CONTROL OF THE FLUAGE OF AN AIRCRAFT TURBOMACHINE BLADE

Abstract

A method for controlling the creep of a turbine blade (12, 16), particularly for aircraft, the blade (12, 16) being made of a metallic alloy and comprising at least one platform connected to at least one blade (22, 32), the method comprising the following steps: a) determining at least one geometric control parameter of the blade (12, 16), this parameter having a value that may change according to the creep deformations of the blade, b) determining an acceptability threshold for the parameter(s), c) verifying in situ the conformity of the parameter(s) with respect to the corresponding threshold, and d) concluding on the conformity of the blade (12, 16) based on the results of the verification. Figure for the abstract: Figure 5

Description

Title of the invention: CONTROL OF THE CREEP OF AN AIRCRAFT TURBOMACHINE BLADE Technical field of the invention

[0001] The present invention relates to the general field of aeronautics. It relates more particularly to a method for controlling the finishing of a turbomachine blade, in particular an aircraft blade. Technical background

[0002] A turbomachine is known comprising a turbine 10 as shown in [Fig.l].

[0003] The turbine 10 may be a low-pressure turbine arranged downstream of a combustion chamber of the turbomachine and configured to drive a fan shaft of the turbomachine by expanding the gases leaving the combustion chamber.

[0004] The turbine 10 comprises a plurality of stator blades 12 carried by a fixed casing 14, and a plurality of rotor blades 16 carried by rotor disks 18 secured to each other and rotatable about a longitudinal axis A of the turbomachine.

[0005] The rotor discs 12 are connected to a low pressure shaft 20 which is configured to drive the fan shaft, directly or via a reducer for example.

[0006] A rotor blade 16 is shown in [Fig.2] and stator blades 12 are shown in [Fig.3].

[0007] The rotor blade 16 comprises for example a blade 22 extending between two platforms, respectively internal 24 and external 26 (or radially internal and external with reference to the axis A). The external platform 26 is called a heel, and the internal platform 24 makes it possible to connect the blade 22 to a root 28.

[0008] The platforms 24, 26 define between them a part of the gas flow path in the turbine 10. The feet 28 allow the rotor blades 16 to be mounted on the rotor discs 18.

[0009] The heels 26 of the rotor blades 16 have wipers 30 configured to cooperate with abradable linings carried by the casing 14 opposite the rotor blades 16 to ensure sealing in the turbine 10.

[0010] The stator vanes 12 form a turbine distributor which has an annular shape and which is sectorized. Each distributor sector comprises two platforms, respectively internal 34 and external 36, between which extend several vanes 12 or blades 32. [Fig. 3] shows a distributor sector comprising three vanes 12 and so three blades 32.

[0011] The rotor and stator blades 12, 16 are subjected to high stresses partly due to the temperatures to which they are subjected during operation, in particular above 800°C. This “threshold” depends on the material and the loading. At low loading, this threshold will be higher to detect the problem of fining over the long term with respect to the engine operating time.

[0012] Consequently, these blades 12, 16 may exhibit deformations due to finning which it is important to monitor.

[0013] The deformation of a part corresponds to the deformation over time of a material subjected to a constant stress and a given temperature. This deformation can lead to the rupture of the part.

[0014] The finishing is characterized by three main states illustrated in [Fig.4]:

[0015] (I) Primary creep: at the initial stage, the deformation rate e is relatively high, but decreases with increasing time and strain due to the material experiencing an increase in creep strength or strain hardening,

[0016] (II) Secondary creep: sometimes called steady-state creep, the strain rate e is constant, i.e. the plot becomes almost linear. The strain rate becomes almost constant at the beginning of the secondary stage. This is due to the balance between work hardening and annealing (thermal softening). This creep stage is the best understood. Steady-state creep is often the longest creep phase, and

[0017] (III) Tertiary creep: this corresponds to an intergranular rupture by decohesion of the grain boundaries which will lead to necking and the appearance of porosities in the material. It is when this state becomes too present in the critical zones that the part begins to crack and finally breaks.

[0018] Numerical models for simulating blade creep are known which make it possible to theoretically determine a blade lifespan. However, the lifespan estimate does not correspond to the actual lifespan of the blades, and by extension of the turbine, due to the conservatisms integrated into the calculation approach.

[0019] Creep ruptures are linked to structural phenomena but their origin can be linked to local phenomena which are difficult to quantify by direct approach (elongation measurement of the order of a few hundredths of a millimeter for example).

