Abstract:
The present embodiment sufficiently ensures the ignition stability and the flame-holding property of an afterburner while suppressing a reduction in the efficiency of an aircraft engine. A flame holder is disposed directly downstream of an injection hole of a fuel injector in a liner. The flame holder comprises: a ring-shaped annulus flame-holding member which is provided on the inner circumferential surface of the liner and is capable of propagating a flame in the circumferential direction; and a plurality of radial flame-holding members which are radially disposed inwards of the annulus flame-holding member and are capable of propagating the flame in the radial direction. A guide ring is provided inwards of the radial flame-holding members, and a ring-shaped guide channel that guides a fuel-containing mixed gas in the downstream direction is formed between the outer peripheral surface of the guide ring and the inner peripheral surface of the annulus flame-holding member.

Description:
CROSS REFERENCE TO RELATED APPLICATIONS 
       [0001]    This application is a continuation application of international application No. PCT/JP2015/062670 filed Apr. 27, 2015, which claims priority to Japanese Patent Application No. 2014-093557 filed Apr. 30, 2014, each of which is hereby incorporated by reference in their entity. 
     
    
     BACKGROUND 
       [0002]    1. Field 
         [0003]    The present embodiment relates to an afterburner or the like which supplies fuel to a gas mixture of combustion gas discharged from a core flow path of an aircraft engine and air discharged from a fan flow path of the aircraft engine for burning (afterburning) to increase thrust of the aircraft engine. 
         [0004]    2. Description of Related Art 
         [0005]    Various studies and developments are being made on afterburners in recent years according to demand for increasing thrust of aircraft engines. The configuration of a general afterburner is briefly described below. 
         [0006]    In the outlet side (the rear side) of an engine case of an aircraft engine, a rear duct (an exhaust duct) is provided. Within the rear duct, a cylindrical liner is provided, through which a gas mixture flows. The rear duct is provided with a fuel injector which injects fuel liquid in the liner. Downstream of the fuel injector in the rear duct, an igniter (an ignition plug) is provided. The igniter ignites the gas mixture containing the fuel in the liner. Downstream of the fuel injector in the liner, a flame holder which holds flame is provided. The flame holder includes: plural radial frame holding members which are provided in a radial fashion within the liner and are capable of propagating flame in the radial direction; and an annulus flame holding member which is situated away from the inner circumferential surface of the liner, concentrically with the plural radial flame holding members and is capable of propagating flame in the circumferential direction. 
         [0007]    The conventional arts relating to the present embodiment are disclosed in Patent Documents 1 and 2. 
       PATENT DOCUMENTS 
       [0008]    Patent Document 1: JP 2008-8606 A 
         [0009]    Patent Document 2: JP H9-4511 A 
       SUMMARY 
       [0010]    The diffused state (diffusion properties) of the fuel injected from the fuel injector in the liner depends on the operating state (operation conditions) of the aircraft engine. In a certain operating state of the aircraft engine, the fuel cannot be supplied to an ignition place of the igniter at proper evaporated state and proper concentration at the same time or just before the gas mixture containing the fuel is ignited by the igniter. In such a case, it is difficult to ignite the gas mixture containing the fuel with the igniter, and the ignition stability of the afterburner is lowered. 
         [0011]    Moreover, to increase the flame propagation in the flame holder and ensure sufficient flame holding capability (burning stability) of the afterburner, the flame holder needs to include not only plural radial flame holding members capable of propagating flame in the radial direction but also the annulus flame holding member capable of propagating flame in the circumferential direction as described above. When the annulus flame holding member is situated at a position away from the inner circumferential surface of the liner, the annulus flame holding member interferes with the main flow of the gas mixture in the liner, and pressure loss (thrust loss) in the liner is increased. The engine efficiency of the aircraft engine is therefore reduced. 
         [0012]    Such an aircraft engine therefore has a problem that it is difficult to ensure sufficient ignition stability and flame holding capability of the afterburner while preventing a decrease in engine efficiency of the aircraft engine. 
         [0013]    Accordingly, an object of the present embodiment is to provide an afterburner having a new configuration and the like which are capable of solving the aforementioned problem. 
