US20140360197A1

US20140360197A1 - Afterburner and aircraft engine

Info

Publication number: US20140360197A1
Application number: US14/470,286
Authority: US
Inventors: Shinji Tanaka; Katsuyoshi Takahashi; Jun Hosoi
Original assignee: IHI Corp
Current assignee: IHI Corp
Priority date: 2012-03-02
Filing date: 2014-08-27
Publication date: 2014-12-11
Also published as: WO2013129648A1; CA2865543C; EP2821627A4; EP2821627A1; EP2821627B1; JP2013181473A; CA2865543A1

Abstract

An aircraft engine includes an afterburner which has a flame stabilizer. The flame stabilizer maintains a flame generated from a mixed gas of a combustion gas and air. The flame stabilizer includes multiple gutters each configured to generate a flame stabilization area for the flame on its downstream side. Each gutter is formed from: a curved apex section having a stagnation point; and flat plate-shaped side surface sections integrally formed on the respective two sides of the apex section. Each gutter has a V-shaped cross-sectional shape which is opened to the downstream side. At least one through-hole is formed only in each side surface section of each gutter.

Description

CROSS REFERENCE TO RELATED APPLICATIONS

This application is a continuation application of International Application No. PCT/JP2013/055665, filed on Mar. 1, 2013, which claims priority to Japanese Patent Application No. 2012-046319, filed on Mar. 2, 2012, the entire contents of which are incorporated by references herein.

BACKGROUND OF THE INVENTION

1. Field of the Invention
The present invention relates to: an afterburner configured to increase thrust of an aircraft engine by burning a mixed gas of a combustion gas (a high-temperature gas) discharged from a core passage of the aircraft engine and air (low-temperature air) discharged from a fan passage of the aircraft engine again while supplying fuel to the mixed gas; and the like.
2. Description of the Related Art
Various research and development programs on afterburners have been conducted in recent years to meet demands for aircraft engines with higher thrust. A configuration and the like of a general afterburner are as follows.
An aircraft engine includes: an engine case; an outer duct provided to an outlet portion (rear portion) of the engine case; and a liner, a fuel injector and an igniter provided inside the outer duct. The liner is shaped like a cylinder, and allows a mixed gas to flow through the liner. The fuel injector injects fuel inside the liner. The igniter is provided in the rear (on the downstream side) of the fuel injector inside the outer duct. The igniter ignites the mixed gas, inclusive of the fuel, inside the liner. The aircraft engine further includes a flame stabilizer (flame holder). The flame stabilizer is provided in the rear of the fuel injector inside the liner, and holds the flame. The flame stabilizer includes multiple gutters which are arranged radially. Each gutter forms a flame stabilization area on its downstream side (immediately downstream side).
While the aircraft engine is in operation, the fuel injector continues injecting the fuel inside the liner, and the igniter continues igniting the mixed gas inclusive of the fuel. For this reason, the flame is formed on the downstream side (in the rear) of the flame stabilizer inside the liner, and the mixed gas is burned again with the fuel. Thereby, more thermal energy can be injected into the combustion gas inside the liner, and the thrust of the aircraft engine can be accordingly increased.
It should be noted that conventional techniques related to the present invention are disclosed in Japanese Patent Application Laid-Open Publication Nos. H 09-112345, H09-119346, and H06-137213 as well as JING-TANG YANG et al., Combustion and Flame, 99,288-294(1994).

