DE102024106894A1 - Method for producing a deformable blade of a turbomachine, deformable blade, use of a nickel-rich shape memory material, turbomachine, and computer program product - Google Patents

Abstract

The invention relates to a method for producing an airfoil (10) of a turbomachine (100), comprising: optimising a geometric profile of a trailing edge (12) and/or leading edge (14) of a trailing edge (12) and/or leading edge (14), said trailing edge (12) and/or leading edge (14) of an airfoil (10) of a selected turbo component (50, 51, 52, 54), said trailing edge (12) and/or leading edge (14) comprising a shape memory material at least in one region, for at least one state (42, 44) deflected relative to an initial state (40) of the airfoil (10), the operating temperatures and/or mechanical loads of which, under real operating conditions, lie at least in regions at the trailing edge (12) and/or leading edge (14) within a permissible operating range of a predetermined shape memory material; Manufacturing the blade (10) in the initial state (40), wherein the trailing edge (12) and/or leading edge (14) is manufactured at least partially from the predetermined shape memory material; conditioning the blade (10) for the at least one, in particular the at least two deflected states (42, 44). Furthermore, the invention relates to a deformable blade (10) of a turbo component, a turbomachine (100) with at least one deformable blade (10), and a computer program product.

Description

The invention relates to a method for producing a deformable blade of a turbomachine, in particular a steam turbine or a gas turbine, a deformable blade of a turbo component of a turbomachine, in particular a steam turbine or a gas turbine, a turbomachine with at least one deformable blade, a use of a nickel-rich shape memory material, and a computer program product.

State of the art

Turbines extract energy from a fluid flow and convert it into shaft power. Turbines can be internal combustion engines, such as gas turbines, but in principle also heat engines, such as steam turbines. In the following, reference is made only to gas turbines.

In gas turbines, blade wheels are driven by the hot flow of a fluid, often by several rotors arranged one behind the other, which are attached to one or more shafts. To reduce the swirl of the flow caused by rotation and to influence the pressure conditions, a row of non-rotating stator blades is usually located between each pair of rotors, with the stator blades oriented opposite to those of the rotors.

Upstream of the turbine, in the direction of flow, is a combustion chamber. In the case of aircraft engines, a compressor is located upstream of the combustion chamber. The compressor is fundamentally similar in design to the turbine, i.e., with several stages of alternating rotors and stators. Unlike the turbine, the compressor's job is to compress the flow of the fluid, usually atmospheric air. This increases the fluid's density, pressure, temperature, and velocity.

The energy required for this comes from the turbine, whose shaft(s) drive the compressor rotors. Fuel is added to the compressed flow in the combustion chamber, and this mixture is ignited. This creates the hot gas, which is then fed into the turbine.

Rotors and stators can be summarized under the term turbo components.

In certain applications, particularly in aviation, but also in ground vehicles and stationary systems in the energy sector, the turbo components are designed to function as efficiently as possible under multiple operating conditions. Operating conditions can differ, for example, due to different speeds, combustion temperatures, ambient temperatures, mass flows, etc. This often requires a compromise between meeting various requirements as best as possible overall and making compromises in meeting individual requirements. For example, aviation engines must be as efficient as possible in cruise mode, while simultaneously delivering very high maximum power for takeoff.

Efficiency can be increased by varying the geometry of the rotor and stator blades so that they are closer to their aerodynamic optimum at different operating points. In particular, a variable effective angle of attack, i.e., the angle between the flow and the straight line connecting the leading edge and trailing edge of the blade, influences efficiency at each individual operating point.

In rotors, geometric variability can be integrated into the construction material of the rotor blades, for example, using shape memory alloys or piezoelectric materials. These react to external stimuli, such as changes in temperature or applied electrical voltages, with component deformation.

The US 2023/0160307 A1 discloses deformable rotor blades for turbomachinery, comprising a blade root and a blade portion having a gate formed with a deformable material. The material is a shape memory alloy that can change its shape due to a thermal stimulus. The thermal stimulus can be achieved by heating with an electric current or by introducing a temperature-controlled fluid into a pipe of the blade.

In the US 8,657,561 B2 A rotor blade is described, comprising a core blade portion and a trailing edge portion, wherein the trailing edge portion is attached to the core blade portion along a connection interface. The trailing edge portion comprises a shape memory material operable to effect selective deformation of the trailing edge portion between a first state in which a trailing edge of the trailing edge portion follows a substantially smooth profile and a second state in which the trailing edge is perturbed. The direction of the disturbance can be oblique or perpendicular to the flow direction over the blade. The rotor blade can be a fan or propeller for a gas turbine engine.

The US 11,549,383 B2 discloses a variable geometry aerodynamic blade system comprising a blade with a leading edge and a trailing edge. A shape memory alloy component is incorporated into the blade to aerodynamically reposition one or both of the leading and trailing edges. A heating element interacts with the shape memory alloy component to provide heating for the transition between an austenitic and a martensitic phase.

The shape memory alloy component reacts to the heating element to change a camber or twist of the blade in response to a control signal. A control system is operatively connected to the heating element, the control system receiving a command signal and outputting the control signal in response to the command signal.

Disclosure of the invention

The object of the invention is to provide an improved method for producing a deformable blade of a turbomachine, in particular a gas turbine or a steam turbine.

A further object of the invention is to provide a deformable blade of a turbo component of a turbomachine, in particular a gas turbine or a steam turbine, produced by the improved method.

A further object of the invention is to provide a use of a material for a deformable blade of a turbo component of a turbomachine.

A further object of the invention is to provide a computer program product for carrying out the improved method.

The objects are achieved by the features of the independent claims. Advantageous embodiments and advantages of the invention emerge from the further claims, the description, and the drawings.

According to one aspect of the invention, a method for producing a deformable blade of a turbomachine, in particular a gas turbine or a steam turbine, is proposed, at least comprising: optimizing a geometric profile of a trailing edge and/or leading edge of a trailing edge and/or leading edge of a blade of a selected turbo component, said trailing edge and/or leading edge comprising a shape memory material at least in one or more regions, for at least one, in particular at least two deflected states relative to an initial state of the blade, the operating temperatures and/or mechanical loads of which under real operating conditions lie at least partially at the trailing edge and/or leading edge within a permissible operating range of a predetermined shape memory material; manufacturing the blade in the initial state, wherein the trailing edge and/or leading edge is manufactured at least partially from the predetermined shape memory material; and conditioning the blade for the at least one, in particular the at least two deflected states.

The proposed method describes the design and manufacture of deformable blades for turbo components of turbomachinery using shape memory materials. The fact that all or at least individual areas of the blades are made of shape memory material, where the areas of shape memory material achieve their desired shape based on the respective operating temperature without additional actuators or stimuli, especially passively, results in various advantages. Optionally, the blades can be made entirely of shape memory material.

For example, a state assumed by the blade can be an initial state, and a deflected state can be assumed by deforming the shape memory material, so that the initial state can be resumed afterwards. Alternatively, conditioning can be performed starting from the initial state in such a way that a first deflected state and a second deflected state can be assumed, whereby the initial state only exists during the manufacture of the blade and does not necessarily have to be resumed during operation.

