Monitoring the creep of an aircraft turbomachine blade

Abstract

Method for monitoring the creep of a turbomachine blade (12, 16), in particular of an aircraft, the blade (12, 16) being made of a metal alloy and comprising at least one platform connected to at least one airfoil (22, 32), the method comprising the following steps: a) determining at least one geometric monitoring parameter of the blade (12, 16), this parameter having a value liable to evolve as a function of the deformations by creep 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, and d) concluding on the conformity of the blade (12, 16) as a function of the results of the verification.

Description

[0001] DESCRIPTION

[0002] TITLE: CONTROL OF THE CREEP OF AN AIRCRAFT TURBOMACHINE BLADE

[0003] Technical field of the invention

[0004] The present invention relates to the general field of aeronautics. It relates more particularly to a method for controlling the creep of a turbomachine blade, in particular an aircraft blade.

[0005] Technical background

[0006] The technical background includes in particular documents EP-A1 - 3 176 561, EP-A1 - 3 171 127 and DE-A1 -10 2008 037412.

[0007] A turbomachine is known comprising a turbine 10 as shown in FIG. 1.

[0008] 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.

[0009] 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 discs 18 secured to each other and rotatable about a longitudinal axis A of the turbomachine.

[0010] 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.

[0011] A rotor blade 16 is shown in Figure 2 and stator blades 12 are shown in Figure 3.

[0012] 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. The platforms 24, 26 define between them a part of the gas flow vein in the turbine 10. The roots 28 make it possible to mount the rotor blades 16 on the rotor disks 18.

[0013] 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.

[0014] 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. Figure 3 shows a distributor sector comprising three vanes 12 and therefore three blades 32.

[0015] The rotor and stator blades 12, 16 are subjected to high stresses partly due to the temperatures to which they are subjected during operation, particularly 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 creep over the long term with respect to the engine operating time.

[0016] Therefore, these blades 12, 16 may exhibit creep deformations which it is important to monitor.

[0017] Creep 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.

[0018] Creep is characterized by three main states illustrated in Figure 4:

[0019] (I) Primary creep: At the initial stage, the strain rate s is relatively high, but decreases with increasing time and strain due to the material experiencing an increase in creep strength or strain hardening,

[0020] (II) Secondary creep: sometimes called steady-state creep, the strain rate s 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 (III) Tertiary creep: this corresponds to intergranular failure 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 critical zones that the part begins to crack and ultimately break. Numerical models for simulating blade creep are known which make it possible to theoretically determine a blade life.However, the lifetime estimate does not correspond to the actual lifetime of the blades, and by extension of the turbine, due to the conservatisms built into the calculation approach.

[0021] 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).

[0022] Today, endoscopies can be performed 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 affected areas and in particular the appearance of tertiary creep in a generalized manner announcing the imminent failure of the part. Thus, it is not possible to predict the moment at which the appearance of creep initiation could be expected.

[0023] Thus, for parts that perish in creep, it is necessary to develop a methodology to quantify the creep state and estimate the actual potential of the part. Since creep is characterized by a local deformation of the material, the creep 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.

[0024] Creep is not currently monitored. One possible solution would be to measure characteristic distances between parts using cameras used in endoscopy.

[0025] However, the camera may not have enough distance traveled back in the motor to accurately measure large distances without using Fish Eye technology. This allows for panoramic vision and therefore extends 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.

[0026] Another way would be to make visual indicators (marker marks) and measure the distance between these indicators and the tips 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.

[0027] The present invention aims to provide a solution to this problem, which is simple, effective and economical.

[0028] Summary of the invention

[0029] 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: a) determining at least one geometric parameter for controlling the blade, this parameter having a value capable of changing as a function of the creep deformations of the blade, b) determining an acceptability threshold for the or each parameter, c) verifying in particular in situ the conformity of the or each parameter with respect to the corresponding threshold, preferably by 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 by detecting the possible presence of one or more clearances between the template and this part of the blade on which the template is positioned,and (d) conclude on the conformity of the blade based on the results of the verification.,

[0030] The invention makes it possible to simply and reliably control the creep level of a blade. To this end, a compliance threshold is predetermined for one or more geometric parameters, and the blade is checked according to 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. The control or verification is here 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.

[0031] A compliant blade can continue to be used in a turbine. Conversely, a non-compliant blade must be scrapped and therefore replaced with a new one in the turbine.

