The invention is directed to a method for cleaning a component from powder residues of an additive layering method using a cleaning device, wherein a machine plate and the component arranged thereon are excited during a cleaning process by a vibration actuator of the cleaning device with a set resonance frequency of the machine plate to carry out a mechanical vibration. It is provided that the machine plate is excited by predefined vibration movements of the at least one vibration actuator to the predefined mechanical vibration, wherein the predefined vibration movements of the at least one vibration actuator occur in parallel to a main plane of the machine plate. The invention also relates to a cleaning device for cleaning an at least partially additively manufactured component, in particular a component of a turbomachine.
B08B 7/02 - Cleaning by methods not provided for in a single other subclass or a single group in this subclass by distortion, beating, or vibration of the surface to be cleaned
B08B 13/00 - Accessories or details of general applicability for machines or apparatus for cleaning
B33Y 30/00 - ADDITIVE MANUFACTURING, i.e. MANUFACTURING OF THREE-DIMENSIONAL [3D] OBJECTS BY ADDITIVE DEPOSITION, ADDITIVE AGGLOMERATION OR ADDITIVE LAYERING, e.g. BY 3D PRINTING, STEREOLITHOGRAPHY OR SELECTIVE LASER SINTERING - Details thereof or accessories therefor
B33Y 40/20 - Post-treatment, e.g. curing, coating or polishing
The present invention relates to a method of calibrating a thermal FEM-model of a turbomachine, the method comprising: i) creating a meta-model of the thermal FEM-model; ii) feeding, for a set of thermal variables, a set of input values to the meta-model for a calculation in the meta-model; iii) obtaining a set of output values from the calculation in the meta-model for the set of input values; iv) comparing the set of output values to measurement data of the turbomachine.
A module for a turbomachine, including a guide blade arrangement, a seal carrier that is situated radially within an inner platform of the guide blade arrangement, seal carrier walls, namely, a first seal carrier wall and a second seal carrier wall, and a sliding body as well as a connecting element. The seal carrier walls have a multipart design with respect to one another, and the second seal carrier wall is fastened to the first seal carrier wall and to the seal carrier via the connecting element. The sliding body holds the seal carrier walls at a distance from one another in such a way that they mutually axially delimit a clearance in which the guide blade arrangement engages via a radially inwardly extending guide pin. A radially inner section of the second seal carrier wall has frictional contact relative to the seal carrier.
METHOD FOR AVOIDING RESONANCE DAMAGE DURING CLEANING OF AN AT LEAST PARTLY ADDITIVELY MANUFACTURED COMPONENT, CLEANING DEVICE, MASS ELEMENT, AND SYSTEM
The invention is directed to a method for cleaning powder residues of an additive layer build-up method away from an at least partly additively manufactured component by a cleaning device, wherein a machine plate and the component arranged thereon are excited to mechanical oscillation during a cleaning process by a vibration actuator of the cleaning device with a set resonant frequency of the machine plate. According to the invention, before the cleaning process is carried out, a resonant frequency of the machine plate is set to the set resonant frequency by an arrangement of a mass element on a securing element of the machine plate.
The present invention relates to a method for inspecting a component, in particular a component of a turbomachine (1), including the steps of: capturing (S2) at least one X-ray or CT image of the component (10) using an image-capturing device (20); providing (S21) metadata about the component (10), the metadata including, in particular, a component type, a running time of the component (10), a number of remaining life cycles, and/or a repair history; classifying, by a machine learning system (30), the component (10) into a “serviceable” category or a “non-serviceable” category based on the image captured by the image-capturing device (20) and the provided metadata.
A geared turbofan engine includes a plurality of turbine stages, wherein for each stage (i) of the turbine, an inner radius Ri has a maximum deviation between +1.5% and −3% as compared to the average inner radius of the inner blade platforms of the plurality of stages. The engine further includes a fan, the fan coupled to the turbine stages via a gear.
F02K 3/06 - Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low-pressure outputs, for augmenting jet thrust, e.g. of double-flow type with front fan
7.
BRAZE ALLOY MIX FOR APPLICATION IN A METHOD FOR BRAZING A COMPONENT, ADDITIVE ALLOY, BRAZING METHOD, AND COMPONENT
The invention relates to a braze alloy mix for application in a method for brazing a component that has a nickel-based superalloy as base material, wherein the braze alloy mix comprises the following powders in a predetermined mixing ratio: a powder of a first braze alloy, a powder of a second braze alloy, a powder of a third braze alloy, and a powder of an additive alloy.
A cooling system (10) for an aircraft is provided, in particular for cooling at least one element of an electric propulsion system (24) of the aircraft, the cooling system (10) being at least partially disposed within a wing (12) of the aircraft and including at least two heat exchangers (16, 42) interconnected via at least one cooling circuit (11), the first heat exchanger (16) being operatively connected to the at least one element to be cooled of the electric propulsion system (24), and the second heat exchanger (42) being disposed within at least one wing (12) and being operatively connected to an upper and/or lower wing shell (46, 48) of the wing (12). A method for cooling at least one element of an electric propulsion system (24) of an aircraft is also provided.
B64D 33/08 - Arrangement in aircraft of power plant parts or auxiliaries not otherwise provided for of power plant cooling systems
B60L 58/33 - Methods or circuit arrangements for monitoring or controlling batteries or fuel cells, specially adapted for electric vehicles for monitoring or controlling fuel cells for controlling the temperature of fuel cells, e.g. by controlling the electric load by cooling
A method for automatically generating an optimized maintenance plan for a fleet of aircraft engines, includes the steps of: acquiring input data on a plurality of engines and providing an existing initial maintenance plan or creating an initial maintenance plan based on the acquired input data. A total maintenance effort for the fleet resulting in an application of the initial maintenance plan is then determined. Next, the engines are sorted into a defined order according to at least one criterion and at least one optimization strategy or heuristic stored as an algorithm in a computer program is applied to each of the engines in the defined order. Next, an optimized maintenance plan for the engines or output data comprising an estimated total maintenance effort of the optimized maintenance plan is output.
A rotor blade (20) for a rotor blade assembly (10) of a turbomachine (1) is provided, having an inner rotor blade platform (40) which extends axially from the rotor blade (20) with respect to a longitudinal turbomachine axis (2) and has two opposite circumferential end faces (41) and a free axial end (42) whose cross section is radially inwardly and radially outwardly bounded by circular arcs of two concentric circles. The inner rotor blade platform (40) has a cross section of connection (45) with the rotor blade (20) which is bounded radially inwardly by an inner connecting line (46) and radially outwardly by an outer connecting line (47). Each of the connecting lines (46, 47) has a central portion (61) having a convex curvature.
The invention is directed to an engine that has a fan, a compressor with a high-pressure compressor, and a combustion chamber. The high-pressure compressor of the engine has a mean stage pressure ratio and an overall pressure ratio formed between the fan and the combustion chamber.
F02C 3/06 - Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor the compressor comprising only axial stages
The invention relates to a compressor for an engine, wherein the compressor has compressor stages arranged in succession in a flow direction of the compressor and each compressor stage has a rotating blade cascade and a guide vane cascade arranged downstream of the rotating blade cascade and the rotating blade cascade and the guide vane cascade each have an aspect ratio.
F02C 3/06 - Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor the compressor comprising only axial stages
13.
GUIDE VANE DEVICE, MOUNTING TOOL, AS WELL AS TURBOMACHINE AND METHOD FOR MOUNTING AND DISMANTLING THE GUIDE VANE DEVICE
The invention is directed to a guide vane device for a guide vane adjustment of a turbomachine, comprising a guide vane, which has a shaft extending along an adjustment axis of the guide vane, as well as an adjusting lever, which forms a connecting element for connection to an outer end of the shaft, wherein, at its outer end, the shaft has a three-dimensional front surface with at least one beveled region, the connecting element surrounds, at least in sections, the outer end of the shaft in a sleeve-like manner and is joined to the outer end in a form-fitting manner, and the guide vane device forms an interference fit between the connecting element and the outer end of the shaft.