Optionally, the method may comprise selecting at least one turbo component whose operating temperatures and/or mechanical loads under real operating conditions are at least partially at a trailing edge and/or leading edge of an airfoil within a permissible operating range of a predetermined shape memory material. Furthermore, the method may optionally comprise performing an aerodynamic evaluation of at least one, in particular at least two, depending on a Operating temperature of the areas of the blade, deflected states of the trailing edge and/or leading edge of the blade of the selected turbo component compared to an initial state.

In particular, prior to selecting the turbo components, a geometric profile of blades of turbo components of the turbomachine can be defined; as well as a determination of operating temperatures and/or mechanical loads of the blades of the turbo components in a simulation of real operating conditions of the turbomachine can be provided.

The shape-memory material sections can be seamlessly integrated into the profiles of existing blades, eliminating gaps that can lead to pressure losses, turbulence, and other fluid mechanics disadvantages, unlike other mechanical adjustment mechanisms of turbo components or their trailing edges. Furthermore, complex mechanical components are not required, which increase weight, take up space, are potentially prone to failure, and require intensive maintenance. This also allows for a space-saving solution, which is particularly advantageous for compressor stages close to a gas turbine combustion chamber, where installation space is often very limited.

There's no need for an adjustment mechanism, which would necessarily rotate, especially with rotors. This significantly simplifies the technical implementation.

Manufacturing blade trailing or leading edges from shape memory material also offers advantages over the use of functional materials in rotors, where only a thin outer layer of functional material allows for limited deformability. Furthermore, the functionality is durable, even at elevated temperatures.

The proposed method can be advantageously used for deformable blades of compressors, low-pressure turbines and power turbines with sufficiently low temperatures in the areas of civil and military aircraft engines, stationary gas turbines in the energy sector, steam turbines in the energy sector or in process engineering.

Applications in these areas can benefit from higher efficiency, as geometric variability means fewer compromises have to be made both at the design point and during partial load or peak load operation.

According to the proposed method, after the blade profile has been designed, those turbo components are selected in the preliminary design phase for which the actual thermal operating conditions of the component allow the use of a given shape memory material with the so-called two-way shape memory effect. This refers to the phenomenon that, under certain circumstances, a component made of shape memory material can assume two states. The transition from one state to the other occurs through a change in the thermal and mechanical loading of the component. One state is characterized by the shape memory material being in the austenitic phase. This is converted into a martensitic phase by a martensitic transformation typical of shape memory materials (decreasing temperature and/or increasing mechanical loading), resulting in the other state of the component. This transformation of the material and the associated change in shape of the component are cyclically approximately reversible with the two-way shape memory effect.

For example, a nickel-rich shape memory material with the main components Ni, Ti, as well as Hf and/or Zr, in particular with a nominal composition Ni _50.3 Ti _29.7 Hf _20-x Zr _x with 0≤x≤20 can be provided and thus a proportion x of zirconium can be determined at which the use of the shape memory alloy is advantageous.

This depends on the temperatures and mechanical loads, such as gas pressure and rotation, since the transformation temperatures of the transitions between the martensitic and austenitic phases in the phase diagram of the shape memory material are load-dependent. It is therefore advantageous to have at least an approximate knowledge of the expected mechanical loads. This can be determined through calculations or experience.

According to a favorable embodiment of the method, at least one of the following criteria can be used to select the at least one turbo component: a maximum operating temperature of the blade of the turbo component under real operating conditions is lower than a final conditioning temperature; there are operating points at which the operating temperature of the blade is higher than an austenite final temperature; there are operating points at which the operating temperature of the blade is smaller than a martensite end temperature of the shape memory material.

For an airfoil to which at least one of these criteria applies, aerodynamic evaluations can then be performed with the at least one, in particular the at least two deflected states of the trailing edge and/or leading edge. The aerodynamic evaluation can be performed using a computer-aided evaluation, an analytical calculation, or empirical values. For example, a first state is to be optimized for the hot operating points and a second state for the cold operating points.

According to a favorable embodiment of the method, the optimization of the geometric profile of the trailing edge and/or leading edge for the at least one, in particular the at least two deflected states can be performed iteratively until an iteration termination criterion is reached. The blade can be formed from shape memory material only in certain regions or completely.

If necessary, the thermodynamic modeling in the conceptual design must be iteratively adapted to this variability, which would also require a repetition of the preliminary design and the previous step of blade design.

According to a favorable embodiment of the method, the optimization of the geometric profile of the trailing edge and/or leading edge of the blade can be carried out under a condition that with the at least one, in particular the at least two deflected states, less than a predetermined limit value, in particular less than 5%, of an additional local strain is generated.

When optimizing the geometry, it is important to consider that, ideally, the transition between the states generates less than 5% additional local strain. The limit of 5% additional local strain may be lower depending on the desired lifetime.

According to a favorable embodiment of the method, cavities can be provided in the profile of the blade, in particular cavities with a smooth-edged, particularly circular, elliptical, or oval, cross-section in the profile of the blade. This ensures that the limit value for additional local strain is maintained.

According to a favorable design of the process, a nickel-rich shape memory material with the main components Ni, Ti, as well as Hf and/or Zr, in particular with a nominal composition Ni _50.3 Ti _29.7 Hf _20-x Zr _x with 0≤x≤20 can be selected.

Fundamentally, materials must be selected for all turbo components that can withstand the thermal and mechanical stresses over a selected minimum service life. Due to the elevated temperatures in the engine, the shape memory alloy family with the nominal composition Ni _50.3 Ti _29.7 Hf _20-x Zr _x can be advantageous. At least in the initial state before manufacture of the airfoil, this consists of 50.3% nickel, 29.7% titanium, 20-x% hafnium, and x% zirconium, where 0≤x≤20. The percentages are atomic percent, i.e. they refer to the number of corresponding atoms. The shape memory material can exhibit deviations from the nominal composition Ni _50.3 Ti _29.7 Hf _20-x Zr _x due to impurities as well as changes in composition during the manufacturing process after manufacture of the airfoil. Fortunately, the nickel content is higher than that of titanium even after the manufacturing process.

The shape memory material with the nominal starting composition Ni _50.3 Ti _29.7 Hf _20-x Zr _x with 0≤x≤20 advantageously results in sufficiently high transformation temperatures of the shape memory effect for use in turbo components of compressors and turbines. Sufficiently good structural and functional durability at high temperatures for use in turbo components of compressors and turbines can be advantageously achieved.

Furthermore, a nickel-rich shape memory material such as Ni _50.3 Ti _29.7 Hf _20-x Zr _x can be advantageously selected for this application to create specific material properties. For example, starting with NiTiHf, a partial substitution of Hf with Zr can be useful, as this reduces weight while still maintaining high temperature resistance. Zr is lighter than Hf, but Hf contributes more to temperature resistance than Zr. The exact ratios of the individual components of the alloy can be advantageously optimized for the respective application.

According to a favorable embodiment of the process, the material composition of the shape memory material can be selected so that the proportion x is as large as possible. The proportion x can advantageously be selected as low as necessary and as high as possible in order to, on the one hand, determine the transformation temperatures of the phase diagram of the shape memory material. memory material Ni _50.3 Ti _29.7 Hf _20-x Zr _x and on the other hand to minimize the material weight.

