a number of possible benefits exist. Fuel staging is eliminated since the cross-sectional area distribution required for supersonic combustion is accommodated aerodynamically without regard for wall pressure gradients and boundary-layer separation because the free-jet does not touch the walls of the diffuser and the combustion chamber. Variable exit diameter D 1  must be set to the proper size for a given flight Mach number. The axial distance available for supersonic mixing and combustion includes the subsonic diffuser  804 , combustion chamber  805  and nozzle contraction sections  806  required for ramjet operation. Heat loads, especially localized effects of shock-boundary-layer interactions, are reduced. Reference numeral  880  signifies incoming air being compressed and reference numeral  881  signifies exiting combustion gases. 
       FIG. 8  is a perspective view  800  of the dual-mode combustor  899  in the ramjet mode.  FIG. 8  illustrates the frusto-conical inlet contraction section  801 , the cylindrical inlet passageway  802 , the diffuser section  804 , the combustion chamber  805 , the contraction section  806  and the variable diameter nozzle throat  807 . Reference numeral  807  signifies the variable nozzle throat at the joining point of the contraction section  806  and the expansion section  808  in the ramjet mode. In the scramjet mode, reference numeral  807  also signifies the variable nozzle throat at the joining point of the contraction section  806  and the expansion section  808 . 
       FIG. 8A  is a cross-sectional schematic view  800 A of the dual-mode combustor  899  of  FIG. 8  in the ramjet mode.  FIG. 8A  illustrates substantial differences in construction when compared to Curran U.S. Pat. No. 3,667,233. First, the flame holders  810  are arranged so as to not obstruct the free-jet as illustrated in  FIG. 9A . The flame holders  810  have a central, circular aperture  850  therein. Reference numeral  810 A signifies the flame holders in operation. Reference numeral  830  represents a terminal shock wave and its location as illustrated diagrammatically in  FIG. 8A  is important. Location of the terminal shock wave  830  in the ramjet mode is important and is controlled by the position of the nozzle throat  807  diameter D. Reference numeral  872  signifies heat release within the combustor. 
     There is no thermal throat in the dual-mode combustor  899  because the variable nozzle throat  807  is positioned so as to control the terminal shock wave  830 . 
       FIG. 12  is a plot  1200  of the prior art thermal throat of Curran  1201 , the geometric/nozzle throat  1202  of the dual-mode combustor  899  in ramjet mode, the geometric/nozzle throat  1205  in scramjet mode, and the inlet throat  1203 ,  1203 A of the dual-mode combustor  899  as a ratio of A/A capture area.  FIG. 12A  is a table  1200 A of inlet contraction ratios  1231  as a ratio ((A/A capture area)  1231 ) for a range of flight mach numbers and combustion processes  1230 .  FIG. 12  B is a control system  1200 B for positioning the variable (geometric) nozzle throat  807 .  FIG. 12  indicates a discontinuity or jump  1204  between the ramjet mode plot  1202  and the scramjet mode plot  1205 . 
     A nozzle positioner  1212  drives and moves the arc section  1125  forming the nozzle throat  1108 , to a desired diametrical opening according to an algorithm ( FIG. 12  curves,  1202 ,  1205 ) which is a function of flight Mach number and combustor mode (ramjet or scramjet). The algorithm has a discontinuity at a given flight mach number, in this example, flight Mach number 5.0, transitioning from the ramjet mode to the scramjet mode forming a free-jet  943  from the inlet section  802 , through the subsonic diffuser  804 , through the combustion chamber  805 , through the contraction section, and rejoins the nozzle throat  807  (diameter D 1 ). The ramjet mode includes subsonic operation from about flight Mach number 2.5 up to about flight Mach number 5.0 to 6.0 and the cross-sectional area of the nozzle throat  807  divided by the inlet capture area, A inlet capture area, should follow curve  1202 . 
     The scramjet mode includes supersonic operation from about flight Mach number 5.0 to 6.0 up to about flight Mach number 12.0 and greater. The nozzle positioner divergingly adjusts the nozzle throat diameter (nozzle area) rapidly to a relatively larger diameter between about flight Mach number 5.0 to 6.0 rapidly transitioning from the ramjet mode to the scramjet mode forming a free-jet  943  extending from the inlet section  802  at the location of the radial step  812 ,  812 A to the nozzle throat  807 . The free-jet does not engage the subsonic diffuser  804 . Nor does the free-jet  943  engage the combustion chamber  805 . The free-jet  943  rejoins the nozzle throat  807  as illustrated in  FIG. 9A . 
     Referring to  FIG. 8A  and  FIG. 12 , reference numeral  1201  indicates the algorithm for the position of the nozzle throat  807  (diameter D) as a ratio of the inlet capture area (area=A inlet capture area). Specifically, the nozzle throat area must be positioned on the line  1202  for ramjet mode operation for flight numbers between 2.5 to 5.0. Further, the nozzle throat  807  (diameter D 1 ) in the scramjet mode must be positioned on the line  1205  for the scramjet mode operation for flight numbers between 5.0 and 12.0. Reference numeral  1204  represents the transition between the ramjet mode (pursuant to curve or algorithm  1202 ) and the scramjet mode (pursuant to curve or algorithm  1205 ). Operation between the modes is switched back and forth between the curves  1202 ,  1205 . 
     Referring to  FIG. 8A  and  FIG. 12 , the location of the shock wave  830  is important. If the nozzle throat area ratio is positioned below the line  1201  in  FIG. 12 , the engine will un-start as the shock wave moves leftwardly and is expelled from the engine in order to spill air around and past the inlet capture area. Similarly, if the nozzle throat area ratio is positioned above the line  1202  in  FIG. 12 , the engine may prematurely transition to the scramjet mode if the flight Mach number is sufficiently high. Transition to the scramjet mode is accomplished by rapidly changing the nozzle throat ratio (A/A inlet capture area) from curve  1202  to curve  1205  in combination with radially oriented step  1203  which causes the free-jet to separate from the diffuser surface and the combustion chamber. The flame holders  810  have no function. 
       FIGS. 12 and 12A  also indicate that the diameter of the cylindrical inlet  802  changes as a function of ramjet mode (see curve  1203 ), and also cylindrical inlet  802  changes as a function of scramjet mode (see curve  1203 A).  FIG. 12A  indicates that the inlet contraction ratio (A inlet capture area/A inlet cylinder) increases as flight Mach number increases in the ramjet mode up to about flight Mach number 6.0. Further,  FIG. 12A  indicates that the inlet contraction ratio increases as flight Mach number increases in the scramjet mode up to about flight Mach number 12.0.  FIG. 12  reference numerals  1203 ,  1203 A represent the inverse of this data, in other words, the inlet throat diameter ratios (A inlet cylinder/A capture area) are the inverse of the previously defined contraction ratio. 
     As a general rule the nozzle throat  1202  and the inlet throat  1203  decrease with increasing flight Mach number in ramjet mode. Similarly, as a general rule the geometric/nozzle throat  1205  and the inlet throat  1203 A decrease with increasing flight Mach number in scramjet mode. 