An airfoil array segment for an airfoil array of a turbomachine including a platform having a platform surface and an upstream-side platform edge, as well as at least two airfoils, whose leading and trailing edges define an inter-airfoil strip, the platform surface having a trough with a bottom configured such that the bottom is a global minimum and a radial position of the trough decreases relative to a reference surface in the circumferential direction from a suction side of one of the airfoils toward the pressure side of the adjacent other airfoil toward the bottom and from there increases in the circumferential direction, at most up to the radial position of the reference surface, and a radial position of the trough decreasing relative to a reference surface axially in the downstream direction toward the bottom and from there increasing axially in the downstream direction, at most up to the radial position of the reference surface, and the platform surface reaching at most a radial position of the reference surface, the reference surface corresponding to an uncontoured platform surface.
A heat engine, in particular an aircraft engine, having a first compressor for supplying a combustion chamber of the heat engine with air and a first turbine arranged downstream of the combustion chamber for driving the first compressor, wherein the heat engine also has at least one steam supply line for supplying steam from a steam source into the combustion chamber. The heat engine also has a steam supply device, which has a second compressor and is designed to compress the working gas further by the second compressor as a function of a mass flow conducted through the steam supply line, before the working gas flows into the combustion chamber.
The invention relates to a device and a method for cooling a heat exchanger of a fuel cell of a flight propulsion drive outside of a flight phase of an aircraft. The flight propulsion drive has an air line with at least one compressor for feeding compressed air for the operation of the fuel cell. A flow amplifier is arranged in the region of the heat exchanger, and is configured to guide air onto a cooling surface of the heat exchanger, whereby thermal energy is diverted from the cooling surface of the heat exchanger.
B64D 33/08 - Arrangement in aircraft of power plant parts or auxiliaries not otherwise provided for of power plant cooling systems
H01M 8/04007 - Auxiliary arrangements, e.g. for control of pressure or for circulation of fluids related to heat exchange
H01M 8/04111 - Arrangements for control of reactant parameters, e.g. pressure or concentration of gaseous reactants using a compressor turbine assembly
17.
BLADE FOR A TURBOMACHINE INCLUDING BLADE TIP ARMOR AND AN EROSION PROTECTION LAYER, AND METHOD FOR MANUFACTURING SAME
blade for a turbomachine is provided. The blade at its blade tip (4) includes blade tip armor (5), and an erosion protection layer (11) above the blade tip armor. For the blade, the erosion protection layer in the area of the blade tip has a layer thickness in the range of 5 µm to 100 µm, in particular 10 µm to 50 µm.
The present invention relates to an exhaust-gas treatment device for an aircraft engine, comprising an exhaust-gas channel, through which an exhaust gas of the aircraft engine flows, and a first cooling unit for cooling with ambient air, characterized by a second cooling unit, which is downstream of the first cooling unit with respect to an exhaust-gas flow in the exhaust-gas channel.
F01N 3/02 - Exhaust or silencing apparatus having means for purifying, rendering innocuous, or otherwise treating exhaust for cooling, or for removing solid constituents of, exhaust
F02C 7/141 - Cooling of plants of fluids in the plant of working fluid
F02C 7/16 - Cooling of plants characterised by cooling medium
19.
GUIDE VANE RING AND ROTOR BLADE RING FOR A TURBOFAN ENGINE
A guide vane ring and/or a rotor blade ring for a low-pressure turbine of an aircraft gas turbine of a turbofan engine includes a plurality of airfoils disposed consecutively in a direction of rotation. The plurality of airfoils includes a first airfoil and a second airfoil adjacent to the first airfoil in the direction of rotation. The first and second airfoils are disposed relative to each other such that a ratio e/t, viewed in a conical section, results from
a smallest free flow cross-section e taken on the pressure side of the first airfoil from its trailing edge to a suction side of the second airfoil, and
a pitch t of the airfoils, such that e/t is less than 0.31, (e/t<0.31).
F01D 9/04 - Nozzles; Nozzle boxes; Stator blades; Guide conduits forming ring or sector
F02K 3/06 - Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low-pressure outputs, for augmenting jet thrust, e.g. of double-flow type with front fan
20.
GUIDE VANE ASSEMBLY FOR A TURBOMACHINE, COMPRESSOR MODULE, TURBOMACHINE, AND METHOD FOR PRODUCING A GUIDE VANE ASSEMBLY
The invention relates to a guide vane assembly for a turbomachine, comprising a guide vane, which has a guide vane airfoil; and a guide vane holder. The guide vane is mounted in the guide vane holder such that the guide vane can be moved about an axis of rotation. For this purpose, the guide vane has at least one axle element, which is inserted into the guide vane holder in such a way that an outer lateral surface of the axle element faces an inner lateral surface of the guide vane holder. A protective coating is applied to at least parts of the guide vane airfoil. A protective coating is applied to at least one of the lateral surfaces. The invention also relates to a compressor module, a turbomachine, and a method for producing a guide vane assembly.
The present invention relates to a method for testing a component by a penetrant test, in which a penetrant is introduced onto a surface of the component; at least one portion of the penetrant is removed from the surface by blasting the surface with a solid blasting material; the surface is optically detected after the blasting.
An aircraft system includes, among other things, an aircraft and a gas turbine engine coupled to the aircraft. The gas turbine engine includes a propulsor section including a propulsor, a compressor section, a turbine section including a first turbine and a second turbine, and a gear reduction between the propulsor and the second turbine. The second turbine includes a number of turbine blades in each of a plurality of rows of the second turbine. The second turbine blades operating at least some of the time at a rotational speed. The number of blades and the rotational speed being such that the following formula holds true for a majority of the blade rows of the second turbine: 5500 Hz≤(number of blades×speed)/60 sec≤10000 Hz. The gas turbine engine is rated to produce 15,000 pounds of thrust or more.
F04D 29/66 - Combating cavitation, whirls, noise, vibration, or the like; Balancing
F02C 3/107 - Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor with two or more rotors connected by power transmission
F02K 3/04 - Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low-pressure outputs, for augmenting jet thrust, e.g. of double-flow type
G06F 30/17 - Mechanical parametric or variational design
F02C 3/04 - Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor
F02K 3/06 - Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low-pressure outputs, for augmenting jet thrust, e.g. of double-flow type with front fan
F01D 5/06 - Blade-carrying members, e.g. rotors - Details thereof, e.g. shafts, shaft connections
F01D 25/24 - Casings; Casing parts, e.g. diaphragms, casing fastenings
F02C 7/36 - Power transmission between the different shafts of the gas-turbine plant, or between the gas-turbine plant and the power user
F04D 25/04 - Units comprising pumps and their driving means the pump being fluid-driven
The present invention relates to an aircraft comprising at least one wing, at least one flight propulsion drive, and a retainer, particularly an engine pylon, which interconnects the wing and the flight propulsion drive. The aircraft comprises at least one heat exchanger for cooling exhaust gas of the fight propulsion drive and/or at least one water removal channel having at least one removal apparatus for removing water from exhaust gas of the flight propulsion drive, especially after the exhaust gas has flowed through the heat exchanger. The removal apparatus is disposed on, more particularly in, the retainer or is connected to the wing by means of the retainer, and/or the flight propulsion drive is fastened to the retainer by means of at least one flight propulsion drive suspension means, and the heat exchanger is fastened, independently thereof, to the retainer by means of at least one heat exchanger suspension means.
The invention relates to a blisk segment for a gas turbine, comprising at least one first blade having an airfoil, a leading edge, a trailing edge, a blade root, a suction side and a pressure side, a pedestal, and a first fillet having a device for influencing crack growth. The robustness of the gas turbine is improved in accordance with the invention in that, at the leading edge of the first blade, at least on the first fillet, a first surface structure is arranged and interacts with the crack-influencing device for influencing crack growth and for flow influencing.
F01D 5/34 - Rotor-blade aggregates of unitary construction
25.