In a favorable configuration of the process, the airfoil can be manufactured using an additive manufacturing process. The airfoil can be manufactured using 3D printing with a previously identified alloy. An initial state of the airfoil profile can be selected as the mold, and the previously identified cavities can be incorporated as needed.

According to a favorable embodiment of the method, the initial state of the airfoil can be adjusted such that, after conditioning at the appropriate temperatures, the leading edge and/or trailing edge of the airfoil assumes at least one, in particular both, deflected states. In this way, the desired aerodynamic behavior of the airfoil can be achieved under the actual operating temperatures of the airfoil.

According to a favorable embodiment of the method, the conditioning of the blade can comprise at least the following steps: clamping the blade in its initial state on the suction side and the pressure side; applying a tool from the suction side or the pressure side of the blade; carrying out a temperature treatment by increasing a temperature of the blade from an initial temperature lower than a martensite final temperature to a final temperature higher than an austenite final temperature of the shape memory material, followed by lowering the temperature to the initial temperature; cyclically carrying out the temperature treatment several times.

The component can be conditioned according to the method described above. Depending on which directional deflection of the trailing edge and/or leading edge is advantageous for the particular blade, a working direction of the tool can be selected, and the initial state and the at least one, in particular the two, deflected states can be defined. The initial state must be selected such that, after conditioning, the at least one, in particular the two, deflected states are assumed at the corresponding operating temperatures. The initial state differs from the second deflected state due to the plastic deformations introduced at certain points in the microstructure during training.

According to a favorable design of the method, the applied force and/or the deformation generated by the blade can be monitored. This allows the desired conditions of the blade to be advantageously adjusted during conditioning.

According to a favorable design of the process, the tool can be controlled to the suction or pressure side of the blade during the heat treatment. This allows the desired conditions of the blade to be advantageously adjusted during conditioning.

According to a favorable embodiment of the method, averaged operating temperatures of the blades can be determined in the simulation for real operating conditions to determine operating temperatures and/or mechanical loads on the blades of the turbo components. The same temperature is assumed for all blades of a turbo component.

According to a favorable embodiment of the process, the blade can be a stator blade or a rotor blade of the turbo component. Shape memory material can advantageously be used on both stationary and rotating components of the turbo component to create a deformable area at the operating temperature.

According to a further aspect of the invention, a deformable blade of a turbo component of a turbomachine, in particular a gas turbine, is proposed, manufactured according to a method described above, comprising at least a center section, a leading edge, and a trailing edge, wherein the leading edge adjoins the center section upstream of a flow direction during intended use, and the trailing edge adjoins the center section in the flow direction. The leading edge and/or the trailing edge are formed at least partially from a shape memory material.

The fact that all or at least individual areas of the blade are formed from shape memory material, in which the areas of shape memory material achieve their desired shape due to the respective operating temperature without additional actuators or stimuli, results in various advantages.

The shape-memory material sections can be seamlessly integrated into the profiles of existing blades, eliminating gaps that can lead to pressure losses, turbulence, and other fluid mechanics disadvantages, unlike other mechanical adjustment mechanisms of turbo components or their trailing edges. Optionally, the blades can be made entirely of shape-memory material.

Furthermore, it eliminates the need for complex, manufactured mechanical components, which increase weight, take up space, are potentially prone to failure, and require intensive maintenance. This also allows for a space-saving solution, which is particularly advantageous for the front compressor stages of a gas turbine, facing away from the flow direction, where installation space is often very limited.

There's no need for an adjustment mechanism, which would necessarily rotate, especially with rotors. This significantly simplifies the technical implementation.

Manufacturing trailing and/or leading edges of blades from shape memory material also offers advantages over the use of functional materials in rotors, where only limited deformability is possible if only a thin outer layer of functional material is used. Furthermore, the functionality is durable even at elevated temperatures. Advantageously, the blades can also be made entirely of shape memory material.

The shape memory material can be used advantageously for deformable blades of compressors, low-pressure turbines and power turbines with sufficiently low temperatures in the areas of civil and military aircraft engines, stationary gas turbines in the energy sector, steam turbines in the energy sector or in process engineering.

Applications in these areas can benefit from higher efficiency, as geometric variability means fewer compromises have to be made both at the design point and during partial load or peak load operation.

According to a favorable design of the deformable blade, the shape memory material can be nickel-rich with the main components Ni, Ti, as well as Hf and/or Zr, in particular with a nominal composition Ni _50.3 Ti _29.7 Hf _20-x Zr _x with 0≤x≤20.

Fundamentally, materials must be selected for all turbo components that can withstand the thermal and mechanical stresses over a selected minimum service life. Due to the elevated temperatures in the engine, materials from the shape memory alloy family with the nominal starting composition Ni _50.3 Ti _29.7 Hf _20-x Zr _x can be advantageous. This consists of 50.3% nickel, 29.7% titanium, 20-x% hafnium, and x% zirconium, where 0≤x≤20. The percentages are atomic percent (at%), i.e. they refer to the number of corresponding atoms. The shape memory material can exhibit deviations from the nominal composition Ni _50.3 Ti _29.7 Hf _20-x Zr _x due to impurities as well as changes in composition during the manufacturing process after the airfoil has been produced.

The shape memory material with the nominal composition Ni _50.3 Ti _29.7 Hf _20-x Zr _x , where 0≤x≤20, advantageously results in sufficiently high transformation temperatures for the shape memory effect for use in turbo components of compressors and turbines. Sufficiently good structural and functional durability at high temperatures for use in turbo components of compressors and turbines can be advantageously achieved.

Furthermore, a nickel-rich shape memory material such as Ni _50.3 Ti _29.7 Hf _20-x Zr _x can be advantageously selected for this application to create specific material properties. For example, starting with NiTiHf, a partial substitution of Hf with Zr can be useful, as this reduces weight while still maintaining high temperature resistance. Zr is lighter than Hf, but Hf contributes more to temperature resistance than Zr. The exact ratios of the individual components of the alloy can be advantageously optimized for the respective application.

According to a favorable design of the deformable airfoil, the trailing edge and/or leading edge can each have at least one, in particular at least two, deflected states of the trailing edge and/or leading edge relative to an initial state, depending on the operating temperature of the trailing edge and/or leading edge. For example, the first state can be assumed for the hot operating points and the second state for the cold operating points. In this way, the desired aerodynamic behavior of the airfoil can be achieved under the actual operating temperatures of the airfoil.

According to a favorable design of the deformable airfoil, the at least one, in particular the at least two deflected states of the trailing edge and/or leading edge can be conditioned by means of a tool after production. In this way, the desired aerodynamic behavior of the airfoil can be achieved under the actual operating temperatures of the airfoil.

According to a favorable design of the deformable airfoil, cavities can be formed in the airfoil profile, in particular cavities with a smooth-edged, particularly circular, elliptical, or oval cross-section in the airfoil profile. This can ensure that the limit value for additional local strain is maintained.

In a favorable embodiment, the deformable airfoil can be manufactured using an additive manufacturing process. The airfoil can be manufactured using 3D printing with a previously identified alloy. An initial state of the airfoil profile can be selected as the mold, and the previously identified cavities can be incorporated as needed.

