Gas turbine and method for improving a gas turbine

Abstract

The invention relates to a gas turbine (1), at least comprising: a rotor (2), a combustion region (7), a compressor (4), upstream of the combustion region (7), as part of the rotor (2) and a hot gas part (10), downstream of the combustion region (7), as part of the rotor (2), wherein the compressor (4) has compressor discs for compressor blades, wherein the hot gas part (10) has turbine discs for turbine blades, wherein at least one compressor disc of the compressor (4) has a material that differs from that of the other compressor discs and / or at least one turbine disc of the hot gas part (10) has a material that differs from that of the other turbine discs.

Description

[0001] Gas turbine and method for improving a gas turbine

[0002] The invention relates to an improved gas turbine and a method for improving a gas turbine.

[0003] It is the constant task of the specialist to increase the efficiency and service life of a gas turbine.

[0004] A compressor, a combustion area and the hot gas part can be improved.

[0005] One possibility is to increase the degree of compression of the air by using an improved compressor.

[0006] However, this has an impact on all components of the gas turbine.

[0007] It is therefore an object of the invention to show changes which take up and adjust these effects of an improved compressor on the gas turbine.

[0008] The object is achieved by a gas turbine according to claim 1 and a method according to claim 23.

[0009] The subclaims list further advantageous measures which can be combined in any desired manner to achieve further advantages.

[0010] It shows

[0011] Figure 1 schematically shows a gas turbine,

[0012] Figure 2 a compressor blade,

[0013] Figure 3 a combustion chamber,

[0014] Figure 4 , 5 an exhaust housing and

[0015] Figure 6 shows an assembled rotor.

[0016] The description and figures represent only exemplary embodiments of the invention.

[0017] The task is solved by a

[0018] Gas turbine 1, at least comprising: a rotor 2, a combustion region 7, a compressor 4 arranged upstream of the combustion region 7 as part of the rotor 2 and a hot gas part 10 arranged downstream of the combustion region 7 as part of the rotor 2, wherein the compressor 4 has compressor disks 13, 16 for compressor blades 40, wherein the hot gas part 10 has turbine disks 22, 23 for turbine blades 25, wherein at least one compressor disk 16 of the compressor 4 has a different material than the other compressor disks 13 and / or at least one turbine disk 23 of the hot gas part 10 has a different material than the other turbine disks 22.

[0019] An advantageous embodiment is when the gas turbine has a compressor 4 with at least twelve compressor stages, in particular a compressor 4 with fifteen compressor stages, wherein at least the last compressor stage, in particular the last two compressor stages, have a different material for the compressor disks 16 than the front compressor disks 13.

[0020] An advantageous embodiment of the gas turbine is wherein the front compressor disks 13 or wherein at least the four compressor disks 13 directly in front of either the last or the penultimate and last compressor disk 16 have a first steel with (in wt.%): Carbon (C), in particular 0.20% < C < 0.35%, Chromium (Gr), in particular 1.0% < Gr < 2.5%,

[0021] Molybdenum (Mo), especially 0.20% < Mo < 0.55%

[0022] Vanadium (V), in particular 0.03% < V < 0.20%, nickel (Ni), in particular 3.0% < Ni < 4.4%, whereby proportions of silicon (Si), manganese (Mn), phosphorus (P), sulphur (S), arsenic (As), antimony (Sb) and / or tin (Sn) are to be kept as small as possible, in particular the first steel consists of them.

[0023] An advantageous embodiment of the gas turbine is in which either the last or the penultimate and last compressor disk 16 has a second steel and, compared to the first steel of the other compressor disks 13, has a low proportion of carbon (C), a significantly higher proportion, in particular at least 100% higher, of chromium (Cr), a higher proportion of molybdenum (Mo) and the deliberate addition of niobium (Nb), tungsten (W), nitrogen (N) and manganese (Mn).

[0024] An advantageous embodiment of the gas turbine is wherein the second steel comprises (in wt.%): carbon (C), in particular 0.08% < C < 0.19%, chromium (Cr), in particular 9.5% < Cr < 11.0%, molybdenum (Mo), in particular 0.80% < Mo < 1.5%, vanadium (V), in particular 0.10% < V < 0.30%, nickel (Ni), in particular 0.4% < Ni < 1.2%, tungsten (W), in particular 0.7% < W < 1.3%, manganese (Mn), in particular 0.02% < Mn < 0.70%, nitrogen (N), in particular 0.035% < N < 0.070% niobium (Nb), in particular 0.02% < Nb 0.08%, in particular the second steel thereof consists.

[0025] An advantageous embodiment is when the gas turbine has torque disks 19 between the compressor 4 and the hot gas part 10, wherein the at least one torque disk 19, in particular wherein all torque disks 19, have the same material as the last compressor disk 16 of the compressor 4.

[0026] An advantageous embodiment is when the gas turbine in the hot gas part 10 has at least 3 stages I, II, III, IV, in particular four stages I, II, III, IV, wherein at least the first two turbine disks 22 of the stages I, II, III, IV, in particular the first three turbine disks 22 of the stages I, II, III, IV have the same material and are different from the material of the turbine disk 23 either of the third and fourth stage III, IV or of the last, fourth stage IV, depending on which stages of the stages I, II and optionally III have the same material.

[0027] An advantageous embodiment is when the gas turbine in the hot gas part 10 has at least 3 stages I, II, III, IV, in particular four stages I, II, III, IV, wherein all turbine disks 22, 23 of the stages I, II, III, IV have the same material.

[0028] An advantageous embodiment of the gas turbine is that at least the front turbine disks 22 have the same material as the material of the torque disks 19.

[0029] An advantageous embodiment is when the gas turbine in the hot gas part 10 has four stages I, II, III, IV, wherein a stage I, II, III, IV consists of stationary turbine blades VI, V2, V3, V4 and unsteady turbine blades B1, B2, B3, B4, wherein the turbine blades VI, B1, V2, B2 of the first two stages I, II and in particular the stationary turbine blades V3 of the third stage III, especially the stationary turbine blades V3 and the unsteady turbine blades B3 of the third stage III, have a different material than the unsteady turbine blades B3 of the third stage

[0030] III or at least the turbine blades V4, B4 of the fourth stage IV, depending on the material selection of the first stages.

[0031] An advantageous embodiment is if the gas turbine has an SX or column-like structure for the first, unsteady turbine blades Bl.

