Method for carrying out method for implementing energy conversion installation service measures, 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 to an energy conversion plant.

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

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

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

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

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

[0007] The subclaims list further advantageous measures which can be combined with each other as desired to achieve further advantages.

[0008] The Figure 1 a gas turbine, Figure 2 a combustion chamber, Figure 3 a turbine 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 surfaces, 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 15, Figures 18, 19 a housing with the introduction of an additional seal, Figures 20, 21 a burner with modified guide vanes of the swirler, Figure 22 a modified base of a steam turbine, Figure 23 a device for monitoring the combustion dynamics, Figure 24 shows a power generation plant.

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

[0010] The Figure 1 shows an example of a gas turbine engine 100 in a longitudinal section.

[0011] The gas turbine engine 100 has inside a rotor 103 with turbine blade 120, which is rotatably mounted about a rotation axis 102 and is also referred to as turbine rotor.

[0012] Along the rotor 103, there follow an intake housing 104, a compressor 105, a combustion chamber 110, for example a toroidal one, in particular an annular combustion chamber, with several coaxially arranged burners 107, a turbine 108 and the exhaust housing 109.

[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 flow direction of a working medium 113, a row of guide vanes 115 in the hot gas duct 111 is followed by a row of rotor blades 125 formed from rotor blades 120.

[0016] The guide vanes 130 are attached to a gas turbine casing 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 ( Fig. 24 ) or a work machine (not shown).

[0018] During operation of the gas turbine 100, air 135 is drawn in by the compressor 105 through the intake casing 104 and compressed. The compressed air provided at the turbine end of the compressor 105 is fed to the burners 107 in a combustion chamber 110, where it is mixed with a fuel. The mixture is then combusted in the combustion chamber 110 to form the working medium 113. From there, the working medium 113 flows along the hot gas duct 111 past the guide vanes 130 and the rotor blades 120. At the rotor blades 120, the working medium 113 expands, transmitting momentum, so that the rotor blades 120 drive the rotor 103, which drives the driven machine coupled to it.

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

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

[0021] Likewise, substrates of the components can have a directed structure, i.e. they are single-crystalline (SX structure) or have only longitudinally directed grains (DS structure).

[0022] For example, iron-, nickel- or cobalt-based superalloys are used as materials for the components, in particular for the turbine blades 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] Likewise, the blades 120, 130 can have anti-corrosive coatings: MCrAlX; M is at least one element from the group consisting of cobalt (Co) and nickel (Ni); X is an active element and represents yttrium (Y) and / or tantalum (Ta) and / or at least one rare earth element, 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] A thermal insulation layer may also be present on the MCrAlX and may consist, for example, of ZrO 2 , Y 2 O 3 -ZrO 2 , ie 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 secured to a mounting ring 140 of the stator 143.

[0027] Figure 2 shows a combustion chamber 110 of a gas turbine. The combustion chamber 110 is designed, for example, as a so-called annular combustion chamber, in which a plurality of burners 107 arranged circumferentially around a rotation 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 an annular structure positioned around the rotation axis 102.

[0028] To achieve a comparatively high level of efficiency, the combustion chamber 110 is designed for a comparatively high working medium temperature of approximately 1273K to 1873K. To enable a comparatively long operating time 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 side facing the working medium, with an inner lining formed from heat shield elements 155. Each heat shield element 155 made of an alloy is equipped with a particularly heat-resistant protective layer (MCrAlX layer and / or ceramic coating) on ​​the working medium side or is made of high-temperature-resistant material (solid ceramic bricks).

[0029] These protective layers of the metallic heat shield elements 155 can be similar to those of turbine blades. For example, MCrAlX means: M is at least one element from the group consisting of iron (Fe), cobalt (Co), and nickel (Ni). X is an active element and represents 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.

[0030] On the MCrAlX there may be a ceramic thermal barrier layer, for example, consisting of ZrO 2 , Y 2 O 3 -ZrO 2 , ie it is not, partially or completely stabilized by yttrium oxide and / or erbium oxide, ytterbium oxide and / or hafnium oxide.

[0031] Many coating processes are conceivable, e.g., atmospheric plasma spraying (APS), LPPS, VPS, or CVD. The thermal barrier coating can contain porous grains with micro- or macrocracks for improved thermal shock resistance.

[0032] Refurbishment means that heat shield elements 155 may need to be freed of protective coatings after use (e.g., by sandblasting). This is followed by the removal of corrosion and / or oxidation layers or products. If necessary, cracks in the heat shield element 155 are also repaired. The heat shield elements 155 are then recoated and reused.

