JPH08144852A - Aircraft engine cooling fluid control system - Google Patents

Abstract

(57) [Abstract] [Purpose] To provide a cooling fluid control device that can control an air flow according to the load of an aircraft engine to improve the performance at a low load and can be easily implemented. [Composition] In an aircraft engine in which a part of a bypass flow is sent as cooling air to an annular cooling flow path between a combustion duct and a combustion liner, and fuel is injected into the combustion liner for combustion, the upstream end of the combustion liner is used. A thermally actuated shutter that changes the amount of air introduced into the annular cooling passage of the bypass flow by thermal displacement based on heat transfer is disposed in the section.

Description

Detailed Description of the Invention

[0001]

BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to a cooling fluid control device for an aircraft engine, and more particularly to improving the overall performance of the engine by controlling the cooling fluid for the combustion burner in an afterburner for an aircraft engine or a ram combustor. It is intended.

[0002]

2. Description of the Related Art FIG. 3 shows a structural example of an aircraft engine (gas turbine engine) having an afterburner.

In the aircraft engine 1, a fan 2 that takes in air, a compressor 3 that compresses the taken-in air, a combustor 4 that mixes and burns the compressed air with fuel, and the combustor 4 are provided. A turbine 5 for driving the fan 2 and the compressor 3 by the combustion gas of 1. and an afterburner 6 for recombusting the combustion gas and combusting the newly added fuel.

In the afterburner 6, a flame stabilizer 7 having a triangular cross-section or the like to form a turbulent flow region X downstream for flame holding, and a fuel nozzle 8 for ejecting fuel.
A spark plug 9 is provided, and the combustion gas from the afterburner 6 is ejected from the exhaust nozzle 12 via the inside of the combustion liner 11 inside the combustion duct 10 to change the thrust.

A relatively low temperature bypass flow (fan flow) branched from the fan 2 is provided in the afterburner 6 portion.
13 and a relatively high temperature core flow 1 that operates the turbine 5.
4 and is sent. Among these, a part (for example, about 10%) of the air flow of the bypass flow 13 is sent to the annular cooling flow path 15 between the combustion duct 10 and the combustion liner 11,
By ejecting air little by little onto the inner surface of the combustion liner 11, it is possible to reduce the influence of heat on the combustion duct 10 when the afterburner 6 is operating (at the time of high load). That is, as shown by modeling in FIG. 4, when the temperature of the combustion liner 11 when the afterburner 6 is not operating at a low load is, for example, 150 ° C., the afterburner 6 is t.
If it is operated at 1 o'clock, the temperature rises rapidly to, for example, 500 ° C. at t 2 after an extremely short time.
Large thermal displacement and thermal stress are applied to the combustion liner 11. For this reason, the annular cooling flow path 15 is based on the time of high load.
By setting the amount of air to be sent to the air conditioner, an air flow is constantly formed in the annular cooling flow passage 15 to ensure the soundness of the aircraft engine 1.

[0006]

However, with such a structure, the afterburner 6 is provided in the annular cooling passage 15.
The air flow must be shaped to be matched at maximum temperature with or without (or ram combustor) operation. Therefore, for example, at the time of low load when the afterburner 6 is not operated, energy loss occurs due to the formation of the airflow in the annular cooling flow path 15, and the performance of the aircraft engine 1 at normal time (at low load). Will be reduced.

The present invention has been made in view of these problems, and provides a cooling fluid control device which can control the air flow according to the load to improve the performance at low load and can be easily implemented. Is intended.

[0008]

A part of a bypass flow is sent as cooling air to an annular cooling flow path formed between a combustion duct and a combustion liner inside thereof, and fuel is supplied into the combustion liner. As an aircraft engine that, when injected, burns by using the air in the bypass flow and the core flow, at the upstream end of the combustion liner, to the annular cooling flow path of the bypass flow by thermal displacement due to heat transfer from the vicinity of the combustion liner. The heat-actuated shutter that changes the amount of introduced air is used. As a thermally actuated shutter, a technique of dividing into multiple parts in the circumferential direction, a technique of integrally attached to a combustion liner, a technique of bimetal, and a shape memory alloy that operates in a plurality of temperature ranges including at least a low temperature transformation phase and a high temperature transformation phase. Is added. The circumferentially divided portions of the heat-actuated shutter are selected to be in an overlapped state, a state in which some air circulation is allowed, or the like.