SUMMARY OF THE INVENTION

As described above, the flame stabilization area is formed on a downstream side of each gutter. The flow inside the flame stabilization area has a relatively low velocity. In contrast, a flow passing between each neighboring two gutters has a high velocity because the flow area is narrowed. For this reason, as shown in FIG. 7, the high-velocity flow passing between the two neighboring gutters and the low-velocity flow inside the flame stabilization area are mixed together on the immediately downstream side of the flame stabilizer. The mixture of the two flows causes a large pressure loss. The pressure loss lowers a rate of increase in thrust of the aircraft engine, and makes it difficult to enhance the engine performance of the aircraft engine to a high level. Incidentally, in FIG. 7, reference sign “F” indicates a frontward direction (upstream direction), reference sign “R” indicates a rearward direction (downstream direction), and reference sign “C” indicates a circumferential direction.
An object of the present invention is to provide an afterburner and the like which can solve the foregoing problem.
A first aspect of the present invention provides an afterburner configured to increase thrust of an aircraft engine by burning a mixed gas of a combustion gas discharged from a core passage of the aircraft engine and air discharged from a fan passage of the aircraft engine again while supplying fuel to the mixed gas. The afterburner includes: an outer duct provided to an outlet portion (rear portion) of an engine case of the aircraft engine; a cylindrical liner provided inside the outer duct, configured to allow the mixed gas to flow through the liner; a fuel injector configured to inject the fuel into the liner; an igniter configured to ignite the mixed gas including the fuel inside the liner; and a flame stabilizer (flame holder) disposed in the rear (on the downstream side) of the fuel injector inside the liner, including a plurality of gutters arranged radially and each configured to form a flame stabilization area on a downstream side (immediately downstream side) of the gutter, and configured to hold a flame. Here, each gutter is formed from a curved apex section and flat plate-shaped side surface sections integrally formed on two sides of the apex section, and the gutter has a V-shaped cross-sectional shape which is opened to a downstream side of the gutter. Moreover, at least one through-hole is formed only in each side surface section of the gutter.
It should be noted that in the scope of claims and the specification of the application concerned, the term “provided” means that a state of being provided indirectly through another member as well as a state of being provided directly. Similarly, the term “arranged” means a state of being arranged through another member as well as a state of being arranged directly. Furthermore, the term “downstream side” means a downstream side in view of a flowing direction of a main stream.
A second aspect of the present invention provides an aircraft engine configured to produce thrust by exhausting a mixed gas of a combustion gas and air in a rearward direction, which includes the afterburner of the first aspect.
The present invention can sufficiently reduce the pressure loss on the downstream side of the flame stabilizer. Thereby, the present invention can increase the rate of increase in thrust of the aircraft engine, and can enhance the engine performance of the aircraft engine to a high level.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a cross-sectional view of a gutter of an embodiment of the present invention taken along the I-I line of FIG. 2.

FIG. 2 is a magnified view of a part indicated with an arrow II in FIG. 5.

FIG. 3 is a magnified perspective view of a part indicated with an arrow III in FIG. 5.

FIGS. 4A, 4B and 4C are partial side views of gutter portions in an afterburner of the embodiment of the present invention, respectively.

FIG. 5 is a schematic sectional side view of an aircraft engine of the embodiment of the present invention.

FIG. 6 is a diagram showing a relationship between a predetermined mass flow rate while the aircraft engine is in operation and a pressure coefficient on the downstream side of a flame stabilizer.

FIG. 7 is a plan development of a vicinity of multiple gutters in a general afterburner.

