JPH09105356A - Afterburner for aircraft engine - Google Patents

Abstract

PROBLEM TO BE SOLVED: To accelerate the evaporation and mixing of fuel by adding a scoop, guiding a turbine exhaust current along an inclined surface, to a ring gutter to be arranged in the crossed state of this turbine exhaust current, and setting up an inner fuel spray nozzle at the upstream position of this ring gutter for fuel injection. SOLUTION: An aircraft engine after burner to be set up downstream a combustor and a turbine is provided with a frame holder 9 which has a ring gutter 13 to be set up crosswise with a turbine exhaust current, and double leg gutters 14 being plurally set up at an interval in the circumferential direction to the ring gutter 13, and set up crosswise with a fan air current. In addition, a ring scoop 15, guiding the turbine exhaust current along an inclined surface, is installed in a nearby position of this inclined surface of the ring gutter 14, and a core flow inlet port 21 is installed in an outer surface of this scoop 15. Subsequently, three fuel spray nozzles 22 to 24 are set up at the upstream side of the ring gutter 13, and each outer position of the core flow inlet port 21 and the scoop 15, respectively.

Description

DETAILED DESCRIPTION OF THE INVENTION

[0001]

BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to an afterburner for an aviation engine, and more particularly to a technique for improving combustion performance in the afterburner.

[0002]

2. Description of the Related Art As a related technology of an afterburner for an aircraft engine, Ishikawajima-Harima Technical Report Vol. 31 No. 2 (Heisei 3.3)
An example of a turbofan engine with an afterburner as shown in FIGS. 3 to 4 is introduced. In these figures, reference numeral 1 denotes an aero engine, 2 denotes a fan, 3 denotes a compressor, 4 denotes a combustor, 5 denotes a turbine, 6 denotes an afterburner, 7
Is a fuel injector, 8 is a spark plug, 9 is a flame stabilizer, 1
0 is a liner, 11 is a duct, 12 is an exhaust nozzle, A is a turbine exhaust flow, and B is a fan air flow.

As shown in FIG. 4, the flame stabilizer 9 is formed in a gutter shape having a V-shaped cross section, so that even when a low-temperature fan airflow B flows in, it is stable and efficient. It can be burned well.

FIG. 5 shows an example of an improved flame stabilizer 9. In the example of FIG.
In addition, a double leg gutter 14 and a scoop 15 are added, and when operated as an afterburner 6, satisfactory results are obtained in terms of combustion efficiency, pressure loss rate, combustion stability, ignition performance, overall air-fuel ratio, and the like. It is reported to have been obtained.

[0005]

The present invention has been made based on the technology of an afterburner using the flame stabilizer of the example shown in FIG. 5, and promotes the evaporation and mixing of fuel and further improves the combustion performance and ignition performance. It is intended to improve.

[0006]

SUMMARY OF THE INVENTION An annular gutter is disposed intersecting with a turbine exhaust flow, and is disposed at a circumferential distance from the annular gutter so as to intersect with a fan air flow.
A heavy leg gutter, a scoop disposed annularly at a position near the inclined surface of the annular gutter to guide the turbine exhaust flow along the inclined surface, and a part of the turbine exhaust flow disposed on the outer surface of the scoop and having a double leg gutter. A technique is employed that includes a core inlet that is interposed therebetween and an inner fuel spray nozzle that is disposed upstream of the annular gutter and inside the scoop and ejects fuel along the inclined surface of the annular gutter. In this case, the double leg gutta is arranged at least outside the annular gutta. A fuel spray nozzle for ejecting fuel into the core inlet is disposed at the core inlet.
An outer fuel spray nozzle for ejecting fuel along the outer surface of the scoop is provided at a position outside the scoop. The outer fuel spray nozzle is disposed on the same circle as the fuel spray nozzle.

[0007]