Depending on the design of the deformable blade, the blade can be a stator blade or a rotor blade of the turbo component. Shape memory material can be advantageously used on both stationary and rotating components of the turbo component to create a deformable area at the operating temperature.

According to a further aspect of the invention, a turbomachine with at least one deformable blade is proposed, at least comprising a turbo component with the at least one blade, which is manufactured according to the method described above.

The proposed turbomachinery can feature deformable blades for compressors, low-pressure turbines, and power turbines with sufficiently low temperatures. Such a turbomachinery can be advantageously used in civil and military aircraft engines, stationary gas turbines in the energy sector, steam turbines in the energy sector, or in process engineering.

Applications in these areas can benefit from higher efficiency, as geometric variability means fewer compromises have to be made both at the design point and during partial load or peak load operation.

According to a further aspect of the invention, the use of a nickel-rich material with the main components Ni, Ti, as well as Hf and/or Zr, in particular with a nominal composition Ni _50.3 Ti _29.7 Hf _20-x Zr _x with 0≤x≤20, is proposed as a shape memory material for a deformable blade of a turbo component of a turbomachine, in particular in regions of a trailing edge and/or leading edge of the deformable blade.

The shape memory material can be used advantageously for deformable blades of compressors, low-pressure turbines and power turbines with sufficiently low temperatures in the areas of civil and military aircraft engines, stationary gas turbines in the energy sector, steam turbines in the energy sector or in process engineering.

According to a further aspect of the invention, a computer program product is proposed, comprising instructions which, when the program is executed by a computer, cause the computer to carry out steps of a method for producing deformable blades of a turbomachine, in particular a gas turbine or a steam turbine, at least comprising: optimising a geometric profile of a trailing edge and/or leading edge of a trailing edge and/or leading edge of a blade of a turbo component, said trailing edge and/or leading edge having a shape memory material in at least one or more regions, for at least one, in particular at least two, deflected states relative to an initial state of the blade, the operating temperatures and/or mechanical loads of which under real operating conditions lie at least partially at the trailing edge and/or leading edge within a permissible operating range of a predetermined shape memory material.

The computer program product advantageously enables the steps for designing and optimizing the profile of the deformable blades of a turbomachine, as well as for selecting a turbo component suitable for use with a given shape memory material. The program can execute these steps either semi-automatically or automatically.

Optionally, at least one turbo component can be selected whose operating temperatures and/or mechanical loads under real operating conditions are at least partially within a permissible operating range of a predetermined shape memory material at a trailing edge and/or leading edge of an airfoil. Furthermore, an aerodynamic evaluation of at least two deflected states of the trailing edge and/or leading edge of the airfoil of the selected turbo component relative to an initial state can be performed, depending on an operating temperature of the airfoil regions.

The aerodynamic evaluation can be performed using computer-aided analysis, analytical calculations, or empirical values. For example, a first state is optimized for the hot operating points and a second state for the cold operating points.

drawing

Further advantages will become apparent from the following description of the drawings. The figures illustrate exemplary embodiments of the invention. The figures, the description, and the claims contain numerous features in combination. Those skilled in the art will also expediently consider the features individually and combine them into useful further combinations.

• 1 eine schematische Darstellung einer als Flugzeugstriebwerk ausgebildeten Turbomaschine;
• 2 eine Schnittdarstellung der als Flugzeugstriebwerk ausgebildeten Turbomaschine nach 1;
• 3 ein Schaufelprofil eines Schaufelblatts mit verformbarer Hinterkante nach einem Ausführungsbeispiel der Erfindung und entsprechender Variation der Richtung der abfließenden Strömung;
• 4 ein Schaufelprofil eines Schaufelblatts mit verformbarer Vorderkante nach einem weiteren Ausführungsbeispiel der Erfindung;
• 5 ein Phasendiagramm eines Formgedächtnismaterials im Spannungs-Temperatur-Raum;
• 6 eine Konditionierung des Schaufelprofils eines Schaufelblatts mittels einer Einspannung und einem Werkzeug nach einem Ausführungsbeispiel der Erfindung;
• 7 ein Ausschnitt einer isometrischen Darstellung eines Stators einer Turbokomponente;
• 8 eine Statorschaufel mit einer verformbaren Hinterkante nach einem Ausführungsbeispiel der Erfindung in einer Draufsicht;
• 9 die Statorschaufel nach 8 in einem Querschnitt;
• 10 ein Ausschnitt einer isometrischen Darstellung eines Rotors einer Turbokomponente;
• 11 eine Rotorschaufel mit einer verformbaren Hinterkante nach einem Ausführungsbeispiel der Erfindung in einer Draufsicht; und
• 12 die Rotorschaufel nach 11 in einem Querschnitt.

Examples include:

• 1 a schematic representation of a turbomachine designed as an aircraft engine;
• 2 a sectional view of the turbomachine designed as an aircraft engine according to 1 ;
• 3 a blade profile of an airfoil with a deformable trailing edge according to an embodiment of the invention and corresponding variation of the direction of the outflowing flow;
• 4 a blade profile of an airfoil with a deformable leading edge according to a further embodiment of the invention;
• 5 a phase diagram of a shape memory material in the stress-temperature space;
• 6 a conditioning of the blade profile of an airfoil by means of a clamping device and a tool according to an embodiment of the invention;
• 7 a section of an isometric view of a stator of a turbo component;
• 8 a stator blade with a deformable trailing edge according to an embodiment of the invention in a plan view;
• 9 the stator blade 8 in a cross-section;
• 10 a section of an isometric view of a rotor of a turbo component;
• 11 a rotor blade with a deformable trailing edge according to an embodiment of the invention in a plan view; and
• 12 the rotor blade 11 in a cross section.

Embodiments of the invention

In the figures, components of the same type or function similarly are designated by the same reference numerals. The figures are merely examples and are not to be construed as limiting.

Before describing the invention in detail, it should be noted that it is not limited to the specific components of the device, as these components may vary. The terms used herein are intended solely to describe particular embodiments and are not intended to be limiting. Furthermore, when the singular or indefinite articles are used in the description or claims, this also refers to the plural of these elements, unless the overall context clearly indicates otherwise.

The directional terminology used below, including terms such as "left," "right," "top," "bottom," "before," "behind," "after," and the like, is intended solely to enhance understanding of the figures and is in no way intended to limit the scope of the invention. The components and elements depicted, as well as their design and use, may vary according to the considerations of a person skilled in the art and may be adapted to specific applications.

1 shows a schematic representation of a turbomachine 100 designed as an aircraft engine. The flow direction 70 of a flow medium is indicated by an arrow.

The design of turbomachinery typically begins with the development of a machine concept and the thermodynamic modeling of its components. To quantify fundamental quantities such as temperature, pressure, and mass flow under real operating conditions, boundary conditions are assumed and established numerical or analytical methods are used. This allows the quantities to be determined at various points in the machine, e.g., at the beginning and end of each component (with respect to the flow direction).

1 roughly outlines a possible schematic structure of an exemplary turbomachine 100 using the example of an aircraft engine. The turbomachine 100 can actually be constructed differently and be an engine for another aircraft, watercraft, or land vehicle or also represent a stationary power machine such as a gas turbine or a steam turbine.

