Cooling system for hot parts of an aircraft engine, and aircraft engine equipped with such a cooling system

Abstract

The cooling system (30) for an aircraft engine (10) comprises a channel (32) that draws off cold air in the secondary air flow (200), and a heat exchanger (34) located in the channel (32) and in which hot air circulates. The channel (32) comprises: a supply pipe (322) and an evacuation pipe (326) fixed to the nacelle (324), an intermediate box (324) located between the supply pipe (322) and the evacuation pipe (326), fixed to the engine (10), and in which the heat exchanger (34) is placed. Application to cooling of hot parts (22) in an aircraft engine (10).

Description

TECHNICAL DOMAIN

[0001] This invention relates to the technical field of cooling systems for hot parts of aircraft engines.

[0002] More particularly it relates to a cooling system comprising a set of heat exchangers, to cool hot parts of an aircraft engine such as high pressure turbine blades in this aircraft engine.

[0003] It also relates to an aircraft engine equipped with such a cooling system.

STATE OF PRIOR ART

[0004] It is known that heat exchangers can be installed in an aircraft engine to cool hot parts of an aircraft engine.

[0005] Document FR 2 400 618 discloses a turbo-fan type of aircraft engine using an air/air type cooling system, and an associated cooling method. Hot parts such as fixed and mobile blades of the high pressure turbine are cooled by cooling air originating from part of the primary air drawn off at the outlet from the compressor or between compressor stages. This air cooling hot parts is itself firstly cooled before passing over the parts to be cooled, by passing inside pipes in a heat exchanger itself installed in a colder air current. This colder air current originates from part of the fan dilution air, or secondary air. It is drawn off from the fan duct, and more precisely in an annular flow passage delimited on one side by the gas generator and on the other side by a casing that surrounds part of the length of the gas generator. This part of the dilution air drawn off through the casing enters a diffuser section in which the dynamic air pressure is largely recovered, and is then transferred through the heat exchanger where it absorbs heat from cooling air drawn off from the compressor. Then once it has cooled the air that cools hot parts, this past of the dilution air is returned to the fan duct, its static pressure being adjusted to the static pressure existing in the fan duct.

[0006] Document U.S. Pat. No. 5,269,135 discloses an air/fuel type cooling system that comprises at least one heat exchanger in which fuel circulates. This heat exchanger is located in a stream in which colder air drawn off on the upstream side in the fan duct circulates and is restored on the downstream side in the fan duct. The channel is partly delimited by the inner wall of the fan duct itself.

[0007] The cooling systems that have just been described have a number of disadvantages.

[0008] A first disadvantage is related to engine maintenance, particularly of the gas generator. If there is no cooling system, maintenance is done by opening a cowling of the nacelle to access the engine directly, and particularly the fuel injectors. In the presence of a cooling system according to prior art, and particularly if the casing is in position around the gas generator or a channel fixed to the nacelle, it becomes difficult to access some parts of the engine for maintenance.

[0009] A second disadvantage of systems according to prior art is due to the fact that colder air used to cool fluid circulating in the heat exchanger is restored on the downstream side inside the fan duct. Therefore its pressure is adjusted to the pressure inside the fan duct at this location. There is then a risk that air could circulate in the inverse direction or not circulate at all, which would make the heat exchanger inoperative.

SUMMARY OF THE INVENTION

[0010] This invention is intended to provide a solution to the disadvantages of the systems of prior art.

[0011] According to a first aspect of the invention, the cooling system for hot parts of an aircraft engine is applicable to an aircraft engine housed in a nacelle, a primary air flow passing inside the engine and a secondary air flow passing around the engine inside the nacelle. The cooling system comprises at least one channel that draws off cold air in the secondary air flow and at least one heat exchanger located in the channel and in which hot air from the primary air flow circulates to be cooled before reaching hot parts to cool them. Said at least, one channel comprises the following three parts:

[0012] a supply pipe on the upstream side of said at least one heat exchanger, said supply pipe being fixed to the nacelle,

[0013] an evacuation pipe located on the downstream side of said at least one heat exchanger, said evacuation pipe being fixed to the nacelle,

[0014] an intermediate box located between the supply pipe and the evacuation pipe in which said at least one heat exchanger is located, said intermediate box being fixed to the engine.