DESCRIPTION OF THE PREFERRED EMBODIMENTS

The present invention has been made with a result of an analysis using the below-described three-dimensional steady viscosity CFD (Computational Fluid Dynamics) taken into consideration.
This analysis used an assumed gutter having a V-shaped cross-sectional shape which was opened to a downstream side of the gutter and including multiple through-holes formed only in its two side surface sections. The analysis found a relationship between a ratio Qb/Qa of mass flow rates of mixed gases (a predetermined ratio of mass flow rates in an aircraft engine in operation) and a pressure coefficient on the downstream side of the flame stabilizer. In this respect, Qa denotes amass flow rate of a mixed gas flowing toward the gutter in the aircraft engine in operation, and Qb denotes a mass flow rate of a mixed gas introduced into the gutter through the multiple through-holes in the gutter. FIG. 6 shows the result of the analysis. As shown in the graph, it was learned that as the ratio Qb/Qa of the mass flow rates became larger (became equal to 0.10 or more, preferably), the pressure loss on the downstream side of the flame stabilizer could be reduced sufficiently. In this graph, one scale mark on the vertical axis represents 0.1. FIG. 6 shows that a smaller value of the pressure coefficient on the vertical axis means a smaller pressure loss on the downstream side of the flame stabilizer. In short, it was learned from this analysis that when the gutter had the above-mentioned shape, the pressure loss on the downstream side of the flame stabilizer could be reduced sufficiently.
A preceding study (see NPL 1), meanwhile, has already showed clearly that when multiple through-holes are formed not only in the two side surface sections of the gutter but also in an apex section of the gutter, a high-velocity mixed gas flows into the gutter and the flame stabilization performance (combustion stabilization performance) of the flame stabilizer accordingly deteriorates to a large extent.
Next, referring to FIGS. 1 to 5, descriptions will be provided for an embodiment of the present invention. It should be noted that reference sign “F” indicates a forward direction (upstream direction) and reference sign “R” indicates a rearward direction (downstream direction.)
As shown in FIG. 5, an aircraft engine (hereinafter simply referred to as an engine) 1 of the embodiment is an apparatus configured to produce thrust (engine thrust) by discharging a mixed gas of a combustion gas (a high-temperature gas) and air (low-temperature air) in the rearward direction. The engine 1 includes: a cylindrical engine case 3; and a core passage (a main passage) 5 and a fan passage (a bypass passage) 7 formed in the inside of the engine case 3 by partitioning the inside thereof. The core passage 5 is formed in an annular shape (in a hollowed cylindrical shape), and allows the combustion gas to flow through the core passage 5 in the rearward direction. The fan passage 7 is placed outside the core passage 5, and is formed in an annular shape (a hollowed cylindrical shape). The fan passage 7 allows the air (the low-temperature air) to flow through the fan passage 7 in the rearward direction.
The engine 1 includes: a fan 9 disposed in a front portion of the inside of the engine case 3; and an inlet cone 11 disposed in a center portion on the front side of the fan 9. The fan 9 takes air into the core passage 5 and the fan passage 7. The inlet cone 11 guides the air in the rearward direction. The engine 1 further includes: a compressor 13 disposed in the rear of the fan 9; and a combustor 15 disposed in the rear of the compressor 13. The compressor 13 compresses the air which has been taken into the core passage 5. The combustor 15 generates the combustion gas by combusting the air including fuel.
The engine 1 includes: a high-pressure turbine 17 disposed in the rear of the combustor 15; and a low-pressure turbine 19 disposed in the rear of the high-pressure turbine 17. The high-pressure turbine 17 is driven by expansion of the combustion gas from the combustor 15. In conjunction with this, the high-pressure turbine 17 drives the compressor 13. The low-pressure turbine 19 is driven by expansion of the combustion gas. In conjunction with this, the low-pressure turbine 19 drives the fan 9. Furthermore, a tail cone 21 is disposed in the rear of the low-pressure turbine 19. The tail cone 21 guides the combustion gas in the rearward direction.
The engine 1 includes an afterburner 23 disposed in the rear of the engine case 3. The afterburner 23 increases the thrust (the engine thrust) of the engine 1 by: burning again the mixed gas of the combustion gas discharged from the core passage 5 and the air (the low-temperature air) discharged from the fan passage 7 while supplying fuel to the mixed gas. An exhaust nozzle 25 configured to exhaust the mixed gas which has been burned again by the afterburner 23 is disposed in the rear of the afterburner 23.
Next, descriptions will be provided for a concrete configuration of the afterburner 23 of the embodiment.
As shown in FIGS. 2 and 5, the afterburner 23 includes: an outer duct 27 disposed in the rear (in the outlet portion) of the engine case 3; and a cylindrical liner 29 disposed inside the outer duct 27. The outer duct 27 is connected to the exhaust nozzle 25. Furthermore, the liner 29 allows the mixed gas to flow through the liner 29 in the rearward direction. In addition, a mixer 31 is disposed in the rear of the engine case 3. The mixer 31 is placed inside the liner 29, and mixes the combustion gas discharged from the core passage 5 and the air discharged from the fan passage 7. Incidentally, the mixer 31 may have the same configuration as does a publicly-known mixer shown in Japanese Patent Application Laid-Open Publication No. H06-137213 as described above.
The afterburner 23 includes multiple spray bars 33 which are examples of a fuel injector, and the spray bars 33 are provided to the outer duct 27. The spray bars 33 are disposed at intervals in the circumferential direction of the outer duct 27. The tip end-side portion of each spray bar 33 is placed inside the liner 29. Each spray bar 33 injects the fuel into the liner 29. The afterburner 23 further includes an ignition device (an igniter) 35 provided to the outer duct 27 in the rear of the spray bars 33. The igniter 35 ignites (lights off) the mixed gas including the fuel inside the liner 29.