The illustrated turbomachine 100 consists of a so-called fan 50 as a pre-compressor, a compressor 51, a combustion chamber 56, a high-pressure turbine 52 and a low-pressure turbine 54. In the compressor 52, a flow compression and flow acceleration of the flow medium takes place, in the combustion chamber 56 an energy increase of the flow through fuel injection and fuel combustion, in the high-pressure turbine 52 and the low-pressure turbine 54 an energy transfer from the flow of the flow medium into the rotation of the turbine rotors.

The components 50, 51, 52, 54 of the turbomachine 100, which have rotating and/or non-rotating blade rows, are referred to as turbo components.

2 shows a sectional view of the turbomachine 100 designed as an aircraft engine according to 1 . The engine is cut along the engine axis 110 and only one half of the engine is shown.

The engine has a three-stage fan 50. This is followed in the flow direction 70 by another three-stage compressor 51 and the combustion chamber 56. A single-stage high-pressure turbine 52, followed by a two-stage low-pressure turbine 54, is connected to the outlet of the combustion chamber 56.

A bypass duct 57 runs outside the compressor 51 and the turbine stages 52, 54.

According to an advantageous embodiment of the method for producing a deformable blade 10 of a turbomachine 100, in particular a gas turbine, a geometric profile of blades 10 of turbo components 50, 51, 52, 54 of the turbomachine 100 can first be defined. Operating temperatures and/or mechanical loads of the blades 10 of the turbo components 50, 51, 52, 54 can then be determined in a simulation of real operating conditions of the turbomachine 100.

To determine operating temperatures and/or mechanical loads of the blades 10 of the turbo components 50, 51, 52, 54 in the simulation for real operating conditions, averaged operating temperatures of the turbo components 50, 51, 52, 54 can advantageously be determined.

On this basis, at least one turbo component 50, 51, 52, 54 is selected, the operating temperatures and/or mechanical loads of which under real operating conditions are at least partially located at a trailing edge 12 and/or leading edge 14 of an airfoil 10 within a permissible operating range of a predetermined shape memory material.

As a shape memory material, a nickel-rich shape memory material with the main components Ni, Ti, as well as Hf and/or Zr, in particular with a nominal composition before manufacture of the blade Ni _50.3 Ti _29.7 Hf _20-x Zr _x with 0≤x≤20 can be used at least for regions of a trailing edge 12 and/or leading edge 14 of a blade 10 of a turbo component 50, 51, 52, 54 of a turbomachine 100.

Optionally, the blade 10 can also be made entirely of shape memory material.

To select the at least one turbo component 50, 51, 52, 54, at least one of the following criteria can advantageously be used: a maximum operating temperature of the airfoil 10 of the turbo component 50, 51, 52, 54 under real operating conditions is less than a final conditioning temperature 68; there are operating points at which the operating temperature of the airfoil 10 is greater than an austenite final temperature 88; there are operating points at which the operating temperature of the airfoil 10 is less than a martensite final temperature 84 of the shape memory material.

An aerodynamic evaluation of at least one, in particular at least two, deflected states 42, 44 of the trailing edge 12 and/or leading edge 14 of the blade 10 of the selected turbo component 50, 51, 52, 54 relative to an initial state 40, is performed, depending on an operating temperature of the regions of the blade 10. The aerodynamic evaluation can be carried out using a computer-aided evaluation, an analytical calculation, or empirical values.

Thus, a geometric profile of the trailing edge 12 and/or leading edge 14 is optimized for the at least one, in particular the at least two deflected states 42, 44. Advantageously, the optimization of the geometric profile of the trailing edge 12 and/or leading edge 14 of the blade 10 can be carried out under a condition that less than a predetermined limit value, in particular less than 5%, of additional local strain is generated with the at least one, in particular the at least two deflected states 42, 44.

The steps of defining the geometric profile of the blades 10, determining operating temperatures and/or mechanical loads of the blades 10, and selecting at least one turbo component 50, 51, 52, 54 can advantageously be carried out iteratively until a suitable termination criterion of the iteration is reached.

The numerical and analytical methods used for the thermodynamic modeling of turbomachine 100 provide important flow parameters such as temperature, pressure, and mass flow of the flow medium at turbo components 50, 51, 52, and 54 at various operating points. In the following, the flow temperature is of primary interest.

The operating temperatures are highly dependent on the actual assumptions made regarding the turbomachine 100 and the operating conditions prevailing at the operating points. These conditions are usually multifactorial and include, but are not limited to, altitude, speed, ambient temperature, and rotational speed. An operating point can refer to a so-called "take-off" operating point or a "cruise" condition under standard atmosphere. In the case of aircraft engines, however, completely different operating points can also be considered, such as a so-called "supercruise" operating point or a "Mach 1 at sea level" operating point. Non-standard atmospheric conditions can also be considered in the thermodynamic analysis. In these cases, the operating temperatures could differ significantly.

The resulting thermodynamic results can be used to test the engine concept for its suitability for the intended mission, including the generated thrust and the associated fuel consumption. Typically, several concepts are developed, tested, and modified iteratively until a suitable concept is found. The concept and estimation methods also determine how many stages each turbo component should reasonably consist of. The stages are examined in more detail in a preliminary design phase.

In the preliminary design, the components are defined more precisely geometrically. A cross-section of the engine can be viewed in a two-dimensional model.

In compressor 50, the flow is compressed through the combined action of rotors, stators, and the reduction of the flow channel's cross-section. In modern aircraft engines, compression ratios range from 10:1 to 60:1, meaning the maximum flow pressure in the engine is ten to sixty times atmospheric pressure. This compression also increases the flow temperature of the fluid.

After the combustion chamber 56, energy is extracted from the flow by the turbine rotors 52, 54. Furthermore, the flow temperature is reduced by the combined action of the rotors, stators, cooling air inlet, and the enlargement of the cross-section of the flow channel.

Established preliminary design methods, such as analytical, numerical, and empirical procedures, can be used to determine the exact number of stators and rotors in the compressor 50, the high-pressure turbine 52, and the low-pressure turbine 54. Furthermore, the number of blades 10 for each turbo component 50, 51, 52, 54 can also be determined. Once this number has been determined, relevant, established methods can be applied to determine the flow temperatures at each turbo component 50, 51, 52, 54. In fact, the flow temperature profile is complex at small length scales because the precise radial, axial, and angular position of the flow point under consideration is important. For example, the flow temperature differs on the pressure side and the suction side of each blade. However, such small-scale effects are negligible in the preliminary design phase. Advantageously, only one average operating temperature per blade 10 is measured, in particular the same operating temperature at each blade 10 of this turbo component 50, 51, 52, 54.

According to the invention, the temperature variation during operation of the turbo components 50, 51, 52, 54 is used to advantageously change the geometry of the blade profiling.

Advantageously, a computer program product can be used for the design of deformable blades 10, which includes instructions which, when the program is executed by a computer, cause the computer to carry out steps of the proposed method for producing deformable blades 10 of a turbomachine 100, in particular a gas turbine.

3 shows a blade profile of a blade 10 with a deformable trailing edge 12 according to an embodiment of the invention and corresponding variation of the direction of the outflowing flow. The blade 10 can have a stator blade 20 ( 7 ) or a rotor blade 30 ( 10 ) of the turbo component 50, 51, 52, 54.

