EP4183980A1 - Blade for a turbomachine and turbomachine comprising at least one blade - Google Patents

Abstract

The invention relates to a blade (1) for a turbomachine (2), having a blade inner profile (5) and a jet deflection section (6) which adjoins the blade inner profile (5) in a radial direction (r) with respect to a rotor shaft (4). wherein respective blade profiles (7) of the jet deflection section (6) oriented normal to the radial direction (r) have respective centers of gravity (8), the centers of gravity (8) relative to a reference line (9) running linearly along the radial direction (r) by respective axial variation values (dx ) and/or which are shifted by respective tangential variation values (dt) in a tangential direction (t), the respective centers of gravity (8) being connected by a threading line emanating from the blade inner profile (5). It is provided that the threading line (16) is a spatial line, with a total deviation (17) of the spatial line in the blade profiles (7) from the reference line (9) depending on a respective radial distance of the blade profile (7) from a reference point by a Polynomial of the nth order is described, where n is a natural number greater than 1, the threading line being selected and designed such that there is a compressive stress during operation in the region of the leading edge of the blade.

Description

The invention relates to a blade for a turbomachine according to the features of the preamble of claim 1, in particular for an aircraft engine and a turbomachine according to the features of the preamble of claim 8.

According to the state of the art, blades for compressors are mainly optimized to maximize the efficiency and to maintain a corresponding surge limit distance, taking into account the structural mechanical strength requirements.

A disadvantage of the blades optimized according to this is their susceptibility to damage from foreign bodies which are sucked into the turbomachine and impact on the blades. Damage can occur as a result of these impacts, in particular on the leading edges of the blades. There is a risk of foreign objects being sucked in by birds or objects on the runway, in particular during the take-off and landing phases of an aircraft. In the event of damage, these areas limit the further operation of the engine and make unscheduled maintenance of the affected components necessary.

According to the prior art, when designing blades, it is common to generate two-dimensional profiles that are optimized for a flow pattern. These profiles lie in a lateral surface of a cylinder coordinate system related to a turbomachine. In order to generate a 3-dimensional structure of the blade, it is usual to stagger the individual profiles aligned normally with respect to a reference line. The reference line runs essentially radially, starting from a rotor shaft. The reference line runs through respective characteristic points of the respective profiles, such as the respective centroids. It is common here to connect the focal points of the individual profiles by a threading line, which has a linear course along the radial direction.

In order to optimize certain properties, it is common to adapt this course of the threading line in such a way that the course of the threading line deviates from that of the linear reference line. The profiles are shifted so that the focal points are not on the radial and linear reference line. The stress distributions in the blades change under operating conditions due to the shifting of the centers of gravity.

The optimization of blades for turbomachines is disclosed by way of example in the following documents.

In the master's thesis submitted by Pablo Rodriguez-Fernandez to the Universidad Polytecnica de Madrid in 2015 entitled "Development of Shape-Optimization Tools for the Aerodynamic Design of Turbomachinery Blades", a development of shape optimization tools for an aerodynamic design of turbomachinery blades is disclosed. It is provided here to define a blade geometry using parameterization techniques on the basis of B-spline curves, which enable a shape of the blade to be defined locally.

The publication Dow, EA, & Wang, Q (2015). "The implications of tolerance optimization on compressor blade design. Journal of Turbomachinery", 137(10), 101008 . describes the impact of tolerance optimization on compressor blade design. The publication deals with the effects of the geometric variability of blades on the performance of the blades. Geometric variability increases performance variability and degrades mean performance of turbomachine compressor blades. These adverse effects can be mitigated through robust blade geometry optimization or tighter manufacturing specifications. The publication describes the leading edge stall that occurs in some blades when the radius of curvature of the leading edge is reduced beyond a critical value, thereby drastically increasing blade losses. The optimal leading edge geometry therefore depends on the level of manufacturing noise, which illustrates the possible interactions between geometry and tolerance optimization.

The 2011 Concordia University PhD thesis "Three dimensional aero-structural shape optimization of turbomachinery blades" by Sivashanmugam, VK describes the development and implementation of a structural shape optimization module and its integration with an aerodynamic shape optimization module into an automated aero-structural optimization process.

The of Dow, EA Dissertation submitted to the Massachusetts Institute of Technology in 2015 on the topic "Robust design and tolerancing of compressor blades " deals with the robust design of compressor blades.

