Abstract:
A gas turbine engine includes a transient plasma igniter in communication with a continuous detonation wave combustor. A method of operating a gas turbine engine includes maintaining ignition of a continuous detonation wave combustor with a transient plasma igniter.

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
       [0001]    This application claims priority to U.S. Provisional Patent Appln. No. 61/826,296 filed May 22, 2013, which is hereby incorporated herein by reference in its entirety. 
     
    
     BACKGROUND 
       [0002]    The present disclosure generally relates generally to a gas turbine engine architecture, and more specifically to a turbine engine with a continuous detonation wave combustor. 
         [0003]    Gas turbine engines, such as those that power modern commercial and military aircraft, generally include a compressor section to pressurize an airflow, a combustor section to burn a hydrocarbon fuel in the presence of the pressurized air, and a turbine section to extract energy from the resultant combustion gases. 
         [0004]    The compressor section is often relatively long and includes numerous stages to achieve the desired compression ratios. Alternate engine architectures may utilize centrifugal compression technology to reduce the required length, but are of a relatively significant diameter to achieve desired compression ratios. Large diameter gas turbine engine architectures increase weight and frontal area which typically relegates such engine architectures to subsonic applications. 
       SUMMARY OF THE DISCLOSURE 
       [0005]    According to an aspect of the invention, a gas turbine engine is provided that includes a tip turbine engine compressor, a continuous detonation wave combustor, a turbine and a transient plasma ignitor. The continuous detonation wave combustor is in fluid communication with and downstream of the compressor. The turbine is in fluid communication with and downstream of the continuous detonation wave combustor. The transient plasma ignitor is in communication with the continuous detonation wave combustor. 
         [0006]    According to another aspect of the invention, another gas turbine engine is provided that includes a continuous detonation wave combustor and a transient plasma ignitor. The transient plasma ignitor is in communication with said continuous detonation wave combustor. 
         [0007]    According to still another aspect of the invention, a method is provided for operating a gas turbine engine. This method includes maintaining ignition of a continuous detonation wave combustor with a transient plasma igniter. 
         [0008]    The gas turbine engine may include a high bypass fan section upstream of said tip turbine engine compressor. 
         [0009]    The gas turbine engine may include a centrifugal compressor gas turbine engine architecture. 
         [0010]    The gas turbine engine may include a fan-turbine rotor assembly with a multiple of hollow fan blades to provide internal, centrifugal compression of a compressed airflow to the continuous detonation wave combustor. 
         [0011]    The gas turbine engine may include an axial compressor axially forward of the fan-turbine rotor assembly. 
         [0012]    The continuous detonation wave combustor may be radially outboard of the multiple of hollow fan blades. 
         [0013]    Each of the hollow fan blades may include a fan blade core airflow passage generally perpendicular to an axis of rotation of the fan-turbine rotor assembly. 
         [0014]    The gas turbine engine may include a high bypass fan section. 
         [0015]    The method may include internally compressing an airflow within a fan-turbine rotor assembly. The method may also include communicating the airflow from the fan-turbine rotor assembly to the continuous detonation wave combustor. 
         [0016]    The method may include axially compressing the airflow upstream of the fan-turbine rotor assembly. 
         [0017]    The method may include generating a compression ratio of about forty to one (40:1) within the continuous detonation wave combustor. 
         [0018]    The method may include centrifugally compressing an airflow within a high bypass gas turbine engine architecture. 
         [0019]    The method may include centrifugally compressing an airflow within a low bypass gas turbine engine architecture. 
         [0020]    The foregoing features and the operation of the invention will become more apparent in light of the following description and the accompanying drawings. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0021]    Various features will become apparent to those skilled in the art from the following detailed description of the disclosed non-limiting embodiments. The drawings that accompany the detailed description can be briefly described as follows: 
           [0022]      FIG. 1  is a partial sectional perspective view of a tip turbine engine; 
           [0023]      FIG. 2  is a longitudinal sectional view of a tip turbine engine along an engine centerline; 
           [0024]      FIG. 3  is a schematic view of an annular continuous detonation wave combustor; 
           [0025]      FIG. 4  is a partial schematic view of another annular continuous detonation wave combustor for a gas turbine engine architecture; and 
           [0026]      FIG. 5  is a partial schematic view of still another annular continuous detonation wave combustor for a gas turbine engine architecture. 