[0032] 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.

[0033] The method according to the invention may also have one or more of the following characteristics, taken alone or in combination with each other:

[0034] -- the template is preferably rigid or non-deformable,

[0035] - step a) consists of determining a single geometric control parameter;

[0036] - 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, in particular with respect to a reference frame;

[0037] - said angle is measured relative to a plane in which the rest of the platform mainly extends;

[0038] - the parameter or one of the parameters determined in step a) is a curvature of the blade;

[0039] - the parameter or one of the parameters determined in step a) is a profile of a leading or trailing edge of the blade;

[0040] - step c) comprises the sub-steps of: i) measuring a value of the or each parameter directly on the blade, and ii) comparing the or each measured value to said corresponding threshold;

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

[0042] - 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;

[0043] - the blade is part of a turbine module or sub-module;

[0044] -- step a) comprises a sub-step consisting of producing at least one chart from several blade samples;

[0045] -- the abacus or each abacus shows the evolution of the value of a control parameter as a function of the number of cycles of the turbomachine or turbine containing the part;

[0046] -- the blade is a rotor or stator blade;

[0047] -- the blade is formed from a single piece; alternatively, it is formed by assembling pieces;

[0048] -- the said platform is an external platform, or alternatively an internal one;

[0049] -- the dawn has a single platform;

[0050] -- the blade has two platforms, respectively internal and external, between which the blade(s) extend;

[0051] -- the dawn has a single blade;

[0052] -- the dawn has at least two blades.

[0053] Brief description of the figures

[0054] 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:

[0055] [Fig. 1] Figure 1 is a half schematic axial sectional view of an aircraft turbomachine, and in particular of a turbine of this turbomachine; [Fig. 2] Figure 2 is a schematic perspective view of a rotor blade of the turbomachine of Figure 1;

[0056] [Fig. 3] Figure 3 is a schematic perspective view of a stator blade of the turbomachine of Figure 1;

[0057] [Fig. 4] Figure 4 is a graph showing the evolution of the elongation s of a part over time t, when this part is subjected to a constant stress at constant temperature; [Fig. 5] Figure 5 is a block diagram showing steps of a method according to the invention for controlling a blade;

[0058] [Fig. 6a-6c] Figures 6a-6c are schematic views of a blade spoiler, and respectively represent three levels of creep deformation of the blade;

[0059] [Fig. 7] Figure 7 is a schematic view of a spoiler of a blade, and shows geometric parameters for controlling this blade;

[0060] [Fig. 8a-8c] Figures 8a-8c are schematic views of a blade spoiler and a control template mounted on this spoiler, and respectively represent three distinct positions of the template linked to the creep of the blade;

[0061] [Fig. 9] Figure 9 is a graph showing the evolution of the parameter illustrated in Figure 7 as a function of the number of operating cycles of the turbomachine or turbine;

[0062] [Fig. 10] Figure 10 is a schematic sectional view of a blade and shows a geometric parameter for controlling this blade; and

[0063] [Fig. 11] Figure 11 is a schematic view of a blade and shows another geometric parameter for controlling this blade.

[0064] Detailed description of the invention

[0065] Figures 1 to 4 have been described in the above.

[0066] We now refer to figure 5 which illustrates an embodiment of a method according to the invention for controlling the creep of a turbomachine blade, in particular an aircraft blade.

[0067] This blade may be a rotor blade 16 or a stator blade 12 as illustrated in Figures 2 and 3.

[0068] The blade is made of a metal alloy. It can be formed from a single piece or alternatively be made by assembling parts. 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.

[0069] During operation, the gases flowing through the blades and vanes reach high temperatures which can cause creep.

[0070] The method according to the invention comprises the steps of: a) determining at least one geometric parameter for controlling the blade, this parameter having a value likely to change, preferably in a significant and easily quantifiable manner, 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, and d) concluding on the conformity of the blade based on the results of the verification.

[0071] The method according to the invention therefore essentially comprises 4 steps a), b), c) and d).

[0072] Each of these steps can include sub-steps.

[0073] Concerning the first step a), it consists of determining one or more geometric parameters for controlling the blade. The control parameter can be unique insofar as a single parameter could be sufficient to control the creep of the blade, for example in the area most likely to deform by creep during operation.

[0074] As we will see in the following, this parameter can be a distance or an angle for example.

