Abstract:
One embodiment of the present invention is a unique pulse detonation combustion system. Another embodiment is a unique gas turbine engine. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for pulse detonation combustion systems, gas turbine engines, and other machines and engines. Further embodiments, forms, features, aspects, benefits, and advantages of the present application shall become apparent from the description and figures provided herewith.

Description:
CROSS REFERENCE TO RELATED APPLICATIONS 
       [0001]    The present application claims benefit of U.S. Provisional Patent Application No. 61/427,731, filed Dec. 28, 2010, entitled GAS TURBINE ENGINE AND PULSE DETONATION COMBUSTION SYSTEM, which is incorporated herein by reference. 
     
    
     FIELD OF THE INVENTION 
       [0002]    The present invention relates to gas turbine engines and pulse detonation combustion systems. 
       BACKGROUND 
       [0003]    Combustion systems, such as for gas turbine engines and other machines, remain an area of interest. Some existing systems have various shortcomings, drawbacks, and disadvantages relative to certain applications. Accordingly, there remains a need for further contributions in this area of technology. 
       SUMMARY 
       [0004]    One embodiment of the present invention is a unique pulse detonation combustion system. Another embodiment is a unique gas turbine engine. Other embodiments include unique apparatuses, systems, devices, hardware, methods, and combinations for pulse detonation combustion systems, gas turbine engines, and other machines and engines. Further embodiments, forms, features, aspects, benefits, and advantages of the present application shall become apparent from the description and figures provided herewith. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0005]    The description herein makes reference to the accompanying drawings wherein like reference numerals refer to like parts throughout the several views, and wherein: 
           [0006]      FIG. 1  schematically depicts some aspects of a non-limiting example of a gas turbine engine in accordance with an embodiment of the present invention. 
           [0007]      FIG. 2  schematically illustrates some aspects of a non-limiting example of a pulse detonation combustion system in accordance with an embodiment of the present invention. 
           [0008]      FIGS. 3A-3G  schematically illustrates some aspects of a non-limiting example of the operation of the pulse detonation combustion system of  FIG. 2 . 
           [0009]    FIG.  3 D 1  schematically illustrates some aspects of a non-limiting example of a pulse detonation combustion system in accordance with an embodiment of the present invention. 
           [0010]      FIG. 4  is a non-limiting example of a plot of pressure vs. time, measured at a detonation chamber, depicting a deflagration-to-detonation transition process. 
       
    
    
     DETAILED DESCRIPTION 
       [0011]    For purposes of promoting an understanding of the principles of the invention, reference will now be made to the embodiments illustrated in the drawings, and specific language will be used to describe the same. It will nonetheless be understood that no limitation of the scope of the invention is intended by the illustration and description of certain embodiments of the invention. In addition, any alterations and/or modifications of the illustrated and/or described embodiment(s) are contemplated as being within the scope of the present invention. Further, any other applications of the principles of the invention, as illustrated and/or described herein, as would normally occur to one skilled in the art to which the invention pertains, are contemplated as being within the scope of the present invention. 
         [0012]    Referring now to the drawings, and in particular  FIG. 1 , some aspects of a non-limiting example of a gas turbine engine  10  in accordance with an embodiment of the present invention are depicted. In one form, gas turbine engine  10  is an air vehicle propulsion power plant. In other embodiments, gas turbine engine  10  may be an aircraft auxiliary power unit, a land-based engine or a marine engine. In one form, gas turbine engine  10  is a single-spool turbojet engine. In other embodiments, gas turbine engine  10  may be a single or multi-spool turbofan, turboshaft, turbojet, turboprop gas turbine or combined cycle engine. 
         [0013]    Gas turbine engine  10  includes a compressor system  12 , a combustion system  14  and a turbine system  16 . Combustion system  14  is fluidly disposed between compressor system  12  and turbine system  16 . During the operation of gas turbine engine  10 , air is drawn into the inlet of compressor system  12 , pressurized, and discharged into combustion system  14 . Fuel is mixed with the pressurized air in combustion system  14 , which is then combusted. The combustion products are directed into turbine system  16 , which extracts energy in the form of mechanical shaft power to drive compressor system  12 . The hot gases exiting turbine system  16  are directed into a nozzle (not shown), and provide a thrust output of gas turbine engine  10 . 