         [0014]    A first aspect of the present embodiment is an afterburner which supplies fuel to a gas mixture of combustion gas discharged from a core flow path (a main flow path) of an aircraft engine and air discharged from a fan flow path (a bypass flow path) of the aircraft engine for burning (afterburning) to increase thrust of the aircraft engine, the afterburner including: a rear duct (an exhaust duct) provided at the outlet behind of an engine case of the aircraft engine; a cylindrical liner which is provided within the rear duct and allows the gas mixture to flow; a fuel injector which injects the fuel in the form of liquid in the liner; an igniter (an ignition plug) which is provided downstream of the fuel injector (downstream of the injection place of the fuel injector) and ignites (lights) the gas mixture containing the fuel in the liner; a flame holder which is situated directly downstream of the fuel injector and is configured to hold flame, the flame holder including: an annulus flame holding member which is provided on the inner circumferential surface (the inner wall surface) of the liner and propagates flame in the circumferential direction; and a plurality of radial flame holding members which are situated in a radial fashion inside the annulus flame holding member and propagate flame in the radial direction; and a guide ring (an annular guiding member) which is provided inside the radial flame holding members, in which an annular guide flow path is formed between the outer circumferential surface (the outer wall surface) of the guide ring and the inner circumferential surface of the annulus flame holding member and is configured to guide the gas mixture containing the fuel downstream. 
         [0015]    Herein, in the claims and specification of the application, the meaning of “to be provided” includes “to be directly provided” and also “to be indirectly provided with another member interposed”. The meaning of “to be situated” includes “to be directly situated” and also “to be indirectly situated with another member interposed”. The “downstream” refers to downstream in the flowing direction of main flow of combustion gas, air, or gas mixture and means that the direction towards the back of the aircraft engine. The “upstream” refers to upstream in the flowing direction of combustion gas, air or gas mixture and means the direction towards the front of the aircraft engine. The “inner diameter decreases towards the downstream end” means that the inner diameter decreases towards the downstream as a whole and includes a case where the inner diameter does not continually decrease towards the downstream end. Similarly, the “outer diameter decreases towards the downstream end” means that the outer diameter decreases towards the downstream end as a whole and includes a case where the outer diameter does not continually decrease towards the downstream end. Moreover, “the radial direction” refers to the radial direction of the aircraft engine (in other words, the radial direction of the rear duct or the liner). “The axial direction” refers to the axial direction of the aircraft engine (in other words, the axial direction of the rear duct or the liner). 
         [0016]    According to the first aspect, when the fuel is injected from the fuel injector in the liner during operation of the aircraft engine, the air containing the fuel is introduced from the inlet of the guide flow path into the guide flow path and flows through the guide flow path. The air containing the fuel then flows out of the outlet of the guide flow path to the ignition place of the igniter (directly downstream of the annulus flame holding member). The igniter ignites the gas mixture containing the fuel to form flame downstream (directly downstream) of the flame holder while burning the gas mixture containing the fuel in the liner (afterburning). A lot of heat energy is thereby put into the combustion gas within the liner, thus increasing the thrust of the aircraft engine. 
         [0017]    Herein, the air containing the fuel is introduced into the guide path flow from the inlet of the guide flow path and flows out of the outlet of the guide flow path to the ignition place of the igniter. This can prevent diffusion of the fuel in the liner and supply the fuel to the ignition place of the igniter at proper concentration and proper evaporated state. In other words, fuel can be supplied to the ignition place of the igniter at proper concentration and proper evaporated state independently of the operating state (the operation conditions) of the aircraft engine at the same time or just before the gas mixture containing the fuel is ignited by the igniter. 
         [0018]    The flame holder includes the annulus flame holding member capable of propagating flame in the circumferential direction as well as the plural radial flame holding members capable of propagating flame in the radial direction. This enhances flame propagation of the flame holder. Moreover, since the annulus flame holding member is provided on the inner circumferential surface of the liner, the annulus frame member is prevented from interfering with the main flow of the gas mixture within the liner  31 , and pressure loss (thrust loss) in the liner  31  is minimized. 
         [0019]    A second aspect of the present embodiment is an aircraft engine which generates thrust by discharging combustion gas (combustion gas and air) rearward, the aircraft including an afterburner according to the first aspect. 
         [0020]    According to the second aspect, it is possible to exert the same operation as the operation according to the first aspect. 
         [0021]    According to the present embodiment, fuel can be supplied to the ignition place of the igniter at proper concentration and proper evaporated state independently of the operating state of the aircraft engine at the same time or just before the gas mixture containing the fuel is ignited by the igniter. Accordingly, fuel can be stably ignited by the igniter, ensuring sufficient ignition stability of the afterburner. Moreover, the flame propagation of the flame holder is enhanced while the pressure loss in the liner is minimized by preventing the interference between the annulus flame holding member and the main flow of the gas mixture within the liner. This ensures sufficient flame holding capability (burning stability) of the afterburner  25  while preventing a decrease in engine efficiency of the aircraft engine. According to the embodiment of the present disclosure, it is possible to ensure sufficient ignition stability and flame holding capability of the afterburner while preventing a decrease in engine efficiency of the aircraft engine. 