       
    
    
     DETAILED DESCRIPTION 
       [0027]      FIG. 1  schematically illustrates a perspective partial sectional view of a tip turbine engine type gas turbine engine  10 . Although depicted as a high bypass tip turbine engine in the disclosed non-limiting embodiment, it should be understood that the teachings herein may also be applied to other types of turbine engine architectures. 
         [0028]    The engine  10  generally includes an outer nacelle  12 , a rotationally fixed static outer support structure  14  and a rotationally fixed static inner support structure  16 . A multiple of fan inlet guide vanes  18  are mounted between the static outer support structure  14  and the static inner support structure  16 . Each inlet guide vane  18  may include a fixed or variable trailing edge  18 A. 
         [0029]    A nose cone  20  is located along a centerline A of the engine  10  to smoothly direct airflow into an axial tip turbine compressor  22  adjacent thereto. The axial tip turbine compressor  22  is mounted about the engine centerline A axially aft of the nose cone  20 . 
         [0030]    A fan-turbine rotor assembly  24  is mounted for rotation about the engine centerline A axially aft of the axial tip turbine compressor  22 . The fan-turbine rotor assembly  24  includes a multiple of hollow fan blades  28  to provide internal, centrifugal compression of the compressed airflow from the axial tip turbine engine compressor  22  for distribution to a combustor section  30  located within the rotationally fixed static outer support structure  14 . 
         [0031]    A turbine  32  includes a multiple of tip turbine blades  34  (two stages shown) which rotatably drive the hollow fan blades  28  relative a multiple of tip turbine stators  36 , which extend radially inwardly from the static outer support structure  14 . The combustor section  30  is radially outboard of the multiple of hollow fan blades  28  and the axially forward of the turbine  32 . 
         [0032]    With reference to  FIG. 2 , the rotationally fixed static inner support structure  16  includes a splitter  40 , a static inner support housing  42  and a static outer support housing  44  located coaxial to the engine centerline A. 
         [0033]    The axial tip turbine compressor  22  includes the axial compressor rotor  46  from which a plurality of compressor blades  52  extend radially outwardly. The axial tip turbine compressor  22  also includes a compressor case  50  fixedly mounted to the splitter  40 . A plurality of compressor vanes  54  extend radially inwardly from the compressor case  50  between stages of the compressor blades  52 . The compressor blades  52  and compressor vanes  54  are arranged circumferentially about the axial compressor rotor  46  in stages (three stages of compressor blades  52  and compressor vanes  54  are shown in this example). The axial compressor rotor  46  is mounted for rotation upon the static inner support housing  42  through a forward bearing assembly  68  and an aft bearing assembly  62 . 
         [0034]    The fan-turbine rotor assembly  24  includes a fan hub  64  that supports a multiple of the hollow fan blades  28 . Each fan blade  28  includes an inducer section  66 , a hollow fan blade section  72  and a diffuser section  74 . The inducer section  66  receives airflow from the axial tip turbine compressor  22  generally parallel to the engine centerline A and turns the airflow from an axial airflow direction toward a radial airflow direction. The airflow is radially communicated through a core airflow passage  80  within the fan blade section  72  where the airflow is centrifugally compressed. From the core airflow passage  80 , the airflow is turned and diffused toward an axial airflow direction toward the annular combustor  30 . In one disclosed non-limiting embodiment, the airflow is diffused axially forward in the engine  10 ; however, the airflow may alternatively be communicated in alternative or additional directions. 