         [0014]    Referring now to  FIG. 2 , a non-limiting example of some aspects of combustion system  14  are depicted. Combustion system  14  is a pulse detonation combustion system. In one form, pulse detonation combustion system  14  is a pressure gain combustion system that generates a higher post-combustion pressure relative to the pre-combustion pressure. In other embodiments, pulse detonation combustion system  14  may not be a pressure gain combustion system. In one form, the post-combustion thermodynamic properties of a pressure gain pulse detonation combustion system  14  are similar to constant volume combustion processes, which provide superior thermodynamic performance relative to constant pressure combustion processes typically employed, e.g., in gas turbine and rocket engines. Pulse detonation combustion system  14  combusts fuel and oxidizer (e.g., air) in a cyclic fashion to generate pulsed detonations that provide work, such as thrust and high pressure output. Fresh propellant (fuel and oxidizer) is provided in every cycle. The total amount of work generated by pulse detonation combustion system  14  is proportional to the cycle frequency, e.g., the number of pulse detonation waves per second. Hence, it is desirable that pulse detonation combustion system  14  generates each single pulse of detonation wave quickly. 
         [0015]    In one form, combustion system  14  is a valveless pulse detonation combustion system. Although valves with high speed valve actuators to control fuel and oxidizer feed systems may be employed in pulse detonation combustion systems, these valve and control systems unnecessarily increase the cost and decrease the reliability of such systems relative to a valveless pulse detonation combustion system. A valveless pulse detonation combustion system is a pulse detonation combustion system that does not employ mechanical valves to control fuel and oxidizer feed timing. Rather a valveless pulse detonation combustion system employs a gasdynamic valving scheme, absent mechanical valves, that controls feed timing of the reactants (fuel and oxidizer). Although employed in the present example as a gas turbine engine combustor, in other embodiments combustion system  14  may be employed as a thrust producing engine in its own right. In still other embodiments, combustion system  14  may be employed for other purposes, e.g., in rocket systems, power generators, cleaning systems and/or manufacturing systems, e.g., as a detonation gun for a material deposition system. 
         [0016]    In one form, pulse detonation combustion system  14  includes a fuel supply line  18 , an oxidizer supply line  20 , an inlet section  22 , a vortex generator  24 , a detonation chamber  26 , a converging/diverging (CD) nozzle  28 , and a discharge opening  29 . In other embodiments, pulse detonation combustion system  14  may include a discharge opening without a nozzle or with another type of nozzle. In one form, pulse detonation combustion system  14  is an elongate tubular structure. In other embodiments, other suitable shapes may be employed. Pulse detonation combustion system  14  is operative to initiate deflagration combustion and to initiate a deflagration-to-detonation transition (DDT) for each detonation cycle. In one form, pulse detonation combustion system  14  is configured to control fuel and oxidizer supply timing without the use of a mechanical valve, i.e., without the use of one or more mechanical valves to control fuel and oxidizer supply timing. In other embodiments, pulse detonation combustion system  14  may be configured to control fuel and/or oxidizer flow/timing with the use of one or more mechanical valves. In one form, fuel and oxidizer and/or air are supplied separately through fuel and oxidizer supply lines  18  and  20  in order to eliminate or reduce potential flame propagation in the upstream direction into fuel and oxidizer supply lines  18  and  20 . 
         [0017]    In one form, inlet section  22 , vortex generator  24 , detonation chamber  26  and CD nozzle  28  are arranged fluidly in series such that a steady state flow though combustion system  14  results in the same flow direction, e.g., left-to-right in  FIG. 2 , through each of inlet section  22 , vortex generator  24 , detonation chamber  26  and CD nozzle  28 . In other embodiments, inlet section  22 , vortex generator  24 , detonation chamber  26  and CD nozzle  28  may be arranged or oriented differently. 