The choice of blade profile has a significant influence on the performance of compressor 50 and turbine 52, 54. Various geometric characteristics of the profiles have a variety of influences on the flow conditions and indirectly on the entire turbomachine 100.

The deformability of the blade profile causes a deflection of the flow 70, shown with arrows. In 3 The direction from which the flow 70 hits the blade 10 is shown. Also shown is the flow 72, which leaves the blade 10 when the blade 10 is not deformed. The corresponding profile (solid profile contour) is referred to as the initial state 40.

By changing the geometry in the area of the trailing edge 12 (dashed area), the direction of the outflow 74, 76 is changed. This changed profile is referred to as the deflected state 42, 44.

The initial state 40 of the blade 10 can be adjusted so that after conditioning at the corresponding temperatures, the trailing edge 12 of the blade 10 assumes the two deflected states 42, 44.

4 shows a blade profile of an airfoil 10 with a deformable leading edge 14 according to another embodiment of the invention. The undeformed profile is designated by state 40. A deformed state 42 of the leading edge 14 is shown in dashed lines.

Typically, i.e., with only one possible condition, the blade profile is selected to represent an optimum considering the overall mission profile of the engine. Compromises are made between the profile's performance at different operating points.

With shape variability, i.e., with at least one, especially two, possible states 42, 44, the blade profile can be adapted to the operating condition during operation. This reduces the need for compromises in profile optimization during profile design, since the two states 42, 44 can each be closer to the optimum for certain operating points.

According to the invention, the shape variability is achieved using shape memory alloys. The material selection, component manufacturing, and component conditioning are described below.

Fundamentally, materials must be selected for all turbo components 50, 51, 52, and 54 that can withstand the thermal and mechanical stresses over a selected minimum service life. Due to the elevated temperatures in the engine, a material from the shape memory alloy family Ni _50.3 Ti _29.7 Hf _20-x Zr _x can be advantageously selected. This alloy consists of 50.3% nickel, 29.7% titanium, 20-x% hafnium, and x% zirconium, where 0≤x≤20. The percentages are atomic percent, meaning they refer to the number of atoms.

According to the proposed method, after the blade profile has been designed, those turbo components are selected in the preliminary design phase for which the actual thermal operating conditions of the component allow the use of a given shape memory material with the so-called two-way shape memory effect. This refers to the phenomenon that, under certain circumstances, a component made of shape memory material can assume two states. The transition from one state to the other occurs through a change in the thermal and mechanical loading of the component. One state is characterized by the shape memory material being in the austenitic phase. This is converted into a martensitic phase by a martensitic transformation typical of shape memory materials (decreasing temperature and/or increasing mechanical loading), resulting in the other state of the component. This transformation of the material and the associated change in shape of the component are cyclically approximately reversible with the two-way shape memory effect.

In 5 A phase diagram of a shape memory material in stress-temperature space is shown. Stress 90 is plotted against temperature 80. Transition regions between the martensite phase 94 and the austenite phase 96 are shown as oblique lines with corresponding temperatures 82, 84, 86, and 88.

If martensite 94 (cold temperature phase) is heated under constant mechanical load (stress 90), the phase transformation to austenite 96 (high temperature phase) begins at the austenitic initial temperature 86 and is completed when the austenite final temperature 88 is reached. Conversely, if the material in the austenite phase 96 is cooled under constant mechanical load 90, the martensitic transformation begins when the martensite initial temperature 82 is reached and is completed at the martensite final temperature 84. The temperatures 82, 84, 86, 88 are from the chip voltage state, they increase approximately linearly with it.

A shape memory material that can be advantageously selected is, for example, a nickel-rich shape memory material with the main components Ni, Ti, as well as Hf and/or Zr, in particular with a nominal composition of Ni _50.3 Ti _29.7 Hf _20-x Zr _x , where 0≤x≤20. The material composition of the shape memory material can be selected such that the proportion x is as large as possible. The proportion x can advantageously be selected as low as necessary and as high as possible in order to achieve the transformation temperatures of the phase diagram of the shape memory material Ni _50.3 Ti _29.7 Hf _20-x Zr _x , on the one hand, and to minimize the material weight, on the other.

For the proposed shape memory alloy Ni _50.3 Ti _29.7 Hf _20-x Zr _x , for unloaded components and x=0, the martensite end temperature 84 is 210°C and the austenite end temperature 88 is 265°C. For x=20, the martensite end temperature 84 is 150°C and the austenite end temperature 88 is 210°C. The austenite end temperature 88 sometimes rises to just over 400°C under load, while the martensite end temperature 84 only varies by approximately 30-40°C with load.

The shape memory material family Ni _50.3 Ti _29.7 Hf _20-x Zr _x has a number of advantages: sufficiently high temperature resistance (low structural fatigue under thermal-mechanical loading), sufficiently high transformation temperatures 82, 84, 86, 88, sufficiently high fatigue stability of the transformation temperatures 82, 84, 86, 88 and the microstructural transformation behavior (low functional fatigue, ie low loss of the two-way shape memory effect). Weight optimization is possible by choosing the parameter x (Zr content): Zr is lighter than Hf, but leads to a decrease in the transformation temperatures 82, 84, 86, 88. For x=0, the maximum transformation temperatures 82, 84, 86, 88 are reached, which, depending on the stress state, are up to approximately 25% higher in the case of the austenite final temperature 88 than for x=20. For x=20, the weight is reduced by up to 23% compared to x=0.

The blade 10 is manufactured after design in the initial state 40, wherein the trailing edge 12 and/or leading edge 14 is manufactured at least in regions from the predetermined shape memory material.

According to the proposed method for producing a deformable airfoil 10, an additive manufacturing process such as 3D printing can be selected. For this purpose, the shape memory alloy is used in powder form. It should be noted that the components of the shape memory alloy can change during pulverization: On the one hand, foreign substances such as carbon can penetrate as impurities, and on the other hand, certain components can volatilize locally due to the high temperatures (evaporation, sublimation). Volatilization (evaporation, sublimation) can also occur after pulverization, during 3D printing. Due to these effects, for example, the nickel content and/or the titanium content in the finished component can differ from that in the starting material (from the original melt). The ratio formula Ni _50.3 Ti _29.7 Hf _20-x Zr _x can therefore advantageously be related to the finished workpiece.

In the area of the trailing edge 12 and/or leading edge 14 of the blade, cavities can advantageously be provided to increase the flexibility of the trailing edge 12 and/or leading edge 14 and to reduce the local maximum strain during edge bending. Cavities can also be introduced in the remaining area, for example, the center section 11, which can serve to reduce weight.

The cavities can be formed with a smoothly edged cross-section, in particular with a circular or elliptical or oval cross-section in the profile of the blade 10.

The two-way shape memory effect must first be "trained" into a component; this is also referred to as conditioning. This can be done in various ways, but is essentially always a cyclical process. A standard procedure can be chosen, as described in 5 A constant mechanical load is applied to the blade 10 and the temperature is cyclically varied between an initial temperature 66 below the martensite final temperature 84 and a final temperature 68 above the austenite final temperature 88. The initial temperature 66 is set to 20-30°C (ambient temperature) for technical simplicity. The cyclization is in 5 shown schematically with the dashed arrow 92.