The dissertation " Robust design methodologies: Application to compressor blades", submitted 2006 at the University of Southampton by Kumar, A. deals with the application of robust design methods to compressor blades. In the dissertation, the determination of compressor blade geometries is disclosed, which provide a robust performance even with geometric deviations from a given shape.

The publication Fagan EM, De La Torre O, Leen SB, & Goggins J (2018). Validation of the multi-objective structural optimization of a composite wind turbine blade. Composite structures, 204, 567-5t ; relates to a validation of the multi-criteria structural optimization of a wind turbine blade. The publication uses structural characteristics of the blade such as mass, center of gravity and results from static and modal tests to validate finite element analysis predictions for the custom blade

In the chapter Köller U., Van den Toorn B. (2010) Design, calculation and manufacture of compressor blades, published in: Lechner C., Seume J. (eds) Stationary turbomachines. VDI book. Springer, Berlin, Heidelberg . the refinement of blade profiles of a moving blade is disclosed.

It is an object of the invention to provide a blade which is less susceptible to foreign object damage than prior art blades.

The object is achieved according to the invention by a blade according to the features of claim 1 and a turbomachine according to the features of claim 8. Advantageous configurations with expedient developments of the invention are in the respective dependent claims specified, with advantageous configurations of each aspect of the invention being to be regarded as advantageous configurations of the respective other aspects of the invention.

A first aspect of the invention relates to a blade for a turbomachine. The blade has an inner blade profile, which can rest against a rotor element, and an airfoil section, which adjoins the inner blade profile in a radial direction with respect to a rotor shaft. The blade inner profile can be, for example, a profile of a hub, a blade root element or a first profile of the blade section, which is followed by the blade section of the blade in a radially outer direction. The coordinates of the turbomachine can be defined with respect to the rotor shaft of the turbomachine, it being possible for the axial direction to run parallel to a longitudinal direction of the rotor shaft. The blade can also be arranged on a disk of a blisk. A radial direction can run perpendicularly from the rotor shaft. A tangential direction can run in a circumferential direction around the rotor shaft. The blade section can be arranged on the blade inner profile facing away from the rotor shaft in the radial direction. The airfoil portion may comprise a jet turning portion of the airfoil. Respective blade profiles of the blade section, which are aligned normal to the radial direction, can have respective centers of gravity. In other words, in the airfoil section, the blade has blade profiles as cross sections which are oriented perpendicular to the radial direction. Each of the blade profiles has the respective center of gravity. It is provided that the centers of gravity are shifted relative to a reference line running linearly along the radial direction by respective axial variation values in an axial direction parallel to the rotor shaft. Additionally or alternatively, the centroids are shifted with respect to the reference line by respective tangential variation values in a tangential direction. In other words, it is provided that the centers of gravity of the respective blade profiles are shifted in relation to the reference line in the axial direction and/or the tangential direction by the respective axial variation values or the respective tangential variation values. The displacements of the respective centroids in the axial direction are described by the respective axial variation values, which are defined with respect to the reference line linearly extending along the radial direction. The displacements of the respective centers of gravity in the tangential direction are described by respective tangential variation values, which are defined with respect to the reference line running linearly along the radial direction. The respective focal points of the respective blade profiles are indicated by a threading line starting from the blade inner profile tied together. In other words, it is provided that the centers of gravity of the respective blade profiles are arranged on the threading line, which runs starting from the blade inner profile. The threading line can run, for example, starting from a center of gravity of the blade inner profile.

It is provided that the threading line is a spatial line, with a total deviation of the spatial line in the blade profiles from the reference line depending on a respective radial distance of the respective blade profile from the blade inner profile being described by a polynomial of the nth order with n>1. In this case, n is a natural number. In other words, the threading line describes a course of the positions of the centroids of the blade profiles as a function of the radial spacing of the blade profiles from the blade inner profile. The threading line describes the position of the respective center of gravity as a function of the radial distance of the center of gravity from an origin, which can be a center of gravity of a blade profile, the blade inner profile. The threading line is defined by a polynomial of order n, which has a degree greater than 1. According to the invention, the threading line is arranged so that under operating conditions a configuration is realized which effects compressive stress on a leading edge of the blade. In other words, a negative tension can be provided at the leading edge of the blade during operation as a result of the aforesaid course of the threading line. This results in the advantage that crack propagation in the leading edge is prevented because compressive stresses counteract crack propagation. Thus, the area of the blade primarily at risk from foreign object impact (FOD) is less susceptible to damage than is the case with prior art blades.