         [0018]    Fuel supply line  18  is in fluid communication with inlet section  22 . Fuel supply line  18  is a multiple return bend supply line having an inlet  30  oriented in parallel to detonation chamber  26  and CD nozzle  28 . In one form, fuel supply line  18  is operative to conserve the momentum of backflow for thrust. In one form, inlet  30  is oriented in the same direction as CD nozzle  28  and discharge  29  to aid in conserving backflow momentum. In other embodiments, inlet  30  may be oriented differently. Fuel is supplied to inlet section  22  via an injection port  32 . From inlet section  22 , fuel is communicated to vortex generator  24  and detonation chamber  26 . In one form, fuel supply line  18  includes two (2) return bends, where fluid attempting to return into fuel supply line  18  must change directions twice, e.g., as indicated by bend arrows  34  and  36 . In one form, the configuration of fuel supply line  18 , including the number of return bends, is determined based on the supply pressures and mass flow rates of the oxidizer. In one form, the fuel supplied by fuel supply line  18  is supplied in gaseous form. In other embodiments, the fuel may be supplied in liquid form. In one form, the fuel is a hydrocarbon fuel. In other embodiments, other fuels may be employed, e.g., hydrogen. 
         [0019]    Oxidizer supply line  20  is in fluid communication with inlet section  22 . Oxidizer supply line  20  is a multiple return bend supply line having an inlet  38  oriented in parallel to detonation chamber  26  and CD nozzle  28 . In one form, oxidizer supply line  20  is operative to conserve the momentum of backflow for thrust. In one form, inlet  38  is oriented in the same direction as CD nozzle  28  and discharge  29  to aid in conserving backflow momentum. In other embodiments, inlet  38  may be oriented differently. Oxidizer is supplied to inlet section  22  via an injection port  40 , and from inlet section  22  is communicated to vortex generator  24  and detonation chamber  26 . In one form, oxidizer supply line  20  includes two (2) return bends, where fluid attempting to return into fuel supply line  18  must change direction twice, e.g., as indicated by bend arrows  42  and  44 . In one form, the configuration of oxidizer supply line  20 , including the number of return bends, is determined based on the supply pressures and mass flow rates of the oxidizer. In one form, the oxidizer supplied by oxidizer supply line  20  is supplied in gaseous form. In other embodiments, the oxidizer may be supplied in liquid form. In one form, the oxidizer is air. In other embodiments, other oxidizers may be employed, e.g., pure or diluted oxygen. In one form, oxidizer is supplied via oxidizer supply line  20  as a purge gas to help purge combustion system  14  of combustion products after each detonation cycle. In other embodiments, a separate purge gas supply line may be employed. In still other embodiments, no purge gas may be employed. 
         [0020]    Inlet section  22  includes a thrust wall  46 . In one form, inlet section  22  also includes a flame accelerator  48 . In other embodiments, inlet section  22  may not employ a flame accelerator. In one form, inlet section  22  functions as a buffer zone which impedes the backflow of gases upstream into return bend supply lines  18  and  20 . The volume of inlet section  22  determines the impedance of inlet section  22  during the backflow process. The volume of inlet section  22  is tuned (designed, configured) to achieve a desired detonation cycle time based on the required cycle time of combustion system  14  for the particular application. 
         [0021]    Thrust wall  46  is operative to reflect shock waves toward discharge opening  29 . In one form, thrust wall  46  forms an end structure of inlet section  22  opposite discharge opening  29 . Fuel supply line  18  and oxidizer supply line  20  are positioned to discharge fuel and oxidizer into pulse detonation combustion system  14  between thrust wall  46  and vortex generator  24 . In one form, injection port  32  of fuel supply line  18  and injection port  40  of oxidizer supply line  20  are positioned between thrust wall  46  and flame accelerator  48 . In one form, injection ports  32  and  40  are positioned to impinge the fuel against the oxidizer to enhance mixing of the fuel and oxidizer. In other embodiments, outlet  32  and outlet  40  may be positioned in other locations, e.g., other locations in inlet section  22 . 
         [0022]    Flame accelerator  48  is configured to generate internal drag forces in the flow of the fuel/oxidizer mixture and combustion products when combustion and precursor shock waves interact with flame accelerator  48  during the deflagration-to-detonation-transition. In one form, flame accelerator  48  is directionally dependent, i.e., is geometrically structured to yield a directionally-dependent drag coefficient, wherein the drag induced in flow through flame accelerator  48 , e.g., of the oxidizer, the fuel/oxidizer mixture and/or the combustion products, depends on the direction of the flow. In other embodiments, flame accelerator  48  may not be directionally-dependent. In one form, flame accelerator  48  is structured to have a greater drag coefficient for precursor and combustion shockwave propagations in direction  50  than in direction  52 . Flame accelerator  48  is configured to enhance the turbulent mixing of fuel and oxidizer in inlet section  22 , and is configured to amplify precursor shock wave strength. In one form, the geometric shape of flame accelerator  48  is selected based on the bulk flow and wave propagation directions during the different phases of operation of combustion system  14 . In other embodiments, the geometric shape of flame accelerator may also or alternatively be determined based on other parameters. 