METHOD FOR DETERMINING INDIVIDUAL VECTORS FOR OPEN-LOOP AND/OR CLOSED-LOOP CONTROL OF AT LEAST ONE ENERGY BEAM OF A LAYERING APPARATUS, AND LAYERING APPARATUS
The invention relates to a method for determining individual vectors for open-loop and/or closed-loop control of at least one energy beam of a layering apparatus, comprising at least the steps of: providing layer data characterizing at least one component layer of a component to be additively manufactured, on the basis of the layer data, determining individual vectors, according to which at least one energy beam is to be moved relative to a construction and joining zone of the layering apparatus in order to solidify a material powder selectively to the component layer, determining at least one node point of a plurality of individual vectors, and adapting at least one property of at least one individual vector of the at least one node point, the at least one property being selected from a group comprising spatial orientation, radiation sequence in relation to at least one other individual vector, and vector length.
A guide vane for arrangement in a gas duct of a turbomachine, including a guide vane blade and an outer platform which, relative to a longitudinal axis of the turbomachine, is arranged radially on an outside on the guide vane blade. The outer platform delimits the gas duct with an inner wall surface radially to the outside. The inner wall surface has a curved profile with an inflection point when viewed in an axial section, the inner wall surface having a concave curvature in a first axial section which is upstream from the inflection point with respect to a flow through the gas duct, and a convex curvature in a second axial section which is arranged downstream from the inflection point with respect to the flow through the gas duct.
The present invention relates to a lever for adjusting an adjustable vane of a turbomachine, wherein the lever has a fulcrum for the rotatable mounting of the lever about an axis of rotation, a first load arm with a first adjusting connection point for at least indirect connection to an adjustable vane and a first force arm with a first actuating connection point for connection to an adjusting device, wherein the lever further has a first strut, which connects the first adjusting connection point and the first actuating connection point to each other and thereby extends in an arc-shaped manner about the axis of rotation.
F01D 17/16 - Final actuators arranged in stator parts varying effective cross-sectional area of nozzles or guide conduits by means of nozzle vanes
F01D 17/12 - Final actuators arranged in stator parts
F01D 9/04 - Nozzles; Nozzle boxes; Stator blades; Guide conduits forming ring or sector
28.
METHOD FOR DETECTING DEFECTS IN A COMPONENT, METHOD FOR TRAINING A MACHINE LEARNING SYSTEM, COMPUTER PROGRAM PRODUCT, COMPUTER-READABLE MEDIUM, AND SYSTEM FOR DETECTING DEFECTS IN A COMPONENT
Provided is a method for detecting defects, in particular cracks and/or pores, in a component, in particular in a component of a turbomachine, preferably in a component of an engine, the method including the following steps: applying penetrant to at least a sub-region of the component such that the penetrant penetrates into any defects, in particular cracks and/or pores, present in the component; cleaning the surface of the component of penetrant that has not penetrated into defects, in particular cracks and/or pores, of the component; capturing an image, in particular a complete image, of the component; inputting the captured image into a machine learning system trained to detect defects, in particular cracks and/or pores; and detecting defects, in particular cracks and/or pores, in the component by machine learning system on the basis of light emitted and/or reflected by the penetrant in the defects, in particular cracks and/or pores.
A rotor blade for a gas turbine, in particular an aircraft gas turbine, including a blade root, a blade neck that adjoins the blade root in the radial direction, an airfoil that adjoins the blade neck in the radial direction, a radially outer partition wall that forms a radially inner delimiting section of an annular space of a gas turbine, an axially front partition wall and an axially rear partition wall that are connected to the radially outer partition wall so that the partition walls surround the blade neck on three sides, the partition walls protruding beyond the blade neck in the circumferential direction. For placement in a blade root receptacle of a rotor disk, the rotor blade is provided with a blade root protective plate that is situated between the blade root and the rotor disk. The blade root protective plate includes at least one sealing section that extends in the axial direction from the front partition wall to the rear partition wall, and whose radial outer side is situated opposite from the radially outer partition wall when the blade root protective plate is situated at the blade root. One or multiple ribs are situated at the blade neck for supporting the sealing section and are integrally joined to the blade neck.
A stator vane for a turbomachine, including an airfoil having a leading edge and a trailing edge, which are interconnected by a suction side and a pressure side, the airfoil extending substantially in a radial direction between an inner platform and an outer platform, wherein, when viewed in an axial direction of the turbomachine toward the trailing edge, the trailing edge extends so as to be inclined at a first angle to the pressure side radially on an outside in relation to the radial direction at a circumferential position at which the trailing edge meets the outer platform, and wherein the first angle between the trailing edge and a tangent to the outer platform at a transition from the outer platform to the trailing edge is between 72° and 84°.
The present invention relates to a guide vane assembly for a turbomachine, having an adjustable vane, which can be adjusted in order to alter an angle of attack, and an adjusting element, by way of which the adjustable vane can be adjusted, wherein the adjustable vane and the adjusting element are connected to each other in a connecting region, and wherein the adjustable vane and the adjusting element adjoin each other in the connecting region at a break edge.
An adjustable guide vane for a compressor, in particular a high-pressure compressor, of a gas turbine, in particular an aircraft gas turbine is described, the vane comprising a radially outer bearing section, a radially inner bearing section, and a vane section, which extends in the radial direction between the outer bearing section and the inner bearing section, wherein the outer and the inner bearing sections are designed in such a way that the adjustable guide vane can be taken up rotatably about a vane axis in the compressor, and wherein the radially inner bearing section is configured like a journal (cone-shaped) and has a lateral surface that is formed circumferentially about the vane axis, the lateral surface being of convex shape. In this way, it is provided that the radius of curvature of the convex lateral surface is at least double the maximum diameter of the bearing section.
F02C 3/06 - Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor the compressor comprising only axial stages
A connecting structure (10) for load transfer, in particular in a gas turbine (1), including a strut (20) and at least one wall element (30) is provided. The strut (20) at one end is integrally joined to the wall element (30), and the strut (20) and the wall element (30) are enclosed by a fillet (40), at least in areas, and integrally joined to same. An elastic deformation of the involved elements of the structure during the load transfer and/or load absorption is improved in that a root section (50) that is formed by a ridge (56) and that extends from the strut (20) to the wall element (30) is situated on the fillet (40).
The invention relates to a rotor for integration in a turbomachine, with a rotor base body, wherein, on the rotor base body in a rotor outer space of the rotor, at least one blade element or airfoil is mounted, wherein the rotor base body extends in an axial direction of the rotor and comprises a connecting section, which extends in an axial direction of the rotor for connection to at least one further rotor and separates the rotor outer space from a rotor inner space of the rotor, wherein the connecting section comprises, in the axial direction, at a front side, preferably in the axial direction at a downstream-lying front side, a balancing flange, wherein the balancing flange is configured for compensating an imbalance of the rotor and extends in the radial direction in the rotor outer space, at least in sections.
A method for checking at least one subregion of a component, in particular a component of a turbomachine, including at least the steps of a) providing a blank; b) producing at least the subregion from the blank by machining the blank using at least one tool and using at least one force sensor-to record at least one force curve of at least one force acting during machining on the at least one tool; c) checking whether there is at least one deviation-of the at least one force curve from at least one predetermined target curve-of the at least one force curve, the at least one deviation-characterizing at least one material defect-contained in an unmachined segment of the subregion. A checking device for checking at least a subregion of a component is also provided.
B23Q 17/09 - Arrangements for indicating or measuring on machine tools for indicating or measuring cutting pressure or cutting-tool condition, e.g. cutting ability, load on tool
G01N 3/40 - Investigating hardness or rebound hardness
G05B 19/418 - Total factory control, i.e. centrally controlling a plurality of machines, e.g. direct or distributed numerical control (DNC), flexible manufacturing systems (FMS), integrated manufacturing systems (IMS), computer integrated manufacturing (CIM)
36.
METHOD FOR PRODUCING AN IMPACT-RESISTANT COMPONENT, AND CORRESPONDING IMPACT-RESISTANT COMPONENT
A method for producing an impact-resistant component, in particular a component of a turbomachine, such as an aircraft engine, and a corresponding component. The component is produced at least partially by an additive manufacturing method from a powder material in such a way that the component is formed at least in a first region from a material with a first toughness and at least in a second region from a material with a second toughness, the second toughness being greater than the first toughness, and wherein the second region is formed, at least in a part of the component, as a continuous or interrupted layer, preferably parallel to the surface of the component, at a distance from the surface of the component.