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

[0034] The Figure 3 shows a perspective view of a rotor blade 120 or guide vane 130 of a turbomachine extending 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 blade 120, 130 has, along its longitudinal axis, a fastening region 400, an adjacent blade platform 403, a blade airfoil 406, and a blade tip 415.

[0037] As a guide vane 130, it may have a further platform at its vane tip 415 (not shown).

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

[0039] The blade base 183, for example, is designed as a hammer head. Other designs such as a fir tree or dovetail base 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] In conventional blades 120, 130, for example, solid metallic materials, in particular superalloys, are used in all areas 400, 403, 406 of the blade 120, 130.

[0042] 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 from Figure 4 known.

[0043] The blade 120, 130 can be manufactured by a casting process, also by means of directional solidification, by a forging process, by a milling process or combinations thereof.

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

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

[0046] In this process, dendritic crystals are aligned along the heat flow and form either a columnar grain structure, meaning grains that extend the entire length of the workpiece and are commonly referred to as directionally solidified, or a monocrystalline structure, meaning the entire workpiece consists of a single crystal. In these processes, the transition to globulitic (polycrystalline) solidification must be avoided, since undirected growth inevitably leads to the formation of transverse and longitudinal grain boundaries, which negate the beneficial properties of the directionally solidified or monocrystalline component.

[0047] When we generally speak of directionally solidified structures, we mean both single crystals that have no grain boundaries or at most small-angle grain boundaries, and columnar crystal structures that have longitudinal grain boundaries but no transverse grain boundaries. These latter crystalline structures are also referred to as directionally solidified structures.

[0048] Such processes are known from US-PS 6,024,792 and EP 0 892 090 A1.

[0049] Likewise, the blades 120, 130 can have coatings against corrosion or oxidation: in particular MCrAlX; M is at least one element from the group consisting of cobalt (Co) or nickel (Ni); 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 95% of the theoretical density.

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

[0052] A thermal barrier layer may also be present on the MCrAlX, which is preferably the outermost layer, and consists for example of ZrO 2 , Y 2 O 3 -ZrO 2 , ie 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 coating covers the entire MCrAlX layer. Other coating processes are conceivable, e.g., atmospheric plasma spraying (APS), LPPS, VPS, or CVD. The thermal barrier coating can contain porous grains with micro- or macrocracks for improved thermal shock resistance. Therefore, the thermal barrier coating is preferably more porous than the MCrAlX layer.

[0054] Refurbishment means that components 120, 130 may need to be freed of protective coatings after use (e.g., by sandblasting). This is followed by the removal of corrosion and / or oxidation layers or products. If necessary, cracks in components 120, 130 are also repaired. The component 120, 130 is then recoated and reused.

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

[0056] Figure 24 shows an example of an energy conversion system 1 with a system. This arrangement according to Figure 24 can be present in an energy conversion plant in multiple forms or in modified forms.

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

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

[0059] Steam turbines 6 are present in a combined-cycle power plant. An energy conversion plant 1 may also have 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 through 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 measures on an energy conversion plant, wherein the energy conversion plant has 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 is or are replaced by either a new, identical component or new, identical components and / or is or are repaired, and wherein when carrying out these repairs, further measures for extending the service life of machines or their components and / or further measures for optimising (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 may 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 are preferably only those aimed at extending the service life.

[0067] Further measures are preferably only those aimed at optimising the situation.

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

[0069] In the service measures, preferably at least one, in particular at least two identical, or at least two different measures to extend the service life of machines or their components from the group: rotor bearings, burners, compressor blades, compressor housings, turbine blades, gas turbine housings, blade carriers, heat shields or combustion chamber bricks, seals, combustion chamber-turbine transition, cooling and / or monitoring devices.