     Cycle analysis was performed over the flight Mach number range of 2.5 to 12 at a dynamic pressure of 1500 psfa in order to establish the variable geometry requirements for the inlet area and nozzle throat area. For supersonic combustion cases, a constant-pressure combustion process was assumed with ethylene fuel entering at sonic velocity, parallel to the propulsion axis at the diffuser exit static pressure and 10000 R. 
       FIG. 12  presents the variation of inlet and nozzle throat areas with flight Mach number for various operating modes. Of primary interest is the large variation in nozzle throat area required in the low flight Mach number range. The dual-mode ramjet&#39;s thermal throat area must vary by a factor of 4.5 from Mach 2.5 to 5. The required throat area variation for the conventional ramjet is slightly less over the same range. The thermally-choked cases require a larger throat area at a given flight Mach number because of the greater total pressure loss associated with the transonic combustion process. In the dual-mode engine, the axial location of combustion in a diverging flow path is varied. The fuel distribution and flame-holding mechanisms used for axial modulation of the heat release must not interfere with scramjet-mode operation. These are the fundamental issues associated with extension of the dual-mode to low Mach number flight. Also shown in  FIG. 12  is the inlet throat area variation representative of the contraction ratio. Finally, the combustor-exit area variation as a result of constant-pressure supersonic combustion is shown in  FIG. 12 , and represents the free-jet combustor nozzle throat area design values. 
     The area ratio due to combustion of the propulsive stream decreases with flight Mach number as the incoming energy increases. A factor of 2.5 reduction in nozzle throat area is required between Mach 5 and 12. For all modes of operation, the required variations in throat area shown are a function of the inlet mass capture and pressure recovery characteristics assumed, and while representative for the purposes herein, could be reduced by integration, or other inlet design that results in greater spillage and higher recovery at the lower end of the flight Mach number range. Nozzle throat area variation requirements could also be relieved by a reduction in fuel-air ratio at the lower flight Mach numbers at the expense of net thrust. Obviously, limiting the flight Mach number range would also diminish the variable geometry requirements. 
       FIG. 128  is a control system  1200 B for positioning the variable (geometric) nozzle throat  807 .  FIG. 12B  illustrates desired  1206  ramjet mode (A nozzle/A inlet capture area) ratios switched into a controller  1210  when in the ramjet mode. Similarly,  FIG. 12B  illustrates desired  1208  scramjet mode (A nozzle/A inlet capture area) ratios switched into controller  1210  when in the scramjet mode. Controller  1210 , based on any differences between desired and actual (A nozzle/A inlet capture area), outputs corrective action to the nozzle positioner  1212  which then positions  1214  the variable geometric nozzle throat. A nozzle positioner sensor  1216  in combination with interconnecting lines  1215 ,  1217  communicate the actual (A nozzle/A inlet capture area) signal to controller  1210  for comparison to the desired (A nozzle/A inlet capture area) pursuant to curve or algorithm  1202 ,  1204  and  1205 . 
       FIG. 8B  is a quarter-sectional schematic view  800 B of the dual-mode combustor  899  of  FIG. 8  in the ramjet. mode. Supersonic compression  841  occurs in the inlet contraction section  801  leading to the cylindrical inlet passageway  802 . Arrow  842  indicates fuel injected perpendicularly to the variable diameter inlet cylindrical passageway/section  802 . Multimode fuel injector  8421  injects fuel radially into passageway  802 . Reference numeral  844  illustrates a region of subsonic diffusion and fuel mixing and reference numeral  845  illustrates a region of subsonic combustion. Reference numeral  846  illustrates contraction to a choked throat  807  and reference numeral  847  illustrates expansion and exhaust. 
       FIG. 8C  is an enlargement  800 C of a portion of  FIG. 8B  illustrating, diagrammatically, the radial step  803  between the inlet  802  cylinder and the subsonic diffuser  804 .  FIG. 8C  also illustrates the fuel injector  8421  and the injection of fuel  842 . 
     One of the important benefits of the dual-mode combustor  899 , however, is that the combustion chamber  805  can be used for robust, subsonic combustion at low flight Mach numbers. Operation as a subsonic combustion ramjet (ramjet mode) is illustrated in  FIGS. 8 ,  8 A and  8 B. Fuel injection can be accomplished with a single array of injectors upstream in the inlet section  802 . Ignition and flame-holding  810  can be accomplished with an in-stream device as shown in  FIGS. 8 and 9 . 
       FIGS. 8 ,  8 A and  8 B illustrate the subsonic combustion ramjet mode. At the desired flight condition, transition to free jet mode is effected by increasing the nozzle throat  807  area suddenly and inducing separation at the radial step  803  located at the diffuser inlet. The flame-holding array  810  does not extend across the subsonic diffuser  804 . In particular, the flame-holding array includes an aperture  850  therein which accommodates passage of the free-jet therethrough in the scramjet mode. The subsonic diffuser section, sometimes referred to herein as the subsonic diffuser  804 , satisfies the requirements of operation as a diffuser in ramjet mode, and separated operation in free jet mode. 
     In free-jet mode (scramjet mode) the propulsive stream re-joins the nozzle throat section, D 1 , with a minimum of disruption. The combustion chamber pressure equilibrates to near that of the diffuser exit, and will depend on many factors such as the nozzle throat area, A, the rate of fuel entrainment, and the aerodynamics of the re-circulation region. Overall heat load to the combustion chamber walls depends on the temperature in the recirculation region, and the competing effects of low velocity and increased surface area. 
       FIG. 10  is a perspective of the dual-mode combustor  1000  employing rectangular geometry.  FIG. 10  illustrates inlet contraction section  1000 , inlet minimum area  1002 , subsonic diffuser section  1004 , combustion chamber  1005 , nozzle contraction section  1006 , variable nozzle throat at the joining point of the contraction section  1006  and the expansion section  1007 . Step  1003  and the expansion section  1008  are illustrated in  FIG. 10 . All components of the dual-mode combustor  1000  can vary dimensionally. In general the various components in  FIG. 10  are rectangularly shaped. In this example, the nozzle throat would be rectangular and would be adjustable. 
       FIG. 11  is a quarter-sectional diagrammatic view  1100  of the dual mode combustor  899  in the scramjet mode for flight Mach number 8 illustrating a radial step  1121 A, a hinged diffuser section  1122 , a hinged combustion section  1123 , a hinged contraction section  1124 , a hinged nozzle throat/arc section  1125  and a hinged expansion section  1126 .  FIG. 11A  provides dimensional information  1100 A relating to  FIG. 11  including the radius of the engine at different stages thereof and the axial position of different stages thereof. Reference numeral  1101  represents station  1  (end of cylindrical inflow section), reference numeral  1102  represents the beginning of cylindrical combustion chamber, reference numeral  1107  represents station  7  (end of cylindrical combustion chamber), reference numeral  1108  represents station  8  (nozzle throat), reference numeral  1121  represents the cylindrical inflow chamber, reference numeral  1121 A represents the hinge and aft facing step, reference numeral  1122  represents the diffuser section, reference numerals  1122 A,  1123 A,  1127 ,  1128  represent hinges, reference numeral  1123  represents the combustion chamber, reference numeral  1124  represents the contraction section, reference numeral  1125  represents the arc section, reference numeral  1126  represents exhaust section, reference numeral  1129  represents the termination of the exhaust section, reference numeral  1180  represents station zero (air inlet from air inlet contraction device), reference numeral  1180 A represents the multi-mode fuel injectors, and reference numeral  1181  indicates arrows of incoming air. 