[0020] Today, endoscopies can be carried out to check that there is no appearance of creep-initiated cracks in the rotor and stator blades of the turbine. However, endoscopies do not allow us to know the state of creep in the areas concerned and in particular the appearance of tertiary creep in a generalized manner announcing the imminent rupture of the part. Thus, it is not possible to predict the time at which the onset of the fining primer could be expected.

[0021] Thus, for parts that perish during finishing, it is necessary to develop a methodology to quantify the finishing state and estimate the actual potential of the part. Since finishing is characterized by a local deformation of the material, the finishing potential of the part can very often be correlated with the geometric deformation of a part of the part. The challenge is then to establish a stopping criterion based on the measurement of this deformation.

[0022] Today, the finage is not followed. One of the possible solutions would be to measure characteristic distances of the parts using cameras used in endoscopy.

[0023] However, the camera may not have enough distance clearance in the engine to accurately measure large distances without using "Fish Eye" technology. The latter allows for panoramic vision and therefore to extend the field of view. However, due to this technology, the images displayed are distorted, making it impossible to make measurements between geometric elements that are too far apart.

[0024] Another way would be to make visual indicators (marker marks) and measure the distance between these indicators and the ends of the blades. However, over time, corrosion, oxidation, dirt and pollution are likely to erase the visual indicator, and therefore reduce the accuracy of the measurement.

[0025] The present invention aims to provide a solution to this problem, which is simple, effective and economical. Summary of the invention

[0026] The invention relates to a method for controlling a turbomachine blade, in particular an aircraft blade, the blade being made of a metal alloy and comprising at least one platform connected to at least one blade, the method comprising the following steps:

[0027] a) determining at least one geometric parameter for controlling the blade, this parameter having a value likely to change depending on the creep deformations of the blade,

[0028] b) determining an acceptability threshold for the or each parameter,

[0029] c) verify in situ the conformity of the or each parameter with respect to the threshold cor respondent, and

[0030] d) conclude on the conformity of the blade based on the results of the verification.

[0031] The invention makes it possible to simply and reliably control the level of creep of a blade. For this, a compliance threshold is predetermined for one or more geometric parameters, and the blade is checked against these thresholds to verify whether or not it is compliant. The value of the or each parameter is preferably likely to change significantly and easily quantifiable.

[0032] The control or verification here is of the in situ type. This means that it is carried out directly on the blade and not at a distance from this blade, by laser means.

[0033] A compliant blade may continue to be used in a turbine. On the contrary, a non-compliant blade must be scrapped and therefore replaced by a new one in the turbine.

[0034] The method according to the invention can be carried out on a disassembled blade, but also on a blade mounted in a turbine module or sub-module. It is sufficient to have access to the blade within this module or sub-module to carry out the inspection.

[0035] The method according to the invention may also have one or more of the following characteristics, taken alone or in combination with each other:

[0036] - step a) consists of determining a single geometric control parameter;

[0037] - the platform comprises a spoiler on the side of a leading or trailing edge of the blade, the parameter or one of the parameters determined in step a) being an angle formed by this spoiler;

[0038] - said angle is measured relative to a plane in which the rest of the platform;

[0039] - the parameter or one of the parameters determined in step a) is a curvature of the pale ;

[0040] - the parameter or one of the parameters determined in step a) is a profile of an edge blade attack or escape;

[0041] - step c) comprises the sub-steps consisting of:

[0042] i) measuring a value of the or each parameter directly on the blade, and

[0043] ii) comparing the or each measured value to said corresponding threshold;

[0044] - step c) comprises the sub-steps consisting of:

[0045] j) mounting at least one control template directly on the blade, this template having a shape complementary to a part of the blade comprising the parameter to be controlled, and

[0046] jj) checking the positioning of the template on the part of the blade and detecting the possible presence of one or more clearances between the template and this part;

[0047] - the template is mounted on the platform, on a spoiler of the platform, or on a leading or trailing edge of the or each blade;

[0048] - the blade is part of a turbine module or sub-module;

[0049] — step a) comprises a sub-step consisting of producing at least one abacus at from several dawn samples;

[0050] — the abacus or each abacus shows the evolution of the value of a parameter of control based on the number of cycles of the turbomachine or turbine containing the part;

[0051] — the blade is a rotor or stator blade;