[0032] An advantageous embodiment of the gas turbine is that the turbine blades V4, B4 of the fourth stage IV, optionally also the unsteady turbine blades B3 and / or stationary turbine blades V3 of the third stage III, comprise a nickel-based alloy (in wt%): cobalt (Co), in particular 9.0% - 10%, chromium (Cr), in particular 13.0% - 15.0%, tungsten (W), in particular 3.5% - 4.5%, molybdenum (Mo), in particular 3.7% - 4.3%, aluminum (Al), in particular 2.8% - 3.2%, titanium (Ti), in particular 4.8% - 5.5%, boron (B), in particular 0.01% - 0.2%, zirconium (Zr), in particular 0.02% - 0.10%, carbon (C), in particular 0.15% - 0.40%, preferably no tantalum (Ta), preferably no hafnium (Hf), preferably no niobium (Nb). In particular, it consists of:

[0033] An advantageous embodiment of the gas turbine is wherein the material for at least the turbine blades VI, Bl, V2, B2 of the front stages I, II, in contrast to the material of the turbine blades B4, V4, optionally also in contrast to the turbine blades B3, V3 of the third stage III, comprises tantalum (Ta) and / or hafnium (Hf).

[0034] An advantageous embodiment is when the gas turbine has at least three stages I, II, III, IV in the hot gas part 10, in particular has four stages I, II, III, IV in the hot gas part 10, wherein a stage I, II, III, IV consists of stationary turbine blades VI, V2, V3, V4 and unstationary turbine blades B1, B2, B3, B4, wherein the turbine blades VI, B1, V2, B2 of the front stages, optionally the turbine blades V3, B3 of the third stage, have a nickel-based alloy (in wt%): cobalt (Co), in particular 9.0% - 10.5%, chromium (Cr), in particular 8.0% - 9.5%, tungsten (W), in particular 9.0% - 10.5%, molybdenum (Mo), in particular 0.2% - 1.0%, aluminum (Al), in particular 5.0% - 6.0%, titanium (Ti), in particular 0.5% - 1.5%, tantalum (Ta), in particular 2.5% - 3.5%, boron (B), in particular 0.01% - 0.25%, zirconium (Zr), in particular 0.004% - 0.06%, carbon (C), in particular 0.05% - 0.16%, hafnium (Hf), in particular 1.0% - 2.0%, in particular consisting thereof.

[0035] An advantageous embodiment is when the gas turbine has at least three stages I, II, III, IV in the hot gas part 10, in particular has four stages I, II, III, IV in the hot gas part 10, wherein a stage I, II, III, IV consists of stationary turbine blades VI, V2, V3, V4 and unstationary turbine blades B1, B2, B3, B4, wherein the turbine blades VI, B1, V2, B2, V3, B3, V4, B4 of all stages I, II, III, IV have a metallic oxidation protective layer, preferably a metallic NiCoCrAlYRe, NiCoCrAlY, NiCoCrAlYTaSi, NiCoCrAlYTaHf, NiCoCrAlYTaSiHf or NiCoCrAlYTa layer, very preferably all turbine blades 25 have the same metallic layer.

[0036] An advantageous embodiment is when the gas turbine has at least three stages I, II, III, IV in the hot gas part 10, in particular has four stages I, II, III, IV in the hot gas part 10, wherein a stage I, II, III, IV consists of unsteady turbine blades B1, B2, B3, B4 and stationary turbine blades VI, V2, V3, V4, wherein only the turbine blades of the first stages I, II, III have a ceramic coating, in particular the first three stages I, II, III have a ceramic coating.

[0037] An advantageous embodiment of the gas turbine is in which a ceramic coating of the turbine blades of the first two stages I, II: VI, Bl, V2, B2 and optionally of the turbine blades V3, B3 of the third stage III have the same ceramic coating, in particular a segmented, in particular dense (< 10%), ceramic layer with zirconium oxide as a base.

[0038] An advantageous embodiment is when the gas turbine in which a ceramic coating of the turbine blades V3, B3 of the third stage III has a highly porous, ceramic layer with a porosity greater than 12%, in particular greater than 18%, preferably with zirconium oxide as a matrix.

[0039] An advantageous embodiment of the gas turbine is that improved compressor blades 40 are present in at least one compressor stage of the compressor 4, the improved compressor blades 40 being characterized in that a rubbing edge 44 does not run along the edge of the blade contour of the blade 43, but is preferably arranged centrally on the blade tip 43, and in particular runs over the entire length of the blade tip 43.

[0040] An advantageous embodiment of the gas turbine is that only one compressor stage of the compressor 4 , in particular a compressor stage between the fourth and eighth compressor stage of the compressor 4 , has the improved compressor blade 40 according to claim 18 .

[0041] An advantageous embodiment is when the gas turbine , which has a combustion chamber 50 , which has 50 a tiling 8 , which has 8 fully ceramic tiles 52 , 53 , 54 and metallic heat shield elements 55 .

[0042] An advantageous embodiment of the gas turbine is that, viewed in the direction of flow of the hot gas in the combustion chamber 50, it has a first ceramic material K1 for fully ceramic tiles 52, then preferably a second ceramic material K2 for fully ceramic tiles 53, then a third ceramic material K3 for fully ceramic tiles 53 and then preferably at the end the metallic heat shields 55 are present, i.e. only a first and third material can be used. An advantageous embodiment of the gas turbine is wherein the first ceramic material K1 comprises a mixture of corundum and mullite in a ratio of 40:60 to 60:40, in particular a half mixture, wherein the optional second ceramic material K2 comprises a proportion of mullite of 15%-25%, wherein the third ceramic material K3 comprises hardly any mullite, in particular the proportion is less than 5%.

[0043] An advantageous embodiment is when the gas turbine , which has three discharge pipes 30 , can guide the compressed air from the compressor 4 into an exhaust housing 33 .

[0044] An advantageous method for modifying a gas turbine is to have at least one feature of one or more of the above features .

[0045] Furthermore, the method is advantageous in which the installed, completely assembled rotor of an existing gas turbine is replaced by a new, completely assembled rotor (2) with at least one feature of one or more of the previous features.

[0046] A further advantageous method is the installation of a further discharge pipe 30 which directs air from the compressor 4 into an exhaust housing 33.

[0047] A further advantageous method is one in which the gas turbine 1 has an exhaust casing 33 with an annular discharge channel 65, in which struts 63 are present, which are modified by removing a front region of the struts 63 and replacing them with a new, modified profile that has less curvature. A gas turbine 1 has at least: a rotor 2, a compressor 4, a downstream combustion region 7, and a downstream hot gas section 10.

[0048] The compressor 4 has compressor disks 13, 16 for the different compressor stages with respective compressor blades.

[0049] A gas turbine 1 has at least twelve compressor stages, in particular fifteen compressor stages, for the compressor 4 .

[0050] Each stage of a compressor 4 consists of guide and rotor blades in corresponding compressor disks 13 , 16 .