     Also, hinges, H, indicate herein that the geometry of the dual-mode combustor may change around these points between component sections thereof to accommodate flight conditions. Reference numerals  1127  and  1128  signify the interconnection of the arc section  1125  to the contraction section and the expansion section, respectively. In reviewing  FIG. 11  tangency is maintained and required in all examples between the arc sections and the contraction and expansion sections. This means that the hinges are the equivalent of sliding joints. Specifically, although joints  1127 ,  1128  are illustrated diagrammatically as hinges, in fact these diagrammatic “hinges” are limited in their movement such that tangency between the contraction section and the arc section is maintained and the arc section may not bend back or extend such that a line coincident with the contraction section would intersect with the arc section  1125 . Similarly, the hinges illustrated in  FIGS. 13-16 , inclusive, may be considered as sliding joints. 
     Still referring to  FIG. 11 , the hinges diagrammatically indicate that the geometry of the engine changes pursuant to the flight Mach number conditions. Now referring to  FIG. 12A , as a general rule the geometric/nozzle throat  1202  and the inlet throat  1203  decrease with increasing flight Mach number in ramjet mode. See  FIG. 12 . Similarly, as a general rule the geometric/nozzle throat  1205  and the inlet throat  1203 A decrease with increasing flight Mach number in scramjet mode. The axi-symmetric geometry used for the analysis consists of the fixed-length, hinged panels and cylindrical sections is shown in  FIG. 11 . The fixed-length cylindrical inlet section diameter varies with flight Mach number to match the contraction ratio schedule given in  FIG. 12A  with an allowance for fuel injection. A small radial step was placed at station  1  to facilitate flow separation. Generally the radial step is one-tenth the radius of the inlet cylinder. The cylindrical combustor section is sized to accommodate ramjet combustion for the Mach 2.5 flight condition. The nozzle throat is formed by a circular arc of radius equal to one-half that of the inlet capture area. As the required throat area varies with flight condition, the nozzle throat arc length varies such that the contraction and expansion panels maintain tangency. The expansion panel trailing edge is maintained at a fixed radius, giving an exit area equal to twice the inlet capture area. Coordinates for the Mach 8 geometry shown in  FIG. 11  are given in  FIG. 11A . Ethylene fuel enters axially at station  1  ( 1101 ) through injectors  1180 A as illustrated in  FIG. 11 . 
       FIG. 11B  is a view of a receiving joint forming the nozzle throat. Reference numeral  1124 B signifies a nozzle contraction section having a receiving joint  1125 R. Reference numeral  1126 B signifies a nozzle expansion section having a receiving joint  1126 R-receiving. Arc section  1125 B slidingly resides within joints/openings  1124 R,  1126 R such that the rotation of the nozzle contraction section  1124 B and/or the rotation of the nozzle expansion section  1126 R moves the nozzle throat  1108  while maintaining a tangential relationship between the sections  1124 B,  1126 E and the arc section  1125 B. 
       FIG. 13A  is a generalized quarter-sectional diagrammatic view  1300 A of the flight Mach number 2.5 ramjet.  FIG. 13B  is a generalized quarter-sectional diagrammatic view  1300 B of the flight Mach number 3.0 ramjet.  FIG. 13C  is a generalized quarter-sectional diagrammatic view  1300 C of the flight Mach number 4.0 ramjet. 
     All numerical values in  FIGS. 13A-16C , inclusive, are in inches with the radius being indicated on the ordinate (“y”) axis and the axial length indicated on the abscissa (“x”) axis. Also, hinges, H, indicate herein that the geometry of the dual-mode combustor may change around these pivot points between component sections thereof to accommodate flight conditions. Reference numerals H 1  and H 2  signify the interconnection of the arc section to the contraction section and the expansion section, respectively. In reviewing  FIGS. 13A-16C , tangency is maintained and required in all examples between the arc sections and the contraction and expansion sections. 
     The reference numerals used in  FIGS. 13A ,  13 B and  13 C are set forth below. Reference numerals  1301 I,  11311 I,  1321 I represent the respective inlet sections illustrated in  FIGS. 13A ,  13 B and  13 C, respectively. Reference numerals  1301 A,  1311 A,  1321 A represent the arc sections illustrated in  FIGS. 13A ,  13 B and  13 C, respectively. Reference numerals  1301 N;  1311 N,  1321 N represent the variable nozzle throat sections illustrated in  FIGS. 13A ,  13 B and  13 C, respectively. A review of  FIGS. 13A ,  13 B and  13 C, respectively, yields the conclusion that the inlet diametrical section, which is cylindrical, is decreasing in diameter as the flight Mach number is increasing from 2.5 to 4.0 in the ramjet mode while the nozzle throat radius is decreasing with increased flight Mach number. Tangency is maintained in all examples of  FIGS. 13A ,  13 B and  13 C between the arc sections and the contraction and expansion sections. 
       FIG. 14A  is a generalized quarter-sectional diagrammatic view  1400 A of the flight Mach number 5.0 ramjet.  FIG. 14B  is a generalized quarter-sectional diagrammatic view  1400 B of the flight Mach number 5.0 scramjet. Reference numerals  1401 I,  1411 I represent the respective inlet sections illustrated in  FIGS. 14A and 14B , respectively. Reference numerals  1401 A,  1411 A represent the arc sections illustrated in  FIGS. 14A and 14B , respectively. Reference numerals  1401 N,  1411 N represent the variable nozzle throat sections illustrated in  FIGS. 14A and 14B , respectively. A review of  FIGS. 14A and 14B , respectively, yields the conclusion that the inlet diametrical section, which is cylindrical, is slightly increasing in diameter as the engine is transitioning from ramjet flight Mach number 5 to scramjet flight Mach number 5 while the nozzle throat radius is substantially increasing while transitioning from ramjet flight Mach number 5 to scramjet flight Mach number 5. Tangency is maintained in all examples of  FIGS. 14A and 14B  between the arc sections and the contraction and expansion sections. 
       FIG. 15A  is a generalized quarter-sectional diagrammatic view  1500 A of the flight Mach number 6.0 ramjet.  FIG. 15B  is a generalized quarter-sectional diagrammatic view  1500 B of the flight Mach number 6.0 scramjet. Reference numerals  1501 I,  1511 I represent the respective inlet sections illustrated in  FIGS. 15A and 15B , respectively. Reference numerals  1501 A,  1511 A represent the arc sections illustrated in  FIGS. 15A and 15B , respectively. Reference numerals  1501 N,  1511 N represent the variable nozzle throat sections illustrated in  FIGS. 15A and 15B , respectively. A review of  FIGS. 15A and 15B , respectively, yields the conclusion that the inlet diametrical section, which is cylindrical, is slightly increasing in diameter as the engine is transitioning from ramjet flight Mach number 6 to scramjet flight Mach number 6 while the nozzle throat radius is substantially increasing while transitioning from ramjet flight Mach number 6.0 to scramjet flight Mach number 6.0. Tangency is maintained in all examples of  FIGS. 15A and 15B  between the arc sections and the contraction and expansion sections. 