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

[0071] The individual measures are described in more detail below and can be combined as required: A method for modifying a gas turbine engine, which at least comprises: 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, viewed in a flow direction of the gas turbine engine, wherein the rotor bearing is replaced, wherein the new rotor bearing is at least 5% longer or a new rotor bearing with a length of at least 370 mm is installed, in particular wherein the new rotor bearing is a maximum of 500 mm long. A method for modifying a gas turbine engine, which at least comprises: a compressor, a combustion chamber, a hot gas duct with a bladed rotor, at least one burner for the combustion chamber, in which the fuel supplying means, in particular tubes, are at least partially, in particular completely, provided internally with a diffusion coating, in particular are alitated, or the fuel supplying means,in particular pipes with a diffusion coating on the inside, in particular internally aluminized, are installed, Method for modifying a gas turbine engine, which at least has: a compressor, a combustion chamber, a hot gas duct with a bladed rotor, wherein the compressor has a compressor housing which is or will be formed in two parts, and in particular in which an inner compressor housing is installed as a blade carrier made of a first material, in particular steel, very particularly cast steel, and the outer compressor housing as a blade carrier has a second material which is significantly different from the first material, in particular grey cast iron, or in which the inner compressor housing has grey cast iron or is replaced by grey cast iron, Method for modifying a gas turbine engine, which at least has: a compressor, a combustion chamber, a hot gas duct with a bladed rotor, with stages (I, II, III,IV) of rotor blades and guide vanes, wherein more high-temperature-resistant rotor blades and guide vanes are installed, in particular in stages (I, II), which in particular have improved cooling, or in which the rotor blades or guide vanes in the hot gas channel (111) have a directionally solidified microstructure in the form of a columnar solidified microstructure, in particular only the first two stages (I, II), very particularly only the first stage (I), or in which a segmented ceramic layer based on yttrium-stabilized zirconium oxide (HGB) is present or applied to the rotor blades and guide vanes, or in which the rotor blades and guide vanes for the metallic substrate have a single-crystal microstructure, or such are installed, in particular only the first two stages (I, II), or in which the ceramic coating comprises partially stabilized yttrium-stabilized zirconium oxide, 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 step-shaped shoulder which directly adjoins a web of the suction side and thus represents additional material in the recess, wherein a cooling air bore runs through the shoulder from the interior of the rotor blade in order to better cool the blade tip, Method for modifying a gas turbine engine, which has 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, Method for modifying a gas turbine engine, which has at least: a compressor, a combustion chamber, a hot gas duct with a bladed rotor, with stages (I, II, III,IV) of rotor blades 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, Method for modifying a gas turbine engine, which has at least: a compressor, a combustion chamber, a hot gas duct with a bladed rotor, with stages (I, II, III, IV) of rotor blades 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, Method for modifying a gas turbine engine, at least comprising: a compressor, a combustion chamber, a hot gas duct with a bladed rotor, in which, viewed in the flow direction, there is a front plenum and a rear plenum outside the hot gas duct, which have different pressures for technical reasons,wherein the front plenum is located behind the rotor blade of stage (III) and above the guide vane of stage (IV) in the flow direction, and either channels that were present in the middle of stage (IV) between the guide vane and the 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, 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 that only has one such channel. Method for modifying a gas turbine engine, which at least has: a compressor, a combustion chamber with combustion chamber bricks or heat shields, a hot gas channel with a bladed rotor,wherein a gap between a heat shield and a guide vane of stage (I) of the engine is formed with a rounded portion at the upstream end of the heat shield, and the opposite rounded portion of the guide vane of stage (I) is formed with the same portion to avoid an overhang or undercut in the heat shield in which dirt could accumulate or erosion could develop. A method for modifying a gas turbine engine, 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 that leads to lower cooling air consumption, wherein the individual elements of the guide vane carrier have a gap that is labyrinth-shaped or S-shaped, wherein the front element in the flow direction has a first nose and the second element in the flow direction has a second nose formed thereabove.so that an S-shaped gap is formed, whereby the opening of the gap in the hot gas duct is located at the rear in the flow direction, Method for modifying a gas turbine engine, which at least has: 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 that are designed to generate a spoiler effect, Method for modifying a gas turbine engine, which at least has: a compressor, a combustion chamber with combustion chamber bricks, a hot gas duct with a rotor, wherein combustion chamber bricks are installed that receive two separate recesses on two opposite side surfaces of the combustion chamber brick, which serve for engagement by a mechanical clamping from the rear side of the combustion chamber brick, Method for modifying a gas turbine engine, which at least has: 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 then closed again. A method for modifying a gas turbine engine, which at least has 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 than the previous guide vanes, and a trailing edge is rotated relative to the longitudinal axis of the blade. A method for modifying a gas turbine engine, which at least has a compressor, a combustion chamber, a hot gas duct with a bladed rotor, wherein a system for monitoring the combustion dynamics and combustion accelerations of the combustion chamber and burner is subsequently installed.