TURBINE CASING COOLING DEVICE FOR A TURBOMACHINE

Abstract

The invention relates to a device (9) for cooling a turbine casing (7) for a turbomachine, such as for example an aircraft turbojet engine, extending around an axis (X) and comprising air-distribution means configured to take in air and convey it to the casing, characterized in that the air-distribution means comprising at least a first ramp (20a, 20b) and a second ramp (20a, 20b) extending circumferentially about the axis (X) respectively on a first circumferential portion and on a second circumferential portion which are different from each other, each ramp (20a, 20b) comprising air ejection orifices intended to be directed towards the casing in order to cool it, characterized in that it comprises adjustment means (23) capable of adjusting the flow rate of air ejected at the level of the first ramp (20a, 20b) with respect to the flow rate of air ejected at the level of the second ramp (20a, 20b).

Claims

1. A method for managing a cooling device for a turbine casing of an aircraft engine extending along an axis (X), the aircraft engine having a fan downstream of which extends a compressor, a combustion chamber and a turbine comprising a turbine casing, the aircraft engine further having: a primary vein in which a primary flow (F1) circulates, said primary vein passing through in a direction of circulation of the primary flow (F1), the compressor, the combustion chamber and the turbine comprising the turbine casing, a secondary vein into which flows a secondary flow from the fan, wherein the secondary flow is distinct from the primary flow, the secondary vein extending around the primary vein, wherein the aircraft engine further includes a device for cooling the turbine casing, said cooling device having air-distribution means configured to take in air and convey it to the turbine casing, the air-distribution means having at least a first ramp and a second ramp extending circumferentially about the axis (X) respectively on a first circumferential portion and on a second circumferential portion different from each other, each of the first ramp and second ramp comprising air ejection orifices configured to be directed towards the turbine casing in order to cool it, the cooling device further having adjustment means capable of adjusting an air flow rate of air ejected at the first ramp with respect to an air flow rate of air ejected at the second ramp, each of the first ramp and second ramp extending circumferentially about the axis (X) and being located radially outside the turbine casing, the air ejection orifices of the ramps first ramp and second ramp being turned towards the said turbine casing, the method comprising: logging along two different axes accelerations experienced by at least one of an airframe or a propulsion system of an aircraft; determining an eccentricity of a rotor from the logged accelerations; and adjusting the adjustment means so as to adapt the air flow rate ejected at the first ramp to the air flow rate ejected at the second ramp.

2. The method according to claim 1, wherein the adjustment means comprise means for reducing an air flow cross-section arranged in the air-distribution means.

3. The method according to claim 1 wherein the cooling device comprises at least a first collector and a second collector connected to the air-distribution means via a first branch and a second branch respectively, each extending circumferentially in an opposite direction, the first collector and the second collector extending circumferentially from the first collector and the second collector respectively.

4. The method according to claim 3, wherein the means for reducing the air flow cross-section are able to adjust an air flow cross-section of at least one of the first branch or the second branch.

5. The method according to claim 4, wherein the means for reducing the air flow cross-section are able to adjust at least one of the air flow cross-section of the first ramp or the air flow cross-section of the second ramp, at a junction area between said first ramp or second ramp and the first collector or second collector, respectively.

6. The method according to claim 5, wherein the means for reducing the air flow cross-section comprises at least one adjustable damper.

7. The method according to claim 6, wherein said adjustable damper is an iris damper.

Description

BRIEF DESCRIPTION OF THE FIGURES

[0037] FIG. 1 is an axial sectional view of a portion of a turbofan according to one embodiment of the invention;

[0038] FIG. 2 is a cross-sectional view of the cooling device according to the invention,

[0039] FIG. 3 is a perspective view of an iris damper,

[0040] FIG. 4 is a perspective view, in detail, illustrating another embodiment of the device according to the invention.

DETAILED DESCRIPTION

[0041] FIG. 1 illustrates a part of a turbofan engine according to the invention, in particular the low-pressure turbine 1. This comprises a rotor with wheels 2 axially connected to each other by annular flanges 3 and each comprising a disc 4 bearing blades 5.

[0042] Annular rows of stationary vanes 6 are mounted by suitable means at their radially outer ends on a case 7 of the low-pressure turbine 1 between the mobile wheels 2. The fixed blades 6 of each row are joined together at their radially inner ends by annular sectors placed circumferentially end to end.

[0043] As previously mentioned, the primary air flow F1 from the combustion chamber into the primary air vein 8 heats the casing 7.

[0044] In order to ensure the cooling of the casing 7, the turbojet engine has a cooling device 9, best seen in FIG. 2.

[0045] The latter includes air intake and supply means comprising: [0046] a scoop 10 comprising an opening 11 leading, for example, into the secondary vein of the turbojet engine in order to take cold air from it, [0047] a connecting member 11 having a general Y shape comprising an upstream part 12 connected to the scoop 10, and a downstream part comprising a first branch 13 whose function will not be detailed here, and a second branch 14, [0048] a control valve 15 connected downstream of the second branch 14 and capable of being controlled as a function of engine speed and/or flight conditions, for example, so as to adjust the flow rate taken in, [0049] a distribution member 16 formed from one or more parts and comprising an upstream part 17 connected to the outlet of the control valve 15, and at least two downstream branches 18a and 18b extending circumferentially around the axis of the turbojet engine, on either side of the downstream end of the upstream part 17. Each branch 18a, 18b extends for example about 90°.