[0015] Advantageously, the intermediate box has a longitudinal section with an approximately rectangular shaped profile.

[0016] Preferably, the intermediate box has a longitudinal section with a profile approximately in the shape of a trapezium, in which the large base is facing the engine and the small base is facing the nacelle.

[0017] Preferably, the cooling system also comprises an upstream seal between the supply pipe and the intermediate box and a downstream seal between the intermediate box and the evacuation pipe.

[0018] According to the invention, each heat exchanger is associated with:

[0019] at least one inlet duct, that draws in air from the hot primary air flow, and carries it into the heat exchanger to cool it, and

[0020] at least one return duct, that collects cooled air in the heat exchanger and returns it to hot parts of the engine to cool them.

[0021] Preferably, each inlet duct comprises an attachment flange at one of its ends for fixing it onto the engine. Similarly, each return duct comprises an attachment flange at one of its ends for fixing it to the engine. All these attachment flanges make a mechanical attachment of the box onto the engine.

[0022] According to one variant, at least one of the inlet ducts is provided with a valve.

[0023] According to another variant, at least one of the return ducts is provided with a valve.

[0024] According to one particular variant, each heat exchanger is associated with four inlet ducts and four return ducts.

[0025] According to the invention, the outlet from the evacuation pipe of each channel opens up at the outlet from the nacelle ejection nozzle, or beyond it on the downstream side. Thus, air that exits from the evacuation pipe is at atmospheric pressure. Consequently, the risk of recirculation of air in the heat exchanger is eliminated, and pressure losses in the heat exchanger are increased.

[0026] According to one preferred variant, the outlet section of the supply pipe of each channel is larger than its inlet section.

[0027] According to another preferred variant, the outlet section of the evacuation pipe of each channel is smaller than its inlet section.

[0028] Preferably, the cooling system comprises at least two heat exchangers distributed circumferentially around the engine, each heat exchanger being placed in a separate channel. Even more preferably, there are four heat exchangers.

[0029] According to another variant, the cooling system comprises a single heat exchanger extending around the entire circumference of the engine and is placed in a single corresponding annular channel.

[0030] According to a second aspect, the invention relates to an aircraft engine equipped with a cooling system according to the first aspect of the invention.

BRIEF DESCRIPTION OF THE FIGURES

[0031] The invention will be better understood after reading the detailed description given below of particular embodiments of the invention, provided for illustrative purposes and in no way limitative, with reference to the appended figures, wherein:

[0032] FIG. 1 shows a diagrammatic sectional view through an aircraft engine comprising a cooling system according to the invention;

[0033] FIG. 2 is an aft perspective view showing the aft side of an aircraft engine comprising a cooling system according to the invention; and

[0034] FIG. 3 shows another aft perspective view of the engine, an aft part having been removed at the location of the cooling system, showing a section through the layout of the heat exchangers.

DETAILED DESCRIPTION OF PARTICULAR EMBODIMENTS

[0035] FIG. 1 diagrammatically shows an aircraft engine 10 with an axis of revolution 12.

[0036] In a manner known per se, the aircraft engine 10 comprises low pressure compressor stages 14, medium pressure compressor stages 16, high pressure compressor stages 18, a combustion chamber 20 and turbines 22. The aircraft engine 10 is surrounded by a nacelle 24 that terminates with an ejection nozzle 26.

[0037] In a manner known per se, a primary air flow represented by arrows 100 circulates inside the engine 10. It is heated as it passes through the compressor stages 14, 16 and 18 before arriving at the turbines 22.