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, wherein the turbine blade has a root with recesses, wherein the recesses have a larger radius than the previously installed and replaced turbine blades. The following machine types are preferably achieved thereby: Gas turbine engine, at least comprising: 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, viewed in a flow direction of the gas turbine engine, wherein the rotor bearing is at least 370 mm long, in particular a maximum of 500 mm long. Gas turbine engine, at least comprising: a compressor, a combustion chamber, a hot gas duct with a bladed rotor, at least one burner for the combustion chamber.in which the fuel supplying means, in particular pipes, at least partially, in particular completely, have a diffusion coating on the inside, in particular are alitated, Gas turbine engine, at least comprising: a compressor, a combustion chamber, a hot gas duct with a bladed rotor, wherein the compressor has a compressor housing which is designed in two parts and has an inner compressor housing as a blade carrier, which has a first material, in particular steel, very particularly cast steel, and an outer compressor housing as a blade carrier has a second material which is significantly different from the first material, in particular grey cast iron, or in which the inner compressor housing as a blade carrier has grey cast iron or is replaced by grey cast iron, Gas turbine engine, at least comprising: a compressor, a combustion chamber, a hot gas duct with a bladed rotor, with stages (I, II, III,IV) of rotor blades and guide vanes, wherein the turbine rotor blades and guide vanes, in particular of stages (I, II), are more resistant to high temperatures, in particular have improved cooling, or in which the rotor blades or guide vanes in the hot gas channel have a directionally solidified microstructure in the form of a columnar solidified microstructure, in particular only the first two stages (I, II), very particularly only the first stage (I), or in which a segmented ceramic layer based on yttrium-stabilized zirconium oxide is present or applied to the rotor blades and guide vanes, or in which the rotor blades 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 in which the ceramic coating comprises partially stabilized yttrium-stabilized zirconium oxide, 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 that directly adjoins a web on the suction side and thus represents additional material in the recess, wherein a cooling air bore extends through the shoulder from the interior of the rotor blade to better cool the blade tip. Gas turbine engine, 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, in which the rotor and guide blades, in particular of stages (I, II), have cooling holes on the side surfaces of the blade platforms. Gas turbine engine, 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 engine, at least comprising a compressor, a combustion chamber, a hot gas duct with a bladed rotor, with stages (I, II, III, IV) of moving blades and guide vanes, wherein the moving blade of stage (IV) does not need to be cooled, in particular the guide vane of stage (III) is not cooled, gas turbine engine, at least comprising: a compressor, a combustion chamber, a hot gas duct with a bladed rotor, in which, viewed in the flow direction, there is a front plenum and a rear plenum outside the hot gas duct, which have different pressures for technical reasons, wherein the front plenum is located behind the moving blade of stage (III) and above the guide vane of stage (IV) in the flow direction, and a long duct from the rear plenum is located in the blade carrier, which cools the stage (III) from the rear plenum, or gas turbine engine,at least comprising: a compressor, a combustion chamber with combustion chamber bricks or heat shields, a hot gas duct with a bladed rotor having a gap between a heat shield and guide vane of stage (I) of the rotor, wherein a curve at the flow-side end of the heat shield and the opposite curve of the guide vane of stage (I) are designed identically to avoid an overhang or undercut in the heat shield in which dirt could accumulate or erosion could develop, gas turbine engine, 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 that leads to lower cooling air consumption, wherein the individual elements of the guide vane carrier have a gap that is labyrinth-shaped or S-shaped,wherein the front element in the flow direction has a first nose and the second rear element 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 at the rear in the flow direction, Gas turbine engine, at least comprising: a compressor, a combustion chamber with heat shields or combustion chamber bricks, a hot gas duct with a bladed rotor, wherein the combustion chamber is designed at the flow-side end of the heat shields or combustion chamber bricks such that they generate a spoiler effect, Gas turbine engine, at least comprising: a compressor, a combustion chamber with combustion chamber bricks, a hot gas duct with a bladed rotor, wherein two separate depressions are formed on two opposite side surfaces of the combustion chamber brick,which serve for engagement by a mechanical clamping from the back of the combustion chamber brick, Gas turbine engine, at least comprising: 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, in particular in the region of the guide vane recesses, Gas turbine engine, at least comprising: a compressor, a combustion chamber with combustion chamber bricks, a hot gas duct with a bladed rotor, wherein the combustion chamber is arranged a burner 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, Gas turbine engine, at least comprising a compressor, a combustion chamber, a hot gas duct with a bladed rotor, wherein a system is installed,which monitors the combustion dynamics and accelerations originating from the combustion and the combustion chamber, gas turbine engine in a combined cycle power plant, at least comprising a compressor, a combustion chamber, a hot gas duct, with a bladed rotor and a steam turbine, wherein an exhaust gas of the gas turbine engine is indirectly used for steam generation in a steam turbine, wherein the steam turbine has blades with a blade root, wherein the recesses have a larger radius.