[0075] Step a) may comprise a sub-step consisting of producing a chart from several blade samples, this chart showing, for example, the revolution 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.

[0076] 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 such that, when the value of this angle is less than the threshold, the blade creep is not significant, and when the value of this angle is greater than the threshold, the blade creep 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 creep is not significant, and when the value of this distance is less than the threshold, the blade creep is too great and this blade risks cracking or breaking.

[0077] Step c) therefore consists of checking the conformity of the parameter(s) 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).

[0078] This step c) can be carried out on a dismantled blade or nozzle sector, or on a blade which is mounted in a turbine module or sub-module;

[0079] Finally, step d) allows us to conclude on the conformity of the blade and therefore its level of creep, based on the results of step c).

[0080] 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.

[0081] 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.

[0082] As seen in the drawings, this spoiler 42 is cantilevered and is a privileged zone of deformation by creep of the blade 12, 16.

[0083] Centrifugal forces, gas forces, and even pretorsion assembly of the blades have a greater influence on the cantilevered portions of the part, which can generate local stress concentrations. For example, the upstream spoiler of the root is cantilevered and generates a local stress concentration at the top of the blade leading edge under the effect of centrifugal force. These local stresses can lead to cracks appearing in the critical zone or even a creep rupture at the tip of the blade from the leading edge towards the trailing edge, along the root. This geometric deformation due to creep is observed at the upstream spoiler of the root, which has an upward deflection. Once correlated with the creep state in the blade, it is the measurement of the spoiler deflection that makes it possible to establish a control parameter for the part.Figures 6a to 6c show for example three distinct levels of creep of the spoiler 42 of an external platform 26, 36 of a blade 12, 16. In Figure 6a, the spoiler 42 has not undergone any creep. In Figure 6b, the spoiler 42 has begun to deform by creep 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.

[0084] 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 creep.

[0085] The angle a can thus be a control parameter within the meaning of the invention.

[0086] 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 (figure 7).

[0087] Figure 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 a1 of the spoiler is lower than the threshold as, the blade can continue to be used because the creep of the blade is not yet too significant. In the case where the angle a2 of the spoiler is higher than the threshold as, the blade must be scrapped and replaced because it risks cracking at any time.

[0088] According to one embodiment of the invention (figure 4), step c) comprises the sub-steps consisting of: i) measuring a value of the or each parameter, such as the angle a or the distance H, directly on the blade, and ii) comparing the or each measured value with said corresponding threshold.

[0089] 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.

[0090] 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. According to an alternative embodiment of the invention, step c) comprises the sub-steps consisting of: 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 jj) controlling 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. This alternative is illustrated in Figures 8a to 8c.

[0091] 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 spoiler 42, a second zone Z2 at the rest of the platform, and a third zone at the junction of the zones Z1 and Z2. The template 44 is here intended to be engaged under the platform 26, 36, at its junction with the blade 22, 32. In the case of a rotor blade 16 (figure 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 may be engaged over only part or all of the circumferential extent of the nozzle and may therefore extend in front of the leading edges 38 of several blades 32 or behind the trailing edges 40 of several blades 40.

[0092] 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.

[0093] 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.

[0094] 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 zones Z2, Z3 and a clearance J appears at the level of zone Z1, this clearance J being visible to an operator. The existence of this clearance is detected in sub-step jj) and allows an operator to declare the blade as non-compliant.

[0095] Alternatively, the parameter which could be controlled in step c) could be: - the curvature of the blade 22, 32 which is likely to evolve locally with the creep of the blade (figure 10),

[0096] - the profile of the leading edge 38 or trailing edge 40 of the blade 22, 32 (figure 11).

[0097] Depending on the control parameter(s) chosen, one or more jigs are mounted on one of the platforms, on one of the spoilers, on the leading or trailing edge of the blade(s), etc.

[0098] The present invention makes it possible to scrap blades as late as possible, while ensuring the healthy operation of 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 impacting the customer, fewer parts to change because the risk of parts breaking 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 removals.

Claims

CLAIMS 1. Method for controlling the creep 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 by detecting the possible presence of one or more clearances (J) between the template (44) and this part of the blade on which the template (44) is positioned, and d) conclude on the conformity of the blade (12,16) depending 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. 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. Method according to claim 1 or 2, in which 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. Method according to one of the preceding claims, in which the blade (12, 16) is part of a turbine module or sub-module.

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