Method for carrying out service measures on an energy conversion installation, and energy conversion installation

Abstract

The invention relates to a method for optimizing service measures for energy generation plants.

Description

[0001] The invention relates to a method for carrying out service measures on an energy conversion plant and an energy conversion plant.

[0002] There are various requirements for gas turbines in energy conversion plants.

[0003] These can be machines for baseload power supply or for balancing load changes, especially due to renewable energies whose input into the power grid can vary. Requirements can include different locations, cooling options, fuels, etc.

[0004] Furthermore, there are also different requirements regarding the desired service intervals or modifications and improvements due to different initial models.

[0005] It is therefore the purpose of the invention to demonstrate modifications that correspond to different operating conditions or customer requirements.

[0006] The problem is solved by a method according to claim 1 and an energy conversion plant according to claim 10, wherein a corresponding existing gas turbine is provided or modified accordingly or newly manufactured.

[0007] The dependent claims list further advantageous measures which can be combined arbitrarily to achieve further advantages.

[0008] The show Figure 1 a gas turbine, Figure 2 a combustion chamber, Figure 3 a turbine rotor blade, Figure 4 a list of superalloys, Figure 5 a gas turbine in cross-section with an improved bearing, Figure 6 a cross-section of a gas turbine with improved burners, Figure 7 a compressor casing in cross-section with a gas turbine, Figure 8 a cross-section of a gas turbine with a guide vane and rotor blade area in each stage, Figure 9 a turbine blade with cooling holes on the side faces, Figure 10 a cooled tip of a turbine blade, Figure 11 a blade carrier of a rear turbine stage, Figure 12 a transition from combustion chamber brick to guide vane, Figure 13 a sealing arrangement of a guide vane carrier, Figure 14 combustion chamber bricks in a combustion chamber with spoiler effect, Figure 15 a combustion chamber brick, Figures 16, 17 each a gap between two combustion chamber bricks according to Figure 15Figures 18 and 19 show a housing with the insertion of an additional seal, Figures 20 and 21 show a burner with modified guide vanes of the swirler, Figure 22 shows a modified base of a steam turbine, Figure 23 shows a device for monitoring the combustion dynamics, and Figure 24 shows a power generation plant.

[0009] The figures and description represent only exemplary embodiments of the invention.

[0010] The Figure 1 shows, as an example, a gas turbine machine 100 in a longitudinal section.

[0011] The gas turbine machine 100 has inside a rotor 103 rotatably mounted around a rotational axis 102 with turbine blade 120, which is also referred to as a turbine runner.

[0012] Along the rotor 103, an intake housing 104, a compressor 105, a combustion chamber 110 (e.g., a torus-shaped combustion chamber, in particular a ring combustion chamber) with several coaxially arranged burners 107, a turbine 108 and the exhaust housing 109 follow one another.

[0013] The annular combustion chamber 110 communicates with a preferably annular hot gas channel 111. There, for example, four turbine stages 112 connected in series form the turbine 108.

[0014] Each turbine stage 112: I, II, III, IV is preferably formed from two blade rings.

[0015] Viewed in the direction of flow of a working medium 113, in the hot gas channel 111 a row of guide vanes 115 is followed by a row of guide vanes 125 formed from guide vanes 120.

[0016] The guide vanes 130 are attached to a gas turbine housing 138 of a stator 143, whereas the rotor blades 120 of a rotor blade row 125 are attached to the rotor 103, for example by means of a turbine disk 133.

[0017] A generator 5 is coupled to the rotor 103 ( Fig. 24 ) or a working machine (not shown).

[0018] During operation of the gas turbine 100, air 135 is drawn in and compressed by the compressor 105 through the intake housing 104. The compressed air supplied at the turbine-side end of the compressor 105 is directed to the burners 107 in a combustion chamber 110 and mixed there with a fuel. The mixture is then combusted in the combustion chamber 110, forming the working fluid 113. From there, the working fluid 113 flows along the hot gas channel 111 past the guide vanes 130 and the rotor blades 120. At the rotor blades 120, the working fluid 113 expands, transferring momentum so that the rotor blades 120 drive the rotor 103, which in turn drives the machine coupled to it.

[0019] The components exposed to the hot working medium 113 are subject to thermal stresses during the operation of the gas turbine 100. The guide vanes 130 and rotor blades 120 of the first turbine stage 112, viewed in the direction of flow of the working medium 113, are subjected to the highest thermal stresses, along with the heat shield elements lining the combustion chamber 110.

[0020] To withstand the temperatures prevailing there, they can be cooled using a coolant.

[0021] Similarly, the substrates of the components can have a directional structure, i.e., they are single-crystal (SX structure) or have only longitudinally oriented grains (DS structure).

[0022] For example, iron-, nickel- or cobalt-based superalloys are used as material for the components, especially for the turbine blade 120, 130 and components of the combustion chamber 110.

[0023] Such superalloys are preferably known from EP 1 204 776 B1, EP 1 306 454, EP 1 319 729 A1, WO 99 / 67435 or WO 00 / 44949 or in Figure 4 listed.

[0024] The blades 120 and 130 can also have anti-corrosion coatings: MCrAlX; M is at least one element from the cobalt (Co) and nickel (Ni) group, X is an active element and represents yttrium (Y) and / or tantalum (Ta) and / or at least one rare earth element or hafnium (Hf) or iron (Fe). Such alloys are known from EP 0 486 489 B1, EP 0 786 017 B1, EP 0 412 397 B1 or EP 1 306 454 A1.

[0025] MCrAlX may have a thermal insulation layer, and consists, for example, of ZrO 2 , Y 2 O 3 -ZrO 2 , i.e., it is not, partially or completely stabilized by yttrium oxide and / or calcium oxide and / or magnesium oxide and / or erbium oxide and / or ytterbium oxide.

[0026] The guide vane 130 has a guide vane root (not shown here) facing the gas turbine casing 138 of the turbine 108 and a guide vane tip opposite the guide vane root. The guide vane tip faces the rotor 103 and is fixed to a mounting ring 140 of the stator 143.

[0027] Figure 2 shows a combustion chamber 110 of a gas turbine.

[0028] The combustion chamber 110, for example, is designed as a so-called annular combustion chamber, in which a multitude of burners 107 arranged circumferentially around a rotational axis 102 open into a common combustion chamber 110, generating flames 156. For this purpose, the combustion chamber 110 as a whole is designed as a ring-shaped structure positioned around the rotational axis 102.

[0029] To achieve a comparatively high efficiency, the combustion chamber 110 is designed for a relatively high working medium temperature of approximately 1273 K to 1873 K. To enable a comparatively long service life even under these operating parameters, which are unfavorable for the materials, the combustion chamber wall 153 of the combustion chamber 110 is provided on its working medium-facing side with an inner lining formed from heat shield elements 155. Each heat shield element 155, made of an alloy, is equipped on the working medium side with a particularly heat-resistant protective layer (MCrAlX layer and / or ceramic coating) or is made of high-temperature-resistant material (solid ceramic bricks).

[0030] These protective layers of the metallic heat shield elements 155 can be similar to those of turbine blades; thus, for example, MCrAlX means: M is at least one element from the group iron (Fe), cobalt (Co), nickel (Ni), X is an active element and stands for yttrium (Y) and / or silicon (Si) and / or tantalum (Ta) and / or at least one rare earth element or hafnium (Hf) and / or iron (Fe). Such alloys are known from EP 0 486 489 B1, EP 0 786 017 B1, EP 0 412 397 B1 or EP 1 306 454 A1.