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

[0073] The Figure 5 shows how Figure 1 a gas turbine engine 100 with the 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] However, for a longer service life of the gas turbine engine 100, a rotor bearing 31 of the rotor 103 in the flow direction 11 of the gas turbine engine 100 at the beginning of the compressor 105 and near the intake casing 104 has a length of at least 370 mm and in particular a maximum length of 500 mm or is designed to be at least 5% longer in the event of an upgrade or revision in order to achieve a lower surface pressure.

[0076] This means that the installed bearing does not need to be replaced during servicing or at one of the next service intervals, or it can be used until the end of the service life of the gas turbine engine 100.

[0077] It can also be replaced as part of a major service project if replacing a bearing requires only a short amount of time and can be performed in parallel or even made easier. A similar situation 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 the fuels, especially gas or oil, can vary locally.

[0079] This also 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] For a longer service life, the fuel supply means such as pipes, in particular also of the burner 107', in particular of gas, are at least partially, in particular completely, provided internally with a diffusion coating, in particular alitated, i.e. an internal alitation (or chroming, ...) is then used here.

[0082] The interior coating can also be carried out with the products in the installed state.

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

[0084] This means that the service life of the 107` burner can be individually adapted to the respective operating conditions.

[0085] Combustion chambers 111 refer to well-known systems such as annular combustion chambers or CANs. Lifespan / Efficiency

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

[0087] The materials of the compressor housings 19', 19", especially if they are one-piece, are generally made of the same first material, in particular gray cast iron. For an improved modification, the inner compressor housing 19" is made as a guide vane carrier from a significantly different second material, in particular from a cast steel.

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

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

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

[0091] Such a blade 120, 130 preferably has a directionally solidified structure SX, DS in the form of a columnar solidified microstructure, such as alloys with the addition DS in Figure 4 .

[0092] Another type of blade 120, 130 has a single-crystalline microstructure in the substrate such as alloy in Figure 4with the suffix SX or CMSX .... In particular, only the first stage I has a DS structure and, more particularly, only the guide vane of stage I.

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

[0094] A ceramic coating (TBC) based on partially stabilized YSZ (yttrium-stabilized zirconia) has a porosity of 1214%.

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

[0096] Another type of blade 120, 130 consists of a directionally solidified structure DS in the substrate, ie 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, in particular starting from Figure 3 , but in which 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 provided singly or multiple times on one, two, three, or all four side surfaces 404, as required. Optionally, cooling holes 405 can also be provided on the blade tip 415 (shown only schematically). Cooling air holes 418 are also provided on the blade airfoil 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 ) extend at an angle different from 90° to the side surface 404 of the blade platform 403 and / or have a diffuser.

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

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

[0101] The blade tip 500 has two outwardly extending webs 503, 505, which, viewed in cross-section, enclose a recess 504. The original recess 504 is indicated by dashed lines 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 directly adjoins the web 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 better cool 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 mounted in the area of ​​this blade carrier 50.

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

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

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

[0109] Figure 12 shows the transition from a last 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 a gap 64 is present between the heat shield 155 and the guide vane 130.

[0110] The curve 72 at the downstream end of the heat shield / combustion chamber brick 155 and the opposite curve 75 of the stage I guide vane 130 are identical. This is intended to prevent an overhang or undercut in the heat shield / combustion chamber brick 155, in which dirt could accumulate or erosion could occur. 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 lower cooling air consumption is achieved by the fact that the front element 81 in the flow direction 11 has a first nose 82 and the rear second element 83 in the flow direction 11 has a second nose 85 formed above it, so that an S-shaped gap 80 is formed and the nose 82 of the front element 81 forms part of the hot gas duct 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 stones 601, 604, 610, which lead to a spoiler effect.

[0115] The combustion chamber stones 601, 602 and 603, 605 are arranged in series and in the circumferential direction as seen in the flow direction 11.

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

[0117] The modified combustion chamber bricks 603, 604, 605, preferably made of solid ceramic, preferably in the two last 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 increasingly thicker in the flow direction 11, resulting in a spoiler effect. Near the rotor hub, preferably only one combustion chamber brick 604 is designed to be spoiler-shaped, whereas further radially away at the outer end of the combustion chamber 110, at least the penultimate and last row of combustion chamber bricks 603, 605, viewed together, have a gradual increase in thickness in the flow direction.