[0009]

The bypass flow and the core flow inside the bypass flow through the aircraft engine according to the flight mode. Most of the bypass flow is sent to the inside of the combustion liner, and a part of the bypass flow passes through the heat-actuated shutter. Then, it is sent to the annular cooling flow path to cool the combustion duct and the combustion liner. In a flight mode that does not require fuel injection inside the combustion liner (when the load is low), the heat transfer amount from the combustion liner and its inside becomes small, and the heat-actuated shutter closes the annular cooling flow path, and By suppressing the amount of air introduced, the occurrence of pressure loss inside the annular cooling passage is reduced, and the performance of the aircraft engine is improved. When the afterburner is operating or when the ramjet engine of the combined cycle engine is operating (when the load is high), fuel is injected into the combustion liner and burned, and the temperature rapidly increases after an extremely short time. Rise to. When the heat-actuated portion of the bimetal or shape memory alloy in the heat-actuated shutter is heated by the conduction heat from the vicinity of the combustion liner and the heat transfer from the combustion gas, etc., thermal displacement occurs and the heat-actuated shutter. Opens the annular cooling passage, thereby increasing the amount of air introduced into the inside of the annular cooling passage. By setting the amount of air to be sent into the annular cooling flow path based on the time of high load, the soundness of the aircraft engine under high load can be secured. When the heat-actuated shutter is divided into a plurality of pieces in the circumferential direction and the divided portions are in the overlapped state, not only the air shielding property at the divided portions is improved, but also the thermal displacement is smoothed. The amount of air required to cool the combustion liner at a low load is ensured by allowing the annular cooling passage to allow some air flow.

[0010]

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS An embodiment of a cooling fluid control device for an aircraft engine according to the present invention will be described below with reference to FIGS.

Even in the aircraft engine 1 of the embodiment, the combustion duct 10 and the combustion liner 1 located inside thereof are provided.
1, a part of the bypass flow 13 is sent as cooling air into the annular cooling flow path 15 formed between the bypass flow 13 and the annular cooling flow path 15. A heat-actuated shutter 21 for adjusting the amount of air introduced into the device is arranged.

As shown in FIGS. 1 and 2, the heat-actuated shutter 21 burns a plurality of heat-actuated flaps 22 which are circumferentially divided so that adjacent portions are overlapped. The liner 11 is integrally attached to the upstream end of the liner 11 by a fastener 23 such as a bolt / nut or a rivet or by welding.

The tip (upstream end) of the heat-operated flap 22 is slightly discharged from the inner surface of the combustion duct 10 at a low load when the fuel is not jetted into the combustion liner 11 and the jetted fuel is not burned. The gap G is formed so that a small amount of air is allowed to flow.

Further, each of the heat-operated flaps 22 receives the heat directly transferred from the combustion liner 11 or the heat transfer from the combustion gas or the like to generate a total or partial heat displacement, thereby forming an annular shape. The upstream opening of the cooling flow path 15 is rapidly opened from the state shown in FIG. 1 (a) to the state shown in FIG. 1 (b), and Fe-N specified in JIS C-2530 is used.
A joint of an i-Mn alloy and a Fe-Ni alloy, or F
A high temperature bimetal such as one obtained by joining an e-Ni-CR alloy and an Fe-Ni alloy is applied.

Further, the heat actuating flap 22 can be made of a shape memory alloy, and in that case, it is operated in a plurality of temperature regions including at least a low temperature transformation phase and a high temperature transformation phase. Is set, for example Ti-
When a leaf spring shape made of Ni alloy is used and attached to the combustion liner 11 and transformed into a high temperature memory shape, most of the annular cooling flow path 15 is opened and transformed into a low temperature memory shape. , The annular cooling flow path 15 is set to be closed.

In the cooling fluid control device configured as described above, the bypass flow 13 and the core flow 14 are inserted depending on the flight mode, and when attention is paid to the portion of the bypass flow 13, most of it is the combustion liner. It is sent inside 11 and a part is sent into the annular cooling flow passage 15 in accordance with the operating state of the thermal shutter 21 to cool the combustion duct 10 and the combustion liner 11.

When the flight mode is in the low Mach state (under load), the combustion liner 11 is brought into a relatively low temperature state, and the thermal operation shutter 2 is operated from the combustion liner 11 and the inside thereof.
1 has a small amount of heat transfer to each heat-actuated flap 22 and the heat-actuated shutter 21 closes most (or a part) of the annular cooling flow passage 15 so that the annular cooling flow passage 1
5, the amount of air introduced into the interior of the annular cooling passage 15 is suppressed.
By limiting the amount of air inside the chamber, the occurrence of pressure loss is reduced. In addition, the performance of the aircraft engine is improved by increasing the amount of air inside the combustion liner 11 by the amount that the amount of air introduced into the annular cooling flow passage 15 is suppressed.

When the flight mode is switched to the high Mach state (high load), fuel is injected into the combustion liner 11 and burned, and as described above with reference to FIG. The temperature of the internal atmosphere of the combustion liner 11 rapidly rises with time.