[0051] The design of compressor 4 is developed and adjusted in such a way that the thermal and mechanical load on at least the last compressor stage, in particular on the last two compressor stages in compressor 4, increases.

[0052] At least the last compressor disk 16 of the compressor 4 therefore preferably has a different material than the remaining, front compressor disks 13 or than the at least four compressor disks 13 up to the penultimate and / or last compressor disk 16.

[0053] Preferably, the last two compressor disks 16 of the compressor 4 have a different material than the remaining, front compressor disks 13 or than the at least four compressor disks 13 up to the penultimate and / or last compressor disk 16.

[0054] The different material achieves increased strength, particularly at higher temperatures. This is achieved through modification of the alloy and / or heat treatment. "Different" or "different" here, in steel or in general, means that at least one alloying element is present in greater or lesser quantities, or that the proportion of an alloying element is at least 10%, in particular at least 20%, higher or lower.

[0055] Steel is always at least an iron-carbon alloy.

[0056] Therefore, different materials are deliberately used for the compressor discs 13, 16, whereby the last compressor disc 16 or the last two compressor discs 16 have a higher tensile strength, which is preferably 50 MPa or most preferably 100 MPa higher than that of the front compressor discs 13.

[0057] Examples with 15 stages ( a) , b) ) for compressor 4 or with 13 stages ( c) , d) ) for compressor 4: a)

[0058] 15. Compressor disc: 2. Steel

[0059] 11. - 14. Compressor disc 1. Steel

[0060] 1st - 10th compressor disc 1st steel or different from 1st and 2nd steel. b)

[0061] 14., 15. Compressor disc: 2. Steel

[0062] 10. - 13. Compressor disc 1. Steel

[0063] 1st - 9th compressor disc 1st steel or different from 1st and 2nd steel. c)

[0064] 13. Compressor disc: 2. Steel

[0065] 9. - 12. Compressor disc 1. Steel

[0066] 1st - 8th compressor disc 1st steel or different from 1st and 2nd steel. d)

[0067] 12th, 13th compressor disc: 2nd steel 9th ​​- 11th compressor disc 1st steel

[0068] 1. - 8. Compressor disc 1. steel or different from 1. and 2. steel.

[0069] The material for the front compressor disks 13 or at least the four compressor disks 13 up to the penultimate and / or last compressor disks 16 preferably comprises a first steel (NiCrMoV) with (in wt%): carbon (C), in particular 0.20% < C < 0.35% chromium (Cr), in particular 1.0% < Cr < 2.5%, molybdenum (Mo), in particular 0.20% < Mo < 0.55% and vanadium (V), in particular 0.03% < V < 0.20% as well as nickel (Ni), in particular 3.0% < Ni < 4.4%.

[0070] The proportions of silicon (Si), manganese (Mn), phosphorus (P), sulphur (S), arsenic (As), antimony (Sb) and tin (Sn) should be kept as small as possible.

[0071] In particular, the first steel is made of it. Its tensile strength is 810 MPa - 960 MPa.

[0072] The second steel is different from the first steel and preferably has, compared to the first steel, a low proportion of carbon (C), a significantly higher proportion, in particular at least 100% higher proportion, of chromium (Cr), a higher proportion of molybdenum (Mo) and the deliberate addition of niobium (Nb), tungsten (W), nitrogen (N) and manganese (Mn).

[0073] The other material achieves increased strength, especially at higher temperatures. This is achieved by modifying the alloy.

[0074] At least the last compressor disc 16 and the torque discs 19 comprise the second steel (CrMoWVNbN) (in wt . %) :

[0075] Carbon (C), in particular 0.08% < C < 0.19%

[0076] Chromium (Cr), in particular 9.5% < Cr < 11.0%, molybdenum (Mo), in particular 0.80% < Mo < 1.5%, vanadium (V), in particular 0.10% < V < 0.30%, nickel (Ni), in particular 0.4% < Ni < 1.2%, tungsten (W), in particular 0.7% < W < 1.3%, manganese (Mn), in particular 0.2% < Mn < 0.7%, nitrogen (N), in particular 0.035% < N < 0.07% niobium (Nb), in particular 0.02% < Nb 0.08%.

[0077] In particular, the second steel is made of it.

[0078] Different or other means here in the case of steel or in general that at least one alloying element is present in greater or lesser quantities, or that a proportion of an alloying element is at least 10%, in particular at least 20%, higher or lower.

[0079] In the combustion area 7, i.e. between the compressor 4 and the hot gas part 10, torque discs 19 ("torque disc") are present, which connect the part of the rotor 2 for the compressor 4 with the hot gas part 10.

[0080] Preferably, three torque discs 19 are used.

[0081] Preferably, the same material is used for the, in particular all, torque disks 19 as for the last compressor disk 16 or the last two compressor disks 16.

[0082] At least the last compressor disc 16 and the torque discs 19 therefore preferably also comprise the second steel (CrMoWVNbN).

[0083] The hot gas part 10 has at least three stages, preferably four stages I, II, III, IV.

[0084] Each stage I, II, III, IV of the hot gas part 10 consists of turbine blades 25, stationary guide blades V and unsteady rotor blades B.

[0085] The turbine blades 25 are arranged in front turbine disks 22 and a rear or rear turbine disk 23.

[0086] The turbine disks 22 at least for the first two turbine stages, in particular for the first three turbine stages, have the same material and are preferably different from the material of the turbine disk of the fourth stage in order to cope with the increased thermal loads.

[0087] The material for the torque disks 19 will preferably also be used for the first, first two or three front turbine disks 22.

[0088] The material for the first, first two or three front turbine disks 22 is preferably different from the material of the rear stage III, IV.

[0089] The material for the rear turbine disks 23 of Stage III, IV or only for the rear turbine disk 23 of Stage IV preferably comprises the first steel, but has a higher strength class due to a different heat treatment.

[0090] Different or various here means, in the case of steel, nickel-based superalloys or in general, that at least one alloying element is present in greater or lesser quantities, or that a proportion of an alloying element is at least 10%, in particular at least 20%, higher or lower.

[0091] Also preferably, the front turbine disks 22 may comprise the same material as the rear or rear turbine disks 23.

[0092] It is therefore also possible to use the same material, in particular the second steel, for the compressor disks 16, the torque disks and all turbine disks 22, 23. In the hot gas part 10, the gas turbine has at least three stages, in particular four stages I, II, III, IV, wherein one stage has stationary guide vanes VI, V2, V3, V4 and unsteady rotor blades B1, B2, B3, B4, wherein the turbine blades 25 of the first two stages VI, B1, V2, B2 and optionally also the stationary turbine blades V3 or stationary turbine blades V3 and unsteady turbine blades B3 of the third stage III have a different cast material than that of the turbine blades of the third III and fourth stage IV or than that of the unsteady turbine blades B3 of the third stage and than that of the turbine blades of the fourth stage IV: V4, B4.