[0118] However, this should not be limiting.

[0119] This spoiler effect reduces erosion and also narrows the hot gas flow, which also increases efficiency. life / 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 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 back and a corresponding through opening 42 on the back 43 of the combustion chamber brick 155, which is opposite the top side 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 section through two combustion chamber stones placed next to each other according to Figure 15 along the through opening 42', 42", whereas in Figure 17 a gap between two combustion chamber bricks 155 according to Figure 15 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 introduced to reduce leakage in this area ( Fig. 19 ). Efficiency

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

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

[0129] For guide vane 73', the dashed line indicates the original position of the swirler's guide vane, whereas the dashed line for guide vane 73'' indicates 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 power a generator, but are often also used in combination with steam turbines in a combined cycle power plant.

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

[0132] The combustion stability and dynamics also have a very high influence on the service life of the system, 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 the measures can be carried out more easily together and, if necessary, in parallel.

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

[0135] Considerations such as extending service life, efficiency, higher temperatures and coolant consumption all 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 respective auxiliary units, service contracts are often concluded, which offer the operator of the energy conversion plant a service maintenance contract which includes a guarantee of a certain mileage (service life) for certain performance characteristics with specified service intervals.

[0137] For the service provider, this means that they not only repair (refurbish) what needs to be repaired because it can no longer be used or can only be used for a short time, but also consider taking measures during a service that allow the period to be extended until the next service interval or the service interval after that.

[0138] For example, when replacing turbine blades, new burners with alumina plating can be used, eliminating the need to service the burner components at the next or subsequent service intervals. This avoids downtime.

[0139] Likewise, a service life extension can be carried out during a service measure or a service measure can be brought forward so that the next service measure is significantly shortened, because measures such as bearing removal, turbine blade removal, replacement of combustion chamber bricks, changes to the casing, etc. involve different maintenance times.

[0140] This is also due to 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 operational lifecycle. With its high efficiency at part-load operation and high operational flexibility, the combined cycle gas turbine (CCGT) plant, along with the associated services, will complement the region's fluctuating renewable energy sources as part of this flexible refurbishment.

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

[0143] Long-term, flexible service goes one step further, offering a maintenance program tailored to specific needs and requirements. Whether the power plant operator wants to maximize production by extending the time between inspections, conduct inspections during scheduled plant downtimes, optimize lifecycle costs through condition-based component replacement, or work without predefined inspection schedules, long-term, flexible refurbishment can adapt to specific needs.

Claims

1. A method for carrying out service measures on an energy conversion plant (1), wherein the energy conversion plant (1) has 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 defective components of the at least one machine (5, 6, 100, ...) is or are replaced by either a new, identical component or new, identical components and / or is or are repaired, and wherein, when carrying out these repairs, further measures for extending the service life of machines (5, 6, 100, ...) or their components and / or further measures for optimizing machines (5, 6, 100, ...) or their components are carried out, 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 (103), with stages (I, II, III, IV) of rotor and guide blades, wherein high-temperature-resistant rotor (120) and guide blades (130) are installed, in particular in the stages (I, II), wherein rotor (120) or guide blades (130) are installed in the hot gas duct (111) which have a directionally solidified microstructure in the form of a columnar solidified microstructure, in particular only the first two stages (I, II), very particularly only the first stage (I), and / or wherein a segmented ceramic layer based on yttrium-stabilized zirconium oxide is applied to the rotor (120) and guide blades (130), and / or wherein the rotor (120) and guide blades (130) are installed which have a monocrystalline 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 zirconium oxide, with a porosity of 12±4%, and / or in which the rotor blades (120) or guide vanes (130) in the hot gas channel (111) have a TBC without segmentation on the guide vanes (130) or rotor blades (120).

2. The 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 as well as burners (107) or burner components.

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

6. Method according to one or more of claims 1, 2, 3 or 4, wherein only measures for optimizing machines (5, 6, 100, ...) or their components are carried out as further measures.

7. Method according to one or more of claims 1, 2, 3 or 4, wherein as further measures, measures for extending the service life and measures for optimising machines (5, 6, 100, ...) or components (120, 130, ...) are carried out.

8. Method according to one or more of claims 1, 2, 3, 4, 5 or 7, wherein during the service measures 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: bearings of the rotor (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) are carried out.

