Method and system for manufacturing a moving blade with augmented chord for an aircraft turbomachine

Abstract

A movable turbine blade (33) of an aircraft turbomachine (2) is modified by obtaining (50) a position (H) within the blade height (330) of the blade as a function of a blade deformation in a first natural torsional mode; then by modifying (52) the geometry of the blade section (602) at said position, the modification comprising stretching (520) the section by lengthening a chord (700) connecting the leading edge (BA) of the section to the trailing edge (BF) of the section and reducing (522) the thickness (e) of the stretched section (602') to maintain a constant cross-sectional area with the blade before modification; and finally by connecting the modified section (602'') to the root section (601) and the tip section (600) of the blade before modification. Figure for the abbreviation: Figure 7

Description

Title of the invention: Method and system for manufacturing a moving blade with augmented chord for an aircraft turbomachine. Technical field

[0001] The present invention relates to the general field of mechanical design of a turbine blade or vane of an aircraft turbomachine. Previous techniques

[0002] Climate change is a major concern for many legislative and regulatory bodies worldwide. Indeed, various restrictions on carbon emissions have been, are being, or will be adopted by various states. In particular, an ambitious standard applies to both new types of aircraft and those currently in operation, requiring the implementation of technological solutions to bring them into compliance with current regulations. Civil aviation has been actively contributing to the fight against climate change for several years now.

[0003] Technological research efforts have already led to very significant improvements in the environmental performance of aircraft. The Applicant takes into account the factors impacting all phases of design and development in order to obtain aeronautical components and products that are less energy-intensive, more environmentally friendly, and whose integration and use in civil aviation have moderate environmental impacts, with the aim of improving the energy efficiency of aircraft. Consequently, the Applicant is constantly working to reduce its climate impact by employing methods and operating virtuous development and manufacturing processes that minimize greenhouse gas emissions to the minimum possible in order to reduce the environmental footprint of its activity.

[0004] This sustained research and development work focuses on new generations of aircraft engines, the weight reduction of aircraft, particularly through the materials used and lighter on-board equipment, the development of the use of electrical technologies to provide propulsion, and finally aviation fuels.

[0005] In this context, aircraft turbomachine blades can be improved.

[0006] Blades are found in various engine components, for example in the fan, low-pressure and high-pressure compressors, and low-pressure and high-pressure turbines. One of the constituent parts of these components is the rotor, which is composed of movable blades. Each blade is formed of a blade fixed at both ends to a foot (to the rotor axis) and to a heel or head.

[0007] During the mechanical design of a blade, it is known to take into account the dynamic dimensioning (or frequency placement) of the blade, and more particularly of the blade blade, in particular in order to limit the risks of damage of vibrational origin linked to aerodynamic resonance phenomena (such as wake effects of parts upstream of the blade) or mechanical phenomena (such as imbalances).

[0008] Indeed, the vibrational potential of the blade is directly dependent on the upstream excitations (aerodynamic source) seen by the blade, but also on its dynamic robustness, which is intrinsically linked to the blade geometry and the materials used. The dynamic robustness of the blade depends on its vibrational behavior, which is defined by its natural modes, each natural mode having a natural frequency and a modal deformation.

[0009] During operation, the blades are aerodynamically excited by several flows in the engine which together form an aerodynamic excitation source. If the excitation frequency associated with these flows is equal to the frequency of one of the natural modes, the blade enters into resonance and deforms, potentially leading to the rapid failure of parts.

[0010] This is why the dynamic dimensioning of a blade is typically carried out by constructing its Campbell diagram, that is, by plotting the curves showing the evolution of the frequency of a natural mode of the blade (associated with the same modal shape) as a function of the rotor's rotational speed. The excitation frequency curve(s) of this rotor are then superimposed on the natural mode frequency curves to form the Campbell diagram.

[0011] Aircraft engine blades exhibit several types of natural modes, such as torsional or bending modes. More precisely, the natural modes of a blade, in order of increasing frequency, give the first bending mode (called mode 1F), the first torsional mode (called 1T), and then subsequent bending and torsional modes. The first modes, 1F and 1T, are generally the most at risk because they are the most easily excitable.