The invention also includes developments that result in further advantages.

According to one embodiment of the invention, at least a maximum of the polynomial shape of the threading line may be displaced axially rearwardly, i.e. downstream, relative to a purely radially oriented reference line. All maxima of the threading line can preferably be offset radially to the rear in relation to the reference line.

A development of the invention provides that the respective tangential variation values lie in a range between -5 degrees and +5. In other words, it is provided that the focal points of the respective blade profiles deviate between -5 degrees and 5 degrees from the reference line. The deviation can therefore be in a range from -5 degrees to +5 degrees. In other words, the tangential deviation between the centroids and the reference line can be -5.0°, -4.9°, -4.8°, -4.7°, -4.6°, -4.5°, -4 .4°, -4.3°, -4.2°, -4.1°, -4.0°, -3.9°, -3.8°, -3.7°, -3.6 °, -3.5°, -3.4°, -3.3°, -3.2°, -3.1°, -3.0°, -2.9°, -2.8°, -2.7°, -2.6°, -2.5°, -2.4°, -2.3°, -2.2°, -2.1°, -2.0°, -1 .9°, -1.8°, -1.7°, -1.6°, -1.5°, -1.4°, -1.3°, -1.2°, -1.1 °, -1.0°, -0.9°, -0.8°, -0.7°, -0.6°, -0.5°, -0.4°, -0.3°, -0.2°, -0.1°, 0°, 0.1°, 0.2°, 0.3°, 0.4°, 0.5°, 0.6°, 0.7°, 0.8°, 0.9°, 1.0°, 1.1°, 1.2°, 1.3°, 1.4°, 1.5°, 1.6°, 1.7°, 1.8°, 1.9°, 2.0°, 2.1°, 2.2°, 2.3°, 2.4°, 2.5°, 2.6°, 2.7°, 2.8°, 2.9°, 3.0°, 3.1°, 3.2°, 3.3°, 3.4°, 3.5°, 3.6°, 3.7°, 3.8°, 3.9°, 4.0°, 4.1°, 4.2°, 4.3°, 4.4°, 4.5°, 4.6°, 4.7°, 4.8°, 4.9°, 5.0°

A development of the invention provides that the respective tangential variation values lie in a range between -2.5 degrees and 2.5 degrees. In other words, it is provided that the centers of gravity of the respective blade profiles deviate from the reference line by a range of −2.5° to 2.5°. The deviation can therefore be in a range from -2.5 degrees to +2.5 degrees. In other words, the tangential deviation between the centroids and the reference line can be -2.5°, -2.4°, -2.3°, -2.2°, -2.1°, -2.0°, -1 .9°, -1.8°, -1.7°, -1.6°, -1.5°, -1.4°, -1.3°, -1.2°, -1.1 °, -1.0°, -0.9°, -0.8°, -0.7°, -0.6°, -0.5°, -0.4°, -0.3°, -0.2°, -0.1°, 0°, 0.1°, 0.2°, 0.3°, 0.4°, 0.5°, 0.6°, 0.7°, 0.8°, 0.9°, 1.0°, 1.1°, 1.2°, 1.3°, 1.4°, 1.5°, 1.6°, 1.7°, 1.8°, 1.9°, 2.0°, 2.1°, 2.2°, 2.3°, 2.4°, 2.5°.

A development of the invention provides that the respective axial variation values are between 0 and 20% of a respective chord length of the respective blade profile. In other words, a maximum deviation in the axial direction is at most 20% of the respective chord length of the respective blade profile, which connects a leading edge to a trailing edge of the respective profile. For example, the axial variation values may be 0%, 1%, 2%, 3%, 4%, 5%, 6%, 7%, 8%, 9%, 10%, 11%, 12%, 13%, 14%, 15 %, 16%, 17%, 18%, 19%, 20% of the respective chord length of the respective blade profile.

A development of the invention provides that the respective axial variation values are between 0 and 10% of a respective chord length of the respective blade profile. In other words, a maximum deviation in the axial direction is at most 10% of the respective chord length of the respective blade profile, which has a leading edge with a trailing edge of the respective profile connects. The axial variation values can be, for example, 0%, 1%, 2%, 3%, 4%, 5%, 6%, 7%, 8%, 9%, 10% of the respective chord length of the respective blade profile.