     
    
     
       BRIEF DESCRIPTION OF DRAWINGS 
         [0022]      FIG. 1  is a cross-sectional side view of an afterburner according to an embodiment of the present disclosure, which is an enlarged view of part I indicated by an arrow in  FIG. 4 . 
           [0023]      FIG. 2  is a perspective view of characteristic part of the afterburner according to the embodiment of the present disclosure. 
           [0024]      FIG. 3  is an enlarged view taken along a line III-III in  FIG. 4 . 
           [0025]      FIG. 4  is a cross-sectional side view of the aircraft engine according to the embodiment of the present disclosure. 
       
    
    
     DETAILED DESCRIPTION OF EMBODIMENTS 
       [0026]    A description is given of an embodiment of the present disclosure with reference to  FIGS. 1 to 4 . In the drawings, F indicates the forward direction (the upstream direction); R, the backward direction (the downstream direction); SD, the axial direction; RD, the radial direction; and CD, the circumferential direction. 
         [0027]    As illustrated in  FIG. 4 , an aircraft engine  1  according to the embodiment of the present disclosure is a device which discharges combustion gas (high-temperature gas) G and air (low-temperature air) A rearward to generate thrust (engine thrust). The aircraft engine  1  includes a cylindrical core case (an engine inner cylinder)  3 , inside of which an annular core flow path (a main flow path)  5  is formed. Outside of the core case  3 , a cylindrical engine case (an engine outer cylinder)  7  is situated concentrically with the core case  3 . Between the inner circumferential surface of the engine case  7  and the outer circumferential surface of the core case  3 , an annular fan flow path (a bypass flow path)  9  is formed. 
         [0028]    Within front part of the engine case  7 , a fan  11  is situated, which takes air A into the core flow path  5  and fan flow path  9 . In front of the center of the fan  11 , an inlet cone  13  guiding the air A rearward is situated. Behind the fan  11 , a compressor  15  is situated. The compressor  15  compresses the air A taken into the core flow path  5 . Behind the compressor  15 , a burner  17  is situated. The burner  17  burns the air A containing the fuel to generate combustion gas G. 
         [0029]    Behind the burner  17 , a high-pressure turbine  19  is situated. The high-pressure turbine  19  is driven by expansion of the combustion gas G from the burner  17  and drives the compressor  15  in conjunction with the same. Behind the high-pressure turbine  19 , a low-pressure turbine  21  is provided. The low-pressure turbine  21  is driven by expansion of the combustion gas G and drives the fan  11  in conjunction with the same. Moreover, within rear part of the core case  3 , a tail cone  23  guiding the combustion gas G rearward is provided concentrically with the core case  3 . The tail cone  23  protrudes rearward from the core case  3 . 
         [0030]    In rear part of the engine case  7 , an afterburner  25  is situated. The afterburner  25  supplies fuel to a gas mixture of the combustion gas G discharged from the core flow path  5  and the air (low-temperature air) A from the fan flow path  9  for burning (afterburning) to increase the thrust (engine thrust) of the aircraft engine  1 . Behind the afterburner  25 , an exhaust nozzle  27  is situated. The exhaust nozzle  27  discharges the combustion gas G and air A (the gas mixture of the combustion gas G and air A). 
         [0031]    Next, a description is given of the concrete configuration of the afterburner  25  according to the embodiment of the present disclosure. 
         [0032]    As illustrated in  FIGS. 1 to 3 , a rear duct (an exhaust duct)  29  is situated concentrically with the engine case  7  downstream of (behind) the engine case  7 . Within the rear duct  29 , a cylindrical liner  31  is situated concentrically with the rear duct  29  with plural supports  33  (one of which is illustrated in  FIG. 1 ) interposed therebetween. The liner  31  allows the combustion gas G and air A to flow rearward. The cross section of the liner  31  along the axial direction (the axial direction of the aircraft engine  1 ) may be corrugated. Between the outer circumferential surface (the outer wall surface) of the liner  31  and the inner circumferential surface (the inner wall surface) of the rear duct  29 , an annular cooling flow path  35  is formed, through which a part of the air A discharged from the fan path flow  9  flows as cooling air CA. Moreover, plural penetrating cooling holes  37  are formed evenly across the liner  31 . The plural cooling holes  37  are configured to blow out the cooling air CA along the inner circumferential surface of the liner  31 . To mix the combustion gas G discharged from the core flow path  5  and the air A discharged from the fan flow path  9 A, a publicly known mixer (not illustrated, see JP 2013-181473 A and JP 2012-132630 A, for example) may be situated in rear part of the core case  3 . 