       FIG. 16A  is a generalized quarter-sectional diagrammatic view  1600 A of the flight Mach number 8.0 scramjet.  FIG. 16B  is a generalized quarter-sectional diagrammatic view  1600 B of the flight Mach number 10.0 scramjet.  FIG. 16C  is a generalized quarter-sectional diagrammatic view  1600 C of the flight Mach number 12.0 scramjet. Reference numerals  1601 I,  1611 I,  1621 I represent the respective inlet sections illustrated in  FIGS. 16A ,  16 B and  16 C, respectively. Reference numerals  1601 A,  1611 A,  1621 A represent the arc sections illustrated in  FIGS. 16A ,  16 B and  16 C, respectively. Reference numerals  1601 N,  1611 N,  1621 N represent the variable nozzle throat sections illustrated in  FIGS. 16A ,  16 B and  16 C, respectively. A review of  FIGS. 16A ,  16 B and  16 C, respectively, yields the conclusion that the inlet diametrical section, which is cylindrical, is slightly decreasing in diameter as the flight Mach number is increasing from 8.0 to 10.0 in the scramjet mode while the nozzle throat radius is moderately decreasing with increased flight Mach number. Tangency is maintained in all examples of  FIGS. 16A ,  168  and  16 C between the contraction and expansion sections. 
     Contours of static pressure ratio for flight Mach numbers 5, 8 and 12 in the scramjet mode flight conditions appear in  FIGS. 17A ,  17 B, and  17 C.  FIG. 17A  is an illustration of the pressure contours  1700 A within the engine for the flight Mach number 5.0 scramjet.  FIG. 17B  is an illustration of the pressure contours  1700 B within the engine for the flight Mach number 8.0 scramjet.  FIG. 17C  is an illustration of the pressure contours  1700 C within the engine for the flight Mach number 12.0 scramjet. 
     Referring to  FIG. 17A , pressure ratio contours, P/Pinlet, for the flight Mach number 5.0 scramjet are illustrated and pressure ratio, P/Pinlet,  1701 , has a magnitude of about 1.04 and is located generally in the recirculation zone, forward portion of the combustion chamber. Reference numeral  1731  is a stagnation streamline. When viewing  FIG. 17A , everything leftwardly of stagnation streamline  1731  is in the recirculation zone. Reference numeral  1731 A represents a free-jet streamline. 
     Referring to  FIG. 17B , pressure ratio, P/Pinlet,  1711 , for the flight Mach number 8.0 scramjet, pressure ratio has a magnitude of about 1.32 and is located generally in the recirculation zone of the forward portion of the combustion chamber. When viewing  FIG. 17B , everything leftwardly of stagnation streamline  1732  is in the recirculation zone. Reference numeral  1732 A represents a free-jet streamline. 
     Referring to  FIG. 17C , pressure ratio, P/Pinlet,  1712 , for the flight Mach number 12.0 has a magnitude of about 1.18 and is located generally in the recirculation zone of the forward portion of the combustion chamber. When viewing  FIG. 17C , everything leftwardly of stagnation streamline  1733  is in the recirculation zone. Reference numeral  1733 A represents a free jet streamline. 
     Reviewing  FIG. 17 , the recirculation zone pressure ratios increase from scramjet flight Mach number 5 to 8 and then decrease from between flight Mach number 8 to 12. 
     Contours of Mach number for flight Mach numbers 5, 8 and 12 in the scramjet mode flight conditions appear in  FIGS. 18A ,  18 B, and  18 C. All three cases for flight Mach numbers 5, 8 and 12 exhibit periodic wave structure in the free-jet, and an overall increase in cross-sectional area due to combustion as the jet traverses the combustion chamber. In all cases the free-jet rejoins the nozzle throat contour and expands to the exit area. The free jet drives a primary recirculation zone in the combustion chamber, the center of which moves aft with increasing flight Mach number. Streamlines in the combustion chamber define the recirculation zone. In the inlet section, and continuing in a conical non-influence region of the free-jet, supersonic combustion elevates the pressure to a level higher than that of the reference pressure at the inflow plane. The highest pressure occurs on the axis, followed by an expansion initiated at the jet boundary. In the Mach 8 and 12 cases, the recirculation zone equilibrates to the pressure at the radial step and the bounding streamline issues axially with little initial deflection. In the Mach 5 case, the recirculation zone equilibrates to a lower pressure, causing an initial expansion of the free-jet at the step. All cases show a subsequent divergence of streamlines required to accommodate the continuing supersonic combustion process while matching combustion chamber pressure. This “entry interaction” initiates the repetitive streamline structure characteristic of an under-expanded jet. The severity of the entry interaction depends on the initial rate of mixing and combustion in the free-jet, and its initial pressure with respect to the recirculation zone. The wavelength and shock losses associated with the streamline structure depend on the entry interaction. At the combustor exit, the Mach 5 case approaches a sonic condition, and its wave structure disappears. Streamlines in the Mach 8 case appear to be approximately in phase with the throat geometry, and the streamlines merge smoothly into the minimum area. The Mach 12 case however, exhibits an “exit interaction” as streamlines are forced to converge, resulting in a strong shock wave on the axis. This interaction could obviously be eliminated by reducing the wavelength of the shock structure or moving the throat, but of most benefit from a propulsion standpoint would be to eliminate the periodic streamline structure altogether by mitigating the entry interaction. 
       FIG. 18A  is an illustration  1800 A of the Mach number contours within the engine for the flight Mach number 5.0 scramjet.  FIG. 18B  is an illustration  1800 B of the Mach number contours within the engine for the flight Mach number 8.0 scramjet.  FIG. 18C  is an illustration  1800 C of the Mach number contours within the engine for the flight Mach number 12.0 scramjet. Referring to  FIG. 18A , reference numeral  1801  indicates a magnitude of about Mach 0.0 located in the recirculation zone of the forward portion of the combustion chamber. 
     Referring to  FIG. 18B , reference numeral  1810  indicates a magnitude of about Mach 0.0 located in the recirculation zone in the middle of the combustion chamber. Referring to  FIG. 18C , reference numeral  1821  represents a magnitude of about Mach 0.0 located in the recirculation zone of the aft portion of the combustion chamber. 
       FIG. 19A  is an illustration of the static temperature contours  1900 A within the engine for the flight Mach number 5.0 scramjet.  FIG. 19B  is an illustration of the static temperature contours  1900 B within the engine for the flight Mach number 8.0 scramjet.  FIG. 19C  is an illustration of the static temperature contours  1900 C within the engine for the flight Mach number 12.0 scramjet. 