[0031] The MCrAlX may also have a, for example, ceramic thermal insulation layer and consists of ZrO 2 , Y 2 O 3 -ZrO 2 , i.e. it is not, partially or completely stabilized by yttrium oxide and / or erbium oxide, ytterbium oxide and / or hafnium oxide.

[0032] Many coating processes are conceivable, e.g., atmospheric plasma spraying (APS), LPPS, VPS, or CVD. The thermal barrier layer can have porous, micro- or macro-cracked grains for improved thermal shock resistance. Refurbishment means that heat shield elements 155 may need to be stripped of protective layers after use (e.g., by sandblasting). This is followed by the removal of corrosion and / or oxidation layers or products. Any cracks in the heat shield element 155 are also repaired, if necessary. Finally, the heat shield elements 155 are recoated and reused.

[0033] Due to the high temperatures inside the combustion chamber 110, a cooling system may also be provided for the heat shield elements 155 or their retaining elements. The heat shield elements 155 are then, for example, hollow and may also have cooling holes (not shown) opening into the combustion chamber space 154.

[0034] The Figure 3 The figure shows in perspective a guide vane 120 or guide vane 130 of a turbomachine, which extends along a longitudinal axis 121.

[0035] The turbomachine can be a gas turbine of an aircraft or a power plant for generating electricity, a steam turbine, or a compressor.

[0036] The bucket 120, 130 has a mounting area 400, an adjacent bucket platform 403, a bucket blade 406 and a bucket tip 415 successively along its longitudinal axis.

[0037] As a guide shovel 130, it can have another platform at its shovel tip 415 (not shown).

[0038] In the mounting area 400 a blade root 183 is formed, which is used to attach the rotor blades 120, 130 to a shaft or a turbine disk 133 ( Fig. 1 ) serves.

[0039] The shovel foot 183, for example, is designed as a hammerhead. Other designs, such as a fir tree or dovetail foot, are possible.

[0040] The blade 120, 130 has a leading edge 409 and a trailing edge 412 for a medium flowing past the blade 406.

[0041] For conventional blades 120, 130, solid metallic materials, especially superalloys, are used in all areas 400, 403, 406 of the blade 120, 130, for example.

[0042] Such superalloys are preferably made from EP 1 204 776 B1, EP 1 306 454, EP 1 319 729 A1, WO 99 / 67435 or WO 00 / 44949 or from Figure 4 known.

[0043] The shovel 120, 130 can be manufactured by a casting process, including directed solidification, by a forging process, by a milling process or combinations thereof.

[0044] Workpieces with single-crystal structure or structures are used as components for machines that are exposed to high mechanical, thermal and / or chemical stresses during operation.

[0045] The production of such single-crystal workpieces is achieved, for example, by directed solidification from the melt. These are casting processes in which the liquid metallic alloy solidifies into a single-crystal structure, i.e., a single-crystal workpiece, or in a directed manner.

[0046] In this process, dendritic crystals align along the heat flow, forming either a columnar grain structure (i.e., grains that extend the entire length of the workpiece and are commonly referred to as directionally solidified) or a single-crystal structure (i.e., the entire workpiece consists of a single crystal). In these processes, the transition to globular (polycrystalline) solidification must be avoided, as undirected growth necessarily leads to the formation of transverse and longitudinal grain boundaries, which negate the desirable properties of the directionally solidified or single-crystal component.

[0047] When referring generally to directionally solidified structures, this includes both single crystals, which have no grain boundaries or at most low-angle grain boundaries, and columnar crystal structures, which have grain boundaries running longitudinally but no transverse grain boundaries. These latter crystalline structures are also called directionally solidified structures.

[0048] Such methods are known from US Patent 6,024,792 and EP 0 892 090 A1.

[0049] The blades can also have coatings against corrosion or oxidation, specifically MCrAlX; M is at least one element of the cobalt (Co) or nickel (Ni) group, X is an active element and represents yttrium (Y) and / or tantalum (Ta) and / or at least one rare earth element and / or hafnium (Hf) and / or iron (Fe). Such alloys are known from EP 0 486 489 B1, EP 0 786 017 B1, EP 0 412 397 B1 or EP 1 306 454 A1.

[0050] The density is preferably at 95% of the theoretical density.

[0051] A protective aluminum oxide layer (TGO = thermally grown oxide layer) forms on the MCrAlX layer (as an intermediate layer or as the outermost layer).

[0052] The MCrAlX may also have a thermal insulation layer, which is preferably the outermost layer, and consists, for example, of ZrO 2 , Y 2 O 3 -ZrO 2 , i.e., it is not, partially or completely stabilized by yttrium oxide and / or calcium oxide and / or magnesium oxide and / or erbium oxide and / or ytterbium oxide.

[0053] The thermal barrier layer covers the entire MCrAlX layer. Other coating methods are conceivable, e.g., atmospheric plasma spraying (APS), LPPS, VPS, or CVD. The thermal barrier layer can contain porous grains with micro- or macro-cracks for improved thermal shock resistance. Therefore, the thermal barrier layer is preferably more porous than the MCrAlX layer.

[0054] Refurbishment means that components 120 and 130 may need to have their protective coatings removed after use (e.g., by sandblasting). This is followed by the removal of corrosion and / or oxidation layers or products. Any cracks in component 120 or 130 are also repaired, if necessary. Finally, component 120 or 130 is recoated and put back into service.

[0055] The bucket 120, 130 can be hollow or solid. If the bucket 120, 130 is to be cooled, it is hollow and may also have cooling holes 418 (indicated by dashed lines).

[0056] Figure 24 Figure 1 shows an example of an energy conversion plant 1 with one unit. This arrangement according to Figure 24 It can be present multiple times in an energy conversion plant, or in a modified form.

[0057] The gas turbine 100 is coupled to a generator 5 for power generation via a gearbox 4 or a coupling 4.

[0058] The generator 5 is also connected to a steam turbine 6 via a coupling 2.

[0059] Steam turbines 6 are present in combined cycle power plants. A power conversion plant 1 can also consist of only a gas turbine 100 without a steam turbine 6.

[0060] A condenser 7 is connected to the steam turbine 6, if present. The exhaust gas from the gas turbine 100 flows via a diffuser 8 into a heat recovery system 9, where the hot exhaust air is used to generate steam.

[0061] There is also an exhaust chimney 10.

[0062] The idea is to carry out service work on an energy conversion plant, wherein the energy conversion plant comprises at least the following machines: at least one gas turbine, at least one generator and optionally at least one steam turbine, wherein repairs are carried out on the at least one machine, in particular a defective component or defective components of the at least one machine are replaced or are replaced by either a new, identical component or new, identical components and / or are repaired, and wherein, in carrying out these repairs, further measures to extend the service life of machines or their components and / or further measures to optimize (efficiency) of machines or their components are carried out.

[0063] In particular, the defective components include turbine blades or their coatings and / or Burners or burner components and / or compressor blades or their coatings and / or combustion chamber bricks.

[0064] The defective components can preferably only include turbine blades.

[0065] The defective components may preferably only include turbine blades or their coatings, as well as burners or burner components.

[0066] Further measures will preferably only be implemented to extend service life.

[0067] Further measures will preferably only be implemented to optimize the system.

[0068] Furthermore, measures to extend the service life of components and measures to optimize components can preferably be implemented as additional measures.

[0069] During service measures, preferably at least one, in particular at least two identical, or At least two different measures for extending the service life of machines or their components from the group: rotor bearings, burner, compressor blade, compressor housing, turbine blades, gas turbine housing, blade carrier, heat shields or combustion chamber bricks, seals, combustion chamber - turbine transition, cooling and / or monitoring devices, must be implemented.