         [0035]    A gearbox assembly  90  aft of the fan-turbine rotor assembly  24  provides a speed increase between the fan-turbine rotor assembly  24  and the axial tip turbine compressor  22 . Alternatively, the gearbox assembly  90  could provide a speed decrease between the fan-turbine rotor assembly  24  and the axial compressor rotor  46 . The gearbox assembly  90  is mounted for rotation between the static inner support housing  42  and the static outer support housing  44 . The gearbox assembly  90  includes a sun gear shaft  92  which rotates with the axial tip turbine compressor  22  and a planet carrier  94  which rotates with the fan-turbine rotor assembly  24  to provide a speed differential therebetween. The gearbox assembly  90  may be a planetary gearbox that provides co-rotating or counter-rotating rotational engagement between the fan-turbine rotor assembly  24  and the axial compressor rotor  46 . The gearbox assembly  90  is mounted for rotation between the sun gear shaft  92  and the static outer support housing  44  through a forward bearing  96  and a rear bearing  98 . The forward bearing  96  and the rear bearing  98  are both tapered roller bearings and both hand radial loads. The forward bearing  96  handles the aft axial loads while the rear bearing  98  handles the forward axial loads. The sun gear shaft  92  is rotationally engaged with the axial compressor rotor  46  at a splined interconnection  100  or the like. 
         [0036]    In operation, air enters the axial tip turbine compressor  22 , and is compressed by the three stages of the compressor blades  52  and compressor vanes  54 . The compressed air from the axial tip turbine compressor  22  enters the inducer section  66  in a direction generally parallel to the engine centerline A and is turned by the inducer section  66  radially outwardly through the core airflow passage  80  of the hollow fan blades  28 . The airflow is further compressed centrifugally within the hollow fan blades  28  by rotation of the hollow fan blades  28 . From the core airflow passage  80 , the airflow is turned and diffused axially forward into the annular combustor  30 . The compressed core airflow from the hollow fan blades  28  is mixed with fuel in the combustor section  30  and ignited to form a high-energy gas stream. The high-energy gas stream is expanded over the multiple of tip turbine blades  34  mounted about the outer periphery of the fan-turbine rotor assembly  24  to drive the fan-turbine rotor assembly  24 , which in turn drives the axial tip turbine compressor  22  through the gearbox assembly  90 . Concurrent therewith, the fan-turbine rotor assembly  24  discharges fan bypass air axially aft to merge with the core airflow from the turbine  32  in an exhaust case  106 . A multiple of exit guide vanes  108  are located between the static outer support housing  44  and the rotationally fixed static outer support structure  14  to guide the combined airflow out of the engine  10  to provide forward thrust. An exhaust mixer  110  mixes the airflow from the turbine blades  34  with the bypass airflow through the fan blades  28 . 
         [0037]    With reference to  FIG. 3 , the combustor section  30  includes an annular continuous detonation wave combustor  120 . The continuous detonation wave combustor  120  derives energy from a continuous wave of detonation. In other words, for a detonation engine as compared to a conventional combustor which operates on the deflagration of fuel, the oxygen and fuel combustion process of the continuous detonation wave combustor  120  is effectively an explosion instead of burning. 
         [0038]    A primary difference between deflagration and detonation is linked to the mechanism of the flame propagation. In deflagration, the flame propagation is a function of the heat transfer from the reactive zone to the fresh mixture (generally conduction). The detonation is a shock induced flame, which results in the coupling of a reaction zone and a shock wave. The shock wave compresses and heats the fresh mixture, for an increase above the self-ignition point. On the other side, the energy released by the flame contributes to the propagation of the shock wave. 
         [0039]    By way of further explanation, continuous detonation is a detonation wave propagating around a closed circuit in a continuous manner which globally operates at very high frequency (e.g., typically several kHz) and are dephased so the mean pressure inside the chamber is higher than for typical combustion system. 
         [0040]    The continuous detonation wave combustor  120  generally includes a fuel plenum  122 , an air diffuser  124 , an outer cylindrical wall  126 , and an inner cylindrical wall  128 . The space between air diffuser  124  and the outer cylindrical wall  126  operates as a mixing chamber  130 , and the space between the inner cylindrical wall  128  and outer cylindrical wall  126  servers as a combustion chamber  132 . An annular chamber  134  in the fuel plenum  122  serves as a fuel chamber. In one embodiment, the outer cylindrical wall  126  includes a cooling system  136  (illustrated schematically) to facilitate thermal management. 