[0050] The device 9 further comprises collectors or connecting areas 19a, 19b, here two in number, connected to the corresponding ends of the branches 18a, 18b, each collector 19a, 19b forming an axially extending channel.

[0051] Of course, the number of collectors 19a, 19b may vary, and may be four, for example. The cooling flow through the air intake and supply means is illustrated by arrows in FIG. 2 (referenced F at the scoop 10 and F′, F″ at the branches 18a, 18b respectively). The device 9 further comprises ramps 20a, 20b (or more generally noted 20) formed by curved pipes of circular cross-section, each ramp 20a, 20b extends circumferentially at an angle of about 90°, more precisely of the order of 90° here.

[0052] Each ramp 20 has a proximal end 21 opening into the corresponding collector channel 19 and a closed distal end 22. Each ramp 20 also has orifices facing the casing 7 so that the air taken in through the scoop 10, the member 11, the valve 15 and the distribution member 16 enters the collectors 19a, 19b and then the ramps 20 before emerging through the orifices facing the casing 7, so as to cool it.

[0053] The two collectors 19a, 19b are diametrically opposed, each collector 19a, 19b being associated with a plurality of pairs of ramps 20, namely ramps 20a extending circumferentially on one side and ramps 20b extending circumferentially on the opposite side. Thus, each collector 19a, 19b and the associated opposing ramps 20a, 20b cover an angular range of approximately 180°. In the embodiment shown in the figures, each collector 19a, 19b is associated with several pairs of ramps, for example nine pairs of ramps 20a, 20b. The ramps 20a, 20b of the same pair are located on the same radial plane, the ramps 20a, 20b of different pairs being offset from each other along the X axis of the turbomachine, as seen in FIG. 2.

[0054] The two collectors 19a, 19b and the associated pairs of ramps 20 have substantially identical structures and are arranged diametrically opposite each other.

[0055] In this way, the ramps 20 are located on several radial planes axially offset from each other, the ramps 20 of the same radial plane forming a cooling ring surrounding the casing 7 which extends substantially over the entire periphery of the casing, i.e. substantially 360°.

[0056] At least one of the branches 18a, 18b, or each of said banks 18a, 18b may include means for adjusting the cross-section of the corresponding branch, for example in the form of an iris damper 23, the structure of which is shown in FIG. 3. Such a damper 23 comprises a fixing ring 24 and movable blades 25 whose position can be changed so as to adjust the internal diameter d of the damper 23, delimited by the blades 25.

[0057] The position of such a damper 23 in each of the branches 18a, 18b is schematically shown in FIG. 2. Each damper 23 may be located near the end of the corresponding branch 18a, 18b connected to the upstream portion 17, or near the end of the corresponding branch 18a, 18b connected to the corresponding collector 19a, 19b, for example.

[0058] Each damper 23 may also be located in a middle region of the corresponding branch 18a, 18b.

[0059] According to another embodiment illustrated schematically in FIG. 4, means for adjusting the cross-section 23 may be provided on the ramps 20a, 20b, at their ends 21 connected to the corresponding collector 19a, 19b. As before, such vein adjustment means may be formed by iris dampers 23.

[0060] Such adjusting means 23 allow to adapt the air flow rate ejected at the ramps 20 connected to the collector 19a, with respect to the air flow rate ejected at the ramps 20 connected to the collector 19b. It is therefore possible to cool two circumferential zones of the housing 7 in a differentiated manner from each other. This allows, for example, the casing 7 to contract locally in a selected circumferential area in order to limit the clearance between the rotor blades 26 (FIG. 1) and the abradable rings 27 supported by the casing 7, for example.

[0061] The overall flow rate can also be adjusted by means of the control valve 15.

[0062] The embodiment shown in FIG. 4 allows the wear compensation to be controlled axially. Indeed, for the same load, it appears that the eccentricity of the rotor (and therefore the wear generated) is not axially uniform. By arranging adjustment means downstream of the collector 19a, 19b so as to be able to adjust the flow of cooling air individually in each of the ramps 20, it is possible to compensate for the clearance as accurately as possible and as required, and thus to limit the consumption of cooling flow which penalises the performance of the turbomachine.

[0063] In operation, in the case of a turbomachine with low wear of the abradable rings 27, the cross-sections of the branches 18a, 18b or ramps 20 are reduced or flanged using the aforementioned cross-section adjustment means 23. As wear and tear is observed and a clearance appears locally between the ends of the rotor blades 26 and the corresponding abradable rings 27, some of the ramps 20 may be supplied with a higher air flow rate, by increasing the cross-section of the ramps 20 or the branches 18a, 18b concerned. The increase in the flow of cooling air causes a local contraction of the casing 7 and locally reduces the aforementioned clearance, which makes it possible to restore the performance of the turbomachine close to the nominal performance when the turbomachine is new.

[0064] The clearance check can be performed at the time of a maintenance period, for example every 500 to 1000 flight hours. The cross-sections of the branches 20 or ramps 18a, 18b can then be adapted accordingly.

[0065] The aforementioned local clearance may be calculated on the basis of rotor eccentricity measurements obtained with specific accelerometers or clearance sensors, located at the airframe and/or propulsion system of the aircraft.