[0070] Service measures preferably involve at least one, in particular at least two identical or at least two different measures for optimizing machines or their components from the group: efficiency increase, cooling improvement, burner, compressor blade, compressor housing, turbine blades, gas turbine housing, blade carrier, heat shields or combustion chamber bricks, seals and / or transition combustion chamber - turbine, must be carried out.

[0071] The individual measures are described in more detail below and can be combined as needed, depending on the requirements: Method for modifying a gas turbine machine comprising at least: a compressor, a combustion chamber, a hot gas duct with a bladed rotor, and at least one rotor bearing of the rotor at the beginning of the compressor in a flow direction of the gas turbine machine, wherein the rotor bearing is replaced, wherein the new rotor bearing is at least 5% longer, or that a new rotor bearing at least 370 mm long is installed, in particular wherein the new rotor bearing is a maximum of 500 mm long; Method for modifying a gas turbine machine comprising at least: a compressor, a combustion chamber, a hot gas duct with a bladed rotor, at least one burner for the combustion chamber, wherein the fuel supply means, in particular pipes, are at least partially, in particular completely, internally provided with a diffusion coating, in particular alitized, or the fuel supply means,in particular, pipes with a diffusion coating inside, in particular internally coated, are installed; a method for modifying a gas turbine machine comprising at least: a compressor, a combustion chamber, a hot gas duct with a bladed rotor, wherein the compressor has a compressor housing that is or is designed in two parts, and in particular in which an inner compressor housing is installed as a blade support made of a first material, in particular steel, most especially cast steel, and the outer compressor housing has as a blade support a second material significantly different from the first material, in particular gray cast iron, or in which the inner compressor housing has gray cast iron or is replaced by gray cast iron; a method for modifying a gas turbine machine comprising at least: a compressor, a combustion chamber, a hot gas duct with a bladed rotor, with stages (I, II, III,IV) of impeller and guide vanes, wherein higher temperature-resistant impeller and guide vanes are incorporated, in particular in stages (I, II), which in particular have improved cooling, or wherein the impeller or guide vanes in the hot gas channel (111) have a directionally solidified microstructure in the form of a columnarly solidified microstructure, in particular only the first two stages (I, II), most in particular only the first stage (I), or wherein a segmented ceramic layer based on yttrium-stabilized zirconia (HGB) is present or applied to the impeller and guide vanes, or wherein the impeller and guide vanes for the metallic substrate have a single-crystal microstructure, or such are incorporated, in particular only the first two stages (I, II), or wherein the ceramic coating comprises partially stabilized yttrium-stabilized zirconia with a porosity of 12±4%,or with a TBC without segmentation on the guide or rotor blades, or in which a blade tip in a recess has a stepped shoulder which connects directly to a web on the suction side and thus represents additional material in the recess, wherein a cooling air bore from the interior of the rotor blade extends through the shoulder to better cool the blade tip, a method for modifying a gas turbine machine comprising at least: a compressor, a combustion chamber, a hot gas duct with a bladed rotor, with stages (I, II, III, IV) of rotor and guide blades, in which the rotor and guide blades are installed, in particular in the stages (I, II) which have cooling holes on the side surfaces of the blade platforms, a method for modifying a gas turbine machine comprising at least: a compressor, a combustion chamber, a hot gas duct with a bladed rotor, with stages (I, II, III,IV) of rotor and guide vanes, wherein rotor blades are installed, wherein the blade tip of the rotor blades, in particular of stages (I, II), is cooled, in particular by cooling holes in the blade tip; a method for modifying a gas turbine machine, which at least comprises: a compressor, a combustion chamber, a hot gas duct with a bladed rotor, with stages (I, II, III, IV) of rotor and guide vanes, wherein rotor blades of stage (IV) are installed which are not cooled, in particular the guide vanes of stage (III) are also not cooled; a method for modifying a gas turbine machine, at least comprising: a compressor, a combustion chamber, a hot gas duct with a bladed rotor, in which, viewed in the direction of flow, a front plenum and a rear plenum are located outside the hot gas duct, which, due to technical reasons, have different pressures;wherein the front plenum is located in the direction of flow behind the rotor blade of stage (III) and above the guide vane of stage (IV), and either channels that were present in the middle of stage (IV) between the guide vane and rotor blade and were previously used to cool the turbine guide vanes and rotor blades of stage (IV) or to supply cooling air are closed off, and a new long channel from the rear plenum is subsequently introduced into the blade carrier, or a new blade carrier is provided and installed which has only one such channel, method for modifying a gas turbine machine which has at least: a compressor, a combustion chamber with combustion chamber bricks or heat shields, a hot gas duct with a bladed rotor,wherein a gap between a heat shield and the guide vane of stage (I) of the engine is formed with a rounding at the flow-side end of the heat shield and the opposite rounding of the guide vane of stage (I) being identical in order to avoid an overhang or undercut in the heat shield in which dirt could accumulate or erosion could occur, a method for modifying a gas turbine machine, comprising at least: a compressor, a combustion chamber, a hot gas duct with a bladed rotor, wherein a guide vane carrier is installed or modified with a sealing arrangement of elements which results in lower cooling air consumption, wherein the individual elements of the guide vane carrier have a gap which is labyrinthine or S-shaped, wherein the element leading in the flow direction has a first nose and the second element trailing in the flow direction has a second nose formed above it,so that an S-shaped gap is formed, whereby the opening of the gap in the hot gas duct is located aft in the direction of flow; a method for modifying a gas turbine machine comprising at least: a compressor, a combustion chamber with heat shields or combustion chamber bricks, a hot gas duct with a bladed rotor, wherein combustion chamber bricks are installed which are designed to generate a spoiler effect; a method for modifying a gas turbine machine comprising at least: a compressor, a combustion chamber with combustion chamber bricks, a hot gas duct with a rotor, wherein combustion chamber bricks are installed which accommodate two separate recesses on two opposite side faces of the combustion chamber brick, which serve to engage a mechanical clamping from the rear of the combustion chamber brick; a method for modifying a gas turbine machine comprising at least: a compressor,a combustion chamber with combustion chamber bricks, a hot gas duct with a bladed rotor, a housing part for the hot gas duct, wherein a recess for a seal is inserted into the contact surface and the housing is closed again, a method for modifying a gas turbine machine comprising at least a compressor, a combustion chamber with combustion chamber bricks and burners, a hot gas duct with a rotor, wherein modified guide vanes are installed in the swirler of the burner, which have a smaller opening angle compared to the previous guide vanes, and a trailing edge is twisted with respect to the longitudinal axis of the blade, a method for modifying a gas turbine machine comprising at least a compressor, a combustion chamber, a hot gas duct with a bladed rotor, wherein a system for monitoring the combustion dynamics and combustion accelerations is subsequently installed by the combustion chamber and burner,To reduce or avoid combustion instabilities, a method for modifying a steam turbine directly connected to a gas turbine, wherein the steam turbine has turbine blades, the turbine blade having a root with recesses, the recesses having a larger radius than the previously installed and replaced turbine blades. The following machine types are preferably achieved thereby: a gas turbine machine, comprising at least: a compressor, a combustion chamber, a hot gas duct with a bladed rotor, and at least one rotor bearing of the rotor at the beginning of the compressor in a flow direction of the gas turbine machine, wherein the rotor bearing is at least 370 mm long, in particular at most 500 mm long; a gas turbine machine, comprising at least: a compressor, a combustion chamber, a hot gas duct with a bladed rotor, and at least one burner for the combustion chamber.a gas turbine engine, in which the fuel supply means, in particular pipes, have at least partially, in particular completely, an internal diffusion coating, in particular are alitized, a gas turbine engine, comprising at least: a compressor, a combustion chamber, a hot gas duct with a bladed rotor, wherein the compressor has a compressor housing that is designed in two parts and has an inner compressor housing as a blade support, which has a first material, in particular steel, in particular cast steel, and an outer compressor housing as a blade support has a second material that is clearly different from the first material, in particular gray cast iron, or in which the inner compressor housing as a blade support has gray cast iron or is replaced by gray cast iron, a gas turbine engine, comprising at least: a compressor, a combustion chamber, a hot gas duct with a bladed rotor, with stages (I, II, III,IV) of rotor and guide vanes, wherein the turbine rotor and guide vanes, in particular those of stages (I, II), are more resistant to high temperatures, in particular have improved cooling, or wherein the rotor or guide vanes in the hot gas channel have a directionally solidified microstructure in the form of a columnarly solidified microstructure, in particular only the first two stages (I, II), most particularly only the first stage (I), or wherein a segmented ceramic layer based on yttrium-stabilized zirconia is present or applied to the rotor and guide vanes, or wherein the rotor and guide vanes for the metallic substrate have a single-crystal microstructure, or such are incorporated, in particular only the first two stages (I, II), or wherein the ceramic coating comprises partially stabilized yttrium-stabilized zirconia with a porosity of 12±4%,or with a TBC without segmentation on the guide or rotor blades, or in which a blade tip in a recess has a stepped shoulder which directly adjoins a web of the suction side and thus represents additional material in the recess, wherein a cooling air bore from the interior of the rotor blade extends through the shoulder to better cool the blade tip, gas turbine machine, at least comprising: a compressor, a combustion chamber, a hot gas duct with a bladed rotor, with stages (I, II, III, IV) of rotor and guide blades, wherein the rotor and guide blades, in particular of stages (I, II), have cooling holes on the side surfaces of the blade platforms, gas turbine machine, at least comprising: a compressor, a combustion chamber, a hot gas duct with a bladed rotor, with stages (I, II, III, IV) of rotor and guide blades, wherein the blade tip, in particular of stages (I, II), is cooled,in particular by cooling holes in the blade tip, gas turbine machine, at least comprising a compressor, a combustion chamber, a hot gas duct with a bladed rotor, with stages (I, II, III, IV) of rotor and stator blades, wherein the rotor blade of stage (IV) does not need to be cooled, in particular the stator blade of stage (III) is not cooled, gas turbine machine, at least comprising: a compressor, a combustion chamber, a hot gas duct with a bladed rotor, in which, viewed in the direction of flow, a front plenum and a rear plenum are located outside the hot gas duct, which, for technical reasons, have different pressures, wherein the front plenum is located in the direction of flow behind the rotor blade of stage (III) and above the stator blade of stage (IV), and a long channel from the rear plenum is located in the blade carrier, which cools stage (III) from the rear plenum, or gas turbine machine,at least comprising: a compressor, a combustion chamber with combustion chamber bricks or heat shields, a hot gas duct with a bladed rotor, a gap between a heat shield and the guide vane of stage (I) of the rotor, wherein a rounding at the downstream end of the heat shield and the opposite rounding of the guide vane of stage (I) are identical in order to avoid an overhang or undercut in the heat shield in which dirt could accumulate or erosion could occur; a gas turbine machine, at least comprising: a compressor, a combustion chamber, a hot gas duct with a bladed rotor, a guide vane carrier with a sealing arrangement of elements which results in lower cooling air consumption, wherein the individual elements of the guide vane carrier have a gap which is labyrinthine or S-shaped;wherein the element leading in the direction of flow has a first nose and the second element trailing in the direction of flow has a second nose formed above it, so that an S-shaped gap is formed, whereby the opening of the gap in the hot gas duct is located aft in the direction of flow, gas turbine machine, comprising at least: a compressor, a combustion chamber with heat shields or combustion chamber bricks, a hot gas duct with a bladed rotor, wherein the combustion chamber at the downstream end of the heat shields or combustion chamber bricks is configured to generate a spoiler effect, gas turbine machine, comprising at least: a compressor, a combustion chamber with combustion chamber bricks, a hot gas duct with a bladed rotor, wherein two separate recesses are formed on two opposite side faces of the combustion chamber brick,a gas turbine machine, comprising at least: a compressor, a combustion chamber with combustion chamber bricks, a hot gas duct with a bladed rotor, an upper and lower housing part for the hot gas duct, wherein housing parts have a recess with a seal on the contact surface, particularly in the area of ​​the guide vane recesses; a gas turbine machine, comprising at least: a compressor, a combustion chamber with combustion chamber bricks, a hot gas duct with a bladed rotor, wherein a burner is arranged in the combustion chamber, which has a swirler with guide vanes, wherein the opening angle of the guide vane is reduced and the blade of the guide vane is along the trailing edge of the blade; a gas turbine machine, comprising at least: a compressor, a combustion chamber, a hot gas duct with a bladed rotor, wherein a system is installedthat monitors the combustion dynamics and accelerations originating from the combustion and the combustion chamber, gas turbine engine in a combined cycle power plant, comprising at least a compressor, a combustion chamber, a hot gas channel, with a bladed rotor and a steam turbine, wherein an exhaust gas from the gas turbine engine is indirectly used for steam generation in a steam turbine, wherein the steam turbine has blades with a blade root, the recesses having a larger radius.

