Gas turbine and method for improving gas turbine

Abstract

The invention relates to a gas turbine (1) having at least: a rotor (2); a combustion region (7); a compressor (4) upstream of the combustion region (7) as part of the rotor (2); the rotor (2) has a combustion region (7), and a hot gas section (10) downstream of the combustion region (7) as part of the rotor (2), the compressor (4) has compressor discs for compressor blades, the hot gas section (10) has turbine discs for turbine blades, at least one compressor disc of the compressor (4) has a different material than the remaining compressor discs, and the rotor (2) has a combustion region (7) downstream of the combustion region (7). And / or at least one turbine disk of the hot gas section (10) has a different material than the remaining turbine disks.

Description

Technical Field

[0001] This invention relates to an improved gas turbine and a method for improving the gas turbine. Background Technology

[0002] The task facing those skilled in the art is to improve the efficiency and lifespan of gas turbines. This can be achieved by improving the compressor, combustion zone, and hot air section. One feasible approach is to increase the degree of air compression through an improved compressor. However, this would affect all components of the gas turbine. Summary of the Invention

[0003] Therefore, the object of the present invention is to disclose changes made that withstand or adapt to the effects of improved compressors on gas turbines.

[0004] This task is accomplished by the gas turbine according to claim 1 and the method according to claim 23.

[0005] Further advantageous measures are listed in the dependent claims, which can be combined in any known manner to achieve additional advantages. Attached Figure Description

[0006] As shown in the attached figure: Figure 1 A gas turbine is schematically shown; Figure 2 The compressor blades are shown; Figure 3 The combustion chamber is shown. Figure 4 , Figure 5 The exhaust casing is shown, and Figure 6 The assembled rotor is shown. Detailed Implementation

[0007] The specification and accompanying drawings only illustrate embodiments of the present invention.

[0008] This task is achieved by a gas turbine 1, which has at least: a rotor 2, a combustion zone 7, a compressor 4 located upstream of the combustion zone 7 as part of the rotor 2, and a hot gas section 10 located downstream of the combustion zone 7 as part of the rotor 2, wherein the compressor 4 has compressor disks 13, 16 for compressor blades 40, wherein the hot gas section 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 from the other compressor disks 13, and / or at least one turbine disk 23 of the hot gas section 10 has a different material from the other turbine disks 22.

[0009] An advantageous design is that the gas turbine has a compressor 4 with at least twelve compressor stages, particularly a compressor 4 with fifteen compressor stages, wherein at least the last compressor stage, and especially the last two compressor stages, have compressor disks 16 made of a different material than the earlier compressor disks 13.

[0010] An advantageous design for the gas turbine is that the leading compressor disc 13, or at least four compressor discs 13 immediately preceding the last compressor disc or immediately preceding the penultimate compressor disc and the last compressor disc 16, has a first steel component having (in weight %): Carbon (C), especially 0.20% < C < 0.35%, Chromium (Cr), especially 1.0% < Cr < 2.5%, Molybdenum (Mo), especially 0.20%<Mo<0.55%, Vanadium (V), especially 0.03% < V < 0.20%, Nickel (Ni), especially 3.0% < Ni < 4.4%, The proportions of silicon (Si), manganese (Mn), phosphorus (P), sulfur (S), arsenic (As), antimony (Sb) and / or tin (Sn) are kept as small as possible, and in particular, the first steel is composed of the above-mentioned components.

[0011] An advantageous design for the gas turbine is that the last compressor disc or the penultimate compressor disc and the last compressor disc 16 have a second steel, and the second steel has a lower share of carbon (C), a significantly higher share, particularly at least 100% higher share of chromium (Cr), a higher share of molybdenum (Mo), and intentionally introduced niobium (Nb), tungsten (W), nitrogen (N) and manganese (Mn) compared to the first steel of the other compressor discs 13.

[0012] An advantageous design for a gas turbine is that the second steel component has the following characteristics (in weight %): Carbon (C), especially 0.08% < C < 0.19%, Chromium (Cr), especially 9.5% < Cr < 11.0%, Molybdenum (Mo), especially 0.80%<Mo<1.5%, Vanadium (V), especially 0.10% < V < 0.30%, Nickel (Ni), especially 0.4% < Ni < 1.2%, Tungsten (W), especially 0.7% < W < 1.3%, Manganese (Mn), especially 0.02% < Mn < 0.70%, Nitrogen (N), especially 0.035% < N < 0.070%, Niobium (Nb), especially 0.02% < Nb < 0.08%, Specifically, the second steel is composed of the aforementioned components.

[0013] One advantageous design is that the gas turbine has multiple torque discs 19 between the compressor 4 and the hot gas section 10, wherein at least one torque disc 19, and particularly all of the torque discs 19, are made of the same material as the last compressor disc 16 of the compressor 4.