[0093] Examples are a) VI, Bl, V2, B2 have a different material than V3, B3, V4, B4, where VI, Bl, V2, B2 have the same material and V3, B3, V4, B4 preferably have the same material, b) VI, Bl, V2, B2, V3 have a different material than B3, V4, B4, where VI, Bl, V2, B2, V3 have the same material and B3, V4, B4 preferably have the same material, c) VI, Bl, V2, B2, V3, B3 have a different material than V4, B4 where VI, Bl, V2, B2, V3, B3 have the same material and V4, B4 preferably have the same material.

[0094] The design depends, among other things, but not only, on the operating conditions of the gas turbine 1 or the design of the compressor 4.

[0095] Different material means that the material differs significantly in the content of carbon (C), chromium (Gr), molybdenum (Mo), titanium (Ti), which are at least 20% lower, as well as tungsten (W) and aluminum (Al), whose shares are at least 50% higher, and the material of the front stages also contains tantalum (Ta) and hafnium (Hf).

[0096] The turbine blades V4, B4, if applicable B3, V3 have a nickel-based alloy (in wt%): cobalt (Co), in particular 9.0% - 10% chromium (Cr), in particular 13.0% - 15.0% tungsten (W), in particular 3.5% - 4.5% molybdenum (Mo), in particular 3.7% - 4.3% aluminum (Al), in particular 2.8% - 3.2% titanium (Ti), in particular 4.8% - 5.5% boron (B), in particular 0.01% - 0.2% zirconium (Zr), in particular 0.02% - 0.10% carbon (C), in particular 0.15% - 0.40%, preferably no tantalum (Ta), preferably no hafnium (Hf), preferably no niobium (Nb). In particular, they consist of it.

[0097] The material for at least the front stages VI, Bl, V2, B2, in contrast to the material for B4, V4, B3, V3, contains tantalum (Ta) and hafnium (Hf).

[0098] The turbine blades VI, Bl, V2, B2, possibly B3, V3 have a nickel-based alloy (in wt%): Cobalt (Co), in particular 9.0% - 10.5% Chromium (Cr), in particular 8.0% - 9.5% Tungsten (W), in particular 9.0% - 10.5% Molybdenum (Mo), in particular 0.2% - 1.0% Aluminum (Al), in particular 5.0% - 6.0% Titanium (Ti), in particular 0.5% - 1.5%, Tantalum (Ta), in particular 2.5% - 3.5%, Boron (B), in particular 0.01% - 0.25% Zirconium (Zr), in particular 0.004% - 0.06% Carbon (C), in particular 0.05% - 0.16%, Hafnium (Hf), in particular 1.0% - 2.0%. They are particularly made of this material. The rotor blade Bl preferably has a directional structure. It is therefore not only cast, but simply allowed to solidify in a directional manner using a known temperature control.

[0099] It then has a single-crystalline (SX) or columnar (columnar) structure.

[0100] This can be done with the same material as for the material of the turbine blades VI, V2 or B2

[0101] The turbine blades 25 VI, Bl, V2, B2, V3, B3, V4, B4, preferably of all existing stages I, II, III, IV, have an oxidation protection layer.

[0102] This is preferably a metallic NiCoCrAlYRe, NiCoCrAlY, NiCoCrAlYTaSi, NiCoCrAlYTaHf, NiCoCrAlYTaSiHf or NiCoCrAlYTa (=MCrAlY) layer.

[0103] The turbine blades 25 of at least the first two stages

[0104] I, II) of the hot gas part 10 have a ceramic coating, ie the turbine blades V3, B3, V4, B4 have no ceramic coating (apart from the TGO of NiCoCrAlYRe, NiCoCrAlY NiCoCrAlYTaSi, NiCoCrAlYTaHf, NiCoCrAlYTaSiHf or NiCoCrAlYTa).

[0105] Preferably, the turbine blades 25 of the first two stages I, II and at least the unsteady turbine blades V3 have a ceramic coating, ie the turbine blades V4, B4, optionally B3 have no ceramic coating (apart from the TGO of the MCrAlY).

[0106] Examples are

[0107] • VI, Bl, V2, B2, V3 have a ceramic coating and B3, V4, B4 have no ceramic coating

[0108] • VI, Bl, V2, B2 have a ceramic coating and

[0109] V3, B3, V4, B4 do not have a ceramic coating

[0110] • VI, Bl, V2, B2, V3, B3, V4 have a ceramic coating and B4 has no ceramic coating • VI, Bl, V2, B2, V3, B3 have a ceramic coating and V4, B4 have no ceramic coating.

[0111] Likewise preferably, a first ceramic coating of the first two stages VI, B1, V2, B2 and optionally the turbine blades V3 of the third stage III) can have the same ceramic coating.

[0112] This is then preferably a segmented, in particular dense (< 10%), segmented ceramic layer, preferably based on zirconium oxide.

[0113] The unsteady turbine blades B3, optionally the stationary turbine blades V3 of the third stage then preferably have a different, second ceramic coating than the turbine blades of the previous stages VI, Bl, V2, B2, opt. V3.

[0114] This is then preferably a highly porous ceramic layer with a porosity > 12%, in particular > 18%.

[0115] This porous layer preferably comprises zirconium oxide, in particular in the form of partially stabilized zirconium oxide (PSZ).

[0116] According to Figure 1, the gas turbine 1 preferably has an additional blow-off pipe 30 ("blow off line"). This blow-off pipe 30 extracts compressed air from the compressor 4, in particular at the level of the fifth or sixth stage of the compressor 4, and is preferably only required when starting up the gas turbine 1. It connects the compressor 4 to an exhaust casing 33.

[0117] The gas turbine 1 preferably already has two vent pipes and, after the conversion or as a new part, three vent pipes 30.

[0118] Figure 2 shows a compressor blade 40 of the improved compressor 4.

[0119] The compressor blade 40 has a platform 41 in a known manner.

[0120] The blade 42 extends from the platform 41 of the compressor blade 40 to a blade tip 43.

[0121] The blade tip 43 is preferably flat and has a preferably single leading edge 44 extending from the blade tip 43 .

[0122] The blade tip 43 of the improved compressor blade 40 has a leading edge 44 that does not run along the edge of the blade contour, but rather preferably extends centrally along the blade tip 43 and, in particular, over the entire length of the blade tip 43. This also results from the improved design of the compressor 4.

[0123] Preferably, only one compressor stage of the compressor 4, in particular a compressor stage between the fourth and eighth compressor stage of the compressor 4, has the improved compressor blade 40.