9. Method according to one or more of claims 1, 2, 3, 4, 6 or 8, wherein during the service measures at least one, in particular at least two identical or at least two different measures for optimising machines (5, 6, 100, ...) or their components from the group: increasing efficiency, improving cooling, 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 combustion chamber-turbine transition are carried out.

10. The method according to one or more of claims 1, 2, 3, 4, 5, 6, 7, 8 or 9, wherein the gas turbine engine (100) is modified, wherein the gas turbine engine (100) comprises 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 engine (100), wherein the rotor bearing is replaced, wherein the new rotor bearing (31) is at least 5% longer or a new, at least 370 mm long rotor bearing (31) is installed, in particular wherein the new rotor bearing (31) is a maximum of 500 mm long, and / or wherein the gas turbine engine (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 supplying means, in particular pipes,at least partially, in particular completely, are provided internally with a diffusion coating, in particular alitated, or the fuel supplying means, in particular pipes, are installed with a diffusion coating on the inside, in particular internally alitated, and / or wherein the gas turbine engine (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) has a compressor housing (19) which is or will be formed in two parts (19', 19"), and in particular in which an inner compressor housing (19") is installed as a blade carrier made of a first material, in particular steel, very particularly cast steel, and the outer compressor housing (19') as a blade carrier has a second material that is significantly different from the first material, in particular gray cast iron,and / or in which the inner compressor housing (19") is replaced by grey cast iron, and / or in which a blade tip (415, 500) in a recess (504) has a stepped shoulder (507), which directly adjoins a web (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 rotor blade (120) in order to better cool the blade tip (500), and / or wherein the gas turbine engine (100) has 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 blades, in which the rotor (120) and guide blades (130) are installed, in particular in the stages (I, II), which have cooling holes (399) on the side surfaces (404) of the blade platforms (403), and / or wherein the gas turbine engine (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 moving blades and guide blades, wherein moving blades are installed, wherein the blade tip (415) of the moving 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 engine (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 moving blades and guide blades, wherein moving blades (402) of stage (IV) are installed, which are not cooled, in particular the guide blades (401) of stage (III) are also not cooled, and / or wherein the gas turbine engine (100) comprises at least: a compressor (105), a combustion chamber (110), a hot gas channel (111) with a bladed rotor (103),in which, viewed in the flow direction (11), a front plenum (54) and a rear plenum (57) are present outside the hot gas duct (111), which (54, 57) have different pressures for technical reasons, wherein the front plenum (54) is present in the flow direction (11) behind the rotor blade of the stage (III) and above the guide vane of the stage (IV), and either ducts (53) that were present in the middle of the stage (IV) between the guide vane (401) and the rotor blade (402) and were previously used to cool the turbine guide vanes and rotor blades of the stage (IV) or to supply cooling air are closed, 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 only has one such duct (60), and / or wherein the gas turbine engine (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) has 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) is designed to be the same in order to avoid an overhang or an undercut in the heat shield (155) in which dirt could accumulate or erosion could develop, and / or wherein the gas turbine engine (100) has 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 provided with a sealing arrangement (79) made of elements (81, 83) is installed or modified, 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 (80) is labyrinth-shaped or S-shaped, wherein the front element in the flow direction (11) has a first nose (82) and the second rear element (83) in the flow direction (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 at the rear as seen in the flow direction, and / or wherein the gas turbine engine (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 such that they generate a spoiler effect, and / or wherein the gas turbine engine (100) comprises at least: a compressor (105), a combustion chamber (110) with combustion chamber stones, a hot gas channel (111) with a rotor (103), wherein combustion chamber stones are installed,which accommodate two separate recesses (40', 40'') on two opposite side surfaces of the combustion chamber brick, which serve for engagement by a mechanical clamping from the rear side (43) of the combustion chamber brick, and / or wherein the gas turbine engine (100) comprises at least: a compressor (105), a combustion chamber (110) with combustion chamber bricks, a hot gas channel (111) with a bladed rotor (103), a housing part (550) for the hot gas channel (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 engine (100) comprises at least: a compressor (105), a combustion chamber with combustion chamber bricks and burners, a hot gas channel (111) with a rotor (103), wherein modified guide vanes (73', 73'') are inserted into the swirler of the burner (70) are installed, which have a smaller opening angle than the previous guide vanes,and a trailing edge is rotated relative to the longitudinal axis of the blade, and / or wherein the gas turbine engine (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 of the combustion chamber and burner is subsequently installed 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 has turbine blades (883), wherein the turbine blade (883) has a root (880) with recesses, wherein the recesses (886', 886", 886‴) have a larger radius than the previously installed and replaced turbine blades.