[0012] The first step in the dynamic dimensioning of the blade consists of correctly positioning mode 1F (this choice largely determines the blade geometry, for example, the radii of the leading edge (BA) and trailing edge (BF)). Once mode 1F is positioned, mode 1T is positioned. Mode 1T is a relatively flat mode, meaning that its frequency depends very little on the rotor speed. Mode 1T generally exhibits crossovers (or dynamic coincidences) with the excitation curves in the Campbell diagram.

[0013] Figure 1 schematically represents a Campbell diagram on which the engine speed RPM is found on the x-axis and the frequency F on the y-axis. It shows the frequency situation of a blade by superimposing the frequency of an excitation harmonic EXC and the frequency of a single natural mode MP, such as the 1T mode. This superposition shows at what engine speed RPMO a dynamic coincidence can appear.

[0014] A high-performance blade can therefore be designed, in its geometry and the material used, to limit the risks of dynamic coincidence for usual operating regimes, i.e. with a high rate of occurrence and a long holding time (for example in cruise, takeoff or landing).

[0015] However, the search for new generations of aircraft engines may lead to the development of new architectures that influence the aerodynamic excitation flows within the engine, resulting in a modification of the excitation curves. In the event of a dynamic coincidence between the 1T mode and the wake of the new turbomachine architecture for typical operating conditions, the level of vibrational excitation of the existing blades may become unacceptable. The existing blades are then no longer suitable. Description of the invention

[0016] Consequently, there is a need to remove this dynamic coincidence from the engine's operating range, while maintaining the aerodynamic performance of the blade. To achieve this, the frequency of mode 1T can be modified by adjusting the blade geometry.

[0017] Indeed, the geometry of the blade and the material used are elements that can strongly impact the frequency f of a local eigenmode, as can be seen from the following (simplified) formula: 100181

[0019] where k represents the blade stiffness in the (local) deformation zone and m the mass moving in the zone. Thus, the more the stiffness or mass in the mode deformation zone is modified, the more its mode frequency 1T is modified.

[0020] In this context, it is proposed to modify the geometry of the blade in the main deformation zone, without degradation of aerodynamic performance and without modification of the mass of the blade in order not to modify the static stresses.

[0021] Thus, the invention relates to a method for manufacturing a movable blade for an aircraft turbomachine. The method comprises the following steps: obtaining a position in the blade height of a movable blade, as a function of a blade deformation in a first natural torsional mode, modifying the geometry of a blade section at said position, the modification comprising the following operations: - stretch the blade section by lengthening a chord connecting a leading edge of the section to a trailing edge of the section, and - reduce the thickness of the stretched section to maintain a constant cross-sectional area with the blade before modification, and connect the modified section to a root section and a tip section of the blade before modification, so as to obtain a modified blade.

[0022] By stretching the section, the profile of the blade is preserved, and therefore its aerodynamic performance.

[0023] By further reducing the thickness of the section, the center of gravity of the blade is moved upstream (towards the leading edge), thus reducing the thickness of the blade forward and therefore the specific mass of the area of ​​significant deformation.

[0024] In this way, the frequency situation of the blade is modified by significantly lowering the frequency of the 1T mode. This results in an exit from a situation at risk of resonance.

[0025] Finally, by maintaining a constant cross-sectional area, the mass of the blade remains unchanged. The effect of the modification on the static / mechanical stresses of the blade is therefore small, or even negligible.

[0026] New blades can thus be manufactured / produced from existing blades to adapt to new aircraft turbomachinery architectures.

[0027] Preferably, obtaining the position in the blade height includes determining a position where the deformation of the blade is significant relative to a threshold, or even maximum, in a first natural mode of torsion.

[0028] In particular, the position within the blade height can be located between 45 and 55% of the blade height, preferably at mid-height (or approximately at mid-height, i.e., between, for example, 49% and 51%). Indeed, it has been observed that the area of ​​the blade with the greatest resonant displacement (mode 1T) is often the trailing edge at approximately 50% of the height.

[0029] In one embodiment, the string is lengthened by 5 to 20%, preferably by 10 to 15%, for example by about 10%. These embodiments represent a compromise between a substantial reduction in the frequency of the 1T mode (by 15% for a string lengthening of 10%) and integration constraints (avoiding parts touching) and manufacturing constraints (too angular a connection with the blade foot or head).