During operation of the combustor and the turbine, the turbine exhaust flow is guided by the inner surface of the scoop and flows along the inclined surface of the annular gutta, and the comparison is made based on the sectional shape of the annular gutta near the downstream of the annular gutta. A stagnation zone of high temperature flow is formed. The fuel injected from the inner fuel spray nozzle is sent downstream from the inclined surface of the annular gutter in a state of being mixed with the turbine exhaust flow, but part of the fuel is sent to the stagnation area to maintain the fuel concentration inside. In combination with the high temperature, ignitability is ensured and flame holding is performed.
When a part of the turbine exhaust flow is introduced between the core inlet and the double leg gutter, the hot gas is guided to the outside and the stagnation of the fan air flow near the downstream of the double leg gutter based on the cross-sectional shape of the gutter. While forming the area,
Temperature rise and flame holding are performed. At this time, the fuel ejected from the fuel spray nozzle is sent to the double leg gutter in a state where it is mixed with a part of the turbine exhaust flow, and the temperature is increased and the fuel concentration is maintained. The ignitability of the part and the flame holding are performed. When fuel is supplied from the outer fuel spray nozzle to the vicinity of the outer surface of the scoop, the fuel merges with the fan air flow to mix the fuel and air, and is ignited by the flame of the annular gutter and the double leg gutter. , The combustion range is expanded.

[0008]

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS Referring to FIGS.
An embodiment of an afterburner for an aeroengine according to the present invention will be described.

The afterburner 6 according to the embodiment is disposed downstream of the combustor 4 and the turbine 5 as in the examples of FIGS. 3 and 4. A ring-shaped annular gutter 13 arranged in an intersecting state with the flow A, and a plurality of double leg gutters 14 arranged at intervals in the circumferential direction with respect to the annular gutter 13 and arranged in an intersecting state with the fan air flow B However, in addition to these techniques, the upstream of the annular gutter 13 and the
In connection with the core inlet 21 and the scoop 15 of the heavy leg gutter 14, an inner fuel spray nozzle 22, a fuel spray nozzle 23 and an outer fuel spray nozzle 24 are arranged to partially fulfill the function of the fuel injector 7. .

As shown in FIG. 1, the inner fuel spray nozzle 22 has a plurality of small fuel jets which are arranged upstream of the annular gutter 13 and between the inner and outer scoops 15 to eject fuel. With a hole and an annular gutter 1
The fuel is supplied along the third inclined surface 13a.

In the fuel spray nozzle 23,
The fuel is injected into the core intake port 21, and mainly ejects a part of the turbine exhaust flow A taken in from the core intake port 21 to supply the fuel between the double leg gutters 14.

As shown in FIG. 1, the outer fuel spray nozzle 24 is provided at a position outside the scoop 15 and has a number of small holes for ejecting fuel along the outer surface of the scoop 15. doing. Further, the fuel spray nozzle 23 and the outer fuel spray nozzle 24 are arranged on the same circle, and are constituted by one ring-shaped tubular body formed as a whole as a whole.

In the scoop 15, the outer scoop 15A and the inner scoop 15B are combined to form a slight gap between the outer scoop 15A and the inclined surface 13a of the annular gutter 13, so that the turbine exhaust flow A is mainly controlled. It is set so as to guide along the inclined surface 13a of the annular gutter 13. In the outer scoop 15A, a plurality (a plurality of sets) of the outer leg gutter 14A and the inner leg gutter 14B constituting the double leg gutter 14 are arranged at intervals in the circumferential direction as described above. A part of the taken-in turbine exhaust flow A is set to be guided to a flow path 14a between the outer leg gutter 14A and the inner leg gutter 14B.

In the afterburner for an aircraft engine thus configured, when performing fuel ignition, fuel is supplied only to the inner fuel spray nozzle 22 to perform ignition, combustion and flame holding. That is, when the afterburner 6 is operated in addition to the operation of the combustor 4 and the turbine 5 described above, the fuel supply amount to the fuel injector 7, the inner fuel spray nozzle 22, the fuel spray nozzle 23, and the outer fuel spray nozzle 24 is reduced. In addition to controlling individually, the fuel supply amount required for ignition is calculated, and an appropriate amount of fuel is injected from the inner fuel spray nozzle 22. As shown in FIG.
2 between the inclined surface 13a and the outer scoop 15A and the inner scoop 15B to mix fuel and air, and a part of the mixed flow is downstream of the annular gutter 13 by an arrow in FIG. As shown, let it go around,
The fuel concentration in the stagnant atmosphere downstream of the annular gutter 13 is optimized, and the fuel is ignited by operating the ignition plug 8. Hereinafter, the fuel concentration is maintained, the stagnation phenomenon occurs, and the turbine Flame holding based on the high temperature of the exhaust gas flow A is performed, and the combustion state is continued due to the flame spreading in the circumferential direction of the annular gutter 13.