11. An energy generation plant, after carrying out a method according to one or more of claims 1 to 10, or an energy conversion plant (1), at least comprising: a gas turbine engine (100), at least comprising: 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) as seen in a flow direction (11) of the gas turbine engine (100), wherein the rotor bearing (31) is at least 370 mm long, in particular a maximum of 500 mm long, and / or in which the gas turbine engine (100) at least comprises: 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), in which the fuel supplying means, in particular pipes, at least partially, in particular completely, have a diffusion coating on the inside, in particular are alitated,and / or in which the gas turbine engine (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 (19) is formed in two parts (19', 19") and has an inner compressor housing (19") as a blade carrier, which comprises a first material, in particular steel, very particularly cast steel, and an outer compressor housing (19') as a blade carrier, which comprises a second material which is significantly different from the first material, in particular grey cast iron, or in which the outer compressor housing (19") comprises grey cast iron, and / or in which the gas turbine engine (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 blades (120) and guide vanes (130), in particular of the stages (I,II), are more resistant to high temperatures, in particular have improved cooling, and / or in which the rotor blades (120) or guide vanes (130) in the hot gas channel (111) have a directionally solidified microstructure in the form of a columnar solidified microstructure, in particular only the first two stages (I, II), very particularly only the first stage (I), and / or in which a segmented ceramic layer based on yttrium-stabilized zirconium oxide is present on the rotor blades (120) and guide vanes (130), and / or in which the rotor blades (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 zirconium oxide, with a porosity of 1214%, and / or with a TBC without segmentation on the guide vanes (130) or rotor blades (120), and / or in which a blade tip (415,500) has a stepped shoulder (507) in a recess (504), which directly adjoins a web (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 rotor blade (120) in order to better cool the blade tip (500), and / or in which the gas turbine engine (100) has 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 blades, in which the rotor (120) and guide blades (130), in particular of the stages (I, II), have cooling holes (399) on the side surfaces (404) of the blade platforms (403), and / or in which the Gas turbine engine (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 blades,wherein the blade tip (415, 500), in particular of the stages (I, II), is cooled, in particular by cooling holes (501) in the blade tip (415), and / or in which the gas turbine engine (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 moving blades and guide blades, wherein the moving blade (402) of stage (IV) does not need to be cooled, in particular the guide blade (401) of stage (III) is not cooled, and / or in which the gas turbine engine (100) comprises at least: a compressor (105), a combustion chamber (110), a hot gas duct (111) with a bladed rotor (103), in which, viewed in the flow direction (11), a front plenum (54) and a rear plenum (57) is located outside the hot gas channel (111), which (54, 57) have different pressures for technical reasons,wherein the front plenum (54) is provided 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 provided in the blade carrier (50), which cools the stage (III) from the rear plenum (57), and / or in which the gas turbine engine (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) having 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 the stage (I) are designed the same to avoid an overhang or undercut in the heat shield (155) in which dirt could accumulate or erosion could occur,and / or in which the gas turbine engine (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 up 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 labyrinth-shaped or S-shaped, wherein the front element in the flow direction (11) has a first nose (82) and the second rear element (83) in the flow direction (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 at the rear in the flow direction, and / or in which the gas turbine engine (100) at least comprises: 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 such that they generate a spoiler effect, and / or in which the gas turbine engine (100) has at least: a compressor (105), a combustion chamber (110) with combustion chamber bricks, a hot gas duct (111) with a bladed rotor (103), wherein on two opposite side surfaces of the combustion chamber brick two separate recesses (40', 40") are formed, which serve for engagement by a mechanical clamping from the rear side (43) of the combustion chamber brick, and / or in which the gas turbine engine (100) has 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), in particular in the region of the guide vane recesses, and / or in which the gas turbine engine (100) has at least: a compressor (105), a combustion chamber (110) with combustion chamber bricks, a hot gas duct (111) with a bladed rotor (103), wherein the combustion chamber (110) has a burner (70) arranged therein, 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 in which the gas turbine engine (100) in a combined cycle power plant (1) has 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 of the gas turbine engine (100) is used indirectly 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 engine (100) comprises at least: a compressor (105), a combustion chamber (110), a hot gas duct (111) with a bladed rotor, wherein a system (90) is installed which monitors the combustion dynamics and accelerations emanating from the combustion and the combustion chamber (110).

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