[0124] Figure 3 shows a combustion chamber 50 of the combustion region 7. In the combustion chamber 50, the highly compressed air from the compressor 4 is mixed with the fuel and burned.

[0125] A burner system 51 is available for this purpose.

[0126] The generated hot gas then flows into the hot gas part 10 .

[0127] The combustion chamber 50 has a tiling 8 and is therefore preferably at least largely, in particular completely, provided with ceramic heat shields 52, which preferably consist of solid ceramic.

[0128] Preferably at least two different, preferably three different materials are used for the all-ceramics.

[0129] The ceramic heat shields 52 in the front region, i.e., beginning at the burner system 51, comprise a first material K1, whereas ceramic heat shields 54 near the outflow region comprise a third, erosion-resistant material K3 for the all-ceramic material. In between, a second ceramic material K2 is preferably used for the tiling 8 for the central ceramic heat shields 53.

[0130] In the case of all-ceramic heat shields, different means that the microstructure is clearly distinguishable and / or the chemical composition differs in at least one component by at least 10%, in particular by at least 20%, or at least one chemical element is contained more or less in the ceramic composition.

[0131] The first ceramic material K1 for the ceramic heat shields 52 in the front region preferably comprises a mixture of corundum and mullite in a ratio of 40:60 to 60:40, in particular half.

[0132] The second ceramic material K2 for the ceramic heat shields 53 in the central region preferably contains less mullite. The proportion of mullite is preferably 15%-25%.

[0133] The third ceramic material K3 preferably contains hardly any mullite or the proportion is < 5%.

[0134] The third ceramic material for the ceramic heat shields near the outflow area preferably contains only corundum , with proportions of 1 % < spinel < 8 % .

[0135] The tiling 8 is constructed in several rows of heat shield elements, seen in the direction of flow.

[0136] Examples of material selection for ceramics Kl, K2, K3 of combustion chambers with different numbers of rows. a) 1st row: Kl

[0137] 2nd row: K2

[0138] 3rd row: K3 b) 1st row: Kl

[0139] 2nd row: K3

[0140] 3rd row: K3 c) 1st row: Kl

[0141] 2nd row: K2

[0142] 3rd row: K3

[0143] 4th row: K3 d) 1st row: Kl

[0144] 2nd row: K2

[0145] 3rd row: K2

[0146] 4th row: K3 e) 1st row: Kl

[0147] 2nd row: Cl

[0148] 3rd row: K2

[0149] 4th row: K3 f) 1st row: Kl

[0150] 2nd row: Cl

[0151] 3rd row: K2

[0152] 4th row: K3

[0153] 5th row: K3 g) 1st row: Kl

[0154] 2nd row: K2

[0155] 3rd row: K2

[0156] 4th row: K3

[0157] 5th row: K3

[0158] The end of the combustion chamber 50 or the inlet bowl has a tiling 8 with metallic heat shields 55.

[0159] The metallic heat shields 55 comprise a metal substrate and a ceramic thermal barrier coating system consisting of a metallic bonding layer (MCrAlY) and a ceramic protective layer, as is known from turbine blades. The combustion chamber 50 has thus also been adapted to the preferably improved conditions of the compressor 4 and has at least two different types / material selections for the tiling 8.

[0160] Figure 4 shows an exhaust housing 33 which has struts 63 for an outer annular channel. Since the exhaust gas exhibits a different outflow behavior due to the preferably improved compressor, the struts 63 are adapted during the modernization (Fig. 5). This is done by cutting off a front part, in particular the front half of the hollow structure of the struts 63, and replacing it with a new profile which reduces the curvature.

[0161] Figure 6 shows a rotor 2 which is completely assembled and has the improved or adapted components at least according to Figures 1, 2, 3, 4 and / or 5.

[0162] Thus, the rotor 2 already has in particular the improved compressor disks 16 with the rotor blades of the compressor 4 or the improved compressor blades 40 or the ceramic heat shield elements 53, 55, 56 as part of a tiling 8 of the combustion chamber 50 or the improved turbine disks or the improved turbine blades 25 and optionally already the adapted exhaust housing 33.

[0163] Such a rotor 2 is easy and quick to install, so that the existing rotor does not have to be unstacked, restacked and rebladed on site.

[0164] This significantly reduces downtime for the plant operator.