         [0023]    Vortex generator  24  includes an inlet face  54  and an outlet face  56 . Vortex generator  24  is in fluid communication with inlet section  22 . In one form, a plurality of igniters  58  are positioned between inlet face  54  and outlet face  56 . In other embodiments, only a single igniter  58  may be employed. Each igniter  58  is operative to initiate deflagration combustion of fuel and oxidizer in vortex generator  24  received from fuel supply line  18  and oxidizer supply line  20 . In one form, vortex generator  24  is structured to reduce the bulk flow speed of the fuel and oxidizer so that a deflagration flame may be more readily initiated using igniters  58 . In one form, vortex generator is structured to generate a vortex recirculation zone to mix fuel and oxidizer to reduce DDT time and distance. In one form, inlet face  54  and outlet face  56  include a plurality of discrete openings, which provide flow areas for fuel, oxidizer and combustion products. In one form, the openings are discrete openings, e.g., shaped holes. In other embodiments, the flow areas may be in the form of porosity in inlet face  54  and outlet face  56 , e.g., wherein inlet face  54  and outlet face  56  are formed of a porous material, such as a metal foam and/or a ceramic foam. In other embodiments, the flow areas may be provided by one or more discrete openings in one or both of inlet face  54  and outlet face  56  in addition to foam. In one form, outlet face  56  has a greater flow area than inlet face  54 . 
         [0024]    Multiple return bend fuel supply line  18  and oxidizer supply line  20 , inlet section  22  and vortex generator  24  form a fluid diode  60 . The function of the fluid diode is similar to the function of a semiconductor diode that acts as a one-way gate to electric current flow. The fluid diode offers a low resistance to inflow and a large resistance to backflow in order to emulate a nonreturn valve or check valve. Characteristics of fluid diode  60 , such as operating supply pressures, effective valve opening/closing times and the flow resistance during the backflow process, are determined by the volume, geometry and flow characteristics of multiple bend fuel supply line  18  and oxidizer supply line  20 , inlet section  22  and vortex generator  24 . 
         [0025]    Detonation chamber  26  includes a flame accelerator  62 . Detonation chamber  26  is in fluid communication with vortex generator  24 , and with inlet section  22  via vortex generator  24 . Flame accelerator  62  is configured to generate internal drag forces in the flow of the fuel/oxidizer mixture and combustion products when combustion and precursor shock waves interact with flame accelerator  62  during the DDT process. In one form, flame accelerator  62  is directionally dependent, i.e., is geometrically structured to yield a directionally-dependent drag coefficient, wherein the drag induced in flow through flame accelerator  62 , e.g., of the oxidizer, the fuel/oxidizer mixture and/or the combustion products, depends on the direction of the flow. 
         [0026]    In one form, flame accelerator  62  is structured to have a greater drag coefficient for precursor and combustion shockwave propagations in direction  52  than in direction  50 . Thus, in one form, the directionally-dependent drag coefficients of flame accelerators  48  and  62  yield increased drag in directions extending away from vortex generator  24 . Flame accelerator  62  is configured to enhance the turbulent mixing of fuel and oxidizer in detonation chamber  26 , and is configured to amplify precursor shock wave strength. In one form, the geometric shape of flame accelerator  62  is selected based on the bulk flow and wave propagation directions during the different phases of operation of combustion system  14 . 