As shown in FIGS. 2 and 3, the afterburner 23 includes a flame stabilizer (a flame holder) 37. The flame stabilizer 37 is disposed in the rear (on the downstream side) of the spray bars 33 inside the liner 29, and holds the flame. In addition, the flame stabilizer 37 includes multiple gutters 39 arranged radially. Each gutter 39 generates a flame stabilization area FA on the downstream side (immediately on the downstream side) of the gutter 39. Incidentally, the tip end-side portion (whose illustration is omitted) of the igniter 35 is placed in the flame stabilization area FA.
As shown in FIGS. 1 and 4A, each gutter 39 is formed from: a curved apex section 39 a including a stagnation point P; and flat plate-shaped side surface sections 39 b integrally formed on the two sides of the apex section 39 a. Each gutter 39 has a V-shaped cross-sectional shape which is opened to the downstream side. Furthermore, multiple circular through-holes 41 are formed only in the two side surface sections 39 b of each gutter 39. It should be noted that the shape of each through-hole 41 is not limited to the circular one. In other words, as shown in FIGS. 4B and 4C, each through-hole 41 may be shaped like an ellipse, or like a slit (slot). Otherwise, instead of the multiple through-holes 41, a single slit (slot)-shaped through-hole (whose illustration is omitted) may be formed.
As shown in FIG. 1, a deceleration space (a deceleration section, a mixing space) S for mixing and concurrently decelerating mixed gases introduced through the multiple through holes 41 is formed in each gutter 39. In addition, the multiple through-holes 41 are formed in each gutter 39 in a way that the ratio Qb/Qa of the mass flow rates of the mixed gases is equal to 0.10 or more. As discussed in the beginning, Qa denotes the mass flow rate of the mixed gas flowing toward each gutter 39 while the engine 1 is in operation, and Qb denotes the mass flow rate of the mixed gas introduced into the inside (the deceleration space S) of each gutter 39 through the multiple through-holes 41 in the gutter 39. To put it concretely, in each gutter 39, the total opening area and arrangement pattern of the multiple through-holes 41, the orientation of the multiple through-holes 41 with respect to the thickness direction of the side surface sections 39 b, and the like are set in a way that the ratio (Qb/Qa) of the mass flow rates is equal to 0.10 or more. It should be noted that when the ratio (Qb/Qa) of the mass flow rates exceeds 0.30, it is difficult to sufficiently secure the flame stabilization performance (the combustion stabilization performance) of the flame stabilizer 37.
As shown in FIG. 2, an annular liner cooling passage 43 is formed between the inner surface of the outer duct 27 and the outer surface of the liner 29. The liner cooling passage 43 allows part of the air discharged from the fan passage 7 to flow through the liner cooling passage 43 as cooling air.
Next, descriptions will be provided for the working and effects of the embodiment.
Once the fan 9 and the compressor 13 are driven by an operation of an appropriate starter apparatus (whose illustration is omitted), the air can be taken by the fan 9 into the core passage 5 and the fan passage 7, and the air taken into the core passage 5 can be compressed by the compressor 13. Subsequently, once the high-pressure combustion gas is generated by causing the combustor 15 to combust the air including the fuel, the high-pressure turbine 17 and the low-pressure turbine 19 are driven by the expansion of the combustion gas. Thereby, the compressor 13 and the fan 9 can be driven in conjunction with the high-pressure turbine 17 and the low-pressure turbine 19, respectively. Thereafter, the engine 1 can be put into operation through a series of operations (the driving of the fan 9, the driving of the compressor 13, the combustion by the combustor 15, as well as the driving of the high-pressure turbine 17 and the low-pressure turbine 19). Subsequently, while the engine 1 is in operation, the combustion gas discharged from the core passage 5 and the air discharged from the fan passage 7 are mixed together by the mixer 31, and, as the mixed gas, are exhausted through the exhaust nozzle 25 in the rearward direction. Thereby, the thrust (the engine thrust) of the engine 1 can be produced.
While the engine 1 is in operation, the fuel is injected from the multiple spray bars 33 into the liner 29, and the mixed gas including the fuel is ignited by the igniter 35. Thereby, the flame is produced in the rear (on the downstream side) of the flame stabilizer 37 inside the liner 29, and the mixed gas is burned again with the fuel. By this, thermal energy can be injected into the combustion gas inside the liner 29, and the thrust of the engine 1 can be increased. On the other hand, while the engine 1 is in operation, part of the air discharged from the fan passage 7 flows through the liner cooling passage 43 as the cooling air. Thereby, the liner 29 can be cooled (in the normal operation of the afterburner 23).
In this respect, each gutter 39 has the V-shaped cross-sectional shape which is opened to the downstream side of the gutter 39, and the multiple through-holes 41 are formed only in the two side surface sections 39 b of each gutter 39. Furthermore, in each gutter 39, the total opening area of the multiple through-holes 41 and the like are set in a way that the ratio Qb/Qa of the mass flow rates is in the range of 0.10 to 0.30 inclusive. For these reasons, as learned from the above-discussed result of the analysis, the afterburner 23 of the embodiment is capable of sufficiently reducing the pressure loss on the downstream side of the flame stabilizer 37 while sufficiently securing the flame stabilization performance of the flame stabilizer 37.
Accordingly, the present invention makes it possible to improve the engine performance of the engine 1 to a high level by: sufficiently securing the flame stability performance of the flame stabilizer 37; and additionally improving a rate of increase in thrust of the engine 1.
It should be noted that the present invention is not limited to the foregoing descriptions which have been provided for the embodiment, and can be carried out in various modes. Moreover, the scope of rights covered by the present invention is not limited to the embodiment.