[0165] Even without a prior improvement of the compressor, the individual features can be used to improve a gas turbine, since the individual features offer advantages separately and can also be retrofitted, i.e.: Gas turbine 1, at least comprising: a rotor 2, a combustion region 7 as part of the rotor 2, a compressor 4 upstream of the combustion region 7 as part of the rotor 2 and a hot gas part 10 downstream of the combustion region 7 as part of the rotor 2, wherein the compressor 4 has compressor disks for compressor blades, wherein the hot gas part 10 has turbine disks for turbine blades 25, optionally at least one compressor disk 16 of the compressor 4 has a different material than the other compressor disks 13 and / or optionally at least one turbine disk 23 of the hot gas part 10 has a different material than the other turbine disks 22.Gas turbine according to claim 1, which has a compressor 4 with at least twelve compressor stages, in particular a compressor 4 with fifteen compressor stages, wherein at least the last compressor stage, in particular the two last compressor stages, have a different material for the compressor disks 16 than the front compressor disks 13. Gas turbine according to one or both of claims 1 or 2, wherein the front compressor disks 13 or wherein at least the four compressor disks 13 directly in front of either the last or the penultimate and last compressor disk 16 have a first steel with (in wt. %): carbon (C), in particular 0.20% < C < 0.35%, chromium (Cr), in particular 1.0% < Cr < 2.5%, molybdenum (Mo), in particular 0.20% < Mo < 0.55%, vanadium (V), in particular 0.03% < V < 0.20%, nickel (Ni), in particular 3.0% < Ni < 4.4%, wherein proportions of silicon (Si), manganese (Mn), phosphorus (P), sulfur (S), arsenic (As), antimony (Sb) and / or tin (Sn) are to be kept as small as possible, in particular the first steel consists thereof. Gas turbine according to one or more of claims 1, 2 or 3, in which either the last or the penultimate and last compressor disk 16 have a second steel and, in comparison to the first steel of the other compressor disks 13, have a low proportion of carbon (C), a significantly higher proportion, in particular at least 100% higher, of chromium (Cr), a higher proportion of molybdenum (Mo) and the deliberate addition of niobium (Nb), tungsten (W), nitrogen (N) and manganese (Mn). Gas turbine according to claim 4, wherein the second steel comprises (in wt%): carbon (C), in particular 0.08% < C < 0.19%, chromium (Cr), in particular 9.5% < Cr < 11.0%, molybdenum (Mo), in particular 0.80% < Mo < 1.5%, vanadium (V), in particular 0.10% < V < 0.30%, nickel (Ni), in particular 0.4% < Ni < 1.2%, tungsten (W), in particular 0.7% < W < 1.3%, manganese (Mn), in particular 0.02% < Mn < 0.70%, nitrogen (N), in particular 0.035% < N < 0.070% niobium (Nb), in particular 0.02% < Nb 0.08%, in particular the second steel consists thereof. Gas turbine according to one or more of claims 1, 2, 3, 4 or 5, which has torque disks 19 between the compressor 4 and the hot gas part 10, wherein the at least one torque disk 19, in particular wherein all torque disks 19, have the same material as the last compressor disk 16 of the compressor 4.Gas turbine according to one or more of claims 1, 2, 3, 4, 5 or 6, which has at least 3 stages I, II, III, IV in the hot gas part 10, in particular four stages I, II, III, IV, wherein at least the first two turbine disks 22 of the stages I, II, III, IV, in particular the first three turbine disks 22 of the stages I, II, III, IV, have the same material and are different from the material of the turbine disk 23 of either the third and fourth stage III, IV or the fourth stage IV. Gas turbine according to one or more of claims 1, 2, 3, 4, 5 or 6, which has at least 3 stages I, II, III, IV in the hot gas part 10, in particular four stages I, II, III, IV, wherein all turbine disks 22, 23 of the stages I, II, III, IV have the same material. Gas turbine according to one or more of claims 6, 7 or 8, wherein at least the front turbine disks 22 comprise the same material as the material of the torque disks 19.Gas turbine according to one or more of claims 1, 7, 8 or 9, which has four stages I, II, III, IV in the hot gas part 10, wherein a stage I, II, III, IV consists of stationary turbine blades VI, V2, V3, V4 and unstationary turbine blades B1, B2, B3, B4, wherein the turbine blades VI, B1, V2, B2 of the first two stages I, II and in particular the stationary turbine blades V3 of the third stage III, very particularly the stationary turbine blades V3 and the unstationary turbine blades B3 of the third stage III, have a different material than that of the unstationary turbine blades B3 of the third stage III or at least the turbine blades V4, B4 of the fourth stage IV, optionally wherein the first unstationary turbine blades B1 have an SX or columnar structure.Gas turbine according to claim 10, wherein the turbine blades V4, B4 of the fourth stage IV, optionally also the unsteady turbine blades B3 and / or stationary turbine blades V3 of the third stage III, comprise a nickel-based alloy (in wt%): cobalt (Co), in particular 9.0% - 10%, chromium (Cr), in particular 13.0% - 15.0%, tungsten (W), in particular 3.5% - 4.5%, molybdenum (Mo), in particular 3.7% - 4.3%, aluminum (Al), in particular 2.8% - 3.2%, titanium (Ti), in particular 4.8% - 5.5%, boron (B), in particular 0.01% - 0.2%, zirconium (Zr), in particular 0.02% - 0.10%, carbon (C), in particular 0.15% - 0.40%, preferably no tantalum (Ta), preferably no hafnium (Hf), preferably no niobium (Nb). In particular, it consists of:Gas turbine according to claim 11, wherein the material for at least the turbine blades VI, Bl, V2, B2 of the front stages I, II, in contrast to the material of the turbine blades B4, V4, optionally also in contrast to the turbine blades B3, V3 of the third stage III, comprises tantalum (Ta) and / or hafnium (Hf). Gas turbine according to one or more of claims 1, 10, 11 or 12, which has at least three stages I, II, III, IV in the hot gas part 10, in particular which has four stages I, II, III, IV in the hot gas part 10, wherein a stage I, II, III, IV consists of stationary turbine blades VI, V2, V3, V4 and unstationary turbine blades B1, B2, B3, B4, wherein the turbine blades VI, B1, V2, B2 of the front stages, optionally the turbine blades V3, B3 of the third stage, have a nickel-based alloy (in wt%): cobalt (Co), in particular 9.0% - 10.5%, chromium (Cr), in particular 8.0% - 9.5%, tungsten (W), in particular 9.0% - 10.5%, molybdenum (Mo), in particular 0.2% - 1.0%, aluminum (Al), in particular 5.0% - 6.0%, titanium (Ti), in particular 0.5% - 1.5%, tantalum (Ta), in particular 2.5% - 3.5%, boron (B), in particular 0.01% - 0.25%.

[0166] Zirconium (Zr), especially 0.004% - 0.06%,

[0167] Carbon (C), especially 0.05% - 0.16%,

[0168] Hafnium (Hf), in particular 1.0% - 2.0%, in particular consisting of it. Gas turbine according to one or more of claims 1,

[0169] 10, 11, 12 or 13, which has at least three stages I, II, III, IV in the hot gas part 10, in particular which has four stages I, II, III, IV in the hot gas part 10, wherein a stage I, II, III, IV consists of stationary turbine blades VI, V2, V3, V4 and unstationary turbine blades B1, B2, B3, B4, wherein the turbine blades VI, B1, V2, B2, V3, B3, V4, B4 of all stages I, II, III, IV have a metallic oxidation protection layer, preferably a metallic NiCoCrAlYRe, NiCoCrAlY, NiCoCrAlYTaSi, NiCoCrAlYTaHf, NiCoCrAlYTaSiHf or NiCoCrAlYTa layer, very preferably all turbine blades 25 have the same metallic layer.Gas turbine according to one or more of claims 1 or 10 to 14, which has at least three stages I, II, III, IV in the hot gas part 10, in particular which has four stages I, II, III, IV in the hot gas part 10, wherein a stage I, II, III, IV consists of unsteady turbine blades B1, B2, B3, B4 and stationary turbine blades VI, V2, V3, V4, wherein only the turbine blades of the first stages I,.