         [0041]    A transient plasma igniter  136  (illustrated schematically in  FIG. 3 ) communicates with the combustion chamber  132 . Transient plasma igniters—and may be referred to as pulsed corona discharges—generate multiple streamers of electrons at high energy which readily facilitates stability of the detonation process along the combustion chamber  132 . That is, the transient plasma igniter  136  assists in sustainment of continuous detonation operations in air rather than an oxygen enriched oxidizer supply. Although the transient plasma igniter  136  is schematically illustrated in a single particular location, it should be appreciated that multiple locations as well as other locations for the transient plasma igniter  136  may also be provided. 
         [0042]    In one disclosed non-limiting embodiment, the igniter  136  operates continuously—not just for ignition—to further facilitate stability of the detonation process which continues substantially without interruption, as one or more waves of detonation continuously propagate around the combustion chamber  132 , consuming the air/fuel mixture, while fresh mixture is continually introduced into the combustion chamber  132 . This assists to sustain the detonation wave or waves to continually cycle around the combustion chamber  132 . 
         [0043]    The continuous detonation wave combustor  120  continuously combusts the mixed gas with the one or more detonation waves that propagate normally to the reaction front to generate a rotational flow that facilitates rotation of the turbine  32 . That is, the significant tangential component to the exhaust vector of the continuous detonation wave combustor  120  beneficially increases the motive force to drive the turbine  32 . 
         [0044]    The continuous detonation wave combustor  120  also advantageously provides significant compression ratios, which in one disclosed non-limiting embodiment are on the order of up to forty to one (40:1) to raise a two (2) to three (3) atmospheric pressure from the axial tip turbine compressor  22  to as much as about one hundred twenty (120) atmospheres. This compares to a thirteen to eighteen (13:1-18-1) compression ratio typical of a conventional gas turbine engine combustor sections. 
         [0045]    With the continuous detonation wave combustor  120 , the tip turbine engine architecture is readily scalable for greater speeds and thrust ranges as high operational compression ratios are provided within relatively small engine diameters. Military and supersonic tip turbine engine architectures are thereby facilitated. In other words, the transient plasma igniter  136  stabilizes combustion in the continuous detonation wave combustor  120  and the continuous detonation wave combustor  120  increases compression and makes use of the tangential exhaust within the tip turbine engine architecture to provides a short, small, lightweight propulsion system with good thrust specific fuel consumption. 
         [0046]    Alternate engine architectures such as a centrifugal compressor gas turbine engine architecture  300  with a fan section  302  (see  FIG. 4 ) and a low bypass centrifugal compressor gas turbine engine architecture  400  (see  FIG. 5 ) and others will also benefit herefrom. 
         [0047]    The use of the terms “a” and “an” and “the” and similar references in the context of description (especially in the context of the following claims) are to be construed to cover both the singular and the plural, unless otherwise indicated herein or specifically contradicted by context. The modifier “about” used in connection with a quantity is inclusive of the stated value and has the meaning dictated by the context (e.g., it includes the degree of error associated with measurement of the particular quantity). All ranges disclosed herein are inclusive of the endpoints, and the endpoints are independently combinable with each other. It should be appreciated that relative positional terms such as “forward,” “aft,” “upper,” “lower,” “above,” “below,” and the like are with reference to the normal operational attitude of a vehicle (e.g., aircraft) and should not be considered otherwise limiting. 
         [0048]    Although the different non-limiting embodiments have specific illustrated components, the embodiments of this invention are not limited to those particular combinations. It is possible to use some of the components or features from any of the non-limiting embodiments in combination with features or components from any of the other non-limiting embodiments. 
         [0049]    It should be appreciated that like reference numerals identify corresponding or similar elements throughout the several drawings. It should also be appreciated that although a particular component arrangement is disclosed in the illustrated embodiment, other arrangements will benefit herefrom. 
         [0050]    Although particular step sequences are shown, described and claimed, it should be understood that steps may be performed in any order, separated or combined unless otherwise indicated and will still benefit from the present disclosure. 
         [0051]    The foregoing description is exemplary rather than defined by the features within. Various non-limiting embodiments are disclosed herein, however, one of ordinary skill in the art would recognize that various modifications and variations in light of the above teachings will fall within the scope of the appended claims. It is therefore to be appreciated that within the scope of the appended claims, the disclosure may be practiced other than as specifically described. For that reason the appended claims should be studied to determine true scope and content.