[0030] Preferably, the stretching of the blade section is carried out using an affinity having as its direction the axis of the chord or the axis of a section between two control points of a skeleton of the blade section, that is to say a geometric transformation in the plane (of the section) which provides a homothety about an axis (formed by the chord or locally by a skeleton of the section), without transformation (i.e., identity function) about an axis orthogonal to the axis of the homothety.

[0031] In one embodiment, reducing the thickness includes: identify a master frame in the stretched section (i.e., the greatest thickness, normal to the skeleton of the section), and reduce the thickness of the section at the midship frame by applying a thickness law between the midship frame and each of the leading and trailing edges, until a surface area of ​​the stretched section is equal to a surface area of ​​the section before modification.

[0032] In one embodiment, reducing the thickness involves modifying the upper surface profile of the blade in the stretched section, while maintaining an unchanged lower surface profile. This arrangement preserves the aerodynamic performance of the blade.

[0033] Preferably, the radii of curvature of the leading and trailing edges are kept constant when modifying the geometry of the blade section. This arrangement makes it possible to maintain the aerodynamic performance of the blade.

[0034] In one embodiment, connecting the modified section to the root and tip sections of the blade comprises: determine a center of gravity of the modified section, align the center of gravity with a gravity axis of the blade, and apply a profile matching law between the section aligned with the gravity axis and the root and tip sections of the blade.

[0035] The mechanical stresses of the blade are therefore not modified.

[0036] The invention also relates to a movable turbine blade of an aircraft turbomachine, the geometry of which is modified in accordance with the above method.

[0037] Furthermore, the invention also relates to a manufacturing system for a movable blade for an aircraft turbomachine, the system comprising: a memory for storing a digital model of a movable blade, a microprocessor configured to, from the digital model: obtain a position in the blade height of the movable blade, as a function of a deformation of the blade in a first natural torsional mode, modify the geometry of a blade section at said position, the modification comprising the following operations: - stretch the blade section by lengthening a chord connecting a leading edge of the section to a trailing edge of the section, and - reduce the thickness of the stretched section to maintain a constant cross-sectional area with the blade before modification, and connect the modified section to a root section and a tip section of the blade before modification, so as to obtain a modified digital model of the blade, and use the modified digital model to control a mobile blade production equipment. Brief description of the drawings

[0038] Other objects, features and advantages of the invention will become apparent from the following description, given solely by way of non-limiting example, and made with reference to the accompanying drawings in which:

[0039] [Fig. 1] represents a simplified Campbell diagram identifying a natural mode of a part in relation to an excitation harmonic;

[0040] [Fig.2] illustrates an example of the realization of a turbomachine structure, here a twin-flow, twin-body turbomachine;

[0041] [Fig.3] is a partial longitudinal cross-sectional view of a low-pressure turbine of an aircraft turbomachine;

[0042] [Fig.4] is a view of a blade of a low pressure turbine;

[0043] [Fig. 5] A flowchart of a method for modifying the geometry of a blade aircraft turbomachine mobile, according to an example of an embodiment of the invention;

[0044] [Fig.6] illustrates the modification of the blade according to an example of an embodiment of the invention;

[0045] [Fig.7] illustrates details of modification of a relevant section of the blade according to an example of the implementation of the invention;

[0046] [Fig.8] illustrates, by superimposition, an initial blade section and a blade section modified 602” according to an example of an embodiment of the invention;

[0047] [Fig.9] superimposes the thickness curves of the initial and modified blade sections according to an example of an embodiment of the invention; and

[0048] [Fig. 10] illustrates a system for manufacturing movable blades or blades for movable blades.

[0049] For the sake of clarity, the same elements are designated by the same reference numerals in the different figures. Furthermore, the various figures are not drawn to scale, as is customary in schematic representations. Detailed description of at least one embodiment

[0050] In the following description, when reference is made to absolute positional qualifiers, such as "front," "back," "top," "bottom," "left," "right," etc., or relative positional qualifiers, such as "above," "below," "superior," "inferior," etc., or to orientational qualifiers, such as "horizontal," "vertical," etc., unless otherwise specified, reference is made to the orientation of the figures or to a component / device in a normal operating position. Unless otherwise specified, the expressions "about," "approximately," "roughly," and "in the order of" mean to the nearest 10%, preferably to the nearest 5%.