[0170] 11, III have a ceramic coating, in particular the first three stages I, II, III have a ceramic coating. Gas turbine according to claim 15, wherein a ceramic coating of the turbine blades of the two first stages I, II: VI, B1, V2, B2 and optionally the turbine blades V3, B3 of the third stage III have the same ceramic coating, in particular a segmented, in particular dense (< 10%), ceramic layer with zirconium oxide as a base, and which in turbine blades V3, B3 of the third stage III have a highly porous, a ceramic layer with a porosity greater than 12%, in particular greater than 18%, preferably with zirconium oxide as a matrix.Gas turbine according to one or more of the preceding claims, in which improved compressor blades 40 are present in at least one compressor stage of the compressor 4, the improved compressor blades 40 being characterized in that a squealer edge 44 does not run on the edge of the blade contour of the blade 43, but is preferably arranged centrally on the blade tip 43, and in particular runs over the entire length of the blade tip 43. Gas turbine according to claim 18, in which only one compressor stage of the compressor 4, in particular a compressor stage between the fourth and eighth compressor stage of the compressor 4, has the improved compressor blade 40 according to claim 18. Gas turbine according to one or more of the preceding claims, which has a combustion chamber 50, which has a tiling 8, which has 8 all-ceramic tiles 52, 53, 54 and metallic heat shield elements 55.Gas turbine according to claim 19, wherein, viewed in the direction of flow of the hot gas in the combustion chamber 50, there is a first ceramic material K1 for all-ceramic tiles 52, then preferably a second ceramic material K2 for all-ceramic tiles 53, then a third ceramic material K3 for all-ceramic tiles 53 and then preferably at the end the metallic heat shields 55 are present. Gas turbine according to claim 20, wherein the first ceramic material K1 comprises a mixture of corundum and mullite in a ratio of 40:60 to 60:40, in particular a half mixture, wherein the optional second ceramic material K2 has a mullite content of 15%-25%, wherein the third ceramic material K3 hardly has any mullite, in particular the content is less than 5%.Gas turbine according to one or more of the preceding claims, which has three bleed pipes 30 which can conduct compressed air from the compressor 4 into an exhaust casing 33. Method for modifying a gas turbine, in which an existing gas turbine is modified so that it has at least one feature of one or more of the preceding claims 1 to 21. Method according to claim 23, in which the installed, completely assembled rotor of an existing gas turbine is replaced by a new, completely assembled rotor 2 with at least one feature of one or more of the preceding claims of the gas turbine 1 to 21. Method according to claim 23, in which a further bleed pipe 30 which conducts air from the compressor 4 into an exhaust casing 33 is installed.Method according to one or more of claims 23, 24 or 25, in which the gas turbine 1 has an exhaust casing 33 with an annular discharge channel 65 in which struts 63 are present, which are modified by a front region of the struts 63 being removed and replaced by a new, modified profile which has less curvature.

Claims

Patent claims 1. Gas turbine (1), at least comprising: a rotor (2), a combustion region (7), a compressor (4) upstream of the combustion region (7) as part of the rotor (2) and a hot gas part (10) downstream of the combustion region (7) as part of the rotor (2), wherein the compressor (4) has compressor disks (13, 16) for compressor blades (40), wherein the hot gas part (10) has turbine disks (22, 23) for turbine blades (25), wherein at least one compressor disk (16) of the compressor (4) has a different material than the other compressor disks (13) and / or at least one turbine disk (23) of the hot gas part (10) has a different material than the other turbine disks (22).

2. Gas turbine according to claim 1, which has a compressor (4) with at least twelve compressor stages, in particular a compressor (4) with fifteen compressor stages, wherein at least the last compressor stage, in particular the two last compressor stages, have a different material for the compressor disks (16) than the front compressor disks (13).

3. Gas turbine according to one or both of claims 1 or 2, wherein the front compressor discs (13) or wherein at least the four compressor discs (13) are arranged directly in front of either the last or the second to last and last compressor disc (16) comprise a first steel with (in wt.%): carbon (C), in particular 0.20% < C < 0.35%, chromium (Cr), in particular 1.0% < Cr < 2.5%, molybdenum (Mo), in particular 0.20% < Mo < 0.55%, vanadium (V), in particular 0.03% < V < 0.20%, nickel (Ni), in particular 3.0% < Ni < 4.4%, wherein proportions of silicon (Si), manganese (Mn), phosphorus (P), sulphur (S), arsenic (As), antimony (Sb) and / or tin (Sn) are to be kept as small as possible, in particular the first steel consists thereof.

4. Gas turbine according to one or more of claims 1, 2 or 3, in which either the last or the penultimate and the last compressor disk (16) have a second steel and, compared to the first steel of the other compressor disks (13), has a low proportion of carbon (C), a significantly higher proportion, in particular at least 100% higher, proportion of chromium (Cr), a higher proportion of molybdenum (Mo) and the deliberate addition of niobium (Nb), tungsten (W), nitrogen (N) and manganese (Mn).

5. Gas turbine according to claim 4, wherein the second steel comprises (in wt.%): carbon (C), in particular 0.08% < C < 0.19%, chromium (Cr), in particular 9.5% < Cr < 11.0%, Molybdenum (Mo), in particular 0.80% < Mo < 1.5%, vanadium (V), in particular 0.10% < V < 0.30%, nickel (Ni), in particular 0.4% < Ni < 1.2%, tungsten (W), in particular 0.7% < W < 1.3%, manganese (Mn), in particular 0.02% < Mn < 0.70%, nitrogen (N), in particular 0.035% < N < 0.070% niobium (Nb), in particular 0.02% < Nb 0.08%, in particular the second steel consists thereof.

6. Gas turbine according to one or more of claims 1, 2, 3, 4 or 5, which has torque disks (19) between the compressor (4) and the hot gas part (10), wherein the at least one torque disk (19), in particular wherein all torque disks (19), have the same material as the last compressor disk (16) of the compressor (4).

7. Gas turbine according to one or more of claims 1, 2, 3, 4, 5 or 6, which has at least 3 stages (I, II, III, IV) in the hot gas part (10), in particular has four stages (I, II, III, IV) in the hot gas part (10), wherein at least the first two turbine disks (22) of the stages (I, II, III, IV), in particular the first three turbine disks (22) of the stages (I, II, III, IV) have the same material and are different from the material of the turbine disk either the third and fourth stages (III, IV) or the fourth stage (IV).

8. Gas turbine according to one or more of claims 1, 2, 3, 4, 5 or 6, which has at least 3 stages (I, II, III, IV) in the hot gas part (10), in particular has four stages (I, II, III, IV) in the hot gas part (10), wherein all turbine disks (22, 23) of the stages (I, II, III, IV) have the same material.

9. Gas turbine according to one or more of claims 6, 7 or 8, wherein at least the front turbine disks (22) have the same material as the material of the torque disks (19).

10. Gas turbine according to one or more of claims 1, 7, 8 or 9, which has at least three stages (I, II, III) in the hot gas part (10), in particular has four stages (I, II, III, IV) in the hot gas part (10), wherein one stage (I, II, III, IV) consists of stationary turbine blades (VI, V2, V3, V4) and non-stationary turbine blades (B1, B2, B3, B4), wherein the turbine blades (VI, B1, V2, B2) of the first two stages (I, II) and in particular the stationary turbine blades (V3) of the third stage (III), very particularly the stationary turbine blades (V3) and the non-stationary turbine blades (B3) of the third stage (III), have a different material than those of the non-stationary turbine blades (B3) of the third stage (III). or at least the turbine blades (V4, B4) of the fourth stage (IV), optionally wherein the first unsteady turbine blades (Bl) have an SX or columnar structure.