[0052] — the blade is formed from a single piece; alternatively, it is formed by assembling of parts;

[0053] — said platform is an external platform, or alternatively internal;

[0054] — the dawn has a single platform;

[0055] — the dawn has two platforms, respectively internal and external, between which the blade(s) extend;

[0056] — the dawn has a single blade;

[0057] — the dawn comprises at least two blades. Brief description of the figures

[0058] Other characteristics and advantages of the invention will appear during the reading of the detailed description which follows for the understanding of which reference will be made to the appended drawings in which:

[0059] [Fig-1] [Fig.l] is a half schematic view in axial section of a turbomachine aircraft, and in particular a turbine of this turbomachine;

[0060] [Fig.2] [Fig.2] is a schematic perspective view of a rotor blade of the turbomachine of [Fig.l];

[0061] [Fig.3] [Fig.3] is a schematic perspective view of a stator blade of the turbomachine of [Fig.l];

[0062] [Fig.4] [Fig.4] is a graph showing the evolution of the elongation e of a part over time t, when this part is subjected to a constant stress at constant temperature;

[0063] [Fig.5] [Fig.5] is a block diagram showing steps of a method according to the invention for controlling a blade;

[0064] [Fig.6a-6c] Figures 6a-6c are schematic views of a spoiler of a blade, and respectively represent three levels of deformation by fining of the blade;

[0065] [Fig.7] [Fig.7] is a schematic view of a spoiler of a blade, and shows pa geometric control parameters of this blade;

[0066] [Fig.8a-8c] Figures 8a-8c are schematic views of a spoiler of a blade and of a control template mounted on this spoiler, and respectively represent three distinct positions of the template linked to the finishing of the blade;

[0067] [Fig.9] [Fig.9] is a graph showing the evolution of the parameter illustrated in [Fig.7] depending on the number of operating cycles of the turbomachine or turbine;

[0068] [Fig. 10] [Fig. 10] is a schematic sectional view of a blade and shows a geometric parameter for controlling this blade; and

[0069] [Fig. 11] [Fig. 11] is a schematic view of a blade and shows another geometric parameter for controlling this blade. Detailed description of the invention

[0070] Figures 1 to 4 have been described in the above.

[0071] We now refer to [Fig. 5] which illustrates an embodiment of a method according to the invention for controlling the fining of a turbomachine blade, in particular an aircraft blade.

[0072] This blade may be a rotor blade 16 or a stator blade 12 as illustrated in Figures 2 and 3.

[0073] The blade is made of a metal alloy. It can be formed from a single piece or, alternatively, be produced by assembling pieces. The blade comprises at least one platform connected to at least one blade. In the example shown, the blade comprises two platforms, respectively internal 24, 34 and external 26, 36, between which at least one blade 22, 32 extends.

[0074] In operation, the gases circulating through the blades and vanes reach high temperatures which can cause a fining phenomenon.

[0075] The method according to the invention comprises the steps consisting of:

[0076] a) determining at least one geometric parameter for controlling the blade, this parameter having a value capable of changing, preferably in a significant and easily quantifiable manner, as a function of the deformations by finning of the blade,

[0077] b) determining an acceptability threshold for the or each parameter,

[0078] c) verify in situ the conformity of the or each parameter with respect to the threshold cor respondent, and

[0079] d) conclude on the conformity of the blade based on the results of the verification.

[0080] The method according to the invention therefore essentially comprises 4 steps a), b), c) and d).

[0081] Each of these steps may include sub-steps.

[0082] Concerning the first step a), it consists of determining one or more geometric parameters for controlling the blade. The control parameter may be unique insofar as a single parameter could be sufficient to control the fining of the blade, for example in the area most likely to be deformed by fining in operation.

[0083] As will be seen below, this parameter can be a distance or an angle for example.

[0084] Step a) may comprise a sub-step consisting of producing a chart from several blade samples, this chart showing for example the evolution of the value of a control parameter as a function of the number of cycles undergone by this part. A cycle is understood as an operating cycle of the turbomachine or turbine, and comprises for the aircraft a start, a takeoff, a cruising flight, a landing and an engine shutdown.