The manufactured blade 10 is conditioned for the at least one, in particular the at least two deflected states 42, 44 before commissioning.

In 6 is a conditioning of the blade profile of a blade 10 by means of a clamping device 62, 64 and a tool 60 according to an embodiment of the invention.

First, the blade profile is clamped in its initial state 40 (in which the part is manufactured) by means of the clamps 62, 64 on the suction side 16 and the pressure side 18.

The tool 60 is brought into contact from the suction side 16 (indicated by the arrow), causing the trailing edge 12 to bend. The applied force and/or the resulting component deformation are monitored to ensure the constant mechanical load is maintained by a control system.

The temperature is cyclically increased from the initial temperature 66 to the final temperature 68 and then decreased back to the initial temperature 66. A sequence of initial temperature 66 → final temperature 68 → initial temperature 66 is referred to as a cycle 92. At the final temperature 68, a sufficient holding time must be selected to ensure complete heating of the blade 10. During cycles 92, the trailing edge 12 of the blade 10 continues to yield, so that the tool 60 must be adjusted by the control system.

At the end of conditioning, the deflected state 44 is below the martensite end temperature 84 and the deflected state 42 is above the austenite end temperature 88, since the elastic modulus of austenite 96 is greater than that of martensite 94.

When designing the tool 60, the usual effects of bending, such as elastic springback, can be advantageously taken into account. For this reason, the curvature of the tool 60 is 6 exaggerated.

The blade profile varies depending on the position in the spanwise direction of the blade 10. The tool 60 is adapted accordingly. It also allows for deformations of varying degrees, for example, greater deformation at the blade tip than near the blade root, or vice versa.

Depending on the aerodynamic conditions during the planned operation of the respective blade 10, the tool 60 can alternatively be applied from the pressure side 18. In addition, the designations of the states 40, 42, 44 can be 6 vary from case to case.

In 7 A section of an isometric view of a stator 120 of a turbo component 50, 51, 52, 54 is shown. The stator blades 20 are arranged radially between an outer stator ring 22 and an inner stator ring 24. The stator 120 can, for example, be formed in one piece. A stator blade 20 is provided with reference numerals as an example. The flow direction 70 is indicated.

8 shows a stator blade 20 of a turbo component 50, 51, 52, 54 of a turbomachine 100 with a deformable trailing edge 12 according to an embodiment of the invention in a plan view.

The deformable airfoil 10 of the stator blade 20 comprises a central part 11, a leading edge 14, and a trailing edge 12. The leading edge 14 adjoins the central part 11 upstream of the flow direction 70 when used as intended, and the trailing edge 12 adjoins the central part 11 in the flow direction 70. In the illustrated embodiment, the trailing edge 12 is formed at least in some regions from a shape memory material.

Alternatively, the front edge 14 could also be formed from the shape memory material.

As a shape memory material, a nickel-rich shape memory material with the main components Ni, Ti, as well as Hf and/or Zr, in particular with a nominal composition Ni _50.3 Ti _29.7 Hf _20-x Zr _x with 0≤x≤20 can be used advantageously.

The blade 10 can advantageously be manufactured by means of an additive manufacturing process.

The dashed area of the trailing edge 12 can move perpendicular to the drawing plane due to the two-way shape memory effect, depending on the operating temperature of the component.

The blade 10 is attached to an outer stator ring 22 on an outer circumference and to an inner stator ring 24 on an inner circumference. The outer stator ring 22 and the inner stator ring 24 can be made of NiTi, Ti, Ni, or another suitable material to reduce weight, for example, since no shape memory effect—especially not with high fatigue strength at high temperatures—is required here.

In cross section, the stator blade 20 can be as in 9 The marked area of the trailing edge 12 can move upwards or downwards in this illustration, depending on the operating temperature of the component.

The trailing edge 12 can thus have at least one, in particular at least two, deflected states 42, 44 of the trailing edge 12 and/or leading edge 14 relative to an initial state 40, as shown in 6 The two deflected states 42, 44 of the trailing edge 12 can be conditioned by means of a tool 60 after production.

To increase deformability and reduce weight, this area can have various cavities extending both in the plane of the drawing and perpendicular to it. The cavities can be formed with a circular or elliptical cross-section in the profile of the blade 10.

The middle part 11 of the blade 10 can also have cavities to save weight, which are located both in the plane of the drawing of the 9 as well as perpendicular to it.

In 10 A section of an isometric view of a rotor 130 of a turbo component 50, 51, 52, 54 is shown. The rotor blades 30 are arranged on a blade root 32 and extend radially outward. The rotor 130 can, for example, be formed in one piece. A rotor blade 30 is provided with reference numerals as an example. The flow direction 70 is indicated.

In 11 a rotor blade 30 of a turbo component 50, 51, 52, 54 of a turbomachine 100 with a deformable trailing edge 12 according to an embodiment of the invention is shown in a plan view.

The dashed area of the trailing edge 12 can move perpendicular to the drawing plane due to the two-way shape memory effect, depending on the operating temperature of the component.

The blade root 32 connects the blade 10 to the rotor disk. The blade root 32 can be made of NiTi, Ti, Ni, or another suitable material to reduce weight, since no shape memory effect—especially not with high fatigue strength at high temperatures—is required.

The center section 11 of the blade 10 may move perpendicular to the plane of the drawing, opposite to the trailing edge 12, depending on the component's temperature. These movements are smaller than the movements of the trailing edge 12 due to the lower deformability in this area.

In cross section, the rotor blade 30 can be as in 12 The marked area of the rear edge 12 can move upwards or downwards in this illustration, depending on the operating temperature of the component. To increase the formability and to reduce weight, this area can have various cavities, which are located both in the drawing plane of the 12 as well as perpendicular to it.

The central part 11 of the rotor blade 30 may possibly be opposite to the trailing edge 12 in the plane of the drawing in 12 Move downwards or upwards, depending on the component's temperature. To increase formability and reduce weight, this area can contain cavities that extend both in the plane of this figure and perpendicular to it.

Reference symbol

10

blade

11

Middle part

12

trailing edge

14

leading edge

16

suction side

18

Print page

20

Stator blade

22

outer stator ring

24

inner stator ring

30

rotor blade

32

Blade foot

40

Initial state

42

deflected state

44

deflected state

50

Turbo component, fan

51

Turbo component, compressor

52

Turbo component, turbine

54

Turbo component, turbine

56

combustion chamber

57

bypass channel

60

Tool

62

Clamping suction side

64

Clamping pressure side

66

Initial temperature

68

Final temperature of conditioning

70

Flow direction

72

Flow direction

74

Flow direction

76

Flow direction

80

temperature

82

Martensite onset temperature

84

Martensite final temperature

86

Austenite initial temperature

88

Austenite final temperature

90

Tension

92

Cyclization

94

Martensite phase

96

Austenite phase

100

Turbomachine

110

Engine axis

120

stator

130

rotor

QUOTES CONTAINED IN THE DESCRIPTION

This list of documents submitted by the applicant was generated automatically and is included solely for the convenience of the reader. This list is not part of the German patent or utility model application. The DPMA assumes no liability for any errors or omissions.