A development of the invention provides that the blade is designed as a moving blade. In other words, the blade is a rotating blade that is provided for energy transmission between the turbomachine and a fluid. Accordingly, the blade may be intended to be mounted on a rotor which, in use, rotates about its longitudinal direction, causing stresses in the blade.

A development of the invention provides that respective stagger angles of at least some of the blade profiles of respective moving blades deviate from a nominal stagger angle of a blade inner profile by respective stagger angle variation values, the respective stagger angle variation values being at most 2 degrees. In other words, the respective stagger angles differ by at most 2 degrees from the nominal stagger angle. The stagger angle variation values can range from -2 degrees to +2 degrees. The stagger angle is defined as the angle between the chord of the respective blade profile and the direction of rotation. The nominal stagger angle can be the stagger angle of the inner blade profile and can thus describe the angle between the chord of the inner blade profile and the direction of rotation. For example, the stagger angle variation values may be -2.0°, -1.9°, -1.8°, -1.7°, -1.6°, -1.5°, -1.4°, -1.3 °, -1.2°, -1.1°, -1.0°, -0.9°, -0.8°, -0.7°, -0.6°, -0.5°, -0.4°, -0.3°, -0.2°, -0.1°, 0.0°, 0.1°, 0.2°, 0.3°, 0.4°, 0 .5°, 0.6°, 0.7°, 0.8°, 0.9°, 1.0°, 1.1°, 1.2°, 1.3°, 1.4°, 1 .5°, 1.6°, 1.7°, 1.8°, 1.9°, 2.0°.

über die unteren 90% der radialen Schaufelerstreckung angestrebt. Dieser Zielwert kann zusätzlich zur Anpassung der Fädelungsstrategie auch durch die Relativänderung der Staffelungswinkel naheliegender Profilschnitte zueinander erreicht werdenSince the generation of a compressive stress as the main stress with the highest magnitude cannot always be guaranteed during operation, the adaptation of the threading strategy can also be used in a preferred development to positively influence the ratio of the leading edge stress to the mean stress in individual profile sections of the blade. A voltage ratio is used as the target value for a FOD/DOD resistant design ${\theta }_{\mathit{ratio}}=\raisebox{1ex}{${\mathrm{<figure-callout\; id="1"\; label="\sigma "\; filenames="imgf0004.png"\; state="\{\{state\}\}">\sigma </figure-callout>}}_{1,\mathit{LE}}$}\!\left/ \!\raisebox{-1ex}{${\sigma }_{\mathit{ave}}$}\right.\le 0.5$

is aimed for over the lower 90% of the blade radial extent. In addition to adapting the threading strategy, this target value can also be achieved by changing the staggering angles of nearby profile sections relative to one another

A second aspect of the invention relates to a turbomachine, in particular an aircraft engine, which has at least one blade of the first aspect of the invention.

Additional features and their advantages can be found in the descriptions of the first aspect of the invention.

• FIG. 1 eine schematische Darstellung eines Verlaufs einer Bezugslinie in einer Schaufel nach dem Stand der Technik;
• FIG. 2 zeigt eine schematische Darstellung einer Spannungsverteilung in der in FIG. 1 gezeigten Schaufel nach dem Stand der Technik;
• FIG. 3 zeigt eine schematische Darstellung einer Schaufel; und
• FIG. 4 zeigt eine schematische Darstellung einer Spannungsverteilung in der in FIG. 3 gezeigten Schaufel.

Further features of the invention result from the claims, the figures and the description of the figures. The features and combinations of features mentioned above in the description, as well as the features and combinations of features mentioned below in the description of the figures and/or shown alone in the figures, can be used not only in the combination specified in each case, but also in other combinations, without going beyond the scope of the invention leave. The invention is therefore also to be considered to include and disclose embodiments that are not explicitly shown and explained in the figures, but that result from the explained embodiments and can be generated by separate combinations of features. Versions and combinations of features are also to be regarded as disclosed which therefore do not have all the features of an originally formulated independent claim. Furthermore, embodiments and combinations of features, in particular through the embodiments presented above, are to be regarded as disclosed which go beyond or deviate from the combinations of features presented in the back references of the claims. It shows:

• FIG. 1 a schematic representation of a course of a reference line in a blade according to the prior art;
• FIG. 2 shows a schematic representation of a stress distribution in the in FIG. 1 prior art blade shown;
• FIG. 3 shows a schematic representation of a blade; and
• FIG. 4 shows a schematic representation of a stress distribution in the in FIG. 3 shown shovel.