[0051] Fig. 2 schematically illustrates the structure of a twin-body, double-flow turbomachine.

[0052] The double-flow turbomachine 2 comprises successively, in the direction of air circulation, i.e. from upstream (left in the figure) to downstream (right in the figure), an air inlet 20 and a blower 21, which delivers air on the one hand into a primary channel 22 and on the other hand into a secondary channel 23. By "channel", we mean the volume through which an airflow circulates.

[0053] The airflow circulating in the primary stream 22 passes successively through a low-pressure compressor 24a, a high-pressure compressor 24b, a combustion chamber 25, a high-pressure turbine 26a, and a low-pressure turbine 26b, before being ejected through a primary flow nozzle 220, providing the thrust generated by the turbomachine. Alternatively, the turbomachine 2 may comprise two counter-rotating low-pressure turbines, i.e., turbines that rotate in opposite directions around the longitudinal axis X of the turbomachine.

[0054] Furthermore, the secondary airflow which flows in the secondary vein 23 is ejected separately through a secondary flow nozzle 230, after passing through a series of guide vanes 231.

[0055] Whether in the blower 21, the compressors 24a, 24b and the turbines 26a, 26b, these components include movable blades or rotors. Those of the turbines are driven in rotation by the flow of gases in the primary flow nozzle 220. Those of the blower and the compressors are driven in rotation, in particular, by the energy recovered by the turbines.

[0056] Fig. 3 partially represents in longitudinal section a low-pressure turbine 1 of an aircraft turbomachine, representative of turbine architectures comprising moving blades or rotors, and fixed blades or distributors.

[0057] It is understood that the present invention applies to any other aircraft turbomachine assembly equipped with movable blades.

[0058] Figure 3 more precisely represents five stages of a low-pressure turbine 31. Regarding the orientation with respect to the general direction of gas flow through the turbine, the first stage of the turbine is located upstream and the last stage of the turbine is located downstream.

[0059] The first stage comprises a movable wheel formed of a rotor disk 32 on which a plurality of movable blades 33 are radially mounted. The other stages each comprise a distributor formed of a plurality of fixed blades 34 and a movable wheel placed behind the distributor and formed of a rotor disk 32' on which a plurality of movable blades 33' are radially mounted.

[0060] The discs 32, 32' are fixed to each other by means of bolted links 35, making all the discs fixed in rotation around the longitudinal axis XX of the turbomachine.

[0061] Each disk 32, 32' has at least one annular flange 6 which extends towards the disks 32, 32' of the adjacent floors.

[0062] Figure 4 shows a movable blade 33, which comprises a blade 330 extending radially in a Z direction between a blade tip 331 and a blade root 332, longitudinally between a leading edge BA and a trailing edge BF, and tangentially between an upper surface 334 and an lower surface 333, only the lower surface being visible in the figure. The Z direction is a radial direction of the blade 330 which is equivalent to a direction perpendicular to the X axis of rotation of the rotor blade when the blade 33 is mounted on it.

[0063] A movable blade 33, and more particularly its blade 330, is designed taking into account aerodynamic performance, but also its natural modes and the excitation frequencies related to the upstream gas flow in the engine. In particular, a blade geometry is sought that has a first torsional mode frequency (mode 1T) far removed from the engine's excitation harmonics, i.e., with sufficient dynamic margins with the excitation harmonics.

[0064] The design of this blade geometry can be carried out by computer, resulting in the creation of a digital model of the movable blade 33 and more particularly of its blade 330.

[0065] By way of illustration, a blade modeling by finite elements, for example by 3D triangular or tetrahedral mesh, can be used.

[0066] For environmental reasons related to technological research efforts, it is preferable to reuse existing moving blade geometries in new generations of aircraft engines. However, the evolution of the turbomachine architecture leads to a modification of the excitation frequencies related to the gas flows upstream of the blades.