[0085] The second step is to determine the acceptability threshold of the or each control parameter. For example, for an angle, a threshold could be determined of such that when the value of this angle is less than the threshold, the blade fining is not significant, and when the value of this angle is greater than the threshold, the blade fining is too great and this blade risks cracking or breaking. In another example, for a distance, a threshold could be determined such that when the value of this distance is greater than the threshold, the blade fining is not significant, and when the value of this distance is less than the threshold, the blade fining is too great and this blade risks cracking or breaking.

[0086] Step c) thus consists of verifying the conformity of the or each parameter with respect to the corresponding threshold, directly on the blade. We will see that there are at least two approaches to carrying out this step c).

[0087] This step c) can be carried out on a disassembled blade or distributor sector, or on a blade which is mounted in a turbine module or sub-module;

[0088] Finally, step d) makes it possible to conclude on the conformity of the blade and therefore its level of finishing, based on the results of step c).

[0089] Figures 6a to 9 illustrate a first geometric parameter which can be controlled within the framework of the method according to the invention, and which relates to a platform spoiler.

[0090] Each of the platforms 24, 26 comprises an upstream edge located on the side of a leading edge 38 of the blade 22, 32, and a downstream edge located on the side of a trailing edge 40 of the blade 22, 32 (Figures 2 and 3). Each of these upstream and downstream edges forms a spoiler 42.

[0091] As seen in the drawings, this spoiler 42 is cantilevered and is a preferred zone of deformation by fining of the blade 12, 16.

[0092] Centrifugal forces, gas forces, or even the pretorsion assembly of the blades, have an even greater influence on the portions of the part that are cantilevered, which can generate local stress concentrations. For example, the upstream spoiler of the heel is cantilevered and generates a local stress concentration at the top leading edge of the blade under the effect of centrifugal force. These local stresses can lead to the appearance of cracks in the critical zone or even a failure by fining at the tip of the blade from the leading edge towards the trailing edge, along the heel. This geometric deformation due to fining is observed at the upstream spoiler of the heel which has an upward deflection. Once correlated with the fining state in the blade, it is the measurement of the deflection of the spoiler which makes it possible to establish a control parameter for the part.

[0093] Figures 6a to 6c show for example three distinct levels of fining of the spoiler 42 of an external platform 26, 36 of a blade 12, 16. In Figure 6a, the spoiler 42 has not undergone fining. In Figure 6b, the spoiler 42 has begun to deform by fining by deforming outwards, that is to say on the side opposite the blade 22. In Figure 6c, this deformation is even greater and the blade risks cracking or even breaking.

[0094] It can be seen in these figures 6a-6c that the angle a that the spoiler 42 forms with the rest of the platform 26, 36, and in particular with a plane P passing through the rest of the platform 26, 36, varies and in particular increases with the finage.

[0095] The angle a can thus be a control parameter within the meaning of the invention.

[0096] Instead of the angle a, a distance H between the plane P and the end of the spoiler 42 could be used as a control parameter ([Fig.7]).

[0097] [Fig.9] shows the evolution of the angle a of the spoiler 42 as a function of the number of cycles of the turbomachine. This figure is a chart as described above. The initial angle of the spoiler visible in Figure 6a is noted a0. The compliance threshold of this angle is noted as. In the case where the angle al of the spoiler is less than the threshold as, the blade can continue to be used because the fining of the blade is not yet too great. In the case where the angle a2 of the spoiler is greater than the threshold as, the blade must be scrapped and replaced because it risks cracking at any time.

[0098] According to one embodiment of the invention ([Fig.4]), step c) comprises the sub-steps consisting of:

[0099] i) measuring a value of the or each parameter, such as the angle a or the distance H, directly on the blade, and

[0100] ii) comparing the or each measured value to said corresponding threshold.

[0101] The measuring sub-step i) may be carried out by means of any suitable tool, such as a protractor, a comparator, a ruler, a gauge, etc.

[0102] The comparison sub-step can be done by a computer or simply by the operator measuring the value. This operator will have no difficulty in determining whether the value measured in sub-step i) is lower or higher than the threshold value.

[0103] According to an alternative embodiment of the invention ([Fig.4]), step c) comprises the sub-steps consisting of:

[0104] j) mounting at least one control template 44 directly on the blade, this template having a shape complementary to a part of the blade comprising the parameter to be controlled, and

[0105] jj) checking the positioning of the template 44 on the part of the blade and detecting the possible presence of one or more clearances J between the template and this part.

[0106] This variant is illustrated in Figures 8a to 8c.