FIG. 1 shows a schematic representation of a course of a reference line in a blade according to the prior art. The blade 1 can in particular be a moving blade 1 for a turbomachine 2, in particular for an aircraft engine. The coordinate system 3 can be defined in cylindrical coordinates in relation to a rotor shaft 4 not shown in the figure. An axial direction x can be aligned parallel to a longitudinal direction of the rotor shaft 4 . A radial direction r can be aligned radially from a center of the rotor shaft 4 . A tangential direction t can be aligned in a direction of rotation around the rotor shaft 4 . The blade 1 can be fastened to the rotor shaft 4 . The blade 1 can be provided for radial arrangement on the rotor shaft 4 . A blade inner profile 5 can be adjoined in the radial direction by a jet deflection section 6 which can extend to a tip of the blade 1 at a radial outer end of the blade 1 . The blade 1 can have blade profiles 7 in the jet deflection section 6 , which can be arranged in respective tangential planes with respect to the rotor shaft 4 at predetermined radial distances from the blade inner profile 5 . The blade profiles 7 can have respective centers of mass 8 which can be arranged along a linear and radial reference line 9 . The shape of the blade profiles 7 can differ from one another. The blade 1 may have a leading edge 11 and a trailing edge 10 . The leading edge 11 can be defined in relation to a flow direction in such a way that the gas can impinge on the leading edge 11 when the turbomachine 2 is in operation. Due to the design of the blade 1 according to the prior art for operating efficiency, a stress distribution during operation on the leading edge 11 can be positive, whereby damage and crack propagation can occur particularly on the leading edge 11 in the event of an impact of a foreign object.

A circumferential and axial position of the blade center of gravity 8 in the hub section is defined by the reference line 9, a purely radial threading of all other blade profiles 7 being provided. The focal points 8 are arranged on the reference line 9 . In FIG. 1 a reference configuration for the adjustment is defined.

FIG. 2 shows a schematic representation of a stress distribution in the in FIG. 1 shown blade 1 according to the prior art. A distribution of an amount of a largest is shown Main stress 12 during operation of the blade 1 in the turbomachine 2 at a predetermined operating point. A maximum amount of the main stress 12 can be seen in an inner region 13 of the blade 1, which can be a tensile stress. Except for a trailing edge area 20 of the trailing edge 10, a tensile stress dominates in the blade 1. The trailing edge area 20 of the trailing edge 10 has a compressive stress, which is negative. The negative stress results in the blade 1 being less susceptible to crack propagation in the trailing edge region 20 . Crack propagation can occur, for example, due to impacting foreign bodies. It would therefore be advantageous if the front edge 11 of the blade 1, which is at risk from foreign bodies, would have a compressive stress over an entire length of the jet deflection section 6 or at least a partial length 15 of 90% of the length of the jet deflection section 6, starting from the blade inner profile 5.

FIG. 3 shows a schematic representation of a blade 1. Shown is a blade 1, which compared to that in FIG FIG. 1 The prior art blade 1 shown was modified by varying the locations of the centers of gravity 8 in order to provide a negative principal stress 12 at a leading edge 11 of the blade 1 . In order to provide the compressive stress as the main stress 12 at the leading edge 11, the position of the blade profiles 7 can be varied, with the centers of gravity 8 of the blade profiles 7 being able to be offset with respect to the linear, radially running reference line 9. The displacements can take place along the tangential direction t and/or the axial direction x. As a result of the shifts, the centers of gravity 8 can be shifted by respective axial variation values dx in an axial direction x and/or by respective tangential variation values dt in a tangential direction t in relation to the reference line 9 . The individual focal points 8 can be connected by a threading line 16, which can thus differ from the reference line 9. A course of the threading line 16 can be described by a polynomial of the nth order, where n can be a natural number greater than 1. The threading line 16 can describe a total deviation 17 of the centers of gravity 8 at a respective radial distance depending on the radial distance of the center of gravity 8 from the blade inner profile 5 . Provision can be made for the respective tangential variation values dt to be in a range between −5 degrees and 5 degrees, in particular in a range between −2.5 degrees and 2.5 degrees. In other words, respective centroids 8 can be shifted in the tangential direction between -5 degrees and +5 degrees, in particular between -2.5 degrees and +2.5 degrees with respect to the reference line 9 .