When fuel is injected from the fuel spray nozzle 23 while the flame is maintained by the annular gutter 13, a part of the turbine exhaust flow A is mixed with fuel, and the mixed fluid flows through the double leg gutter 14. The outer leg gutter 14A and the inner leg gutter 14 are guided outward by the path 14a.
It is sent downstream from between B. The mixed flow passing through the flow path 14a is in a state where the fuel concentration and the temperature are increased, and a stagnant area of the flow based on the cross-sectional shape is formed near the downstream of the double leg gutter 14; Ignition is promoted downstream of the heavy leg gutter 14, and flame holding is performed at a plurality of locations.

Next, a scoop 15 is provided from each of the above-described fuel injectors 7 and the outer fuel spray nozzle 24.
When fuel is supplied near the outer surface of the fan, the fan airflow B
The fuel is combined with the air to mix the fuel and the air, and the mixed fluid is ignited by the flame of the annular gutter 13 and the double leg gutter 14, so that the combustion range is reduced as shown in FIG.
It extends throughout the interior of the example liner 10.

[0017]

The afterburner for an aircraft engine according to the present invention has the following excellent effects. (1) By adding a scoop to the annular gutter and arranging an inner fuel spray nozzle upstream of the annular gutter to perform fuel injection, it promotes evaporation and mixing of fuel and utilizes a high-temperature turbine exhaust flow. Thus, the evaporation and mixing of the fuel can be promoted, and the ignitability and the flame holding property of the annular gutter can be improved. (2) A fuel spray nozzle for ejecting fuel into the core inlet is disposed at the core inlet, thereby increasing the fuel concentration and temperature between the double leg gutters and evaporating and mixing the fuel by the double leg gutter. Flame holding properties can be further improved. (3) By arranging the outer fuel spray nozzle at a position outside the scoop, the ignition performance can be improved in combination with the flame holding property. (4) The overall combustion performance can be improved by the above synergistic action.

[Brief description of the drawings]

FIG. 1 is a perspective view of a main portion of an embodiment of an afterburner for an aircraft engine according to the present invention.

FIG. 2 is a front sectional view of FIG.

FIG. 3 is a front sectional view of an example of a turbofan engine with an afterburner, with a part of the description omitted;

FIG. 4 is a front sectional view showing a structural example of a portion of an afterburner of FIG. 3;

5 is a perspective view and a front sectional view of a part of a flame stabilizer of the afterburner of FIG. 3;

[Explanation of symbols]

 DESCRIPTION OF SYMBOLS 1 Aero engine 4 Combustor 5 Turbine 6 After burner 7 Fuel injector 8 Spark plug 9 Flame stabilizer 10 Liner 11 Duct 12 Exhaust nozzle 13 Annular gutter 13a Inclined surface 14 Double leg gutter 14A Outer leg gutter 14B Inner leg gutter 14A Flow path 15A Scoop 15B Inner scoop 21 Core inlet 22 Inner fuel spray nozzle 23 Fuel spray nozzle 24 Outer fuel spray nozzle A Turbine exhaust flow B Fan air flow

Front page continuation (72) Inventor Takeshi Kashiwagi 229 Tonogaya, Mizuho-cho, Nishitama-gun, Tokyo Ishi Kawashima Harima Heavy Industries Ltd., Mizuho Plant (72) Inventor Shizuka Nakano 229 Tonogaya, Mizuho-cho, Nishitama-gun, Tokyo Ishikawajima Harima Heavy Industries Mizuho Factory Co., Ltd.

Claims (4)

[Claims] 1. An annular gutta (13) disposed in a state intersecting with a turbine exhaust flow (A), and a circumferentially spaced interval in a state intersecting with the fan airflow (B) with respect to the annular gutta. A double leg gutta (14), a scoop (15) annularly arranged near the inclined surface (13a) of the annular gutta for guiding the turbine exhaust flow along the inclined surface, and arranged on the outer surface of the scoop. A core inlet (21) for taking in a part of the turbine exhaust flow between the double leg gutta and an inner fuel spray nozzle (22) located upstream of the annular gutta and ejecting fuel along the inclined surface. Afterburner for aero engines characterized by: 2. The afterburner for an aeronautical engine according to claim 1, wherein the core inlet (21) is provided with a fuel spray nozzle (23) for injecting fuel into the core inlet. 3. A position on the outside of the scoop (15),
An afterburner for an aeroengine according to claim 1 or 2, further comprising an outer fuel spray nozzle (24) for ejecting fuel along the outer surface of the scoop. 4. The afterburner for an aeroengine according to claim 3, wherein the outer fuel spray nozzle (24) is arranged on the same circle as the fuel spray nozzle (23).