A method is provided including the steps: —first excitation of the object via a multifrequency signal; —detecting a first response signal of the object at one or multiple measuring points at the object; —transforming the first response signal from a time range into a frequency-dependent range; —selecting one or multiple frequencies, based on the frequency-dependent range; —second excitation of the object based on the selected frequencies; —detecting a second response signal of the object at one or multiple measuring points of the object; —ascertaining a mechanical parameter based on the second response signal.
G01M 15/14 - Testing gas-turbine engines or jet-propulsion engines
G01M 5/00 - Investigating the elasticity of structures, e.g. deflection of bridges or aircraft wings
38.
ADDITIVELY MANUFACTURED INTERMEDIATE CHANNEL FOR ARRANGING BETWEEN A LOW-PRESSURE COMPRESSOR AND A HIGH-PRESSURE COMPRESSOR, AND CORRESPONDING MANUFACTURING METHOD
An intermediate duct (10) for disposition between an outlet of a low-pressure compressor and an inlet of a high-pressure compressor of a turbomachine, in particular of an aircraft engine is provided, the intermediate duct including an outer wall (2) and an inner wall (3) between which are disposed an optional exit stator ring (4) and at least one strut (12) extending radially with respect to a central axis of the intermediate duct (10). The intermediate duct (10) is at least partially manufactured by additive manufacturing. A method for manufacturing such an intermediate duct (10), and a turbomachine having such an intermediate duct (10) are also provided.
A guide blade arrangement (20) for a turbomachine (1), including a guide blade airfoil (22) and a platform (21). The guide blade airfoil (22) is situated at a side (21.1) of the platform (21) facing the gas channel, an opposite side (21.2) of the platform (21) facing away from the gas channel being contoured at least in one area (30.1, 30.2) with elevations (25) and depressions (26) that follow one another in the circumferential direction (23) in relation to a longitudinal axis (2) of the turbomachine (1), and the elevations (25) and depressions (26) at the side (21.2) facing away from the gas channel being set via a platform thickness (31), taken radially in each case, that is variable in the circumferential direction (23) and that repeatedly increases and decreases with a continuous profile.
The present invention relates to a method for testing a component, in particular an aircraft engine, comprising the steps of: determining (S40) a value of a first toleranced parameter (A1; A2) of the component; determining (S50) a value of a second toleranced parameter (E1; ...; E4) of the component; and classifying (S70) the component in a predefined quality class if this value pair lies outside of a predefined tolerance range, the upper and/or lower limit (G) of which for the second parameter depends on the first parameter, in particular linearly, in at least one first permissible value range (Ta1,1) of the first parameter.
Described is a heat-protection element (50) for a gas turbine (10), in particular an aircraft gas turbine, the heat-protection element (50) being adapted to at least partially surround a bearing chamber (60) of the gas turbine (10) and having at least one connecting portion (52) which is disposed in an axially forward region (VB) and connectable or connected by a material-to-material bond to a protective element (54) of a seal carrier, in particular a seal carrier with a carbon seal, at least one supporting portion (58) which is disposed in an axially central region (MB) and adapted to support the heat-protection element (50) radially on the bearing chamber (60), an end portion (64) which is disposed in an axially rearward region (HB) and forms a free end (66) of the heat-protection element (50) and which is configured such that the end portion surrounds (64) the bearing chamber (60) in a contactless manner.
Described is a rotor disk (40) for a compressor (29, 32) of a gas turbine, in particular an aircraft gas turbine (10), the rotor disk having a main body (42), at least one rotor arm (44) projecting from the main body (42) in the axial direction (AR), the rotor arm (44) having, in a sectional view taken in a sectional plane defined by the axial direction (AR) and the radial direction (RR) a beginning portion (44a) merging into the main body (42); an end (44e) portion remote from the main body (42) and forming a kind of free end in the axial direction (AR), the beginning portion (44a) and the end portion (44e) being interconnected by an intermediate portion (44z), characterized in that the intermediate portion (44z) is curved with at least one radius of curvature (Ri, Ra).
The invention relates to an airfoil as well as to a method for producing an airfoil for a turbomachine, comprising a leading edge and a trailing edge joined to each other by a suction side and a pressure side and which, in at least one region, extends in a curved manner from an airfoil root to an airfoil tip, wherein the airfoil tip has a squealer tip, which is arranged at the airfoil tip.
F01D 5/20 - Specially-shaped blade tips to seal space between tips and stator
44.
COMPRESSION SYSTEM FOR A GAS TURBINE, HIGH-PRESSURE COMPRESSOR, COMPRESSION SYSTEM COMPRISING A HIGH-PRESSURE COMPRESSOR, LOW-PRESSURE COMPRESSOR, COMPRESSION SYSTEM COMPRISING A LOW-PRESSURE COMPRESSOR, AND GAS TURBINE
The invention relates to a compression system for a gas turbine, in particular for an aircraft gas turbine, wherein the compression system comprises a flow duct. The flow duct includes cross-sectional areas that are aligned perpendicular to the axial direction along the flow duct length and have the respective predetermined sizes, where the inlet cross-sectional area has a size that is 15.3 to 16.1 times a size of the outlet cross-sectional area, and/or a cross-sectional area arranged at a distance of 40% of the flow duct length from the inlet cross-sectional area has a size that is 5.0 to 5.2 times the size of the outlet cross-sectional area.
A blade has an internal blade profile, and a jet deflection portion adjoining the internal blade profile. Each blade profile of the jet deflection portion has a center of gravity. The centers of gravity are joined by a thread line, which is a space line, starting from the internal blade profile. A total deviation of the space line in relation to the reference line is described by an nth order polynomial based on radial distance from the blade profile to a reference point. The thread line is selected and designed such that a compressive stress is produced in the region of the leading edge of the blade during operation.
A method (100) for classifying an etch indication (11) of a component (10), the method including the steps of: providing a captured image (13) of the at least one etch indication (11); detecting at least one criterion (15) of the etch indication (11) based on the captured image (13); comparing the determined criterion (15) to at least one criteria data set (16) of etch indications (11) that is stored in a database; and classifying the etch indication (11) into at least one predetermined defect class (D) based on the comparison.
The present invention relates to an aircraft having at least one wing, on which at least one propulsion unit is arranged, comprising at least one heat engine, especially a gas turbine, as well as an exhaust gas passage for conducting exhaust gas of the heat engine into and inside the wing.
The invention relates to a guide vane assembly as well as to a method for mounting a guide vane assembly of a turbomachine, including a number of adjustable guide vanes, the guide vane platform of which has a guide vane journal that is mounted in a receiving opening of a housing of the turbomachine, wherein a guide vane head of the guide vane has a bearing pin, which is mounted on an inner ring arranged on a rotor of the turbomachine. In the method, a guide vane journal, which is arranged on a guide vane platform of a guide vane, is brought into a receiving opening of a housing, which is arranged radially with respect to a rotor axis of the turbomachine, and an inner ring is provided on a rotor of the turbomachine.
A brush seal can be used for a gas turbine. The brush seal includes a support ring structure; and at least one bundle of bristles that is arranged in an axial direction on the support ring structure. The support ring structure has a basis portion arranged radially on the outside, and a supporting portion arranged radially on the inside. The supporting portion has a supporting surface that faces the bundle of bristles and supports the bundle of bristles in the axial direction. The supporting portion has a radially inner edge portion that faces a rotor portion of the gas turbine when the brush seal is in an assembled state. Starting from the radially inner edge portion, the supporting portion has a radial supporting-portion length together with a substantially constant axial supporting-portion width, the supporting-portion length being greater than the supporting-portion width by at least a factor of four.