     Referring to  FIG. 19A , reference numeral  1901  indicates a temperature of about 3500° R and reference numeral  1903  indicates a temperature of about 5000° R. Referring to  FIG. 19B , reference numeral  1906  indicates a temperature of about 6000° R. Referring to  FIG. 19C , reference numeral  1910  indicates a temperature of about 9000° R. 
     Temperature contours appear in  FIGS. 19A ,  19 B and  19 C. The effects of combustion are apparent in the individual shear layers. The Mach 5 case shows a degree of stratification that persists into the nozzle throat. The recirculation zone equilibrates to greater than 90% of the ethylene-air theoretical value in the Mach 8 and 12 cases, but is significantly cooler in the Mach 5 case. This is likely due to the two-injector arrangement used in the Mach 5 case, and suggests that the recirculation zone temperature and combustor heat load depend on the fuel injection method, and could be reduced in future design iterations. Exit interaction in the Mach 12 case may also contribute to elevated temperature in the recirculation zone. 
     In order to make a quantitative assessment of the losses in the free-jet combustion process, and their effect on net thrust, mass-averaged axial distributions of pressure, temperature, and velocity were obtained during the analysis. The combustor friction coefficient thus represents the momentum loss associated with the recirculation zone and shock structure in the free-jet. The ideal net thrust per unit airflow is illustrated in  FIG. 20 . 
       FIG. 20  illustrates the ideal net thrust per unit airflow based on use of different computational methods/tools.  FIG. 20  illustrates ideal net thrust per unit of airflow against flight Mach numbers for a conventional ramjet, thermally-choked ramjet and a scramjet. Reference numeral  2001  represents the ideal net thrust for scramjet mode operation. Reference numeral  2003  represents the ideal net thrust for the thermally choked operation such as in Curran et al. Reference numeral  2002  represents the ideal net thrust for the ramjet disclosed herein.  FIG. 20  illustrates a comparison of a thermally choked ramjet to the dual-mode ramjet disclosed herein. The subsonic combustion ramjet disclosed herein is 6-8% more efficient than the thermally-choked or “dual-mode” ramjet as a consequence of lower combustion Mach number. Of greater significance than higher performance however, is the practicality of fuel distribution and flame-holding in the conventional ram burner. 
       FIG. 21  illustrates the mass-averaged static pressure distributions  2100  with the pressure at the nozzle throat station denoted by symbols (supersonic combustor exit) for various flight conditions, to with, scramjet flight Mach numbers 5, 8 and 12. Reference numeral  2101  represents Mach 5 pressure ratio data, reference numeral  2102  represents Mach 8 pressure ratio data, and reference numeral  2103  represents Mach 12 pressure ratio data. Compression due to mixing and combustion in the cylindrical inlet section from station zero to 0.36 feet is evident, as is the subsequent expansion and periodic streamline structure. As the free-jet traverses the combustion chamber, the mean pressure is generally above the inflow value, consistent with the elevated recirculation zone pressures. The Mach 5 pressure distribution shows a damped character as combustion drives the free-jet toward a sonic condition. Of interest is the phase shift and elevated amplitude of the last peak in the Mach 12 case consistent with the exit interaction seen in the pressure contours. Note that the combustor exit pressure (at the minimum area) used for cycle analysis of the Mach 8 and 12 solutions is significantly higher than the inflow, and would cause a discrepancy with cycle analysis assuming combustion at constant pressure. 
       FIG. 21A  illustrates the mass-averaged axial velocity ratio (V/V inlet) distributions  2100 A for various flight conditions, to with, scramjet flight Mach numbers 5, 8 and 12. Reference numeral  2111  represents Mach 5 velocity ratio data, reference numeral  2112  represents Mach 8 velocity ratio data, and reference numeral  2113  represents Mach 12 velocity ratio data. A marked reduction in velocity occurs upstream of the throat station for the Mach 8 and 12 cases, and is more gradual for the Mach 5 case, consistent with the pressure distributions. The loss coefficients used to match the combustor exit velocities are listed in the  FIG. 21A . Shock and viscous losses are represented in these values, and an estimate of their relative contributions to the total is not determined. Shock losses arise from the entry and exit interactions discussed above, and may be reduced by better tailoring of the combustion process, and optimization of the combustion chamber geometry. The viscous loss arises from the momentum required to drive the recirculating flow in the combustion chamber, which presumably is a function of the combustion chamber volume and wetted area. These are determined by the cross-sectional area required at the minimum ramjet Mach number, subsonic diffuser length requirements, and the free jet length required for supersonic mixing and combustion. 
       FIG. 21B  illustrates the mass-averaged temperature distributions  2100 B for scramjet mode flight Mach numbers 5, 8 and 12. Reference numeral  2121  indicate Mach 5 temperature data as a function axial position, reference numeral  2122  indicate Mach 8 temperature data as a function axial position, and reference numeral  2123  represents Mach 12 temperature data as a function of axial position. Temperatures increase with increasing Mach flight numbers. 
       FIG. 23A  illustrates the ethylene mass fraction  2300 A for flight Mach numbers 5, 8 and 12 versus axial position. Reference numeral  2305  signifies the flight Mach number 5, reference numeral  2306  signifies the flight Mach number 8, and reference numeral  2307  signifies the flight Mach number 12. 
     Calculations at various nozzle throat areas were performed in order to evaluate the effect on recirculation zone pressure, entry and exit interactions, and performance at the flight Mach number 8 as illustrated in  FIGS. 11 and 16A .  FIGS. 22A ,  22 B,  22 C and  22 D illustrate static pressure contours for throat areas equal to 110%, 100%, 90% and 80% of the design value.  FIG. 17B  and  FIG. 22B  are identical but different data is presented and discussed in connection with each drawing figure. 
       FIG. 22A  illustrates the static pressure ratio  2200 A for scramjet mode flight Mach number 8 with the variable nozzle throat positioned at 110% of the design operating point. Reference numeral  2201  indicates a stagnation streamline and reference numeral  2202  indicates the pressure ratio of 0.95 located in recirculation zone of the combustion chamber (110% nozzle throat ratio). When viewing  FIG. 22A , everything to the left of stagnation streamline  2201  is in the recirculation zone. Reference numeral  2221 T is the nozzle throat location (110% nozzle throat area ratio). 
       FIG. 22B  illustrates the static pressure ratio  2200 B for scramjet mode flight Mach number 8 with the variable nozzle throat positioned at 100% of the design operating point. Reference numeral  2203  represents a stagnation streamline and reference numeral  2204  indicates a pressure ratio of 1.32 located in the recirculation zone of combustion chamber (100% nozzle throat ratio). When viewing  FIG. 22B , everything to the left of stagnation streamline  2203  is in the recirculation zone. Reference numeral  2223 T is the nozzle throat location (100% nozzle throat ratio). 
       FIG. 22C  illustrates the static pressure ratio  2200 C for scramjet mode flight Mach number 8 with the variable nozzle throat positioned at 90% of the design operating point. Reference numeral  2205  represents a stagnation streamline and reference numeral  2206  is the pressure ratio of 1.60 located in recirculation zone of combustion chamber (90% nozzle throat ratio). When viewing  FIG. 22C , everything to the left and above the stagnation streamline  2205  is in the recirculation zone. Reference numeral  2225 T is the nozzle throat location (90% nozzle throat ratio). 