[0038] In a manner known per se, a secondary air flow represented by the arrows 200 circulates in the nacelle 24 around the engine 10. This secondary air 200 external to the engine 10 is colder than the primary air 100 inside the engine 10.

[0039] According to the invention, the engine 10 comprises a cooling system 30 designed to cool hot parts of the engine 10, for example such as the turbine blades 22. The principle of this cooling system consists of cooling the air that then flows on or in hot parts to be cooled.

[0040] The cooling system 30 comprises at least one channel 32 inside which air circulates (arrows 300, 400) drawn off from the secondary air flow of the nacelle 24 and at least one heat exchanger 34 placed in this channel 32. For example, this heat exchanger may for example be a tube exchanger or a plate exchanger.

[0041] Each channel 32 comprises three successive parts:

[0042] a supply pipe 322 fixed to the nacelle 24 and located on the upstream side with respect to the air flow direction (arrows 100, 200, 300, 400),

[0043] an evacuation pipe 326, fixed to the nacelle 24 and located on the downstream side with respect to the air flow direction (arrows 100, 200, 300, 400),

[0044] an intermediate box 324 fixed to the engine 10 and located between the supply pipe 322 and the evacuation pipe 326.

[0045] The junction 342 between the supply pipe 322 and the intermediate box 324 is made by assuring continuity between the outlet section 323 of the supply pipe 322 and the inlet section 328 of the intermediate box 324, which have approximately the same dimensions for this purpose. This junction 342 comprises a seal (not shown) on the upstream side that fits to said sections.

[0046] Similarly, the junction 346 between the intermediate box 324 and the evacuation pipe 326 is made by assuring continuity between the outlet section 329 of the intermediate box 324 and the inlet section 325 of the evacuation pipe 326, which have approximately the same dimensions for this purpose. This junction 346 comprises a seal (not shown) on the upstream side that fits to said sections.

[0047] Preferably, the end sections of the intermediate box 324 at the corresponding junction 342, 346, have a profile approximately in the shape of a trapezium when viewing the longitudinal section, for which the large base is facing the engine 10 and the small base is facing the nacelle 24.

[0048] The inlet 321 to the channel 32 at the inlet to the supply pipe 322, may be a static air inlet or a dynamic air inlet.

[0049] The outlet 327 from the channel 32 that is the outlet from the evacuation pipe 326, is arranged so that it coincides approximately with the free end of the ejection nozzle from the nacelle 24. Thus, air evacuated through the evacuation pipe 326 is at atmospheric pressure.

[0050] The heat exchanger 34 is located in the intermediate box 324 of the channel 32.

[0051] This intermediate box 324 is facing the engine 10 at a certain distance from it. It is connected to the engine 10 using at least one inlet duct 42 that is used to bring in (arrow 420) air drawn off at the outlet from the compressor 14, 16, 18 to the heat exchanger 34, so that this air is cooled, and at least one return duct 44 that is used to transfer this cooled air to the turbines 22 (arrow 440).

[0052] FIGS. 2 and 3 illustrate an example cooling system 30 according to the invention in more detail, including four channels 32 distributed around the periphery of the nacelle 24. To simplify the figures, the engine 10 is not shown in these figures. The arrows 200 indicate the secondary air flow direction, and therefore indicate the upstream and downstream sides of the channels 32.

[0053] FIG. 2 illustrates an example cooling system according to the invention more particularly showing the channels 32 and their supply pipes 322 and evacuation pipes 326.

[0054] According to the illustrated embodiment of the cooling system, the inlet section 321 of the supply pipes 322 is smaller than their outlet section 323. Thus, air (arrows 300) originating from the nacelle 24 is cooled in supply pipes 322 before reaching the heat exchangers 34 where it is heated on contact with them. Similarly, the inlet section 325 of the evacuation pipes 326 is identical to their outlet section 327. Thus, air (arrows 400) that was heated in contact with the heat exchangers 34 is not accelerated in the evacuation pipes, before exiting towards the outside, so as to limit pressure losses.