F01D 11/12 - Preventing or minimising internal leakage of working fluid, e.g. between stages for sealing space between rotor blade tips and stator using a rubstrip, e.g. erodible, deformable or resiliently biased part
F01D 25/24 - Casings; Casing parts, e.g. diaphragms, casing fastenings
An adjustment lever adjusts a stator vane of a turbomachine. The adjustment lever has: a first connection site, of a plurality of connection sites, the first connection site being configured to join to an adjustment ring; a second connection site, of the connection sites, the second connection stie being configured to join to the stator vane; and a joining member arranged between the first connection site and the second connection site. The joining member is shaped having at least two struts which adjoin at least one of the connection sites.
The invention relates to a rotor for a turbomachine, having at least one blade and having at least one rotor main part, which has at least one recess, in which a blade root of the least one blade is interlockingly received, wherein the blade root comprises at least one depression, in which at least one protrusion of the at least one rotor main part which protrusion delimits the at least one recess in regions is received, wherein the at least one depression is delimited by a first delimiting face on the blade root side and the at least one protrusion is delimited by a second delimiting face on the rotor main part side. At least the first delimiting face has at least one elevation which narrows a gap at least in regions, which extends between the first delimiting face and the second delimiting face.
The invention relates to a monitoring system for a gas turbine, in particular for an aircraft engine. The monitoring system comprises at least one borescope device that is able to be mounted in a borescope opening of a gas turbine housing and has a housing, in which at least one optical sensor device for acquiring images of at least one inner region of the gas turbine is arranged, and an evaluation device that is able to be connected to the at least one borescope device in order to exchange data and is designed to inspect the at least one inner region for the presence of a fault on the basis of the at least one image acquired by way of the sensor device. The invention furthermore relates to a borescope device to an evaluation device and to a gas turbine.
F01D 21/00 - Shutting-down of machines or engines, e.g. in emergency; Regulating, controlling, or safety means not otherwise provided for
G01M 15/14 - Testing gas-turbine engines or jet-propulsion engines
G02B 23/24 - Instruments for viewing the inside of hollow bodies, e.g. fibrescopes
53.
METHOD FOR COATING A COMPONENT OF AN AIRCRAFT ENGINE WITH A WEAR-RESISTANT LAYER, AND COMPONENT FOR AN AIRCRAFT ENGINE WITH AT LEAST ONE WEAR-RESISTANT LAYER
A method for coating a component of an aircraft engine with a wear-resistant layer, wherein the component is first coated at least regionally with a nickel- or cobalt-based alloy and subsequently aluminized. Also disclosed is a method for producing a spray powder for producing a wear-resistant layer of a component of an aircraft engine.
The invention relates to a seal carrier for a turbomachine, in particular a gas turbine, having a carrier base and at least one seal body, wherein the at least one seal body is connected to the carrier base, and wherein the at least one seal body is formed by a plurality of cavities arranged next to one another, in particular uniformly, in the peripheral direction and in the axial direction, wherein the cavities extend out from the carrier base in the radial direction and are delimited by a cavity wall. According to the invention, the seal body has a plurality of damping portions which are designed to locally damp or disrupt the flow of force in the seal body, wherein the carrier base is continuous in the region of the damping portions.
F01D 11/12 - Preventing or minimising internal leakage of working fluid, e.g. between stages for sealing space between rotor blade tips and stator using a rubstrip, e.g. erodible, deformable or resiliently biased part
55.
Rotor arrangement for a gas turbine with inclined axial contact surfaces formed on rotor segments, gas turbine and aircraft gas turbine
A rotor assembly may be used in a gas turbine. The rotor assembly includes rotor segments arranged in succession in an axial direction and interconnected in the axial direction by a tie-rod, and a rotor segment disposed forwardly in the axial direction having a first contact surface and a rotor segment disposed rearwardly in the axial direction having a second contact surface. The first and second contact surfaces are at least partially in contact with each other, are substantially annular in shape, and extend in a radial direction and in a circumferential direction. The first contact surface and/or the second contact surface extend at least partially obliquely relative to the radial direction. An angle is formed between the first contact surface and the second contact surface when viewed in a sectional plane defined by the axial direction and the radial direction.
The present invention relates to an adjustment assembly for the adjustment of adjustable blades or vanes of a turbomachine, having an adjustment ring for coupling to the adjustable blades or vanes of a blade or vane ring and having an adjusting mechanism, which has a coupling rod for coupling to a further blade or vane ring as well as a lever and a push rod, wherein the lever is rotatably mounted at a pivot and has a load arm as well as a force arm, wherein the load arm of the lever is coupled to the push rod and its force arm is coupled to the coupling rod, on different sides of the lever, so that an offset of the coupling rod via the lever and the push rod is converted to a rotation of the adjustment ring around the ring axis thereof.
The invention relates to an airfoil for a compressor of a turbomachine, which extends starting from a blade root between a leading edge and a trailing edge to a blade tip, wherein the leading edge has a leading-edge thickness and the airfoil has a maximum profile thickness, the ratio of which to each other represents a relative leading-edge thickness, and the airfoil has a leading-edge wedge angle.
MANUFACTURING DEVICE FOR ELECTROCHEMICAL MACHINING OF A COMPONENT, IN PARTICULAR A TURBINE COMPONENT, METHOD FOR ELECTROCHEMICAL MACHINING OF A COMPONENT, AND COMPONENT
The invention relates to a manufacturing device for the electrochemical machining of a component, in particular a turbine component, wherein the manufacturing device comprises at least one machining device, which is set up to remove material of the component in accordance with a predetermined electrochemical machining method. It is provided that the manufacturing device comprises at least one cleaning device, which is set up to spray jets of the electrolyte solution onto the component in accordance with a predetermined jet-spraying method in order to remove a residue layer formed on the component during the predetermined electrochemical machining method.
The disclosure is directed to a method comprising the steps: carrying out multiple measurements on a mechanical object, the measurements each differing by one or more parameters influencing the measurement; determining a spectrogram on the basis of the measurement data of the measurements and depending on a predefined parameter of the mechanical object; determining one or more excitations of the mechanical object; reproducing the excitations in the spectrogram.
The invention relates to a machining apparatus for electrochemically removing component layers of a component, having at least one electrode, which is mounted so as to be movable along at least one infeed axis, and having at least one auxiliary electrode, which is mounted so as to be movable along an auxiliary infeed axis, wherein a gap for arranging the component for electrochemically removing the component layers extends between the at least one electrode and the at least one auxiliary electrode. At least the infeed axis and a longitudinal extension direction of the gap enclose an acute angle with each other. The machining apparatus comprises at least one oscillation device (40), which is set up at least to move the at least one electrode in an oscillating manner along the infeed axis and relative to the at least one auxiliary electrode.
Blade root receptacle for receiving a blade root of a rotor blade of a turbomachine. The blade root receptacle, for radially bearing in a form-fitting manner on the blade root, has a supporting flank which, in terms of a rotation axis, at least in proportions faces radially inward, wherein the supporting flank is provided with a convexity which, when viewed in an axially perpendicular section, at least in portions has a convex shape and, also when viewed in an axially parallel section, at least in portions has a convex shape.
The present invention relates to a blade for a gas turbine, in particular of an aircraft engine, having a blade airfoil, which has a blade-root-side first profile section and a blade-tip-side second profile section, which is spaced apart from the first profile section in a radial direction, from the first profile section to the second profile section, by a blade airfoil height, wherein a stagger angle of the blade airfoil changes with a height in the radial direction over the first profile section at least over certain portions, wherein, in a first region between a first height and a second, greater height, the change in the stagger angle over the height does not decrease with increasing height at least over certain portions.
The present invention relates to a rotor blade (20) for arrangement in a gas duct (2) of a turbomachine (1), having a rotor blade airfoil (23), which, viewed in a tangential section, has a blade airfoil profile (24) with a leading edge radius RVK and a rotor blade airfoil thickness d, wherein the blade airfoil profile (24) is thickened, at least in sections, specifically the blade airfoil thickness d is specified, in relation to the front edge radius RVK, such that (2d/Rvk2)−d≤5.5.