[0067] The method according to the invention for manufacturing and producing moving blades proposes a modification of the geometry of the already existing moving blade to significantly lower its frequency of the 1T mode and thus get out of a situation at risk of resonance.

[0068] We will now describe with reference to [Fig.5] such a method 5 of modifying the geometry of a movable blade of an aircraft turbomachine, according to an example of implementation.

[0069] This method can be implemented using a computer in a computer-aided design (CAD) approach, starting from a digital model of the existing blade. The result of the design – a modified digital model of the blade – is used to control equipment or machines for the production of blades and, more generally, moving blades, for example to control the manufacture of a foundry mold or to control a 3D printer.

[0070] In step 50, a position is obtained within the blade height of the movable blade 33 to be modified, based on a deformation of the blade in the first natural torsional mode. The objective of this step is to identify a section of the blade 330 where geometric modifications of the cross-sectional profile have a significant effect on the frequency of the 1T mode. Therefore, the deformation zones of the blade in this 1T mode are taken into account.

[0071] In practice, it has been observed that the blade area with the greatest resonant displacement (generally on the order of one hundredth to one tenth of a millimeter for conventional aircraft turbomachine blades) is the trailing edge BF at approximately 45 to 55% of the blade height, with 0% corresponding to the blade root and 100% to the blade tip. A height H can therefore be chosen within this range of values, for example, at approximately 50% of the total height.

[0072] Furthermore, measurements on an actual blade or computer simulations make it possible to obtain a blade deformation profile. In this case, a height H can be chosen where the blade deformation is significant, i.e., greater than a predefined threshold. A particular case is that in which the height corresponding to the maximum blade deformation 330 is chosen.

[0073] Typically, a computer-aided design tool can define the blade 330 through several blade sections, including a blade tip section at the upper end of the blade, a blade root section at the lower end of the blade, and one or more intermediate blade sections at different height levels.

[0074] [Fig. 6] Section (a) illustrates, for example, a blade 60 defined by three blade sections, including a blade tip section 600, a blade root section 601, and a single intermediate section 602 at mid-height (50%). Of course, other embodiments with a greater number of intermediate sections (for example, every 25%, 20%, 10%, or 5% of the height) can be considered.

[0075] The blade profile 610 between successive sections can follow a profile transition law. For example, the profile transition law can be a linear law. Alternatively, more complex laws for smoothing the profile while avoiding significant geometric changes can be considered.

[0076] The position obtained in the blade height can, for example, correspond to the position of the intermediate section closest to the maximum deformation of the blade (in its BF zone), i.e. at mid-height (H=50%) in the example of [Fig.6].

[0077] Once position H is obtained and consequently the blade cross-section 602 at this position H is known, step 52 consists of modifying the geometry of the blade cross-section, while maintaining the same surface area. The objective of this step is to move the center of gravity of the cross-section upstream (i.e., towards the leading edge BA), without altering its mass or its general shape, which defines its aerodynamic performance. In this way, the frequency of mode 1T can be significantly lowered without modifying the static / mechanical stresses of the blade, thus avoiding a situation at risk of resonance.

[0078] The modification step 52 comprises two successive substeps 520 and 522, illustrated by [Fig.7].

[0079] Substep 520 consists of stretching the section 602 by lengthening the chord 700 of the section profile, that is to say, increasing the length of the segment connecting the leading edge BA of the section to the trailing edge BF of the section. A stretched section 602' is obtained.

[0080] Several stretching functions can be used.

[0081] For example, an affinity can be established with the direction of the chord axis (direction axis) and an affinity ratio X greater than 1, meaning that section 602 undergoes increasing homothety only along the axis XI formed by the chord 700, the component along the axis Y1 perpendicular to the chord remaining unchanged (identity function). Taking the reference frame (O,X1,Y1) of [Fig.7] section (a), the geometric transformation of each point P(x,y) forming the profile of section 602 in this reference frame leads to a point P'(Xx,y) represented on [Fig.7] section (b).

[0082] The affinity (or homothety) ratio X defines the elongation of the chord 700, and therefore of the segment BA-BF. This ratio is between 5 and 20%. Preferably, it is chosen between 10 and 15%, for example 10% as illustrated in the Figure.