[0072] These measures or models are explained in more detail below using drawings and classified with regard to their intention. life

[0073] The Figure 5 shows how Figure 1 a gas turbine machine 100 with compressor 105 and rotor 103 in cross-section.

[0074] In the intake housing 104, air 135 is pumped into the compressor 105, which has a compressor housing 19.

[0075] For a longer service life of the gas turbine machine 100, a rotor bearing 31 of the rotor 103 in the flow direction 11 of the gas turbine machine 100 at the beginning of the compressor 105 and near the intake housing 104 has a length of at least 370mm and in particular a maximum length of 500mm or is designed to be at least 5% longer in the case of upgrade or overhaul in order to achieve a lower surface pressure.

[0076] In the event of a service call or at one of the next service intervals, the built-in bearing does not need to be replaced, or it can be used until the end of the gas turbine machine's lifespan.

[0077] It can also be replaced as part of a major service operation if replacing a bearing requires only a small amount of time and, in particular, can be done in parallel or is even facilitated by doing so. The same applies to a bearing of the rotor 103 in the area of ​​the exhaust housing 109 ( Fig. 24 ). life

[0078] In Figure 6 is the burner 107' of a gas turbine engine 100 starting from Figure 1 , 5 or 7 , 8 , 12 or 14 The corrosive properties of fuels, especially gas or oil that is burned, can vary locally.

[0079] The same applies if petroleum or other fuels are used.

[0080] The burner system with the burner 107 ( Fig. 1 ) is exposed to the highest temperatures.

[0081] To extend the service life, the fuel supply means such as pipes, especially those of the burner 107', particularly those for gas, are at least partially, especially completely, internally coated with a diffusion coating, especially alitized, i.e., internal alitization (or chromizing, ...) is used.

[0082] The internal coating can also be applied using the same methods while the part is installed.

[0083] This increases the service life due to reduced corrosion, but also the efficiency.

[0084] This allows the lifespan of the 107' burner to be individually adapted to the respective operating conditions.

[0085] The term combustion chambers 111 refers to well-known systems such as ring combustion chambers or CANs. Lifespan / Efficiency

[0086] Figure 7 shows a similar arrangement of a cross-section of a gas turbine machine 100 according to Figure 1 , 5 , 6, 8 , 12 , 14 or 24 , but with a now two-part compressor housing 19, which in the end area of ​​the compressor 105 has an inner compressor housing 19'' and an outer compressor housing 19'.