       FIG. 22D  illustrates the static pressure ratio  2200 D for scramjet mode flight Mach number 8 with the variable nozzle throat positioned at 80% of the design operating point. Reference numeral  2207  represents the stagnation streamline and reference numeral  2208  represents the pressure ratio of 1.87 located in recirculation zone of combustion chamber (80% nozzle throat ratio). When viewing  FIG. 22D , everything to the left and above stagnation streamline  2207  is in the recirculation zone. Reference numeral  2227 T is the nozzle throat location (80% nozzle throat ratio). 
     As throat area is reduced, combustion chamber pressure increases, and the period of the streamline structure decreases. As expected, combustion in the inlet section, and a short distance downstream is not affected. Beyond this however, increased pressure increases the rate of combustion, reinforcing the tendency toward shorter wavelengths. Reference numerals  2201 ,  2203 ,  2205  and  2207  represent the streamlines and streamline  2207  (variable nozzle throat at 80% of design value) has a shorter wavelength than streamline  2201  (variable nozzle throat at 110%) or streamline  2203  (variable nozzle throat at 100%). Further, the pressure increase in the combustion chambers is viewed in  FIGS. 22A ,  22 B,  22 C and  22 D as the variable nozzle&#39;s area is reduced. Referring back now to  FIGS. 22A-D , it is evident that the free jet entry conditions range from under-expanded at 110% throat area to over-expanded at 80%, but the streamline structure is never eliminated due to the rapidity of combustion and divergence of streamlines in the inlet region. The severity of the exit interaction depends on synchronization of the streamline structure with the throat geometry. The streamline  2203  associated with the variable nozzle throat at 100% of the design case appears to be in phase and exhibits almost no exit interaction with the nozzle throat. Reference numeral  2223 T represents the variable nozzle throat for the 100% example. Reference numerals  2221 T,  2225 T and  2207 T represent the throats in the examples where the variable nozzle throat is 110%, 90% and 80%, respectively. Interference with the throat is greatest for the 80 and 110% cases which show the strongest interactions. 
       FIG. 23  illustrates the effect of nozzle throat area variation  2300  for scramjet mode flight Mach number 8 on the rate of ethylene fuel depletion. Reference numeral  2301  signifies the effect of throat area variation on ethylene mass fraction (110% nozzle throat ratio), reference numeral  2302  signifies the effect of throat area variation on ethylene mass fraction (100% nozzle throat ratio), reference numeral  2303  signifies the effect of throat area variation on ethylene mass fraction (90% nozzle throat ratio), and reference numeral  2304  signifies the effect of throat area variation on ethylene mass fraction (80% nozzle throat ratio). 
       FIG. 24  illustrates the effect of nozzle throat area variation on mass-averaged static pressure distribution  2400  for scramjet mode flight Mach number 8. Reference numeral  2401  signifies the effect of throat area variation on mass averaged static pressure distribution for the flight Mach number 8 (110% nozzle throat ratio), reference numeral  2402  signifies the effect of throat area variation on mass averaged static pressure distribution for the flight Mach number 8 (100% nozzle throat ratio), reference numeral  2403  signifies the effect of throat area variation on mass averaged static pressure distribution for the flight Mach number 8 (90% nozzle throat ratio), and reference numeral  2404  signifies the effect of throat area variation on mass averaged static pressure distribution for the flight Mach number 8 (80% nozzle throat ratio). 
     Mass-averaged pressure distributions for scramjet mode flight Mach number 8 illustrated in  FIG. 24  also show that as throat area is reduced, the initial pressure rise increases, the period of the streamline structure decreases, and the mean is approximately equal to the recirculation zone pressure. Peak-to-peak amplitude is roughly the same for all cases. Note that for the 100% case, the waveform merges smoothly with the nozzle expansion. The designation A 8  in  FIG. 24  refers to  FIG. 11 , station  8 , reference numeral  1108 . The 110% case shows a slight slope discontinuity just prior to the throat station and the 80 and 90% cases show out-of-phase features at the throat, consistent with the interactions seen in the pressure contours. 
     The effect of throat area variation was to change the combustion chamber pressure and the period of the streamline structure without significantly altering its amplitude. The amplitude of the primary streamline structure is, therefore, most likely dependent on the initial rate of combustion. The exit interaction was affected by the phasing of the shock structure and was nearly eliminated in the 100% throat area case. The adiabatic wall temperature and the gas temperature in the recirculation zone were not significantly affected by throat area variation. 
       FIG. 25  illustrates  2500  the ideal net thrust per unit of airflow plotted against combustor exit pressure ratio and nozzle throat area variation for scramjet mode flight Mach number 8. The ideal net thrust per unit airflow for the example of scramjet flight Mach number 8 for variable nozzle throat opening ratios (80%, 90%, 100% and 110%) is plotted  2501  versus the mass-averaged combustor exit pressure ratio in  FIG. 25 . Reference numeral  2502  represents the ideal net thrust per unit airflow with a Cf of 0.0025. Friction loss coefficients required to match the exit velocities are also listed with the throat area for each point. The 90% variable nozzle throat case exhibits the least momentum loss, the 110% case the greatest, and despite the entry and exit interactions seen in pressure contours for the 80% case, its loss coefficient is slightly less than the 100% case which showed little interaction. This relative insensitivity and lack of correlation of loss coefficient to throat area is not unexpected however, since the amplitude of the basic streamline structure, and presumably the viscous loss component were not significantly affected. Cycle analysis results at the corresponding pressure ratios and with nominal momentum loss are also plotted for reference and to show the basic sensitivity of scramjet net thrust to combustor pressure ratio. 
     REFERENCE NUMERALS 
     Reference numerals  10 - 86  pertain to the prior art.
           10 —aircraft     12 —ramjet combustion engine     14 —inlet scoop     16 —exhaust outlet     17 ,  18 ,  19 —walls     20 —fourth wall     21 —converging inlet cowl passage     22 —diverging supersonic combustion section     24 —substantially uniform cross section subsonic combustion section     26 —exit nozzle     27 —pilot zone recesses     28 —fuel pump     30 —fuel control system     32 —plurality of nozzles     34 —fuel control system     36 —plurality of nozzles     40 —central body     42 —elongated inlet spike     43 —flameholders     44 —exhaust plug     46 —annular member     47 ,  48 —struts     49 —fuel pump     50 —subsonic combustion chamber     51 —fuel control     52 —nozzles  52  in the struts  47       55 —nozzles supplied from fuel ducts     56 —fuel ducts     58 —recesses     60 —supersonic combustion chamber     61 —fuel control     62 ,  64 —nozzles     65 —ducts     70 ,  72 —pumps     74 ,  75 —nozzles     76 ,  78 —fuel control system     80 —supersonic combustion chamber     82 —subsonic chamber     86 —recess pilot zones     600 —cross-sectional view of a prior art dual mode supersonic ramjet engine operating in the scramjet mode     601 —fuel injection nozzle     602 —inlet contraction section     603 —diverging supersonic combustion section     604 —exit nozzle     605 —fuel-air mixture     606 ,  606 A,  880 —incoming air being compressed     607 ,  607 A,  881 —exiting combustion gases     608 —interior wall of engine     700 —cross-sectional view of a prior art dual mode supersonic ramjet engine operating in the thermally-choked ramjet mode     701 —shock train to subsonic ramjet mode     702 —beginning of shock train to subsonic ramjet mode     703 —fuel injector     704 —fuel injector
 
Reference numerals  800  and above pertain to the disclosed and claimed invention.