A turbine module (2) for a turbomachine (1). The turbine module (2) includes a main channel (26) to guide a main flow (36) through the turbine module (2), a rotor blade (21) and a stator vane (22), the stator vane (22) including a stator airfoil (22) and a platform (23), with the stator airfoil (22) arranged downstream of the rotor blade (21) in the main channel (26), and a cavity (30) including an inlet (31) for injecting a part (36.2) of the main flow (36) into the cavity (30), an outlet (32) for a reinjection of the part (36.2) of the main flow (36) from the cavity (30) into the main channel (26), wherein the cavity (30) is arranged at an axial position of the stator vane (20) and is radially offset from the stator airfoil (22).
The present invention relates to a rotating blade for a turbomachine, having a blade element and a shroud, wherein the shroud is profiled at a peripheral side, i.e., when observed in a tangential section, has a contact flank that is oriented axially, at least proportionally, and has a free flank that is also oriented axially, at least proportionally, and is in fact opposite the contact flank, wherein, when observed in the tangential section, an intermediate segment between the contact flank and the free flank has a first curvature in a first transition segment to the contact flank and has a second curvature in a second transition segment to the free flank, wherein the first curvature is greater than the second curvature.
The present invention relates to an apparatus and a method for determining a scaling factor for a strain measurement in a machine element, comprising steps for measuring a strain in a measurement surface portion by means of a strain measuring device; for measuring a displacement of a detection surface portion of the machine element by an optical scanning; for determining a displacement field on a surface of the machine element on the basis of a model of the machine element and the measured displacement of the at least one detection surface portion; for determining a strain field on the surface of the machine element on the basis of the determined displacement field and the model of the machine element; and for determining a scaling factor of the strain measuring device on the basis of the determined strain field and the measured strain in the measurement surface portion.
A rotor blade system includes a blade root, a blade neck adjoining the blade root, an airfoil adjoining the blade neck, a radially outer partition wall, an axially forward partition wall and an axially rearward partition wall, connected to the radially outer partition wall such that the partition walls surround the blade neck on three sides, the partition walls projecting beyond the blade neck in the circumferential direction, and further including a blade root protection plate designed to be arranged on the blade root. It is provided that the blade root protection plate have at least one sealing portion that extends in the axial direction from the forward partition wall to the rearward partition wall and whose radially outer side is disposed opposite the radially outer partition wall when the blade root protection plate is arranged on the blade root.
A turbomachine blade system, in particular for a compressor or turbine stage of a gas turbine, which includes at least one blade, in particular a moving or guide blade, and at least one moving body for reducing the vibrations of this blade, at least one area of a guide for guiding the body and/or at least one area of a supporting structure for resiliently mounting the body and/or at least one area of the body being or becoming generatively manufactured together with at least one area of the blade, in particular of a vane and/or blade root and/or a shroud situated thereon.
The invention relates to a device and a method for electrochemically treating a component, comprising at least one electrode which has at least one working surface with an outer contour that is shaped so as complement the surface of the component to be produced, thereby forming a gap, and comprising at least one contour surface which adjoins said working surface and in which at least one cleaning opening is arranged, a cleaning fluid flowing through said cleaning opening. According to the method, at least one component is provided, a voltage is applied between the component and the at least one electrode during the electrochemical treatment, and the electrode is moved relative to the component.
A bearing chamber housing (20) for bearing a shaft (3) of a turbomachine (1), including a housing outer shell (21) that delimits an oil chamber (33) of the bearing chamber housing (20) radially outwardly in relation to a rotational axis (4) of the shaft (3), and a housing inner shell (22) for bearing the shaft (3). The housing inner shell (22) is radially connected to the housing outer shell (21) via support ribs (23) that in each case extend axially, at least in part, and the housing inner shell (22), the housing outer shell (21), and two support ribs (23) that are next-adjacent to one another jointly delimit a cavity (41) that is axially open at the rear, and thus lead into a rear opening (32). The rear opening (32), viewed in tangential sections, has a clearance (35) in each case that constitutes at least 50% of a circumferential distance (43) between the next-adjacent support ribs (23).
An arrangement reduces oscillation (vibration) of an oscillatory structure. The arrangement has a structure having at least one mode in at least one direction; and an oscillation-reducing device (vibration-reducing device). The oscillation-reducing devices has a housing formed by or provided on the structure, a cavity, and a body configured for making impact contacts with the housing and disposed in the cavity in such a manner that the body is configured to make impact contacts with the housing at least temporarily for as long as the structure is excited in the at least one mode in the at least on direction.
F01D 5/26 - Antivibration means not restricted to blade form or construction or to blade-to-blade connections
F01D 5/16 - Form or construction for counteracting blade vibration
F16F 15/14 - Suppression of vibrations in rotating systems by making use of members moving with the system using freely-swinging masses rotating with the system
The present invention is directed to a method for coating a component, wherein the component has a first and a second surface, and wherein the first and the second surface adjoin each other at an edge, in which method i) first of all, the edge between the first and the second surface is rounded, and ii) subsequently, a coating is applied to the first surface.
B23P 15/02 - Making specific metal objects by operations not covered by a single other subclass or a group in this subclass turbine or like blades from one piece
C25D 9/08 - Electrolytic coating other than with metals with inorganic materials by cathodic processes
C25D 15/02 - Combined electrolytic and electrophoretic processes
The present invention relates to a method for coating a component of a turbomachine in a bath, in which method, the component is partially immersed in the bath containing a coating material; the component is rotated at least intermittently around an axis of rotation, which lies outside of the bath, during the at least partial immersion; the component is at most immersed partially over and beyond the rotation.
A method for investigating an electrolyte solution for processing a component, particularly a component or a component material of an aircraft engine, by near infrared spectroscopy.
G01N 21/359 - Investigating relative effect of material at wavelengths characteristic of specific elements or molecules, e.g. atomic absorption spectrometry using infrared light using near infrared light
G01N 21/3577 - Investigating relative effect of material at wavelengths characteristic of specific elements or molecules, e.g. atomic absorption spectrometry using infrared light for analysing liquids, e.g. polluted water
An arrangement reduces oscillation of an oscillatory structure. The arrangement has: a structure having at least one mode in at least one direction; and an oscillation-reducing device. The oscillation-reducing device includes a housing on the structure; a cavity; and a body. The body is configured to make impact contact with a first surface portion and a second surface portion of an inner wall of the housing and disposed is in the cavity such that the body can make impact contact with the first surface portion and the second surface portion of the inner wall of the housing at least temporarily for as long as the structure is excited in the at least one mode in the at least one direction. The first surface portion or the second surface portion of the inner wall of the housing has a curved profile.
F16F 15/14 - Suppression of vibrations in rotating systems by making use of members moving with the system using freely-swinging masses rotating with the system
F01D 5/16 - Form or construction for counteracting blade vibration
F01D 5/26 - Antivibration means not restricted to blade form or construction or to blade-to-blade connections
A bearing chamber housing for supporting a shaft of a turbomachine is provided, the bearing chamber housing including an additively built-up housing wall which bounds an oil chamber of the bearing chamber housing radially outwardly relative to an axis of rotation of the shaft, the housing wall being built up with an oil duct which has an inlet opening toward the oil chamber for admission of oil from the oil chamber into the oil duct, and which has an outlet opening for discharging the oil from the oil duct, the outlet opening being located at a different axial position and at a different circumferential position than the inlet opening, considered relative to the axis of rotation of the shaft, and the oil duct having an extent with both an axial component and a circumferential component, at least over a portion thereof.
The invention relates to an ECM system comprising a holder for the arranging of at least one component to be machined and at least one machining station for the electrochemical machining of the at least one component at at least one machining position. The ECM system has a positioning apparatus, which is designed to mount, in a predefined position, a machining platform arranged on the positioning apparatus, and the machining platform comprises at least one machining station having at least one electrode arranged thereon, which electrode can be moved along a machining path in order to machine the at least one component that can be arranged on the holder.
A curved contour of the lateral surface of a blade arrangement includes in at least one meridian section on mutually opposite sides of a blade airfoil an intersection point that is closer to the blade airfoil front edge, and an intersection point that is closer to the blade airfoil rear edge, and a best-fit line of least square distances from the curved contour. The curved contour includes first and/or the second contour section which meet specified conditions.