[0087] The materials of the compressor housings 19', 19'', especially if they are one-piece, are generally the same primary material, particularly gray cast iron. For an improved modification, the inner compressor housing 19'', serving as the guide vane support, is manufactured from a significantly different secondary material, particularly cast steel.

[0088] Different between the first and second material means that at least one alloying element differs by 10% in weight fraction and / or at least one other alloying element is present or less present and / or a different manufacturing process was used or a different, distinguishable microstructure is present. Lifespan / Efficiency

[0089] Figure 8 shows in particular the hot gas channel 111 with its stages I, II, III and especially also stage IV.

[0090] Stages I and II are subjected to higher thermal loads compared to stages III and IV. Appropriate modifications to the substrate material are employed here, particularly in the form of directionally solidified alloys (SX, DS) or additional or improved cooling, especially of the blade tip 415.

[0091] Such a shovel 120, 130 preferably has a directed solidification structure SX, DS in the form of a columnar solidification microstructure, such as alloys with the addition of DS in particular. Figure 4 .

[0092] Another type of shovel 120, 130 has a single-crystal microstructure in the substrate, similar to an alloy in Figure 4with the suffix SX or CMSX .... In particular, only the first stage I has a DS structure, and most specifically only the guide vane of stage I.

[0093] In particular, the blades have 120,130 cooling holes on the side surfaces of the blade platform 403, whereby the blade tips 415 are also cooled.

[0094] A ceramic coating (TBC) based on partially stabilized YSZ (yttrium-stabilized zirconia) has a porosity of 12±4%.

[0095] Another type of shovel 120, 130 features a segmented TBC based on yttrium-stabilized zirconia.

[0096] Another type of blade 120, 130 consists of a directed solidification structure DS in the substrate, i.e. in the form of a columnar microstructure and with a TBC based on YSZ without segmentation. life

[0097] Figure 9shows a turbine blade 120, 130, especially starting from Figure 3 , in which, however, cooling holes 399 are provided on the side surfaces 404 of the blade platform 403. The cooling holes 399 on the side surfaces 404 can be present on one, two, three, or all four side surfaces 404, as required, either singly or in multiples. Optionally, cooling holes 405 can also be provided on the blade tip 415 (shown schematically only). Cooling air holes 418 are also provided on the blade 406 in a known manner.

[0098] The orientation and arrangement of the cooling air holes 399, 405, 418 are only schematic. Likewise, the cooling holes 399, 405, 418, 501 ( Fig. 10 ) run at an angle other than 90° to the side surface 404 of the bucket platform 403 and / or have a diffuser.

[0099] Cooling by means of cooling air taken from the compressor reduces efficiency, which may be partially compensated for by the cooling effect. life

[0100] Figure 10 shows a blade tip 415, 500 of a turbine running blade 120, in particular stages I, II.

[0101] The shovel tip 500 has two externally extending ribs 503, 505 which, in cross-section, enclose a recess 504. The original recess 504 is indicated by a dashed line and is rectangular in cross-section.

[0102] According to the invention, the blade tip 500 has a stepped shoulder 507 in the recess 504, which connects directly to the rib 505 of the suction side and thus initially represents additional material in the recess 504. However, a cooling air hole 501 now extends through the shoulder 507 from the interior of the rotor blade 120 to improve cooling of the blade tip 500.

[0103] The cooling air hole 501 is preferably aligned with the longitudinal axis 121 of the turbine blade 120. Efficiency

[0104] The Figure 11 shows a blade carrier 50 with modified cooling air supply.

[0105] Stage IV is attached in the area of ​​this bucket carrier 50.

[0106] Viewed in the direction of flow 11, there is a front plenum 54 and a rear plenum 57 outside the hot gas channel 111, the 54 and 57 having different pressures due to technical reasons. Viewed in the direction of flow 11, the front plenum 54 is preferably located behind the rotor blade of stage III and in the area above the guide vane IV.

[0107] There are two channels 53 in the bucket carrier 50 in the Figure 11The two channels 53, which were used in older models to cool the Stage IV turbine guide vanes and rotor blades, or more generally to supply cooling air, were present in the Stage IV section between the guide vane 401 and the Stage V rotor blade 402. They ran almost perpendicular to, or at an acute angle to, the axis of rotation 102. These two channels 53 are closed, if present, and a new, long channel 60 is subsequently added, extending from the rear plenum 57 into the blade carrier 50. Alternatively, a new blade carrier 50 is provided that has only one such channel 60. The channel 60 runs approximately parallel to the inner surface of the hot gas channel 111.

[0108] Due to the lower pressure, less cooling air is supplied to the guide vane 401. The inlet of the new channel 60 is located downstream of the rotor blade 402 in the axial flow direction through the gas turbine, and not between the guide vane and the rotor blade of stage IV. Less cooling air is consumed, resulting in higher efficiency. Lifespan / Efficiency

[0109] Figure 12 The figure shows the transition from a final heat shield 155 or combustion chamber brick 155 of the combustion chamber 110 to a guide vane 130 of stage I. It can be seen that there is a gap 64 between the heat shield 155 and the guide vane 130.

[0110] The rounded section 72 at the flow-side end of the heat shield / combustion chamber brick 155 and the opposite rounded section 75 of the guide vane 130 of stage I are identical. An overhang or undercut in the heat shield / combustion chamber brick 155, where dirt could accumulate or erosion could occur, is to be avoided. Efficiency

[0111] Figure 13 shows a sealing arrangement 79 of a guide vane carrier 50 ( Fig. 11 ), which leads to lower cooling air consumption.

[0112] The individual elements 81, 83 of the guide vane carrier 79 have a gap 80, which here is labyrinthine or S-shaped. The reduced cooling air consumption is achieved by the fact that the element 81, which is forward in the direction of flow 11, has a first nose 82, and the second element 83, which is aft in the direction of flow 11, has a second nose 85 formed above it, so that an S-shaped gap 80 is formed and the nose 82 of the forward element 81 forms part of the hot gas channel 111.

[0113] Thus, the opening of the gap in the hot gas channel 111 is located at the rear in the direction of flow. Efficiency

[0114] Figure 14 shows a combustion chamber 110 with combustion chamber bricks 601, 604, 610, which lead to a spoiler effect.

[0115] The combustion chamber bricks 601, 602 and 603, 605 are arranged in a row and circumferentially when viewed in the direction of flow 11.

[0116] Viewed in the direction of flow 11, modified combustion chamber bricks 603, 605; 604 are present at the end, with further such combustion chamber bricks arranged in the circumferential direction around the axis of rotation 102.

[0117] The modified combustion chamber bricks 603, 604, 605, preferably made of solid ceramic, preferably located in the last two rows of the combustion chamber 110 (viewed in the flow direction 11) before the inlet or transition to the guide vanes 130 or the first rotor blade row 120 of stage I, are designed to be thicker in the flow direction 11, resulting in a spoiler effect. Near the rotor hub, preferably only one combustion chamber brick 604 is spoiler-shaped, whereas further radially away at the outer end of the combustion chamber 110, at least the penultimate and last rows of the combustion chamber bricks 603, 605 together exhibit a gradual increase in thickness in the flow direction.

[0118] However, this should not be a limitation.

[0119] This spoiler effect leads to reduced erosion and also to a narrowing of the hot gas flow, which also increases efficiency. Lifespan / Efficiency

[0120] Figure 15 shows a combustion chamber brick 155 with a side surface 35, whereas Figures 16, 17 Sectional views of the Figure 15 are.