[0083] The basis of the affinity is a straight line A that intersects the chord 700 perpendicularly (i.e., the segment BA-BF), at a point denoted C. A represents the points invariant under the implemented geometric affinity. For example, C is the leading edge BA, or the trailing edge BF, or the foot of the height passing through the mid-section 701 (at its midpoint, as in the Figure) corresponding to the maximum thickness of the section along a normal to the skeleton (or mean or median line) 702 of the section 602, or the foot of the height passing through the center of gravity CGI of the section 602, or any other location on the chord.

[0084] Each point P is therefore transformed by homothety with ratio X and center the point of intersection between A and the parallel to the chord 700 passing through P.

[0085] As an alternative to a general affinity, a local affinity can be used where the direction of the affinity is the axis defined by two successive control points of the skeleton 702 of the blade section. Thus, each "subsection" of the blade section 602 is modified by its own affinity, the direction of which varies. The affinity ratio X remains the same.

[0086] [Fig. 7] Section (a') illustrates, for example, the blade section 602 with its skeleton 702 on which PCX control points have been positioned. The PCX control points can be defined at regular intervals, for example every 5% or 10% of the skeleton 702 between the leading edge and the trailing edge.

[0087] The blade section is for example constructed by defining two profile points PI, P2 (an intrados point and an extrados point) for each of the control points PC, typically on a normal to the skeleton and at a defined distance (which varies for each PC point).

[0088] A local affinity at the subsection between control point PC; and control point PCj consists of using the axis (PC; PCj) as the direction of the homothety, simply increasing the distance PC; PCj. As illustrated in [Fig.7] section (b'), the distance d of the segment (PC; PCj) is stretched to equal kd in the stretched section 602'. Both points P1 and P2 are maintained at a (defined) distance from their control point PCX.

[0089] By way of illustration, for conventional turbomachine blades, the elongation will be on the order of a few millimeters at most, typically between 0.5 mm and 2 mm.

[0090] The minimum value of 5% ensures a substantial reduction in the frequency of mode 1T as desired for the invention. The maximum value of 20% ensures compliance with integration constraints such as the presence of other parts in the vicinity of the blade 33 and manufacturing constraints such as the absence of angular joints at the blade root or tip.

[0091] Once the stretched section 602' is obtained ([Fig.7] section (b) or (b')), the thickness of the stretched section 602' is reduced in substep 522 to maintain a constant cross-sectional area with the blade before modification. In other words, if the area of ​​the section 602 before modification is S and that of the section 602' after stretching is S'>S, the reduction 522 consists of reducing the area of ​​the stretched section 602' to S.

[0092] For aerodynamic performance reasons related to the blade 33, in some embodiments, a modification of the upper surface profile 334 is preferred, while keeping the lower surface profile 333 unchanged. Other aerodynamic performance improvements can be achieved by modifying only the lower surface profile 333 while keeping the upper surface profile 334 unchanged.

[0093] Similarly, for similar aerodynamic performance considerations, preference is given to a reduction in thickness which does not alter the radii of curvature of the leading edge BA and trailing edge BF.

[0094] Typically, the mid-section 701 of the stretched section 602 can be identified, and then the thickness e of the section at the mid-section 701 can be reduced by applying a thickness law between the mid-section and each of the leading and trailing edges BA. BF (while not modifying the extrados profile in certain embodiments), until a surface S' of the stretched section 602 is obtained equal to a surface S of the section 602 before modification.

[0095] Preferably, the thickness law is that used in the original design of the blade 33 before modification. Of course, other laws may be used.

[0096] A modified section 602'' is obtained, as illustrated in [Fig. 7], section (c) or (c') in which only the intrados profile 333 has been modified. [Fig. 8] also illustrates, by superimposition, the original section 602 and the modified section 602'' resulting from step 52. This figure illustrates the 10% stretching of the section.

[0097] By modifying the geometry of section 602, the center of gravity CGI of section 602 has shifted upstream (towards the leading edge BA, CG2 in [Fig. 7] section (c)), reducing the local mass and stiffness on the trailing edge side BF where the resonant deformation is greatest. The frequency of mode 1T is thus reduced, without altering the static / mechanical stresses of the blade 33 since the mass (surface area) remains unchanged.