11. Gas turbine according to claim 10, wherein the turbine blades (V4, B4) of the fourth stage (IV), optionally also the unsteady turbine blades (B3) and / or stationary turbine blades (V3) of the third stage (III), comprise a nickel-based alloy (in wt%): cobalt (Co), in particular 9.0% - 10%, chromium (Cr), in particular 13.0% - 15.0%, tungsten (W), in particular 3.5% - 4.5%, molybdenum (Mo), in particular 3.7% - 4.3%, aluminum (Al), in particular 2.8% - 3.2%, titanium (Ti), in particular 4.8% - 5.5%, boron (B), in particular 0.01% - 0.2%, zirconium (Zr), in particular 0.02% - 0.10%, carbon (C), in particular 0.15% - 0.40%, preferably no tantalum (Ta), preferably no hafnium (Hf), preferably no niobium (Nb). in particular consists of this.

12. Gas turbine according to claim 11, wherein the material for at least the turbine blades (VI, Bl, V2, B2) of the front stages (I, II) in contrast to the material of the turbine blades (B4, V4), optionally also in contrast to the turbine blades (B3, V3) of the third stage (III), but has tantalum (Ta) and / or hafnium (Hf).

13. Gas turbine according to one or more of claims 1, 10, 11 or 12, which has at least three stages (I, II, III, IV) in the hot gas part (10), in particular which has four stages (I, 11, III, IV), wherein a stage (I, II, III, IV) consists of stationary turbine blades (VI, V2, V3, V4) and unstationary turbine blades (Bl, B2, B3, B4), wherein the turbine blades (VI, Bl, V2, B2) of the front stages, optionally the turbine blades (V3, B3) of the third stage, comprise a nickel-based alloy (in wt%): cobalt (Co), in particular 9.0% - 10.5%, chromium (Cr), in particular 8.0% - 9.5%, tungsten (W), in particular 9.0% - 10.5%, molybdenum (Mo), in particular 0.2% - 1.0%, aluminum (Al), in particular 5.0% - 6.0%, titanium (Ti), in particular 0.5% - 1.5%, tantalum (Ta) , especially 2.5% - 3.5%, boron (B) , in particular 0.01% - 0.25%, zirconium (Zr), in particular 0.004% - 0.06%, carbon (C), in particular 0.05% - 0.16%, hafnium (Hf), in particular 1.0% - 2.0%, in particular consisting thereof.

14. Gas turbine according to one or more of claims 1, 10, 11, 12 or 13, which has at least three stages (I, II, III) in the hot gas part (10), in particular which has four stages (I, 11, III, IV), wherein a stage (I, II, III, IV) consists of stationary turbine blades (VI, V2, V3, V4) and unstationary turbine blades (Bl, B2, B3, B4), wherein the turbine blades (VI, Bl, V2, B2, V3, B3, V4, B4) of all stages (I, II, III, IV) have a metallic oxidation protection layer, preferably a metallic NiCoCrAlYRe, NiCoCrAlY, NiCoCrAlYTaSi, NiCoCrAlYTaHf, NiCoCrAlYTaSiHf or NiCoCrAlYTa layer, very preferably all turbine blades (25) have the same metallic layer.

15. Gas turbine according to one or more of claims 1 or 10 to 14, which has at least three stages (I, II, III) in the hot gas part (10), in particular which has four stages (I, II, III, IV) in the hot gas part (10), wherein one stage (I, II, III, IV) consists of unsteady (Bl, B2, B3, B4) and stationary turbine blades (VI, V2, V3, V4), whereby only the turbine blades of the front stages (I, II, III) have a ceramic coating, In particular, only the turbine blades of the first three stages (I, II, III) have a ceramic coating.

16. Gas turbine according to claim 15, in which a ceramic coating of the turbine blades of the two first stages (I, II: VI, B1, V2, B2) and optionally of the turbine blades (V3, B3) of the third stage (III) have the same ceramic coating, in particular a segmented, in particular dense (< 10%), ceramic layer with zirconium oxide as a base, and / or which in turbine blades (V3, B3) of the third stage (III) have a highly porous, a ceramic layer with a porosity greater than 12%, in particular greater than 18%, preferably with zirconium oxide as a matrix.

17. Gas turbine according to one or more of the preceding claims, in which improved compressor blades (40) are present in at least one compressor stage of the compressor (4), wherein the improved compressor blades (40) are characterized in that a squealer edge (44) does not run along the edge of the blade contour of the blade (43), but rather the squealer edge (44) is preferably arranged centrally on the blade tip (43), and in particular runs over the entire length of the blade tip (43).

18. Gas turbine according to claim 18, in which only one compressor stage of the compressor (4), in particular a compressor stage between the fourth and eighth compressor stage of the compressor (4), the improved compressor blade (40) according to claim 18.

19. Gas turbine according to one or more of the preceding claims, which has a combustion chamber (50) which (50) has a tiling (8) of tiles arranged in rows, which (8) has all-ceramic tiles (52, 53, 54) and metallic heat shield elements (55).

20. Gas turbine according to claim 19, wherein, viewed in the flow direction of the hot gas in the combustion chamber (50), a first ceramic material (K1) for all-ceramic tiles (52), then preferably a second ceramic material (K2) for all-ceramic tiles (53), then a third ceramic material (K3) for all-ceramic tiles (53) and then preferably at the end the metallic heat shields (55) are present.

21. Gas turbine according to claim 20, wherein the first ceramic material (K1) comprises a mixture of corundum and mullite in a ratio of 40:60 to 60:40, in particular a half mixture, wherein the optional second ceramic material (K2) has a proportion of mullite of 15%-25%, wherein the third ceramic material (K3) hardly has any mullite, in particular the proportion is less than 5%.

22. Gas turbine according to one or more of the preceding claims, which has three discharge pipes (30) which discharge compressed air from the compressor (4) into a discharge gas housing (33).

23. A method for modifying a gas turbine, in which an existing gas turbine is modified so that it has at least one feature of one or more of the preceding claims 1 to 21.

24. Method according to claim 23, wherein the installed, completely assembled rotor of an existing gas turbine is replaced by a new, completely assembled rotor (2) having at least one feature of one or more of the preceding claims of the gas turbine 1 to 21.

25. A method according to claim 23 or 24, wherein a further discharge pipe (30) is installed which directs air from the compressor (4) into an exhaust housing (33).

26. Method according to one or more of claims 23, 24 or 25, in which the gas turbine (1) has an exhaust casing (33) with an annular discharge channel (65) in which struts (63) are present which are modified by removing a front region of the struts (63) and replacing it with a new, modified profile which has less curvature.

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