         [0027]    In one form, flame accelerators  48  and  62  are configured such that the drag produced by flame accelerator  62  in detonation chamber  26  is counteracted by the drag produced by the flame accelerator  48  in inlet section  22  during the DDT process. In one form, inlet section  22 , vortex generator  24  and detonation chamber  26  are operative to initiate DDT. In some embodiments, inlet section  22 , vortex generator  24  and detonation chamber  26  are structured to initiate DDT using both flame accelerators  48  and  62 . In other embodiments, i.e., embodiments that do not include flame accelerator  48 , inlet section  22 , vortex generator  24  and detonation chamber  26  are structured to initiate DDT with the aid of directional flame accelerator  62 . In still other embodiments, i.e., embodiments that do not include flame accelerator  62 , inlet section  22 , vortex generator  24  and detonation chamber  26  are structured to initiate DDT with the aid of flame accelerator  48 . 
         [0028]    CD nozzle  28  is in fluid communication with detonation chamber  26 . In one form, CD nozzle  28  is positioned at the end of the detonation chamber  26 . In one form, CD nozzle  28  is operative to pressurize a fresh charge of propellants in the detonation chamber  26  in every cycle of operation, e.g., for each pulse detonation cycle. In one form, CD nozzle  28  is operative to convert propellant chemical energy to kinetic energy. 
         [0029]    Discharge opening  29  is in fluid communication with detonation chamber  26 . Discharge opening  29  is operative to discharge the combustion products from detonation chamber  26 . In one form, discharge opening  29  is the discharge of CD nozzle  28 . In embodiments that do not employ CD nozzle  28  or another nozzle, discharge opening  29  may be the outlet of detonation chamber  26 . 
         [0030]    Referring now to  FIGS. 3A-3G , a non-limiting example of some aspects of the operation of combustion system  14  is described. During the operation of combustion system  14 , fuel, such as a gaseous hydrocarbon fuel, is supplied via the multiple return bend fuel supply line  18 ; and oxidizer, such as oxygen or air, is supplied via multiple return bend supply line  20 . In one form, gaseous fuel and oxidizer/air are employed to promote DDT process in terms of reduction of the DDT distance and/or the DDT time. In other embodiments, a liquid hydrocarbon fuel may be employed, e.g., directly or by preconditioning the fuel, such as by heating it to a flash point in order to reduce induction time. The fuel and the oxidizer are injected into inlet section  22  via injection ports  32  and  40 , and flow into vortex generator  24  and detonation chamber  26 . In one form, fuel injection  32  and oxidizer and air injection port  40  are oriented approximately perpendicular with the axis  64  of combustion system  14 . In this way, the advantage of the return bend supply lines  18  and  20  is obtained by minimizing pressure loss during the backflow phase, e.g., as set forth herein. 
         [0031]    Referring now to  FIG. 3A , fuel  66  and oxidizer  68  are injected into inlet section  22 . Fuel  66  and oxidizer  68  flow past flame accelerator  48 , into vortex generator  24  and detonation chamber  26  under the action of the fuel and oxidizer supply pressures via the multiple return bend supply lines  18  and  20 , respectively. Mixing takes place in inlet section  22 , e.g., due to the action of the flow streams impinging upon each other, yielding a combined propellant stream  70 . Further mixing of the combined propellant flow  70  takes place in flame accelerator  48  (for those embodiments so equipped), resulting from the drag induced into the flowstream by flame accelerator  48 . The propellant flow expands as it flows toward discharge opening  29 . CD nozzle  28  is dimensioned to achieve a desired mass flow rate that is exhausted in each cycle to maintain a desired fill pressure inside combustion system  14 . The total amount of mass charged at each cycle is proportional to thrust production per cycle, and hence it is desirable to keep a high fill pressure inside combustion system  14 . In one form, the fill described here is the pressure inside of detonation chamber  26 . The higher this pressure, the more mass (fresh fuel and oxidizer) that is stored in the chamber leading to more work extraction. Acceptable fill pressure is dependent upon various factors, including, for example, the thermal efficiency of the cycle being produced, e.g., with a pulse detonation combustor-based gas turbine system. In some embodiments, it is desirable to maintain higher fill pressure to load mass into the chamber as much as possible/practical. 