[0121] Figure 15 Figure 1 shows the side surface 35 of a combustion chamber brick 155, as used in a combustion chamber 110, wherein a side surface 35 has two elongated recesses 40, 40' in which a holder engages from the rear and a corresponding through opening 42 to the rear 43 of the combustion chamber brick 155, which is opposite the top 44.

[0122] On the back side 43 of the combustion chamber brick 155, an undercut 41 is present along the side surface with the recesses 40.

[0123] In Figure 16is a cross-section through two combustion chamber bricks placed next to each other according to Figure 15 along the continuous opening 42', 42'' shown, whereas in Figure 17 a gap between two combustion chamber bricks 155 according to Figure 15 shown outside the opening 42 or elongated recess 40.

[0124] This leads to reduced coolant consumption because the cooling effort is lower. Lifespan / Efficiency

[0125] Figure 18 shows a lower turbine housing part 550 with the guide vane recesses 553 and channel 557 in the contact surface 600 for the other, upper housing half.

[0126] In this contact surface 600, one or more additional recesses 630 and a seal are incorporated to reduce leakage in this area ( Fig. 19 ). Efficiency

[0127] Figure 20shows a burner 70 which has a swirler in which air and fuel are mixed together.

[0128] Two different guide vanes, 73' and 73'', of the swirler are in the Figure 21 shown and in a different, modified shallower angle (73') from the first position to a second position and / or a torsion of the guide vane 73'' along a trailing edge to achieve better turbulence.

[0129] The dashed line on guide vane 73' indicates the original position of the swirler's guide vane, whereas the dashed line on guide vane 73'' shows how it is twisted along its longitudinal axis, which runs parallel to the leading edge. Lifespan / Efficiency

[0130] Gas turbines can be operated alone to drive a generator, but are often also used in combination with steam turbines in a combined cycle power plant.

[0131] Due to the higher power output generated by the gas turbine, the performance of a steam turbine 6 must also be adjusted. This is achieved in particular by, as described in Figure 22 It is shown that a blade root of a turbine running blade 883 receives a larger radius in the three recesses 886', 886'', 886‴ of the fir tree base 880 than before. life

[0132] Combustion stability and dynamics also have a very high influence on the system's lifespan, so a control system 90 is installed here which registers the combustion dynamics and acceleration ( Fig 23 ). Service measures

[0133] Some of the service measures can be carried out together because they are easier to perform together and can be carried out in parallel if necessary.

[0134] Some measures can be offered and implemented free of charge to the operator of the plant to extend or even skip the next service interval, or to reduce operating costs through increased efficiency, in which the service provider participates.

[0135] Considerations regarding extended lifespan, efficiency, higher temperatures and coolant consumption play a role.

[0136] When servicing an energy generation plant 1, which has at least one gas turbine 100, one generator 5 and optionally one steam turbine 6 with corresponding auxiliary units, service contracts are often concluded, which offer the operator of the energy conversion plant a service maintenance contract that includes a guarantee of a certain operating time (lifespan) with certain performance characteristics and specified service intervals.

[0137] This means that the service provider not only repairs (refurbishes) what needs to be repaired because it can no longer be used or only for a short time, but also considers measures to be taken during a service that allow the period until the next service interval or until the service interval after that to be extended.

[0138] For example, when replacing turbine blades, new burners with an aluminized coating can also be installed, eliminating the need for servicing the burner components at the next or subsequent service interval. This avoids downtime.

[0139] Similarly, a service measure can be used to extend the service life or to bring forward a service measure, thus significantly shortening the next service measure, because measures such as bearing removal, turbine blade removal, replacement of combustion chamber bricks, changes to the housing, etc. involve different maintenance times.

[0140] This also results from the fact that certain measures can be carried out in parallel.

[0141] The flexible refurbishment includes enhanced remote monitoring and diagnostic capabilities as part of the Omnivise Digital Services portfolio, as well as spare parts deliveries, scheduled maintenance, and performance guarantees throughout the plant's operating life. With its high efficiency at part load and its high operational flexibility, the combined cycle gas turbine power plant, along with its associated services, will complement the fluctuating renewable energy sources in the region as part of this flexible refurbishment.

[0142] Protecting the investments of power plant operators through first-class service for rotating machinery is the key to the philosophy of a successful service provider.

[0143] The long-term, flexible service goes a step further and offers a maintenance program that is individually tailored to specific needs and requirements.

[0144] Whether the power plant operator wants to maximize production by extending the time between inspections, whether inspections should take place during the plant's scheduled downtime, whether life cycle costs should be optimized through condition-based component replacement, or whether work should be free from predetermined inspection dates: Long-term flexible refurbishment can adapt to needs.

Claims

1. A method for carrying out service work on an energy conversion plant (1), wherein the energy conversion plant (1) comprises at least the following machines (5, 6, 100, ...): at least one gas turbine (100), at least one generator (5) and optionally at least one steam turbine (6), wherein repairs are carried out on the at least one machine (5, 6, 100, ...), in particular a defective component or components of the at least one machine (5, 6, 100, ...) are replaced by either a new, identical component or components and / or are repaired, and wherein, in carrying out these repairs, further measures for extending the service life of machines (5, 6, 100, ...) are carried out.) or their components, wherein a blade tip (415, 500) of a running blade (120) has a step-shaped shoulder (507) in a recess (504) which directly connects to a rib (505) of the suction side and thus represents additional material in the recess (504), wherein a cooling air bore (501) extends through the shoulder (507) from the interior of the running blade (120) to better cool the blade tip (500) and / or further measures for optimizing machines (5, 6, 100, ...) or their components are carried out.

2. Method according to claim 1, wherein the defective component or components comprise turbine blades (120, 130) and / or their coatings and / or burners (107) or burner components and / or compressor blades and / or their coatings and / or combustion chamber bricks (155).

3. Method according to one or both of claims 1 or 2, wherein the defective components comprise only turbine blades (120, 130).

4. Method according to one or both of claims 1 or 2, wherein the defective components comprise only turbine blades (120, 130) and / or their coatings and burners (107) or burner components.

5. Method according to one or more of claims 1, 2, 3 or 4, wherein the only further measures taken are measures to extend the service life of machines (5, 6, 100, ...) or their components.

6. Method according to one or more of claims 1, 2, 3 or 4, wherein the only further measures taken are measures to optimize machines (5, 6, 100, ...) or their components.

7. Method according to one or more of claims 1, 2, 3 or 4, wherein further measures are taken to extend the service life and to optimize machines (5, 6, 100, ...) or components (120, 130, ...).

8. Method according to one or more of claims 1, 2, 3, 4, 5 or 7, wherein the service measures include at least one, in particular at least two identical, or at least two different measures for extending the service life of machines (5, 6, 100, ...) or their components from the group: rotor bearings (103), burner (107), compressor blade, compressor housing (19), turbine blades (120, 130), gas turbine housing (138), blade carrier (50), heat shields (155) or combustion chamber bricks, seals, combustion chamber - turbine transition, cooling and / or monitoring devices (90).

9. Method according to one or more of claims 1, 2, 3, 4, 6 or 8, wherein the service measures include at least one, in particular at least two identical or at least two different measures for optimizing machines (5, 6, 100, ...) or their components from the group: efficiency improvement, cooling improvement, burner (107), compressor blade, compressor housing (19), turbine blades (120, 130), gas turbine housing (138), blade carrier (50), heat shields (155) or combustion chamber bricks, seals and / or transition combustion chamber - turbine.