[0055] FIG. 3 only shows the supply pipes 322 and the intermediate boxes 324 of the channels 32.

[0056] This figure shows intermediate boxes 324 in which the heat exchangers 34 are located in more detail, together with inlet ducts 42 and return ducts 44. In the example shown, there are four inlet ducts 42 per intermediate box 324 arranged on the upstream side of the intermediate boxes 324. Similarly, there are four return ducts per intermediate box 324 located on the downstream side of the intermediate boxes 324. The inlet ducts 42 and the return ducts 44 terminate on the engine side with attachment flanges 43 that are provided for fixing together the intermediate boxes 324 to said engine 10.

[0057] Thus, the inlet ducts 42 and the return ducts 44 are also used as attachment means for the intermediate boxes 324, and therefore for the associated heat exchangers 34 on the engine 10.

[0058] The inlet ducts 42 are fixed on each intermediate box 324 at a distributor 46 that is used to supply all tubes or all plates of the heat exchanger(s) 34 located in the intermediate box 324. Air that has just been cooled in contact with the heat exchanger(s) 34 is collected by a header 48 on which the return ducts 44 are fixed.

[0059] In the example illustrated in the figures, the inlet ducts 42 are approximately straight and draw air in directly at the exit from the compressor 14, 16, 18 to bring it to the heat exchangers 34. The return pipes 44 are bent so as to return air that passed through the heat exchangers 34, to a point further downstream at the inlet to turbines 22.

[0060] The cooling system that has just been described has a number of advantages.

[0061] A first advantage is related to the maintenance of some engine parts 10, for example such as fuel injectors (not shown). According to the invention, the supply pipe 322 and the evacuation pipe 326 for each channel 32 are fixed to the nacelle 24, while the intermediate box 324 is fixed to the engine 10. Subsequently, it is preferable to arrange the supply pipe 322 and the evacuation pipe 326 such that they are fixed to an opening cowling (not shown) of the nacelle 24. Thus, when the cowling is opened, these two pipes 322, 326 are lifted at the same time as the cowling, while the intermediate box 324 containing one or several heat exchangers 34 remains fixed to the engine 10. Therefore, it is easy for an operator to access the engine even in the presence of a channel 32, due to the fact that the channel is made from three separate parts 322, 324, 326.

[0062] Since the intermediate box 324 is also held at a distance from the engine 10 due to the presence of the inlet ducts 42 and the return ducts 44, it becomes easy for an operator to access parts of the engine, even in an area located under the intermediate box 324 itself, by passing between the inlet ducts 42 and the return ducts 44. Consequently, all that is necessary to obtain easy access to fuel injectors for maintenance, is to put the intermediate box 324 into position facing the combustion chamber 20.

[0063] Another advantage is related to the trapezoidal shape of the profile of the intermediate box 324. The result of this shape is that when the cowling of the nacelle 24 is closed, the outlet section 323 of the supply pipe, 322 of each channel 32 covers the inlet section 328 of the intermediate box 324. Similarly, the inlet section 325 of the evacuation pipe 326 covers the outlet section 329 of the intermediate box 324. This arrangement makes it possible to make the supply pipe 322 coincide well with the evacuation pipe 326, and with the intermediate box 324 when the cowling is closed, and therefore assure a good seal of the channel 32 at junctions 342 and 346. This seal can be further improved by the presence of seals around the periphery of junctions 342, 346.

[0064] The cooling system that has just been described is a passive cooling system, in other words the air flow drawn off at the high pressure compressor 18 (arrow 420) is proportional to the cooling air flow in the high pressure turbine 22 (arrows 440). Without departing from the scope of the invention, it would be possible to envisage an active cooling system that includes at least one valve (not shown) placed on the inlet ducts 42 or on the return ducts 44. By controlling opening and/or closing of these valves, it would then be possible to improve engine flight performances at the detriment of an increase in its mass, as a function of the different flight phases.