The present invention relates to a method for characterizing a coating, in which method a mass and/or a volume of a basic body is/are measured prior to coating; a mass and a volume of the basic body with the applied coating are measured; for characterizing the coating, a density of the coating is determined from the volume and mass measurements; wherein the volume is optically measured.
G01N 9/02 - Investigating density or specific gravity of materials; Analysing materials by determining density or specific gravity by measuring weight of a known volume
An integrally bladed rotor for a turbomachine, in particular a compressor or turbine stage of a gas turbine, to which at least one separately formed impulse element housing (40; 40′) is fastened by at least one fastening element (30; 30′) which engages for this purpose into an opening (41) of the impulse element housing and into an opening (11) of the rotor, the impulse element housing having at least one cavity (44) in which at least one impulse element (5) is accommodated with play.
A method for producing a rotary disk/blisk for a high-pressure compressor or a high-speed turbine and to a corresponding geared turbofan engine. The method involves providing a Ni base alloy comprising, in % by weight, 15.5-16.5 Cr, 14.0-15.5 Co, 4.75-5.25 Ti, 2.75-3.25 Mo. 2.25-2.75 Al, 1.00-1.50 W, as well as optionally 0.0250-0.0500 Zr, 0.0100-0.0200 B, 0.0100-0.0200 C, remainder Ni. The base alloy is shaped by forging to obtain a preform of the disk/blisk, the final contour thereof being produced by electrical discharge machining or electrochemical machining.
The invention relates to an energy supply device with at least one fuel cell and to a method for operating at least one energy supply device with at least one fuel cell, which has at least one anode that can be supplied with a fuel and at least one cathode that can be supplied with ambient air for generating electrical energy. The proposed energy supply device has a converter device.
H01M 8/043 - Processes for controlling fuel cells or fuel cell systems applied during specific periods
83.
LAYER CONSTRUCTION METHOD AND LAYER CONSTRUCTION DEVICE FOR ADDITIVELY MANUFACTURING AT LEAST ONE COMPONENT REGION OF A COMPONENT, AND COMPUTER PROGRAM PRODUCT AND STORAGE MEDIUM
The layer construction method comprises at least the following steps: a) applying at least one powder layer of a material to at least one construction and joining zone of at least one movable construction platform; b) locally solidifying the material to form a component layer, wherein the material is selectively scanned along scan lines by at least one energy beam and fused; c) lowering the construction platform layer by layer by a predefined layer thickness; and d) repeating the steps a) to c) until the component region is complete.
The layer construction method comprises at least the following steps: a) applying at least one powder layer of a material to at least one construction and joining zone of at least one movable construction platform; b) locally solidifying the material to form a component layer, wherein the material is selectively scanned along scan lines by at least one energy beam and fused; c) lowering the construction platform layer by layer by a predefined layer thickness; and d) repeating the steps a) to c) until the component region is complete.
In step b), a distance hs between at least two central lines of neighboring scan lines in at least one component layer is adjusted in accordance with Formula I
The layer construction method comprises at least the following steps: a) applying at least one powder layer of a material to at least one construction and joining zone of at least one movable construction platform; b) locally solidifying the material to form a component layer, wherein the material is selectively scanned along scan lines by at least one energy beam and fused; c) lowering the construction platform layer by layer by a predefined layer thickness; and d) repeating the steps a) to c) until the component region is complete.
In step b), a distance hs between at least two central lines of neighboring scan lines in at least one component layer is adjusted in accordance with Formula I
0.85≤bsmin/hs≤1.00 (I)
The layer construction method comprises at least the following steps: a) applying at least one powder layer of a material to at least one construction and joining zone of at least one movable construction platform; b) locally solidifying the material to form a component layer, wherein the material is selectively scanned along scan lines by at least one energy beam and fused; c) lowering the construction platform layer by layer by a predefined layer thickness; and d) repeating the steps a) to c) until the component region is complete.
In step b), a distance hs between at least two central lines of neighboring scan lines in at least one component layer is adjusted in accordance with Formula I
0.85≤bsmin/hs≤1.00 (I)
wherein bsmin represents a minimum melt pool width of the scan lines.
B33Y 30/00 - ADDITIVE MANUFACTURING, i.e. MANUFACTURING OF THREE-DIMENSIONAL [3D] OBJECTS BY ADDITIVE DEPOSITION, ADDITIVE AGGLOMERATION OR ADDITIVE LAYERING, e.g. BY 3D PRINTING, STEREOLITHOGRAPHY OR SELECTIVE LASER SINTERING - Details thereof or accessories therefor
B33Y 40/20 - Post-treatment, e.g. curing, coating or polishing
84.
METHOD AND APPARATUS FOR MACHINING COMPONENTS BY MEANS OF ELECTROCHEMICAL MACHINING
The invention relates to a method machining a particularly planar component by means of electrochemical machining, wherein the component has internal stresses resulting particularly from preceding manufacturing steps. In a first step a) of the method, the component to be machined is provided. Subsequently, in step b), at least two tools are provided in the form of electrodes and, in step c), an electrolyte is provided between the component and the at least two electrodes. In step d), a positive voltage is applied to the component and a negative voltage is applied to the at least two electrodes. Thus, in step e), by moving the at least two electrodes along their respective movement paths with respect to the component, electrochemical machining can take place; in the process, the gap between each electrode and the component is flushed with the electrolyte at least intermittently.
An intermediate element is for a blade/rotor disk connection in a rotor of a fluid flow machine. The intermediate element is adapted to a shape of a blade root of a blade and to a blade root slot in a rotor disk for receiving the blade root such that, when arranged between the blade root and rotor disk, the intermediate element prevents contact between the blade root and rotor disk. The intermediate element has, on an outer surface that faces the rotor disk, at least one protrusion to reduce an air flow parallel to an axis of rotation of the rotor between the rotor disk and the intermediate element; and on an inner surface that faces the blade root, a recess that corresponds to the at least one protrusion.
A rotor blade (20) for placement in a gas channel (3) of a turbomachine (1), including a rotor blade airfoil (23) which, in relation to a flow in the gas channel (3), includes a front edge (23a) and a rear edge (23b) downstream therefrom, as well as a suction side (41) and a pressure side (42). The rotor blade airfoil (23) is provided with an inclination toward the suction side (41) over at least one section (45.1) of its radial rotor blade airfoil height (45). The inclination is set in such a way that during operation a centrifugal force bending moment (46), which effectuates the centrifugal force on the rotor blade airfoil (23) due to the inclination, is greater than a gas force bending moment (47) that acts on the rotor blade airfoil (23) due to the circulation around the rotor blade airfoil (23) in the gas channel (3).
Rotor blade (20) to be arranged in a gas conduit (3) of a turbomachine (1), having a rotor blade airfoil (23), which radially inwardly has a chord length Si, radially outwardly has a chord length Sa, and in a radial position
Rotor blade (20) to be arranged in a gas conduit (3) of a turbomachine (1), having a rotor blade airfoil (23), which radially inwardly has a chord length Si, radially outwardly has a chord length Sa, and in a radial position
rx inbetween has a chord length Sx, the chord length S in the radial position rx being at least equal to the chord length Si radially inwardly (Si
A guide vane segment for a gas turbine having at least one guide vane, a radially inner covering strip and a radially outer covering strip, wherein the at least one guide vane extends between the inner covering strip and the outer covering strip. The inner covering strip and/or the outer covering strip having a respective support portion which extends in a circumferential direction (UR) and has a respective axial end face. The support portion comprises on the respective inner covering strip and/or the respective outer covering strip at least one support portion rib which extends in a radial direction, is connected to the support portion, and comprises a rib end face which merges into the axial end face of the support portion.
A gas turbine has a rotor blade. The rotor blade has a blade root connected to an airfoil. The blade root has a root contour with respect to a cross-sectional view. From a lower end of the blade root, the blade root contour has convex contour portions and concave contour portions. From the lower end along the blade root contour between a convex contour portion and an adjoining concave contour portion, there is a contour portion as a flank portion that is load-bearing. From the lower end along the blade root contour between a concave contour portion and an adjoining convex contour portion, there is a contour portion as a flank portion that is not load-bearing in operation. At least one of the concave contour portions has a first arc portion, a second arc portion, and a straight portion disposed between the two arc portions.