     800 —perspective view of dual-mode combustor operating in the ramjet mode     800 A—cross-sectional schematic view of the dual-mode combustor operating in the ramjet mode     800 B—quarter sectional schematic view of the dual-mode combustor operating in the ramjet mode     800 C—enlarged portion of  FIG. 8A  illustrating the radial step and the multimode fuel injector     801 —inlet contraction section     802 —inlet minimum area, variable diameter inlet cylindrical passageway/section     803 —radial step     804 —subsonic diffuser section     805 —combustion chamber     806 —nozzle contraction section     807 —variable nozzle throat at the joining point of the contraction section  806  and the expansion section     808  in the ramjet mode or the scramjet mode     808 —nozzle expansion section     810 —ramjet mode flame holder     812 —beginning of radial step  803       812 A—end of radial step  803       830 —terminal shock waves, position controlled by algorithm governing nozzle throat position     841 —supersonic compression     842 —arrow indicating fuel injected     8421 —multimode fuel injector     844 —subsonic diffusion and fuel mixing     845 —subsonic combustion     845 A—supersonic combustion     846 —contraction to choked throat     847 ,  847 A—expansion     850 —aperture in flame holder  810  for the passage of the free-jet     872 —heat release     899 —dual-mode combustor     900 —perspective view of dual-mode combustor operating in the scramjet mode     900 A—cross-sectional schematic view of the dual-mode combustor operating in the scramjet mode     900 B—quarter sectional schematic view of the dual-mode combustor operating in the scramjet mode     900 C—cross-sectional perspective view of the diffuser illustrating the array of flame holders  810  and a central aperture  850  within the array of flame holders  810       943 —free-jet in the scramjet mode     943 A—supersonic free jet boundary wherein the pressure is approximately equal with that of the recirculation zone     944 —recirculation zone     972 —heat release     1000 —perspective view of a dual-mode combustor using different geometry     1001 —inlet contraction section     1002 —inlet minimum area     1003 —step     1004 —subsonic diffuser section     1005 —combustion chamber     1006 —nozzle contraction section     1007 —variable nozzle throat at the joining point of the contraction section  1006  and the expansion section  1008       1008 —expansion section     1100 —quarter sectional view of the dual-mode combustor in the scramjet mode for flight Mach number 8     1100 A—dimensional information for the quarter sectional view of the dual-mode combustor in the scramjet mode for flight Mach number 8     1100 B—view of receiving joint forming the nozzle throat     1101 —station  1 , end of cylindrical inflow section     1102 —station  2 , beginning of cylindrical combustion chamber     1107 —station  7 , end of cylindrical combustion chamber     1108 —station  8 , nozzle throat     1121 —cylindrical inflow chamber     1121 A—hinge and aft facing step     1122 —diffuser section     1122 A,  1123 A,  1127 ,  1128 —hinge, sliding joint     1123 —combustion chamber     1124 —contraction section     1125 —arc section     1126 —expansion section     1124 B—nozzle contraction section     1126 B—nozzle expansion section     1126 R—receiving joint     1125 B—arc section     1125 R—receiving joint     1129 —termination of expansion section     1180 —station zero, station i, air inlet from air inlet contraction device     1180 A—multi-mode fuel injectors     1181 —arrows representing incoming air     1200 —illustration of flight Mach number versus thermal throat for prior art device, geometric/nozzle throat for dual-mode combustor of present invention in ramjet mode and in scramjet mode as a ratio of inlet capture area, and inlet throat in ramjet mode and scramjet mode as a ratio of inlet capture area     1200 A—table of flight Mach numbers versus inlet contraction ratios, Ac/Ai     1200 B—variable nozzle throat position schematic     1201 —thermal throat of prior art device     1202 —geometric/nozzle throat expressed as a ratio of nozzle throat area to inlet capture area in ramjet mode     1203 —dual mode combustor, inlet throat in ramjet mode     1203 A—dual mode combustor, inlet throat in scramjet mode     1204 —discontinuity/jump of variable nozzle throat position between the ramjet mode  1202  and the scramjet mode  1205       1205 —geometric/nozzle throat expressed as a ratio of nozzle throat area to inlet capture area in scramjet mode     1206 —desired ramjet nozzle throat position as a function of flight Mach number for the ramjet mode     1207 ,  1209 —switch     1208 —desired ramjet nozzle throat position as a function of flight Mach number for the scramjet mode     1210 —controller operating on the difference of desired position of the nozzle throat minus the actual position of the nozzle throat     1211 —output of controller     1212 —nozzle throat positioner     1213 —position signal     1214 —variable geometric nozzle throat     1215 ,  1217 —interconnecting signal transmission lines     1216 —nozzle throat position sensor     1218 —actual nozzle throat position as a function of flight Mach number     1230 —inlet contraction ratio     1231 —combustion process     1300 A—quarter-sectional schematic profile of the dual-mode combustor in the ramjet mode, flight Mach number 2.5     1300 B—quarter-sectional schematic profile of the dual-mode combustor in the ramjet mode, flight Mach number 3     1300 C—quarter-sectional schematic profile of the dual-mode combustor in the ramjet mode, flight Mach number 4     1301 A,  1311 A,  1321 A—arc section     1301 C,  1311 C,  1321 C—combustion chamber     1301 D,  1311 D,  1321 D—diffuser section     1301 E,  1311 E,  1321 E—expansion section     1301 I,  1311 I,  1321 I—inlet section     1301 N,  1311 N,  1321 N—variable nozzle throat section     1301 X,  1311 X,  1321 X—contraction section     1400 A—quarter-sectional schematic profile of the dual-mode combustor in the ramjet mode, flight Mach number 5     1400 B—quarter-sectional schematic profile of the dual-mode combustor in the scramjet mode, flight Mach number 5     1401 A,  1411 A—arc section     1401 C,  1411 C—combustion chamber     1401 D,  1411 D—diffuser section     1401 E,  1411 E—expansion section     1401 I,  1411 I—inlet section     1401 N,  1411 N—variable nozzle throat section     1401 X,  1411 X—contraction section     1500 A—quarter-sectional schematic profile of the dual-mode combustor in the ramjet mode, flight Mach number 6     1500 B—quarter-sectional schematic profile of the dual-mode combustor in the scramjet mode, flight Mach number 6     1501 A,  1511 A—arc section     1501 C,  1511 C—combustion chamber     1501 D,  1511 D—diffuser section     1501 E,  1511 E—expansion section     1501 I,  1511 I—inlet section     1501 N,  1511 N—variable nozzle throat section     1501 X,  1511 X—contraction section     1600 A—quarter-sectional schematic profile of the dual-mode combustor in the scramjet mode, flight Mach number 8     1600 B—quarter-sectional