[0065] In the cooling system that has just been described, each heat exchanger is associated with four inlet ducts and four outlet ducts. Without departing from the scope of the invention, it would be possible that the number of inlet ducts and the number of outlet ducts is different from four, and/or different from each other.

[0066] In the cooling system that has just been described, the outlet from the channel coincides with the free end of the nacelle ejection nozzle. It would be possible to envisage the channel going beyond the free end of the ejection nozzle, in the downstream direction.

Claims

1. Cooling system (30) for hot parts (22) of an aircraft engine (10), the aircraft engine (10) being housed in a nacelle (24), a primary air flow (100) passing inside the engine and a secondary air flow (200) passing around the engine (10) inside the nacelle (24), cooling system (30) characterised in that it comprises at least one channel (32) that draws off cold air (300) in the secondary air flow (200) and at least one heat exchanger (34) located in the channel (32) and in which hot air (420) from the primary air flow (100) circulates to be cooled before reaching (440) hot parts (22) to cool them, and in that said at least one channel (32) comprises the following three parts: a supply pipe (322) on the upstream side of said at least one heat exchanger (34), said supply pipe (322) being fixed to the nacelle (24), an evacuation pipe (326) located on the downstream side of said at least one heat exchanger (34), said evacuation pipe (326) being fixed to the nacelle (24), an intermediate box (324) located between the supply pipe (322) and the evacuation pipe (326) in which said at least one heat exchanger (34) is located, said intermediate box (324) being fixed to the engine (10).

2. Cooling system (30) according to claim 1, characterised in that the intermediate box (324) has a longitudinal section with a profile approximately in the shape of a trapezium, in which the large base is facing the engine (10) and the small base is facing the nacelle (24).

3. Cooling system (30) according to claim 1, characterised in that it also comprises an upstream seal between the supply pipe (322) and the intermediate box (324) and a downstream seal between the intermediate box (324) and the evacuation pipe (326).

4. Cooling system (30) according to claim 1, characterised in that each intermediate box (324) is associated with: at least one inlet duct (42), that draws in air (420) from the hot primary air flow (100), and carries it into the heat exchanger (34) to cool it, and at least one return duct (44), that collects cooled air (440) in the heat exchanger (34) and returns to hot parts (22) of the engine (10) to cool them.

5. Cooling system (30) according to claim 4, characterised in that each inlet duct (42) and each return duct (44) comprises an attachment flange (43) at one of its ends for fixing it onto the engine (10).

6. Cooling system (30) according to claim 5, characterised in that at least one of the inlet ducts (42) is provided with a valve.

7. Cooling system (30) according to claim 5, characterised in that at least one of the return ducts (44) is provided with a valve.

8. Cooling system (30) according to claim 5, characterised in that each intermediate box (324) is associated with four inlet ducts (42) and four return ducts (44).

9. Cooling system (30) according to claim 1, characterised in that the outlet (327) from the evacuation pipe (326) of each channel (32) opens up at the outlet from the nacelle (24) ejection nozzle (26), or beyond it on the downstream side.

10. Cooling system (30) according to claim 1, characterised in that the outlet section (323) of the supply pipe (322) of each channel (32) is larger than its inlet section (321).

11. Cooling system (30) according to claim 1, characterised in that the outlet section (327) of the evacuation pipe (326) of each channel (32) is identical to its inlet section (325).

12. Cooling system (30) according to claim 1, characterised in that it comprises at least two heat exchangers (34) distributed circumferentially around the engine (10), and in that each heat exchanger (34) is placed in a separate channel (32).

13. Cooling system (30) according to claim 1, characterised in that it comprises a single heat exchanger (34) extending around the entire circumference of the engine (10) and in that this heat exchanger (34) is placed in a single corresponding annular channel (32).

14. Aircraft engine (10), characterised in that it is equipped with a cooling system (30) according to any one of claims 1 to 13.