         [0032]    Referring now to  FIG. 3B-3D , igniters  58  provide electrical discharge sparks, which ignite deflagration combustion waves  72  and  74 . Combustion waves  72  and  74  propagate toward CD nozzle  28  and thrust wall  46 , respectively. The combustion waves  72  and  74  interact with flame accelerators  62  and  48 , respectively, which enhances the turbulent mixing process of fuel and oxidizer. The turbulent mixing process enhances the DDT process by increasing the volumetric heat release rate of the mixture as the reaction accelerates the bulk flow of the mixture. The reaction waves or flames  72 ,  74  generate precursor shock waves  76 ,  78 , which are strengthened through interactions with the flame accelerators  48 ,  62 , respectively. The precursor shock waves  76 ,  78  elevate post-shock temperature due to its non-isentropic nature, resulting in a positive feedback mechanism between the reaction waves  72 ,  74  and the precursor shock waves  76 ,  78  to promote further acceleration of the bulk flow, which results in the generation of detonation (shock) waves  80 ,  82 . The flame acceleration process in a deflagration-to-detonation-transition (DDT) creates internal drag forces during the interactions with flame accelerators, which may result in some thrust production loss in some propulsion applications of a pulse detonation combustion system. However, the configuration of combustion system  14  compensates for some of the internal drag loss by cancellation of two counter-propagating waves  72  and  74 , and  76  and  78 , respectively. 
         [0033]    It will be noted that in various embodiments, the flame accelerator ( 48 ,  62  or both) with the same features may also be incorporated into a valved pulse detonation combustor (i.e., a pulse detonation combustor with mechanically actuated valves to control feed timing) or a CVC (Constant Volume Combustor) incorporated into a dynamic pressure exchanger/wave rotor in order to recover drag losses with the flow-direction-dependent drag coefficient feature. Such an example is described in FIG.  3 D 1  (below) where two valves (X 1  and X 2 ) are equipped at both inflow (X 1 ) and outflow (X 2 ) ends of the combustor with the flame accelerator (X 3 ) and the ignition system (X 4 ). As presented in FIG.  3 D 1 , the configuration of the ignition system may change associated with the design of the flame accelerator for the unsteady pressure gain combustors (e.g., pulse detonation combustors). 
         [0034]    Referring now to  FIGS. 3E-3G , detonation wave  82  reflects off thrust wall  46 , yielding a reflected shock wave  84 . A post-detonation or post-shock pressure is increased such that the inlets (injection ports  32 ,  40 ) of the gasdynamic valves are pressurized to be closed, as shown in  FIG. 3E . Mass fluxes of fuel and oxidizer are decelerated or impeded because of the reduced pressure gradient across the return bend supply lines  18  and  20 . Fuel and oxidizer are supplied from the separated supply lines  18  and  20 , and thus flames do not propagate upstream into multiple return bend supply lines  18  and  20  due to their flammability limits (e.g., the combustibility of the mixture is exhausted). In the mean time, detonation wave  80  propagates toward the open end CD nozzle  28 . Reactions behind the detonation wave  80  are completed, and the reflected shock wave  84  continues to attenuate, contributing to a reduction of internal drag production caused by interactions between reflected shock wave  84  and flame accelerator  62 . In one form, flame accelerator  48  is configured to reduce the drag coefficient produced by interactions with the reflected shock wave  84  by virtue of its directionally dependent drag coefficient. After the detonation wave  80  propagates out of discharge opening  29 , the pressure inside of combustion system  14  decreases due to expansion waves initiating a blowdown process  86  resulting from a pressure gradient across combustion system  14 . Once the inside of the combustion system  14  is depressurized enough to create a negative pressure gradient across multiple return bend oxidizer supply line  20 , oxidizer injection port  40 , functioning as a component of a gas dynamic valve, opens to supply oxidizer into inlet section  22 , creating a buffer zone between hot combustion products on the right side of an oxidizer/combustion product interface  88  and a fresh propellant charge (fuel and oxidizer), in order to prevent auto-ignition of the fuel and oxidizer intended for the next detonation cycle. In one form, the supply pressure for the oxidizer is higher than that for the fuel, and hence the oxidizer gasdynamic valve opens before the fuel gasdynamic valve opens. It will be understood that the term, “gasdynamic valve” describes that the injection ports  32  and  40  are effectively opened or closed based on gas dynamics, and does not refer to the use of a mechanical valve. After the pressure at the inlet of the multiple return bend fuel supply line  18  is lower than the fuel supply pressure, the fuel gasdynamic valve (injection port  32 ) opens, which results in fuel mixing with oxidizer, flowing into the inlet section  22 , the vortex generator  24  and the detonation chamber  26 .  FIG. 3G  depicts contact surfaces  90  and  92 , which represent boundaries or interfaces between different types of mass flow inside combustion system  14 . A contact surface is a boundary between two different gases in terms of temperature and density, but the same in pressure and velocity. The incoming propellant (fuel and oxidizer) is located behind (to the left of) the contact surface  90 , oxidizer is located between contact surfaces  90  and  92 , and the hot combustion products of the previous detonation cycle are located to the right side of the contact surface  92 . 