         [0033]    The rear duct  29  is provided with plural hollow stick-shaped fuel injectors  39  which inject liquid fuel within the liner  31 . The fuel injectors  39  are situated at intervals in the circumferential direction (in a predetermined circumferential direction). The plural fuel injectors  39  are connected to a fuel supply source (not illustrated) which supplies the liquid fuel. The top of each fuel injector  39  penetrates the liner  31 . At the top of each fuel injector  39 , an injection hole  41  is formed. The injection hole  41  can inject the liquid fuel. Downstream of the injection holes  41  (injection positions) of the fuel injectors  39  in the rear duct  29 , plural igniters (ignition plugs)  43  are situated at intervals in the circumferential direction. The igniters  43  ignite (light) the gas mixture containing the fuel within the liner  31 . The top of each igniter  43  penetrates the liner  31 . Instead of the plural hollow stick-shaped fuel injectors  39  provided for the rear duct  29 , a hollow annular fuel injector (not illustrated) which injects liquid fuel in the liner  31  may be provided on the inner wall surface of the liner  31 . 
         [0034]    Directly downstream of the injection holes  41  of the fuel injectors  39  in the liner  31 , a flame holder  45  to hold flame is provided. The frame holder  45  includes an annulus flame holding member  49  and plural radial flame holding members  51 . The annulus flame holding member  49  is provided on the inner circumferential surface (the inner wall surface) of the liner  31  with plural supports  47  (one of which is illustrated in  FIG. 1 ) interposed therebetween and is capable of propagating flame in the circumferential direction. The plural radial flame holding members  51  are situated in a radial fashion inside the annulus flame holding member  49  and are capable of propagating flame in the radial direction. The annulus flame holding member  49  includes plural circular flame holding segments  53  separated in the circumferential direction. The annulus flame holding member  49  is capable of forming a flame holding area (a low-speed area) FA downstream (directly downstream) thereof. The inner diameter of the annulus flame holding member  49  decreases towards the downstream end thereof. The base end (the outside end in the radial direction) of each radial flame holding member  51  is fixed to the rear duct  29 . The radial flame holding members  51  are capable of forming the flame holding area FA downstream (directly downstream) thereof. The cross section of each radial flame holding member  51  along the direction orthogonal to the radial direction has a V shape opened downstream. The annulus frame holding member  49  may be composed of plural circular flame holding segments (not illustrated) separated in the circumferential direction. 
         [0035]    Inside the radial flame holding members  51 , a guide ring (an annular guide member)  53  is provided concentrically with the radial flame holding member  49  with plural supports  55  (one of which is illustrated in  FIG. 1 ) interposed therebetween. The outer diameter of the guide ring  53  decreases towards the downstream end thereof. The upstream end (the edge on the upstream side) of the guide ring  53  is located upstream of the injection holes  41  of the fuel injectors  39 . Between the outer circumferential surface (the outer wall surface) of the guide ring  53  and the inner circumferential surface of the annulus flame holding member  49 , an annular guide flow path  57 , which guides the gas mixture containing fuel downward, is formed. Herein, an inlet  57   i  of the guide flow path  57  is located near the injection holes  41  of the fuel injectors  39  while an outlet  57   o  of the guide flow path  57  is located near the ignition places of the igniters  43  (directly downstream of the annulus flame holding member  49 ). The guide ring  53  may be composed of plural circular guide segments (not illustrated) separated in the circumferential direction. 
         [0036]    Subsequently, a description is given of the operation and effects of the embodiment of the present disclosure. 
         [0037]    When a proper starter device (not illustrated) is operated to drive the fan  11  and compressor  15 , the air A is taken into the core flow path  5  and fan flow path  9  by the fan  11 , and the air A taken into the core flow path  5  is compressed by the compressor  15 . Next, the air A containing fuel is burned by the burner  17  to generate high-pressure combustion gas. By expansion of the combustion gas, the high-pressure turbine  19  and low-pressure turbine  21  are driven to drive the compressor  15  and fan  11  in conjunction. Moreover, the series of operations (drive of the fan  11 , drive of the compressor  15 , burning by the burner  17 , drive of the high-pressure turbine  19 , and drive of the low-pressure turbine  21 ) are successively performed to operate the aircraft engine  1 . Accordingly, during the operation of the aircraft engine  1 , the combustion gas G having passed through the core flow path  5  and the air A having passed through the fan path flow  9  are discharged rearward from the exhaust nozzle  27  to generate thrust of the aircraft engine  1  (engine thrust). The air A discharged from the exhaust nozzle  27  covers the combustion gas G discharged from the exhaust nozzle  27  (normal operation of the aircraft engine  1 ). 