For the axial variation values dx, it can be provided that the amounts of the respective axial variation values dx amount to at most 20%, in particular at most 10% of a respective length of a profile chord 18 of the respective blade profile. The chord 18 of a respective blade profile 7 can connect a profile leading edge of the blade profile 7 to a trailing edge of the blade profile 7 . In other words, the axial variation dx is no more than 20% of the length of the respective chord 18 of the respective blade profile 7. The reference line 9 and the threading line 16 can run at least through a first of the centers of gravity 8, which is the center of gravity 8 of the blade inner profile 5 can act. The center of gravity 8 of the blade inner profile 5 can thus be a common origin 19 of the reference line 9 and the threading line 16 . The respective blade profiles 7 can have stagger angles which describe an angle between the profile chord 18 of the respective blade profile 7 and the direction of rotation t. The stagger angles can differ from a nominal stagger angle A by stagger angle variation values dA, which can describe an angle between the profile chord 18 of the blade inner profile 5 and the circumferential direction t. In other words, the stagger angles of at least some of the blade profiles 7 can differ from the nominal stagger angle A by respective stagger angle variation values dA. With a stagger angle variation value dA of 0 degrees, the chord 18 of the respective blade profile 7 is aligned parallel to the chord 18 of the blade inner profile 5 .

The stagger angles of at least some of the blade profiles 7 can deviate from the nominal stagger angle A by the respective stagger angle variation values dA, with the respective stagger angle variation values dA being at most 2 degrees. In other words, the respective stagger angles differ by at most 2 degrees from the nominal stagger angle. The stagger angle variation values can range from -2 degrees to +2 degrees.

The threading strategy of the focal points 8 of the profile sections/blade profiles 7 is thus adapted in the axial direction x and the circumferential direction t in such a way that the main stress 12 with the greatest magnitude in the critical profile area 13, the blade leading edge 11, is a compressive stress. In addition, the stagger angle of at least some of the blade profiles 7 can be varied by the respective stagger angle variation values dA.

The adaptation of the threading, specified by the threading line 16 relative to the reference configuration, includes an optimization of the threading course (polynomial of the nth order) with the aim of influencing the tension at the leading edge. A parameter space for the shift in the center of gravity of the blade profiles 7 can be specified by a maximum circumferential variation dt of +/-5 degrees and/or a maximum axial variation dx of +/-20% chord length of the respective blade profile 7 .

FIG. 4 shows a schematic representation of a stress distribution in the in FIG. 3 shown shovel. It can be seen that the leading edge region 14 of the leading edge has a negative voltage. In this area, a compressive stress is the local principal stress. This results in the advantage that the leading edge is more resistant to impacts from foreign bodies than the in FIG. 1 shovel shown.

Compressor blading is currently being designed in an interdisciplinary manner, mainly with the focus on maximizing efficiency and the surge limit distance while complying with the structural-mechanical strength requirements (utilization < 60% Goodman Ratio).

Geometries that result from this procedure (prior art) often have areas that are sensitive to foreign object damage (FOD) or domestic object damage (DOD). During operation of the components, damage occurs more frequently, particularly on the leading edges of the blades. In the event of damage, these areas are limiting for the further operation of the engine and make unscheduled maintenance of the affected components necessary.

During the interdisciplinary design process of the compressor blades, the stress distribution in critical component areas can be influenced in a targeted manner. In order to make the region of the blade leading edge, which is often affected by damage, more robust, an adaptation of the previously used threading strategy of the individual blade profile sections is proposed. The aim of this adapted threading strategy (centre of gravity position of the individual blade profiles adjusted axially and in the direction of rotation) is to load the blade leading edge in the operating area in such a way that the main stress in the critical profile area with the greatest amount is compressive stress. Such a pronounced stress field is counterproductive for rapid crack propagation in the component and thus represents a means of increasing the blade robustness against damage.