10. A method according to one or more of claims 1, 2, 3, 4, 5, 6, 7, 8 or 9, wherein the gas turbine machine (100) is modified, the gas turbine machine (100) comprising at least: a compressor (105), a combustion chamber (110), a hot gas duct (111) with a bladed rotor (103), and at least one rotor bearing of the rotor (103) at the beginning of the compressor (105) as seen in a flow direction (11) of the gas turbine machine (100), wherein the rotor bearing is replaced, wherein the new rotor bearing (31) is at least 5% longer or a new rotor bearing (31) at least 370 mm long is installed, in particular wherein the new rotor bearing (31) is a maximum of 500 mm long, and / or wherein the gas turbine machine (100) comprises at least: a compressor (105), a Combustion chamber (110), a hot gas channel (111) with a bladed rotor (103), at least one burner (107') for the combustion chamber (110), in which the fuel supply means, in particular pipes,at least partially, in particular completely, internally provided with a diffusion coating, in particular alitized, or the fuel supply means, in particular pipes with a diffusion coating on the inside, in particular internally alitized, are installed, and / or wherein the gas turbine machine (100) comprises at least: a compressor (105), a combustion chamber (110), a hot gas duct (111) with a bladed rotor (103), wherein the compressor (105) comprises a compressor housing (19) which is or is formed in two parts (19', 19''), and in particular wherein an inner compressor housing (19'') is installed as a blade support made of a first material, in particular steel, most especially cast steel, and the outer compressor housing (19') as a blade support made of a second material, in particular gray cast iron, which is clearly different from the first material,and / or in which the inner compressor housing (19'') is replaced by gray cast iron, and / or in which the gas turbine machine (100) comprises at least: a compressor (105), a combustion chamber (110), a hot gas duct (111) with a bladed rotor (103), with stages (I, II, III, IV) of rotor and guide vanes, wherein more temperature-resistant rotor (120) and guide vanes (130), in particular in stages (I, II), are installed, which in particular have improved cooling, and / or in which rotor (120) or guide vanes (130) are installed in the hot gas duct (111) which have a directionally solidified microstructure in the form of a columnarly solidified microstructure, in particular only the first two stages (I, II), most in particular only the first stage (I), and / or in which a segmented ceramic layer based on yttrium-stabilized zirconia is applied to the rotor- (120) and guide vanes (130) are applied,and / or in which the rotor (120) and guide vanes (130) are installed, which have a single-crystal microstructure for the metallic substrate, in particular only the first two stages (I, II), and / or in which the ceramic coating comprises partially stabilized yttrium-stabilized zirconia with a porosity of 12 ± 4%, and / or in which a TBC is present in the hot gas duct (111) of the rotor (120) or guide vanes (130) without segmentation on the guide vanes (130) or rotor blades (120), and / or in which the gas turbine machine (100) comprises at least: a compressor (105), a combustion chamber (110), a hot gas duct (111) with a bladed rotor (103), with stages (I, II, III, IV) of rotor and guide vanes, in which the rotor (120) and guide vanes (130) are installed will, in particular in stages (I, II), have cooling holes (399) on the side surfaces (404) of the bucket platforms (403),and / or wherein the gas turbine machine (100) comprises at least: a compressor (105), a combustion chamber (110), a hot gas duct (111) with a bladed rotor (103), with stages (I, II, III, IV) of rotor and guide vanes, wherein rotor blades are installed, wherein the blade tip (415) of the rotor blades, in particular of stages (I, II), is cooled, in particular by cooling holes (501) in the blade tip (415), and / or wherein the gas turbine machine (100) comprises at least: a compressor (105), a combustion chamber (110), a hot gas duct (111) with a bladed rotor (103), with stages (I, II, III, IV) of rotor and guide vanes, wherein rotor blades (402) of stage (IV) are installed which are not cooled, in particular the guide vanes (401) of stage (III) are also not cooled become, and / or wherein the gas turbine machine (100) at least comprises: a compressor (105), a combustion chamber (110),a hot gas duct (111) with a bladed rotor (103), in which, viewed in the direction of flow (11), a front plenum (54) and a rear plenum (57) are located outside the hot gas duct (111), the (54, 57) having different pressures due to technical reasons, the front plenum (54) being located in the direction of flow (11) behind the rotor blade of stage (III) and above the guide vane of stage (IV), and either ducts (53) that were located in the middle of stage (IV) between the guide vane (401) and rotor blade (402) and were previously used to cool the turbine guide vanes and rotor blades of stage (IV) or to supply cooling air are closed off, and a new long duct (60) from the rear plenum (57) is subsequently introduced into the blade carrier (50), or a new blade carrier (50) is provided and installed which has only one such duct (60) exhibitsand / or wherein the gas turbine machine (100) comprises at least: a compressor (105), a combustion chamber (110) with combustion chamber bricks or heat shields (155), a hot gas duct (111) with a bladed rotor (103), wherein a gap (64) between a heat shield (155) and guide vane (130) of stage (I) of the engine (103) is formed with a rounding (72) at the downstream end of the heat shield (155) and the opposite rounding (75) of the guide vane (130) of stage (I) being identical in order to avoid an overhang or undercut in the heat shield (155) in which dirt could accumulate or erosion could occur, and / or wherein the gas turbine machine (100) comprises at least: a compressor (105), a combustion chamber (110), a hot gas duct (111) with a bladed rotor (103) wherein a guide vane carrier (79) is fitted or modified with a sealing arrangement (79) consisting of elements (81, 83) which results in a lower cooling air consumption,wherein the individual elements (81, 83) of the guide vane carrier (79) have a gap (80) which is labyrinthine or S-shaped, wherein the element leading in the direction of flow (11) has a first nose (82) and the second element (83) trailing in the direction of flow (11) has a second nose (85) formed above it, so that an S-shaped gap (80) is formed, whereby the opening of the gap in the hot gas duct (111) is located aft in the direction of flow, and / or wherein the gas turbine machine (100) comprises at least: a compressor (105), a combustion chamber (110) with heat shields (155) or combustion chamber bricks, a hot gas duct (111) with a bladed rotor (103), wherein combustion chamber bricks (601, 604, 610) are installed which are designed to generate a spoiler effect, and / or wherein the The gas turbine machine (100) at least comprises: a compressor (105), a combustion chamber (110) with combustion chamber bricks,a hot gas duct (111) with a rotor (103), wherein combustion chamber bricks are installed which accommodate two separate recesses (40', 40'') on two opposite side faces of the combustion chamber brick, which serve for engagement by a mechanical clamping from the rear (43) of the combustion chamber brick, and / or wherein the gas turbine machine (100) comprises at least: a compressor (105), a combustion chamber (110) with combustion chamber bricks, a hot gas duct (111) with a bladed rotor (103), a housing part (550) for the hot gas duct (111), wherein a recess (630) for a seal is inserted into the contact surface (600) and the housing is closed again, and / or wherein the gas turbine machine (100) comprises at least: a compressor (105), a combustion chamber with combustion chamber bricks and burners, a hot gas duct (111) with a rotor (103), wherein modified Guide vanes (73', 73'') are installed in the swirler of the burner (70),which have a smaller opening angle compared to the previous guide vanes, and a trailing edge is twisted with respect to the longitudinal axis of the blade, and / or wherein the gas turbine machine (100) comprises at least: a compressor (105), a combustion chamber (110), a hot gas duct (111) with a bladed rotor (103), wherein a system (90) for monitoring the combustion dynamics and combustion accelerations is subsequently installed from the combustion chamber and burner in order to reduce or avoid combustion instabilities, and / or wherein the steam turbine (6) is modified, which is directly or indirectly connected to a gas turbine, wherein the steam turbine comprises turbine blades (883), wherein the turbine blade (883) has a foot (880) with recesses, wherein the recesses (886', 886'', 886‴) have a larger radius compared to the previously installed and to be replaced turbine blades.