[0098] Figure 9 shows a diagram superimposing the thickness (in mm) of the original 602 and modified 602 sections as a function of their position (in mm) along the X-axis of the motor (the origin being the leading edge BA). This diagram shows a reduction of approximately 5% in the greatest thickness (between the main frames 701) as well as a displacement of the main frame upstream (BA), from an initial position located at approximately 37% of the total length along the motor axis to a final position located at approximately 23% of the total length along the motor axis. A displacement of approximately 14 points is thus obtained. The center of gravity is consequently also displaced upstream by approximately 10 to 20%.

[0099] Returning to [Fig. 5], once the modified section 602 is obtained, step 54 consists of connecting the modified section to the root section 601 (at 0% height) and the tip section 602 (at 100% height) of the blade 330 before modification. That is, the root and tip sections remain identical to those they were in the original blade 330.

[0100] Connecting consists of linking the section profiles using a profile connecting law, typically the same profile connecting law that was used to establish the 610 profile of the original blade. Of course, other laws may be used.

[0101] To perform the connection, the different sections are superimposed (at their respective heights) with their centers of gravity aligned. Section (a) [Fig. 6] shows the alignment of the centers of gravity (represented by crosses, including CGI of section 602) of the sections forming the original blade 60, on the axis of gravity 699.

[0102] Also, in embodiments, during step 54, the center of gravity CG2 of the modified section 602” ([Fig.6] section (b)) is determined and aligned with the axis of gravity 699 ([Fig.6] section (c)). This alignment moves the modified section 602” further downstream, since the center of gravity CG2 is moved upstream. Only then is the profile matching law applied between the sections to obtain the modified blade.

[0103] Aligning the centers of gravity makes it possible not to modify the mechanical stresses of the blade.

[0104] As shown schematically in [Fig.6] section (c), the geometric profile of the blade (in solid lines 610) is modified compared to the geometric profile of the initial blade (in dashed lines), due to the modification made in section 602.

[0105] This gives us a modified numerical model of a blade or moving blade corresponding to the modified blade 60'.

[0106] This digital model is then used in step 56 in the production of blades, for example for the manufacture of foundry molds or to control a 3D printer. The digital model is therefore used to control a movable blade production equipment.

[0107] A plurality of blades 33 with blade 60' thus modified are manufactured and installed on a rotor in the blower 21, the compressors 24a, 24b and / or the turbines 26a, 26b.

[0108] Figure 10 illustrates a system 10 for manufacturing movable blades or blades for movable blades. The system includes blade design equipment 11 and blade production equipment 12.

[0109] The blade design equipment 11 can be a simple computer with a microprocessor 110 and non-volatile memory 111. Typically, a computer has other components (for example, volatile memory, communication interface, input / output interface, etc.) not shown here for reasons of brevity.

[0110] The memory 111 stores an initial digital model 112 of a moving blade whose geometry is modified according to the teachings of the present invention.

[0111] The microprocessor 110 is configured to implement these teachings, and in particular the process of [Fig.5], a process in which a modified digital model 113 is produced by modifying the geometry of the blade defined in the initial digital model 112. The modified digital model 113 is stored in memory 111.

[0112] The modified digital model 113 is communicated to the blade production equipment 12 by any means. For example, the microprocessor 110 controls the equipment 12 via a communication network and provides it with the modified digital model 113 to produce new moving blades, for example to manufacture or configure foundry molds corresponding to the modified digital model or the control of a 3D printer. New blades and moving vanes are thus manufactured.

[0113] The invention makes it possible to drastically reduce (by about 15%) the frequency of the first torsion mode of the moving blades of an aircraft turbomachine, and therefore the associated coincidence regime. The frequency of mode 1T is thus sufficiently low that coincidence does not occur in the usual operating ranges of the engine with the new holding architecture (for example, during cruise, takeoff, or landing).

[0114] In particular, it has been observed that a 10% increase in the chord length of the mid-section of the 330 blade, combined with a reduction in thickness at the mid-section while maintaining the original thickness law, leads to a 15% decrease in the frequency of mode 1T. This decrease is sufficient to break out of dynamic coincidence under normal engine operating conditions.