Claims

1. An afterburner configured to increase thrust of an aircraft engine by burning a mixed gas of a combustion gas discharged from a core passage of the aircraft engine and air discharged from a fan passage of the aircraft engine again while supplying fuel to the mixed gas, the afterburner comprising:

an outer duct provided to an outlet portion of an engine case of the aircraft engine;

a cylindrical liner provided inside the outer duct, and configured to allow the mixed gas to flow through the liner;

a fuel injector configured to inject the fuel into the liner;

an igniter configured to ignite the mixed gas including the fuel inside the liner; and

a flame stabilizer disposed in a rear of the fuel injector inside the liner, including a plurality of gutters arranged radially and each configured to form a flame stabilization area on a downstream side of the gutter, and configured to hold a flame, wherein

each gutter is formed from a curved apex section and flat plate-shaped side surface sections integrally formed on two sides of the apex section, and has a V-shaped cross- sectional shape which is opened to a downstream side of the gutter, and

at least one through-hole is formed only in each side surface section of the gutter.

2. The afterburner according to claim 1, wherein a deceleration space for mixing and concurrently decelerating the mixed gas introduced through the at least one through-hole is defined inside the gutter.

3. An aircraft engine configured to produce thrust by exhausting a mixed gas of a combustion gas and air in a rearward direction, comprising the afterburner according to claim 1.