11. Energy generation plant, according to a method according to one or more of claims 1 to 10, or an energy conversion plant (1), comprising at least: a gas turbine machine (100), comprising at least: a compressor (105), a combustion chamber (110), a hot gas duct (111) with a bladed rotor (103), and at least one rotor bearing (31) of the rotor (103) at the beginning of the compressor (105) in a flow direction (11) of the gas turbine machine (100), wherein the rotor bearing (31) is at least 370 mm long, in particular at most 500 mm long, and / or wherein the gas turbine machine (100) comprises at least: a compressor (105), a combustion chamber (110), a hot gas duct (111) with a bladed rotor (103), at least one burner (107') for the combustion chamber (110), wherein the Fuel supply means, in particular pipes, have at least partially, in particular completely, an internal diffusion coating, in particular are alitized,and / or in which the gas turbine machine (100) comprises at least: a compressor (105), a combustion chamber (110), a hot gas duct (111) with a bladed rotor (103), wherein the compressor (105) comprises a compressor housing (19) which is formed in two parts (19', 19'') and has an inner compressor housing (19'') as a blade carrier, which has a first material, in particular steel, most especially cast steel, and an outer compressor housing (19') as a blade carrier, which has a second material that is clearly different from the first material, in particular gray cast iron, or in which the outer compressor housing (19'') has gray cast iron, and / or in which the gas turbine machine (100) comprises at least: a compressor (105), a combustion chamber (110), a hot gas duct (111) with a bladed rotor (103), with stages (I, II, III, IV) of running and Guide vanes, wherein the turbine rotor (120) and guide vanes (130),in particular the stages (I, II), are more resistant to high temperatures, in particular have improved cooling, and / or in which the impeller (120) or guide vanes (130) in the hot gas channel (111) have a directionally solidified microstructure in the form of a columnarly solidified microstructure, in particular only the first two stages (I, II), most especially only the first stage (I), and / or in which a segmented ceramic layer based on yttrium-stabilized zirconia is present on the impeller (120) and guide vanes (130), and / or in which the impeller (120) and guide vanes (130) for the metallic substrate have a single-crystal microstructure, in particular only the first two stages (I, II), and / or in which the ceramic coating comprises partially stabilized yttrium-stabilized zirconia with a porosity of 12±4%, and / or with a TBC without segmentation on the guide vanes. (130) or running vanes (120),and / or in which a blade tip (415, 500) in a recess (504) has a stepped shoulder (507) which connects directly to a web (505) of the suction side and thus represents additional material in the recess (504), wherein a cooling air bore (501) extends from the interior of the rotor blade (120) through the shoulder (507) to better cool the blade tip (500), and / or in which the gas turbine machine (100) comprises at least: a compressor (105), a combustion chamber (110), a hot gas duct (111) with a bladed rotor (103), with stages (I, II, III, IV) of rotor and guide vanes, wherein the rotor (120) and guide vanes (130), in particular those of stages (I, II), have cooling holes (399) on the side surfaces (404) of the blade platforms (403) and / or in which the gas turbine machine (100) at least comprises: a compressor (105), a combustion chamber (110), a hot gas duct (111) with a bladed rotor (103), with stages (I,II, III, IV) of rotor and guide vanes, wherein the blade tip (415, 500) of stages (I, II) is cooled, in particular by cooling holes (501) in the blade tip (415), and / or wherein the gas turbine machine (100) comprises at least: a compressor (105), a combustion chamber (110), a hot gas duct (111) with a bladed rotor (103), with stages (I, II, III, IV) of rotor and guide vanes, wherein the rotor blade (402) of stage (IV) does not need to be cooled, in particular the guide vane (401) of stage (III) is not cooled, and / or wherein the gas turbine machine (100) comprises at least: a compressor (105), a combustion chamber (110), a hot gas duct (111) with a bladed rotor (103), wherein, viewed in the direction of flow (11), a front Plenum (54) and a rear plenum (57) outside the hot gas channel (111) are present, the (54, 57) having different pressures due to technical reasons,wherein the front plenum (54) is located in the flow direction (11) behind the rotor blade of stage (III) and above the guide vane of stage (IV), and a long channel (60) from the rear plenum (57) is located in the blade carrier (50), which cools stage (III) from the rear plenum (57), and / or wherein the gas turbine machine (100) comprises at least: a compressor (105), a combustion chamber (110) with combustion chamber bricks or heat shields (155), a hot gas channel (111) with a bladed rotor (103) comprising a gap (64) between a heat shield (155) and guide vane (130) of stage (I) of the rotor (103), wherein a rounding (72) at the flow-side end of the heat shield (155) and the opposite rounding (75) of the guide vane (130) of stage (I) are equal are designed to avoid an overhang or undercut in the heat shield (155) where dirt could accumulate or erosion could occur,and / or in which the gas turbine machine (100) comprises at least: a compressor (105), a combustion chamber (110), a hot gas duct (111) with a bladed rotor (103), a guide vane carrier (79) with a sealing arrangement (79) made of elements (81, 83) which leads to a lower cooling air consumption, wherein the individual elements (81, 83) of the guide vane carrier (79) have a gap (80) which is labyrinthine or S-shaped, wherein the element leading forward (11) in the direction of flow has a first nose (82) and the second element (83) trailing in the direction of flow (11) has a second nose (85) formed above it, so that an S-shaped gap (80) is formed, whereby the opening of the gap in the hot gas duct (111) is located aft in the direction of flow, and / or in which the gas turbine machine (100) comprises at least: a compressor (105), a combustion chamber (110) with heat shields (155) or combustion chamber bricks,a hot gas duct (111) with a bladed rotor (103), wherein the combustion chamber (110) at the flow-side end of the heat shields (155) or combustion chamber bricks (601, 604, 610) are designed to generate a spoiler effect, and / or wherein the gas turbine machine (100) comprises at least: a compressor (105), a combustion chamber (110) with combustion chamber bricks, a hot gas duct (111) with a bladed rotor (103), wherein two separate recesses (40', 40'') are formed on two opposite side faces of the combustion chamber brick, which serve to engage a mechanical clamping from the rear (43) of the combustion chamber brick, and / or wherein the gas turbine machine (100) comprises at least: a compressor (105), a combustion chamber (110) with combustion chamber bricks, a hot gas duct (111) with a bladed rotor (103), an upper and lower housing part (550) for the hot gas channel (111),wherein housing parts (550) have a recess (630) with a seal on the contact surface (600), particularly in the area of ​​the guide vane recesses, and / or wherein the gas turbine machine (100) comprises at least: a compressor (105), a combustion chamber (110) with combustion chamber bricks, a hot gas duct (111) with a bladed rotor (103), wherein a burner (70) is arranged in the combustion chamber (110) which has a swirler with guide vanes (73', 73''), wherein the opening angle of the guide vane is reduced and the blade of the guide vane (73'') is along the trailing edge of the blade, and / or wherein the gas turbine machine (100) in a combined cycle power plant (1) comprises at least: a compressor (105), a combustion chamber (110), a hot gas duct (111) with a bladed rotor (103) and a steam turbine (6) wherein an exhaust gas from the gas turbine engine (100) is indirectly used for steam generation in a steam turbine,wherein the steam turbine (6) has blades (883) with a blade root, wherein the recesses (886', 886'', 886‴) have a larger radius, and / or wherein the gas turbine machine (100) has at least: a compressor (105), a combustion chamber (110), a hot gas duct (111) with a bladed rotor, wherein a system (90) is installed that monitors the combustion dynamics and accelerations originating from the combustion and the combustion chamber (110).

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