[0107] The template 44 chosen has a shape complementary to the platform and the spoiler 42, for example when the parameter to be controlled is at the threshold value. The template 44 defines support points on the platform in predetermined zones, and for example a first zone Z1 at the level of the spoiler 42, a second zone Z2 at the level of the rest of the platform, and a third zone at the level of the junction of the zones Z1 and Z2. The template 44 is here intended to be engaged under the platform 26, 36, at the level of its junction with the blade 22, 32. In the case of a rotor blade 16 ([Fig.2]), the template 44 can be engaged on one side or the other of the leading edge 38 or trailing edge 40, or on both sides. In the case of a stator blade 12 or a turbine nozzle ([Fig.3]), the template 44 can be engaged on only part or on the entire circumferential extent of the nozzle and can therefore extend in front of the leading edges 38 of several blades 32 or behind the trailing edges 40 of several blades 40.

[0108] In the case of Figure 8a, the spoiler 42 is similar to that of Figure 6a. The template 44 is supported at the level of the zones Z1, Z2 and there is deliberately no support at the level of the zone Z3.

[0109] In the case of Figure 8b, the spoiler 42 is similar to that of Figure 6b. The template 44 is supported at all zones Z1, Z2, Z3.

[0110] In the case of Figure 8c, the spoiler 42 is similar to that of Figure 6c. The template 44 is supported at the level of the zones Z2, Z3 and a clearance J appears at the level of the zone Z1, this clearance J being visible to an operator. The existence of this clearance is detected in the sub-step jj) and allows an operator to declare the blade as non-compliant.

[0111] Alternatively, the parameter that could be controlled in step c) could be:

[0112] - the curvature of the blade 22, 32 which is likely to evolve locally with the finishing of dawn ([Fig. 10]),

[0113] - the profile of the leading edge 38 or trailing edge 40 of the blade 22, 32 ([Fig.l 1]).

[0114] Depending on the control parameter(s) chosen, one or more templates are mounted on one of the platforms, on one of the spoilers, on the leading or trailing edge of the blade or one of the blades, etc.

[0115] The present invention makes it possible to scrap the blades as late as possible, while ensuring healthy operation of the turbines and turbomachines. It thus allows significant financial savings because there are fewer parts to change because their use is maximized, fewer early turbine removals by the customer, fewer parts to change because the risk of breakage of parts and debris being released is significantly reduced, etc. In conclusion, this invention makes it possible to improve the robustness of turbomachines and turbines and to reduce the risk of unplanned engine removal.

Claims

Claims

1. Method for controlling the fining of a blade (12, 16) of a turbomachine, in particular an aircraft, the blade (12, 16) being made of a metal alloy and comprising at least one platform connected to at least one blade (22, 32), the method comprising the following steps: a) determining at least one geometric parameter for controlling the blade (12, 16), this parameter having a value which changes as a function of the creep deformations of the blade, b) determining an acceptability threshold for the or each parameter, c) verifying in situ the conformity of the or each parameter with respect to the corresponding threshold, by mounting at least one control template (44) directly on the blade (12, 16), this template (44) having a shape complementary to a part of the blade (12, 16) comprising the parameter to be controlled, and detecting the possible presence of one or more clearances (J) between the template (44) and this part of the blade on which the template is positioned (44),and d) conclude on the conformity of the blade (12, 16) based on the results of the verification.,

2. Method according to claim 1, in which step a) consists of determining a single geometric control parameter.

3. Method according to claim 1 or 2, in which the platform (26, 36) comprises a spoiler (42) on the side of a leading edge (38) or trailing edge (40) of the blade (22, 32), the parameter or one of the parameters determined in step a) being an angle (a) formed by this spoiler (42).

4. A method according to claim 3, wherein said angle (a) is measured relative to a plane (P) in which the remainder of the platform (26, 36) mainly extends.

5. A method according to claim 1 or 2, wherein the parameter or one of the parameters determined in step a) is a curvature of the blade (22, 32).

6. Method according to claim 1 or 2, in which the parameter or one of the parameters determined in step a) is a profile of a leading edge (38) or trailing edge (40) of the blade (22, 32).

7. Method according to one of the preceding claims, in which the template (44) is mounted on the platform, on a spoiler (42) of the platform (26, 36), or on a leading edge (38) or trailing edge (40) of the or each blade (22, 32).

8. A method according to any preceding claim, wherein the blade (12, 16) is part of a turbine module or sub-module.

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