         [0035]    Referring to  FIG. 4 , a non-limiting example of a plot  94  of pressure vs. time, measured in detonation chamber  26  after the DDT process, is depicted. The pressure history is normalized by the theoretical Chapman-Jouguet detonation pressure P CJ  at an initial condition. In the example of  FIG. 4 , the initial condition was close to ambient condition (approximately 14.7 psia and approximately 540° R). In other embodiments, other initial conditions may be employed, e.g., depending upon the needs of the particular application. The time scale is normalized by τ=tc/L where t, c and L represent time, the speed of sound at the pressure conditions in combustion system  14  and the length of the detonation chamber  26 , respectively. A typical pressure spike from the detonation wave is observed at τ=2.2, followed by approximately unity pressure and a decrease of pressure during the blowdown process. The combustion products are exhausted out of detonation chamber  26 , vortex generator  24  and inlet section  22  during the blowdown process, due to the pressure gradient across the combustion system  14 . Once the pressure inside of the combustion system  14  sufficiently decreases during the blowdown process, mass fluxes of fuel and oxidizer start to flow to inlet section  22  due to pressure gradient between the supply pressures of fuel and oxidizer, and the pressure inside inlet section  22 . Characteristics of the gasdynamic valves, such as mass flow rate, and the valve effective opening/closing time are dependent on configuration of the combustion system  14 ; e.g., including length and volume of combustion system  14 , and the number of the return bends in supply lines  18  and  20 . 
         [0036]    Embodiments of the present invention include a pulse detonation combustion system, comprising: an inlet section having a first flame accelerator; a fuel supply line in fluid communication with the inlet section; an oxidizer supply line in fluid communication with the inlet section; a vortex generator having an inlet face and an outlet face, wherein the vortex generator is in fluid communication with the inlet section; an igniter coupled to the vortex generator and positioned between the inlet face and the outlet face of the vortex generator, wherein the igniter is operative to initiate deflagration combustion of fuel and oxidizer received from the fuel supply line and the oxidizer supply line; a detonation chamber having a second flame accelerator, wherein the detonation chamber is in fluid communication with the inlet section via the vortex generator; and a discharge opening in fluid communication with the detonation chamber and operative to discharge combustion products, wherein the inlet section, the first flame accelerator, the vortex generator and the second flame accelerator are operative to initiate a deflagration to detonation transition. 
         [0037]    In a refinement, the inlet face of the vortex generator has a first flow area; wherein the outlet face of the vortex generator has a second flow area; and wherein the second flow area is greater than the first flow area. 
         [0038]    In another refinement, the vortex generator is structured to reduce the bulk flow speed of the fuel and oxidizer. 
         [0039]    In yet another refinement, at least one of the first flame accelerator and the second flame accelerator are structured to have a directionally-dependent drag coefficient. 
         [0040]    In still another refinement, both the first flame accelerator and the second flame accelerator are structured to have directionally-dependent drag coefficients. 
         [0041]    In yet still another refinement, the directionally-dependent drag coefficients have increased drag in directions extending away from the vortex generator. 
         [0042]    In a further refinement, the pulse detonation combustion further comprises a converging-diverging nozzle in fluid communication with the detonation chamber. 
         [0043]    In a yet further refinement, the pulse detonation combustion system is configured to control fuel and oxidizer supply timing without the use of a mechanical valve. 