When the temperature rises rapidly, the radiant heat from the combustion gas, the conductive heat directly passing through the combustion liner 11, etc. heat the entire heat-operated flaps 22 and the heat-operated portions,
The thermal displacement indicated by the solid line from the chain double-dashed line in FIG. 1B is generated in each heat-actuated flap 22 to open the annular cooling passage 15, so that the amount of air introduced into the annular cooling passage 15 is 1
It becomes larger as shown by the arrows in (a) and (b). At this time, the amount of air sent into the annular cooling flow path 15 is set on the basis of high load, and the combustion liner 11 is cooled with an appropriate amount of air to ensure the soundness of the aircraft engine at high load. Be done.

[Other Embodiments] The cooling fluid control device according to the present invention can employ the following techniques. a) When the heat-actuated shutter 21 is divided into a plurality of pieces in the circumferential direction and the divided portions are in an overlapped state, slits or holes are formed in the heat-actuated flap 22 in order to obtain the minimum necessary air insertion amount at low load. To form. b) In the above case, the tip of the thermally actuated flap 22 should be in intimate contact with the inner surface of the combustion duct 10. c) A slit is formed at a divided portion of the heat-actuated flap 22 to ensure a necessary minimum air insertion amount. d) In the case where the plurality of thermally actuated flaps 22 are formed of a shape memory alloy or the like, the thermal displacement start temperature is set to a plurality of stages. e) Adopt a structure in which a portion of the thermally actuated flap 22 necessary for thermal displacement is directly attached to the combustion liner 11 and the thermally actuated flap 22 drives the opening / closing portion.

[0021]

According to the cooling fluid control device for an aircraft engine of the present invention, the following excellent effects are exhibited. (1) At the upstream end of the combustion liner, a heat-actuated shutter that changes the amount of air introduced into the annular cooling passage of the bypass flow by heat displacement due to heat transfer is arranged, so that the annular cooling flow at low load is provided. The amount of air introduced into the passage can be reduced, energy loss in the annular cooling passage can be suppressed, and performance can be improved especially when the load is low. (2) By adjusting the opening degree of the annular cooling flow path based on the thermal displacement, the air flow can be controlled according to the load of the aircraft engine, and the engine performance at all Mach can be improved. (3) The thermal actuating shutter has a structure in which the thermal actuating shutter is divided into a plurality of parts in the circumferential direction, whereby each thermal displacement can be facilitated and the accuracy at the time of control can be improved. (4) By mounting the heat-actuated shutter integrally with the combustion liner, it is possible to quickly perform thermal displacement when the temperature changes, and to secure the required amount of air introduction. (5) The application of bimetals and shape memory alloys can facilitate the implementation and enhance applicability in afterburners for aircraft engines and ram combustors.

[Brief description of drawings]

FIG. 1 is a front sectional view of essential parts showing an embodiment of an aircraft engine cooling fluid control device according to the present invention.

FIG. 2 is a side view, omitting a part of the description, showing one embodiment of the cooling fluid control device for an aircraft engine according to the present invention.

FIG. 3 is a front cross-sectional view showing a structural example of an aircraft engine (gas turbine engine) having an afterburner.

FIG. 4 is a model diagram of a temperature rise when the afterburner in FIG. 3 operates.

[Explanation of symbols]

 1 Aircraft Engine 2 Fan 3 Compressor 4 Combustor 5 Turbine 6 Afterburner 7 Flame Retainer 8 Fuel Nozzle 9 Spark Plug 10 Combustion Duct 11 Combustion Liner 12 Exhaust Nozzle 13 Bypass Flow (Fan Flow) 14 Core Flow 15 Annular Cooling Flow Path 21 Thermal shutter 22 Thermal flap 23 Fastener G Gap

Claims (5)

[Claims] 1. An annular cooling flow path (15) formed between a combustion duct (10) and a combustion liner (11) inside thereof.
In the aircraft engine (1), a part of the bypass flow (13) is sent as cooling air, and when the fuel is injected into the combustion liner, the bypass flow and the core flow air are used for combustion. A heat-actuated shutter (21) is arranged at the upstream end of the combustion liner to change the amount of air introduced into the annular cooling passage of the bypass flow by thermal displacement based on heat transfer from the vicinity of the combustion liner. Aircraft engine cooling fluid control device. 2. The cooling fluid control device for an aircraft engine according to claim 1, wherein the heat-actuated shutter (21) is divided into a plurality of pieces in the circumferential direction. 3. A cooling fluid control system for an aircraft engine according to claim 1, wherein the heat-actuated shutter (21) is integrally attached to the combustion liner (11). 4. The heat-actuated shutter (21) is composed of bimetal.
A cooling fluid control device for an aircraft engine as described. 5. The aircraft engine according to claim 2, wherein the heat-actuated shutter (21) is a shape memory alloy that operates in a plurality of temperature ranges including at least a low temperature transformation phase and a high temperature transformation phase. Cooling fluid control device.