[0115] An advantage linked to lowering the frequency of mode 1T is also that the excitation regime is lowered and therefore the excitation pressure at the origin of the aerodynamic forces seen by the blade is reduced: the vibratory potential of the blade is thus improved.

[0116] Of course, the present invention is not limited to the embodiments described above by way of example; it extends to other variants.

[0117] In particular, if the invention is described above in connection with the modification of a single section (at mid-height in some examples), it can be implemented for the modification of a plurality of blade sections (for example at 25%, 50%, 75% of height, or 45%, 50% and 55% of height to be more precise in the area of ​​high deformation) before connecting the sections together.

Claims

Demands

1. A method for manufacturing a movable blade (33) for an aircraft turbomachine (2), comprising the following steps: obtaining (50) a position (H) in the blade height (330, 60) of a movable blade (33), as a function of a deformation of the blade in a first natural torsion mode, modifying (52) the geometry of a blade section (602) at said position, the modification comprising the following operations: - stretching (520) the blade section (602) by lengthening a chord (700) connecting a leading edge (BA) of the section to a trailing edge (BF) of the section, and - reducing (522) the thickness (e) of the stretched section (602') to maintain a constant cross-sectional area (S) with the blade before modification, and connecting (54) the modified section (602") to a root section (601) and a tip section (600) of the blade before modification, so as to obtain a modified blade (60').

2. Method according to claim 1, wherein obtaining (50) the position (H) in the blade height comprises determining a position where the deformation of the blade is significant relative to a threshold, or even maximum, in a first natural torsion mode.

3. A method according to claim 1 or 2, wherein the position in the blade height is located between 45 and 55% of the blade height, preferably at mid-height.

4. A method according to any one of the preceding claims, wherein the rope (700) is lengthened by 5 to 20%, preferably by 10 to 15%, for example by about 10%.

5. Method according to claim 4, wherein the stretching (520) of the blade section (602) is carried out using an affinity having direction along the axis (XI) of the chord (700) or the axis of a section between two control points (PC) of a skeleton (702) of the blade section.

6. A method according to any one of the preceding claims, wherein reducing (522) the thickness (e) comprises: identifying a midframe (701) in the stretched section (602'), and reducing the thickness of the section at the midframe by applying a thickness law between the midframe and each of the leading edges (BA) and trailing edges (BF), until a surface area of ​​the stretched section is equal to a surface area of ​​the section before modification.

7. A method according to any one of the preceding claims, wherein reducing (522) the thickness (e) includes modifying an extrados profile (334) of the blade (330, 60) in the stretched section (602'), while keeping an intrados profile (333) of the blade unchanged.

8. A method according to any one of the preceding claims, wherein the radii of curvature of the leading (BA) and trailing (BF) edges are kept constant during the modification (52) of the geometry of the blade section (602).

9. A method according to any one of the preceding claims, wherein connecting (54) the modified section (602”) to the root (601) and tip (600) sections of the blade comprises: determining a center of gravity (CG2) of the modified section (602”), aligning the center of gravity with a gravity axis (699) of the blade, and applying a profile matching law between the section aligned with the gravity axis and the root and tip sections of the blade.

10. Movable blade (33) of an aircraft turbomachine, the geometry of which is modified by a method according to any one of claims 1 to Q

11. d y. System (10) for manufacturing a movable blade (33) for an aircraft turbomachine (2), the system comprising: a memory (111) for storing a digital model (112) of the movable blade (33), a microprocessor (110) configured to, from the digital model (112): obtain a position (H) in the blade height (330) of the movable blade, as a function of a deformation of the blade in a first natural torsion mode, modify the geometry of a blade section (602) at said position, the modification comprising the following operations: - stretching the blade section (602) by lengthening a chord (700) connecting a leading edge (BA) of the section to a trailing edge (BF) of the section, and - reducing the thickness (e) of the stretched section (602') to maintain a constant cross-sectional area with the blade before modification, and connect the modified section (602”) to a foot section (601) and a tip section (602) of the blade before modification, so as to obtain a modified digital model (113) of the blade blade, and use the modified digital model to drive a mobile blade production equipment (12).

Images