über die unteren 90% der radialen Schaufelerstreckung angestrebt. Dieser Zielwert kann zusätzlich zur Anpassung der Fädelungsstrategie auch durch die Relativänderung der Staffelungswinkel naheliegender Schaufelprofile zueinander erreicht werden.Since the generation of a compressive stress, which is the largest main stress in terms of magnitude, cannot always be guaranteed during operation, the adjustment of the threading line should also be used to positively influence the ratio of the leading edge stress to the mean stress in individual blade profiles of the blade. A voltage ratio is used as the target value for a FOD/DOD resistant design ${\theta }_{\mathit{ratio}}=\raisebox{1ex}{${\mathrm{<figure-callout\; id="1"\; label="\sigma "\; filenames="imgf0004.png"\; state="\{\{state\}\}">\sigma </figure-callout>}}_{1,\mathit{LE}}$}\!\left/ \!\raisebox{-1ex}{${\sigma }_{\mathit{ave}}$}\right.\le 0.5$

is aimed for over the lower 90% of the blade radial extent. In addition to adapting the threading strategy, this target value can also be achieved by changing the stagger angles of nearby blade profiles relative to one another.

Reference list:

1

shovel

2

flow machine

3

coordinate system

4

rotor shaft

5

blade inner profile

6

beam deflection section

7

blade profile

8th

main emphasis

9

reference line

10

trailing edge

11

leading edge

12

main stress

13

Area

14

leading edge area

15

part length

16

threading line

17

total variance

18

chord

19

origin

20

trailing edge area

21

blade root element

t

tangential direction

German

Tangential Variation Value

right

radial direction

x

axial direction

dx

axial variation value

A

nominal blade angle

there

blade angle variation

Claims (9)

Blade (1) for a turbomachine (2), in particular for an aircraft engine, having a blade inner profile (5) and a jet deflection section (6) which adjoins the blade inner profile (5) in a radial direction (r) with respect to a rotor shaft (4). , wherein respective blade profiles (7) of the jet deflection section (6) aligned normal to the radial direction (r) have respective centers of gravity (8), the centers of gravity (8) are shifted relative to a reference line (9) running linearly along the radial direction (r) by respective axial variation values (dx) in an axial direction (x) parallel to the rotor shaft (4),
and or the centroids (8) are shifted with respect to the reference line (9) by respective tangential variation values (dt) in a tangential direction (t), the respective centers of gravity (8) being connected by a threading line emanating from the inner blade profile (5), characterized in that the threading line is a spatial line, with a total deviation (17) of the spatial line in the blade profiles (7) from the reference line (9) depending on a respective radial distance of the respective blade profile (7) from a reference point being described by a polynomial of the nth order , where n is a natural number greater than 1 and that the threading line is selected and designed in such a way that during operation there is a compressive stress in the area of the leading edge of the blade. Blade (1) according to Claim 1, characterized in that the threading line is designed in such a way that a maximum of the polynomial form of the threading line is shifted backwards axially in relation to a threading line which runs purely radially. Blade (1) according to claim 1 or 2,
characterized in that
the respective tangential variation values (dt) are in a range between -5 degrees and +5, preferably between -2.5 degrees and 2.5 degrees. Blade (1) according to one of Claims 1 to 3,
characterized in that
a respective one of the axial variation values (dx) lies in a range between 0 and 20 percent of a respective chord length (18) of the respective blade profile (7). Blade (1) according to one of Claims 1 to 3,
characterized in that
a respective one of the axial variation values (dx) lies in a range between 0 and 10 percent of a respective chord length (18) of the respective blade profile (7). Blade (1) according to one of Claims 1 to 5,
characterized in that
the blade (1) is designed as a moving blade. Blade (1) according to one of the preceding claims,
characterized in that
respective stagger angles of at least some of the blade profiles (7) deviate from a nominal stagger angle (A) of the blade inner profile (5) by respective stagger angle variation values (dA), the respective stagger angle variation values (dA) being in a range between -2 degrees and 2 degrees. Blade according to one of the preceding claims, characterized in that the course of the threading line is selected such that a ratio $\raisebox{1ex}{${\sigma }_{1,\mathit{LE}}$}\!\left/ \!\raisebox{-1ex}{${\sigma }_{\mathit{ave}}$}\right.$

of the stress in the operating state at the leading edge (σ _{1 ,LE} ) to the average stress in individual profile sections ( σ _ave ) over the lower 90% of the radial blade extent is less than or equal to 0.5 Turbomachine (2), having at least one blade (1) according to one of Claims 1 to 8.

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