schematic profile of the dual-mode combustor in the scramjet mode, flight Mach number 10     1600 C—quarter-sectional schematic profile of the dual-mode combustor in the scramjet mode, flight Mach number 12     1601 A,  1611 A,  1621 A—arc section     1601 C,  1611 C,  1621 C—combustion chamber     1601 D,  1611 D,  1621 D—diffuser section     1601 E,  1611 E,  1621 E—expansion section     1601 I,  1611 I,  1621 I—inlet section     1601 N,  1611 N,  1621 N—variable nozzle throat section     1601 X,  1611 X,  1621 X—contraction section     1700 A—pressure ratio, P/Pinlet, for the flight Mach number 5.0     1700 B—pressure ratio, P/Pinlet, for the flight Mach number 8.0     1700 C—pressure ratio, P/Pinlet, for the flight Mach number 12.0     1701 —pressure ratio, P/Pinlet, about 1.04 located generally in the forward portion of the combustion chamber     1711 -pressure ratio, P/Pinlet, about 1.32 located generally in the forward portion of the combustion chamber     1721 —pressure ratio, P/Pinlet, about 1.18 located generally in the forward portion of the combustion chamber     1731 ,  1732 ,  1733 —stagnation streamline     1731 A,  1732 A,  1733 A—free-jet streamline     1800 A—Mach number contours for the flight Mach number 5.0     1800 B—Mach number contours for the flight Mach number 8.0     1800 C—Mach number contours for the flight Mach number 12.0     1801 —about Mach 0.0, located in the recirculation zone of the forward portion of the combustion chamber     1810 —about Mach 0.0, located in the recirculation zone in the middle of the combustion chamber     1821 —about Mach 0.0, located in the recirculation zone of the aft portion of the combustion chamber     1900 A—static temperature contours for the flight Mach number 5.0     1900 B—static temperature contours for the flight Mach number 8.0     1900 C—static temperature contours for the flight Mach number 12.0     1901 —3500° R     1903 —5000° R     1906 —6000° R     1910 —9000° R     2000 —ideal net thrust per unit airflow over various flight Mach numbers     2001 —scramjet mode net thrust     2002 —conventional, prior art, net thrust in the ramjet mode     2003 —Curran (prior art) ramjet mode net thrust     2100 —mass averaged pressure distributions for scramjet flight mach numbers 5, 8 and 12     2100 A—mass averaged axial velocity distributions for scramjet flight mach numbers 5, 8 and 12     2100 B—mass averaged temperature distributions for scramjet flight mach numbers 5, 8 and 12     2101 —Mach 5 pressure ratio data as a function of axial position     2102 —Mach 8 pressure ratio data as a function of axial position     2103 —Mach 12 pressure ratio data as a function of axial position     2111 —Mach 5 axial velocity ratio data as a function of axial position     2112 —Mach 8 velocity ratio data as a function of axial position     2113 —Mach 12 velocity ratio data as a function of axial position     2121 —Mach 5 temperature data as a function of axial position     2122 —Mach 8 temperature data as a function of axial position     2123 —Mach 12 temperature data as a function of axial position     2200 A—static pressure plot for variable area nozzle throat position at 110% of design point for the flight Mach number 8     2200 B—static pressure plot for variable area nozzle throat position at 100% of design point for the flight Mach number 8     2200 C—static pressure plot for variable area nozzle throat position at 90% of design point for the flight Mach number 8     2200 D—static pressure plot for variable area nozzle throat position at 80% of design point for the flight Mach number 8     2201 —stagnation streamline     2202 —pressure ratio of 0.95 located in recirculation zone of combustion chamber (110% nozzle throat ratio)     2203 —stagnation streamline line     2204 —pressure ratio of 1.32 located in recirculation zone of combustion chamber (100% nozzle throat ratio)     2205 —stagnation streamline line     2206 —pressure ration of 1.60 located in recirculation zone of combustion chamber (90% nozzle throat ratio)     2207 —stagnation streamline line     2208 —pressure ratio of 1.87 located in recirculation zone of combustion chamber (80% nozzle throat ratio)     2221 T—nozzle throat location (110% nozzle throat ratio)     2223 T—nozzle throat location (100% nozzle throat ratio)     2225 T—nozzle throat location (90% nozzle throat ratio)     2227 T—nozzle throat location (80% nozzle throat ratio)     2300 —effect of throat area variation on ethylene mass fraction for the flight Mach number 8     2300 A—ethylene mass fraction for scramjet mode flight Mach numbers 5, 8 and 12 versus axial position     2301 —effect of throat area variation on ethylene mass fraction (110% nozzle throat ratio)     2302 —effect of throat area variation on ethylene mass fraction (100% nozzle throat ratio)     2303 —effect of throat area variation on ethylene mass fraction (90% nozzle throat ratio)     2304 —effect of throat area variation on ethylene mass fraction (80% nozzle throat ratio)     2305 —Mach flight number 5 axial position and ethylene mass fraction     2306 —Mach flight number 8 axial position and ethylene mass fraction     2307 —Mach flight number 12 axial position and ethylene mass fraction     2400 —effect of throat area variation on mass averaged static pressure distribution for the flight Mach number 8     2401 —effect of throat area variation on mass averaged static pressure distribution for the flight Mach number 8 (110% nozzle throat ratio)     2402 —effect of throat area variation on mass averaged static pressure distribution for the flight Mach number 8 (100% nozzle throat ratio)     2403 —effect of throat area variation on mass averaged static pressure distribution for the flight Mach number 8 (90% nozzle throat ratio)     2404 —effect of throat area variation on mass averaged static pressure distribution for the flight Mach number 8 (80% nozzle throat ratio)     2500 —ideal net thrust per unit airflow as a function of nozzle throat pressure ratio, Pnozzle/Pinlet     2501 —net thrust per unit airflow for the current free-jet disclosed herein     2502 —net thrust per unit airflow with a Cf of 0.0025.   A=Cross-sectional area   Cf=Friction coefficient   D=Nozzle throat diameter ramjet mode   D 1 =Nozzle throat diameter scramjet mode   H=Hinge/sliding joint   H 1 =First Arc Hinge/sliding joint   H 2 =Second Arc Hinge/sliding joint   M=Mach number   P=Pressure   r=Radial distance   x=Axial distance   Z=Altitude       

     SUBSCRIPTS 
     
         
         
           
               0 =Freestream 
               1 =Cylindrical inflow section exit station 
               2 =Combustion chamber inlet station 
               7 =Combustion chamber exit station 
               8 =Nozzle throat station 
             C=Inlet capture area 
             i=Inflow station 
             min=Minimum 
             T=Total 
           
         
       
    
     Those skilled in the art will readily recognize that the invention has been set forth by way of example only and that changes may be made to the examples without departing from the spirit and the scope of the claims which follow herein below.