         [0038]    When fuel is injected from the plural fuel injectors  39  in the liner  31  during operation of the aircraft engine  1 , the air A containing the fuel is introduced from the inlet  57   i  of the guide flow path  57  into the guide flow path  57  and flows through the guide flow path  57 . The air A containing the fuel then flows out of the outlet  57   o  of the guide flow path  57  to the ignition places of the plural igniters  43  (directly downstream of the annulus flame holding member  49 ). The plural igniters  43  ignite the gas mixture containing the fuel to form flame downstream (directly downstream) of the flame holder  45  while burning the gas mixture containing the fuel in the liner  31  (afterburning). A lot of heat energy is thereby put into the combustion gas within the liner  31 , thus increasing the thrust of the aircraft engine  1 . 
         [0039]    On the other hand, a part of the air A discharged from the fan flow path  9  flows through the cooling flow path  35  as the cooling air CA during operation of the aircraft engine  1 . The liner  31  is thereby convection-cooled. Moreover, the cooling air CA having contributed to the convection cooling of the liner  31  is blown out from the plural cooling holes  37  and forms the film cooling layer (not illustrated) covering the inner circumferential surface of the liner  31 . The liner  31  is thereby film-cooled. The cooling air CA having contributed to the film cooling of the liner  31  becomes a part of the gas mixture (normal operation of the afterburner  25 ). 
         [0040]    Herein, the air A containing the fuel is introduced into the guide path flow  57  from the inlet  57   i  of the guide flow path  57  and flows out of the outlet  57   o  of the guide flow path  57  to the ignition places of the plural igniters  43 . This can prevent diffusion of fuel in the liner  31  and supply the fuel to the ignition places of the plural igniters  43  at proper concentration and proper evaporated state. The upstream end of the guide ring  53  is located upstream of the injection holes  41  of the fuel injectors  39  in particular. This promotes the air A containing the fuel to be introduced into the guide flow path  57 , further preventing diffusion of the fuel within the liner  31 . In other words, fuel can be supplied to the ignition places of the plural igniters  43  at proper concentration and proper evaporated state independently of the operating state (the operation conditions) of the aircraft engine  1  at the same time or just before the gas mixture containing the fuel is ignited by the plural igniters  43 . 
         [0041]    The flame holder  45  includes the annulus flame holding member  51  capable of propagating flame in the circumferential direction as well as the plural radial flame holding members  51  capable of propagating flame in the radial direction. This enhances flame propagation of the flame holder  45 . Moreover, since the annulus flame holding member  49  is provided on the inner circumferential surface of the liner  31 , the annulus frame member  49  is prevented from interfering with the main flow of the gas mixture within the liner  31 , and pressure loss (thrust loss) in the liner  31  is therefore minimized (characteristic operation of the afterburner  25 ). 
         [0042]    According to the embodiment of the present disclosure, fuel can be supplied to the ignition places of the plural igniters  43  at proper concentration and proper evaporated state independently of the operating state of the aircraft engine  1  at the same time or just before the gas mixture containing the fuel is ignited by the plural igniters  43 . Accordingly, fuel can be stably ignited by the plural igniters  43 , ensuring sufficient ignition stability of the afterburner  25 . Moreover, the flame propagation of the flame holder  45  is enhanced while the pressure loss in the liner  31  is minimized by preventing the interference between the annulus flame holding member  49  and the main flow of the gas mixture within the liner  31 . This ensures sufficient flame holding capability (burning stability) of the afterburner  25  while preventing a decrease in engine efficiency of the aircraft engine  1 . According to the embodiment of the present disclosure, it is possible to ensure sufficient ignition stability and flame holding capability of the afterburner  25  while preventing a decrease in engine efficiency of the aircraft engine  1 . 
         [0043]    The present disclosure is not limited to the description of the above embodiment and can be embodied in various modes. The scope of the present disclosure is not limited to these embodiments. 
         [0044]    While embodiments have been exemplified with the help of the drawings, many modifications and changes are apparent to those skilled in the art.