         [0044]    Embodiments of the present invention include a pulse detonation combustion system, comprising: a fuel supply line; an oxidizer supply line separate from the fuel supply line; an inlet section in communication with the fuel supply line and the oxidizer supply line; a vortex generator having an inlet face and an outlet face, wherein the vortex generator is in fluid communication with the inlet section; an igniter coupled to the vortex generator and positioned between the inlet face and the outlet face of the vortex generator, wherein the igniter is operative to initiate deflagration combustion of fuel and oxidizer received from the fuel supply line and the oxidizer supply line; a detonation chamber in fluid communication with inlet section via the vortex generator; a flame accelerator having a directionally dependent drag coefficient; and a discharge opening in fluid communication with the detonation chamber and operative to discharge combustion products, wherein the inlet section, the vortex generator and the flame accelerator are operative to initiate a deflagration to detonation transition. 
         [0045]    In a refinement, the fuel supply line and the oxidizer supply line are multiple return bend supply lines. 
         [0046]    In another refinement, the fuel supply line and the oxidizer supply line have inlets oriented in parallel to the detonation chamber. 
         [0047]    In yet another refinement, the pulse detonation combustion system further comprises a thrust wall that is operative to reflect a shock wave toward the discharge opening. 
         [0048]    In still another refinement, the fuel supply line and the oxidizer supply line are positioned to discharge the fuel and the oxidizer into the pulse detonation combustion system between the thrust wall and the vortex generator. 
         [0049]    In yet still another refinement, the thrust wall forms an end structure of the inlet section. 
         [0050]    In a further refinement, the pulse detonation combustion system is configured to control fuel and oxidizer supply timing without the use of a mechanical valve. 
         [0051]    Embodiments of the present invention include a gas turbine engine, comprising: a compressor; a turbine; and a pulse detonation combustion system fluidly disposed between the compressor and the turbine, including: a fuel supply line; an oxidizer supply line separate from the fuel supply line; an inlet section in communication with the fuel supply line and the oxidizer supply line; a vortex generator having an inlet face and an outlet face, wherein the vortex generator is in fluid communication with the inlet section; an igniter coupled to the vortex generator and positioned between the inlet face and the outlet face of the vortex generator, wherein the igniter is operative to initiate deflagration combustion of fuel and oxidizer received from the fuel supply line and the oxidizer supply line; a detonation chamber in fluid communication with inlet section via the vortex generator; a flame accelerator having a directionally dependent drag coefficient; and a discharge opening in fluid communication with the detonation chamber and operative to discharge combustion products, wherein the inlet section, the vortex generator and the flame accelerator are operative to initiate a deflagration to detonation transition. 
         [0052]    In a refinement, a volume of the inlet section is tuned to achieve a desired detonation cycle time. 
         [0053]    In another refinement, the gas turbine engine further comprises a converging-diverging nozzle in fluid communication with the detonation chamber. 
         [0054]    In yet another refinement, the flame accelerator is structured to amplify a precursor shock wave strength. 
         [0055]    In still another refinement, the flame accelerator is structured to perform turbulent mixing of the fuel and the oxidizer. 
         [0056]    In yet still another refinement, the pulse detonation combustion system is configured to control fuel and oxidizer supply timing without the use of a mechanical valve. 
         [0057]    In a further refinement, the vortex generator includes a recirculation zone for mixing fuel and oxidizer. 
         [0058]    In a yet further refinement, the flame accelerator is positioned in the detonation chamber. 
         [0059]    In a still further refinement, the flame accelerator is positioned in the inlet section. 
         [0060]    In a yet still further refinement, the gas turbine engine further comprises an other flame accelerator positioned in the detonation chamber. 
         [0061]    While the invention has been described in connection with what is presently considered to be the most practical and preferred embodiment, it is to be understood that the invention is not to be limited to the disclosed embodiment(s), but on the contrary, is intended to cover various modifications and equivalent arrangements included within the spirit and scope of the appended claims, which scope is to be accorded the broadest interpretation so as to encompass all such modifications and equivalent structures as permitted under the law. Furthermore it should be understood that while the use of the word preferable, preferably, or preferred in the description above indicates that feature so described may be more desirable, it nonetheless may not be necessary and any embodiment lacking the same may be contemplated as within the scope of the invention, that scope being defined by the claims that follow. In reading the claims it is intended that when words such as “a,” “an,” “at least one” and “at least a portion” are used, there is no intention to limit the claim to only one item unless specifically stated to the contrary in the claim. Further, when the language “at least a portion” and/or “a portion” is used the item may include a portion and/or the entire item unless specifically stated to the contrary.