Cited patent literature

• US 2023/0160307 A1 [0010]
• US 8,657,561 B2 [0011]
• US 11,549,383 B2 [0012]

Claims (23)

A method for producing a deformable blade (10) of a turbomachine (100), in particular a gas turbine or a steam turbine, at least comprising: - optimizing a geometric profile of a trailing edge (12) and/or leading edge (14) of a blade (10) of a selected turbo component (50, 51, 52, 54) comprising a shape memory material at least in one or more regions for at least one, in particular at least two, deflected states (42, 44) relative to an initial state (40) of the blade (10), whose operating temperatures and/or mechanical loads under real operating conditions are at least partially within a permissible operating range of a predetermined shape memory material at the trailing edge (12) and/or leading edge (14); - Manufacturing the blade (10) in the initial state (40), wherein the trailing edge (12) and/or leading edge (14) is manufactured at least partially from the predetermined shape memory material; - Conditioning the blade (10) for the at least one, in particular the at least two deflected states (42, 44). Procedure according to Claim 1 , wherein at least one turbo component (50, 51, 52, 54) is selected according to at least one of the following criteria: - a maximum operating temperature of the blade (10) of the turbo component (50, 51, 52, 54) under real operating conditions is less than a final conditioning temperature (68); - there are operating points at which the operating temperature of the blade (10) is greater than an austenite final temperature (88); - there are operating points at which the operating temperature of the blade (10) is less than a martensite final temperature (84) of the shape memory material. Procedure according to Claim 1 or 2 , wherein the optimization of the geometric profile of the trailing edge (12) and/or leading edge (14) for the at least one, in particular the at least two deflected states (42, 44) is carried out iteratively until a termination criterion of the iteration is reached. Method according to one of the preceding claims, wherein the optimization of the geometric profile of the trailing edge (12) and/or leading edge (14) of the blade (10) takes place under a condition that with the at least one, in particular the at least two deflected states (42, 44), less than a predetermined limit value, in particular less than 5%, of an additional local strain is generated. Method according to one of the preceding claims, wherein cavities are provided in the profile of the blade (10), in particular cavities with a smoothly edged, in particular a circular or elliptical or oval, cross-section in the profile of the blade (10). Method according to one of the preceding claims, wherein a nickel-rich shape memory material with the main components Ni, Ti, and Hf and/or Zr is selected, in particular with a nominal composition Ni _50.3 Ti _29.7 Hf _20-x Zr _x with 0≤x≤20. Procedure according to Claim 6 , whereby the material composition of the shape memory material is chosen so that the proportion x is as large as possible. Method according to one of the preceding claims, wherein the blade (10) is manufactured by means of an additive manufacturing process. Method according to one of the preceding claims, wherein the initial state (40) of the blade (10) is adapted such that after conditioning at the corresponding temperatures, the leading edge (14) and/or trailing edge (12) of the blade (10) assumes at least one, in particular the two deflected states (42, 44). Method according to one of the preceding claims, wherein the conditioning of the airfoil (10) comprises at least the following steps: - clamping the airfoil (10) in its initial state (40) on the suction side (16) and the pressure side (18); - applying a tool (60) from the suction side (16) or the pressure side (18) of the airfoil (10); - performing a temperature treatment by increasing a temperature of the airfoil (10) from an initial temperature (66) lower than a martensite final temperature (84) to a final temperature (68) higher than an austenite final temperature of the shape memory material, followed by lowering the temperature to the initial temperature (66); - cyclically performing the temperature treatment several times. Procedure according to Claim 10 , wherein an applied force and/or a generated deformation of the blade (10) are monitored. Procedure according to Claim 10 or 11 , wherein the tool (60) is guided in a controlled manner to the suction side (16) or pressure side (18) of the blade (10) during the temperature treatment. Method according to one of the preceding claims, wherein, in order to determine operating temperatures and/or mechanical loads of the blades (10) of the turbo components (50, 51, 52, 54), averaged operating temperatures of the blades (10) are determined in the simulation for real operating conditions. Method according to one of the preceding claims, wherein the blade (10) is a stator blade (20) or a rotor blade (30) of the turbo component (50, 51, 52, 54). Deformable blade (10) of a turbo component (50, 51, 52, 54) of a turbomachine (100), in particular a gas turbine or a steam turbine, produced by a method according to one of the preceding claims, at least comprising a central part (11), a leading edge (14) and a trailing edge (12), wherein the leading edge (14) adjoins the central part (11) upstream of a flow direction (70) when used as intended and the trailing edge (12) adjoins the central part (11) in the flow direction (70), wherein the leading edge (14) and/or the trailing edge (12) are formed at least in regions from a shape memory material. Deformable blade according to Claim 15 , wherein the shape memory material is nickel-rich with the main components Ni, Ti, as well as Hf and/or Zr, in particular with a nominal composition Ni _50.3 Ti _29.7 Hf _20-x Zr _x with 0≤x≤20. Deformable blade according to Claim 15 or 16 , wherein the trailing edge (12) and/or leading edge (14) each have at least one, in particular at least two, deflected states (42, 44) of the trailing edge (12) and/or leading edge (14) relative to an initial state (40), depending on an operating temperature of the trailing edge (12) and/or leading edge (14), in particular wherein the at least one, in particular the at least two deflected states (42, 44) of the trailing edge (12) and/or leading edge (14) are conditioned by means of a tool (60) after production. Deformable blade according to one of the Claims 15 until 17 , wherein cavities are formed in the profile of the blade (10), in particular cavities with a smooth edge, in particular a circular or elliptical or oval cross-section in the profile of the blade (10). Deformable blade according to one of the Claims 15 until 18 , which is manufactured using an additive manufacturing process. Deformable blade according to one of the Claims 15 until 19 , wherein the blade (10) is a stator blade (20) or a rotor blade (30) of the turbo component (50, 51, 52, 54). Turbomachine (100) with at least one deformable blade (10) according to one of the Claims 15 until 20 , at least comprising a turbo component (50, 51, 52, 54) with the at least one blade (10) which is produced according to the method according to one of the Claims 1 until 14 is manufactured. Use of a nickel-rich material with the main components Ni, Ti, and Hf and/or Zr, in particular with a nominal composition Ni _50.3 Ti _29.7 Hf _20-x Zr _x with 0≤x≤20, as a shape memory material for a deformable blade (10) of a turbo component (50, 51, 52, 54) of a turbomachine (100), in particular in regions of a trailing edge (12) and/or leading edge (14) of the deformable blade (10). A computer program product comprising instructions which, when executed by a computer, cause the computer to perform steps of a method for producing deformable blades (10) of a turbomachine (100), in particular a gas turbine or a steam turbine, at least comprising: - optimizing a geometric profile of a trailing edge (12) and/or leading edge (14) of a blade (10) of a turbo component (50, 51, 52, 54) comprising a shape memory material in at least one or more regions for at least one, in particular at least two, deflected states (42, 44) relative to an initial state (40) of the blade (10), the operating temperatures and/or mechanical loads of which, under real operating conditions, lie at least partially at the trailing edge (12) and/or leading edge (14) within a permissible operating range of a predetermined shape memory material.