LAYER BUILDING PROCESS AND LAYER BUILDING APPARATUS FOR THE ADDITIVE MANUFACTURE OF AT LEAST ONE WALL OF A COMPONENT, AS WELL AS COMPUTER PROGRAM PRODUCT AND STORAGE MEDIUM
The invention relates to a layer building process for the additive manufacture of at least one wall region of a component including applying at least one powder layer of a material to at least one building-up and joining zone of at least one movable building platform, carrying out a first solidifying step, in which the material is irradiated selectively with at least one energy beam, wherein irradiation parameters of the at least one energy beam are set so a molten bath is produced and a defect-affected wall region of the wall is produced, without applying a further powder layer, carrying out a second solidifying step, in which the defect-affected wall region produced in the first solidifying step is irradiated selectively with the at least one energy beam, lowering the building platform layer by layer by a predefined layer thickness, and repeating the steps above one or more times
B33Y 30/00 - ADDITIVE MANUFACTURING, i.e. MANUFACTURING OF THREE-DIMENSIONAL [3D] OBJECTS BY ADDITIVE DEPOSITION, ADDITIVE AGGLOMERATION OR ADDITIVE LAYERING, e.g. BY 3D PRINTING, STEREOLITHOGRAPHY OR SELECTIVE LASER SINTERING - Details thereof or accessories therefor
B33Y 80/00 - Products made by additive manufacturing
91.
ROTOR DISK AND ROTOR BLADE FOR A GAS TURBINE COMPRESSOR OR TURBINE STAGE OF AN AEROENGINE
A method for producing a rotor disk or a rotor blade for a gas turbine compressor stage or turbine stage of an aeroengine, wherein at least one blade groove of the rotor disk for arrangement of a blade foot of a rotor blade for fastening the rotor blade to the rotor disk, or a blade foot of the rotor blade for arrangement in a blade groove of a rotor disk for fastening the rotor blade to the rotor disk is fabricated using an electrochemical material removal and in the axial direction has a profile which is curved once or more.
The present invention relates to a guide vane, in particular an outlet guide vane and/or a guide vane for a compressor stage of a gas turbine, wherein the vane has a vane blade with a first vane blade portion and a second vane blade portion, and the first vane blade portion can be reversibly rotated in relation to the second vane blade portion about an axis of rotation from a first position to a second position, wherein, in at least one profile portion of the vane blade, the axis of rotation is arranged outside of a profile of the first vane blade portion, and/or a profile of the first vane blade portion has a suction side with a first contour portion and a second contour portion, which, in particular, is adjoined thereto, and a profile of the second vane blade portion has a pressure-side contour portion.
The present invention relates to a method for manufacturing a plate of a turbomachine having a plurality of receiving grooves, which are formed on the cylindrical peripheral surface of the plate in order to receive blade roots of blades of the turbomachine. The receiving grooves extend in a straight line from one face of the plate to the other face of the plate, such that a straight line delimits the receiving groove from each point on the cross section of the receiving groove on one of the faces to the corresponding point on the cross section on the other face. A circular plate is provided and a plurality of receiving grooves are introduced into the plate simultaneously on opposing regions relative to a central axis of rotation of the plate.
A method for examining a component with the steps:
Determining an actual value of a first parameter of a first characteristic of the component;
Determining an actual value of the first parameter of at least one further characteristic of the component;
Determining a first statistical characteristic variable of these actual values of the first parameter; and
Classifying the component on the basis of the first statistical characteristic variable and the actual values of the first parameter,
wherein the component is classified in a predetermined quality class if the first statistical characteristic variable is outside a predetermined first characteristic variable range or at least one of these actual values of the first parameter is outside a predetermined first range of values.
A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a propulsor section, a geared architecture, a high spool and a low spool. The high spool includes a high pressure compressor and a high pressure turbine. The low spool includes a low pressure compressor and a low pressure turbine. At least one stage of the turbine section includes an array of rotatable blades and an array of vanes. A ratio of the number of vanes to the number blades is greater than or equal to 1.55. A mechanical tip rotational Mach number of the blades is greater than or equal to 0.5 at an approach speed.
F02K 3/06 - Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low-pressure outputs, for augmenting jet thrust, e.g. of double-flow type with front fan
F02C 3/107 - Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor with two or more rotors connected by power transmission
F02C 7/36 - Power transmission between the different shafts of the gas-turbine plant, or between the gas-turbine plant and the power user
96.
METHOD FOR CARRYING OUT MAINTENANCE WORK ON A COMPLEX STRUCTURAL COMPONENT
The invention relates to a method for carrying out maintenance work on a complex structural component, in particular, a gas turbine, by means of a virtual work environment. The virtual work environment has a display device and a computing device, said computing device being connected to a memory device, in which the characteristics of the complex structural component, together with its component parts and their assembly sequence, are stored. During the method, at least one component part needing maintenance, or a predefined modification state of the structural component is selected in the virtual work environment From this step, the disassembly or assembly steps required for disassembling the component parts of the structural component are calculated and displayed, in order to obtain access to the at least one component part needing maintenance, or to achieve the selected modification state.
The invention relates to a method for producing a welding wire that includes the steps of providing a hollow wire, through at least part of which at least one cavity extends; producing the welding wire by introducing a welding material containing titanium aluminide or at least one nickel-based superalloy into the at least one cavity, the at least one cavity being evacuated or being filled with a protective gas before, during and/or after the introduction of the welding material, and the hollow wire being formed from nickel if the welding material contains the at least one nickel-based superalloy. Further aspects of the invention relate to a welding wire and to a component having at least one component region obtained by hardfacing using at least one such welding wire.
B23K 35/40 - Making wire or rods for soldering or welding
B22F 7/08 - Manufacture of composite layers, workpieces, or articles, comprising metallic powder, by sintering the powder, with or without compacting of composite workpieces or articles from parts, e.g. to form tipped tools with one or more parts not made from powder
B23K 35/02 - Rods, electrodes, materials, or media, for use in soldering, welding, or cutting characterised by mechanical features, e.g. shape
B23K 35/30 - Selection of soldering or welding materials proper with the principal constituent melting at less than 1550°C
B23K 35/36 - Selection of non-metallic compositions, e.g. coatings, fluxes; Selection of soldering or welding materials, conjoint with selection of non-metallic compositions, both selections being of interest
A gas turbine component, in particular a stator vane cluster (10), having at least one radial flange (12, 13) with a radial slot (20) for spoke-type centering of a gas turbine element, in particular an inner ring (100), the radial slot (20) having a first slot flank (21), a slot base (23) adjoining the same, and a second slot flank (22) adjoining the same, the slot base (23) being asymmetric with respect to a mid-plane (M) which extends radially and perpendicularly to the circumferential direction and centrally between the two slot flanks (21, 22).
The invention relates to a method and a set of balance weights for balancing a rotor particularly of a turbomachine, on which a plurality of balance weights, which are distributed over the periphery, can be arranged for balancing an imbalance. A set of balance weights, in particular plate-shaped weights, with a predetermined number of groups of balance weights, is used for balancing the rotor. Each of the balance weights of every group of balance weights has an identical weight that is assigned to this group and a shape assigned to this group, which differs in basic shape and/or thickness of the balance weights from the balance weights of the other groups.
A sealing apparatus for a turbomachine, including a stator vane component, which includes an inner shroud element and a flow-directing element connected to the inner shroud element. The sealing apparatus includes a sealing component, which has a seal-carrier ring element coupled to the stator vane component. The seal-carrier ring element includes at least one ring body element and at least one projection, which is connected in one piece to the at least one ring body element, protrudes from the at least one ring body element in the radial direction of the sealing apparatus, and is inserted into at least one opening, which extends through the inner shroud element. Other aspects relate to a seal-carrier ring element for a sealing apparatus, and to a turbomachine which includes at least one sealing apparatus and/or at least one seal-carrier ring element.
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