[0014] An advantageous design is that the gas turbine has at least three stages I, II, III, IV in the hot gas section 10, and particularly at least four stages I, II, III, IV, wherein at least the first two turbine disks 22 of stages I, II, III, IV, and especially the first three turbine disks 22 of stages I, II, III, IV, are made of the same material and are made of a different material than the turbine disks 23 of the third and fourth stages III and IV, or different material than the turbine disk of the final fourth stage IV, depending on which stages I, II, and optionally III have the same material.

[0015] An advantageous design is that the gas turbine has at least three stages I, II, III, IV in the hot gas section 10, and particularly at least four stages I, II, III, IV, wherein all turbine disks 22, 23 of stages I, II, III, IV are made of the same material.

[0016] An advantageous design for a gas turbine is that at least the forward turbine disk 22 is made of the same material as the torque disk 19.

[0017] An advantageous design is that the gas turbine has four stages I, II, III, and IV in the hot gas section 10, wherein stages I, II, III, and IV are composed of stationary turbine blades V1, V2, V3, and V4 and non-stationary turbine blades B1, B2, B3, and B4. The turbine blades V1, B1, V2, and B2 of the first two stages I and II, and especially the stationary turbine blade V3 of the third stage III, and particularly the stationary turbine blade V3 and non-stationary turbine blade B3 of the third stage III, have either a material different from the non-stationary turbine blade B3 of the third stage III, or a material different from at least one turbine blade V4, B4 of the plurality of turbine blades in the fourth stage IV, depending on the material selection of the preceding stages.

[0018] One advantageous design is that the gas turbine has an SX structure or a columnar structure for the first non-stationary turbine blade B1.

[0019] An advantageous design for the gas turbine is that the turbine blades V4 and B4 of the fourth stage IV, and optionally the non-stationary turbine blade B3 and / or the stationary turbine blade V3 of the third stage III, are made of a nickel-based alloy having a composition in weight percent: Cobalt (Co), especially 9.0%–10%, Chromium (Cr), especially 13.0%–15.0%, Tungsten (W), especially 3.5%–4.5%, Molybdenum (Mo), especially 3.7%–4.3%, Aluminum (Al), especially 2.8%–3.2%, Titanium (Ti), especially 4.8%–5.5%, Boron (B), especially 0.01%–0.2%, Zirconium (Zr), especially 0.02%–0.10%, Carbon (C), especially 0.15%–0.40%, Preferably, it does not contain tantalum (Ta), preferably it does not contain hafnium (Hf), and preferably it does not contain niobium (Nb). In particular, the nickel-based alloy is composed of the above-mentioned components.

[0020] An advantageous design for a gas turbine is that the turbine blades V1, B1, V2, B2 of at least the first two stages are made of a different material than the turbine blades B4, V4, and optionally also different from the turbine blades B3, V3 of the third stage III, but instead have tantalum (Ta) and / or hafnium (Hf).

[0021] An advantageous design is that the gas turbine has at least three stages I, II, III, and IV in the hot gas section 10, and more particularly, at least four stages I, II, III, and IV in the hot gas section 10, wherein stages I, II, III, and IV are composed of stationary turbine blades V1, V2, V3, and V4 and non-stationary turbine blades B1, B2, B3, and B4, wherein the turbine blades V1, B1, V2, and B2 of the earlier stages, and optionally the turbine blades V3 and B3 of the third stage, are made of a nickel-based alloy, which has (in weight %): Cobalt (Co), especially 9.0%–10.5%, Chromium (Cr), especially 8.0%–9.5%, Tungsten (W), especially 9.0%–10.5%, Molybdenum (Mo), especially 0.2%–1.0%, Aluminum (Al), especially 5.0%–6.0%, Titanium (Ti), especially 0.5%–1.5%, Tantalum (Ta), especially 2.5%–3.5%, Boron (B), especially 0.01%–0.25%, Zirconium (Zr), especially 0.004%–0.06%, Carbon (C), especially 0.05%–0.16%, Hafnium (Hf), especially 1.0%–2.0%, Specifically, it is composed of the above-mentioned components.

[0022] An advantageous design is that the gas turbine has at least three stages I, II, III, and IV in the hot gas section 10, and more particularly, at least four stages I, II, III, and IV in the hot gas section 10, wherein stages I, II, III, and IV are composed of stationary turbine blades V1, V2, V3, and V4 and non-stationary turbine blades B1, B2, B3, and B4, wherein all turbine blades V1, B1, V2, B2, V3, B3, V4, and B4 of stages I, II, III, and IV have a metal oxide protective layer, preferably a metallic NiCoCrAlYRe layer, NiCoCrAlY layer, NiCoCrAlYTaSi layer, NiCoCrAlYTaHf layer, NiCoCrAlYTaSiHf layer, or NiCoCrAlYTa layer, and most preferably, all turbine blades 25 have the same metallic layer.

[0023] An advantageous design is that the gas turbine has at least three stages I, II, III, and IV in the hot gas section 10, and more particularly, at least four stages I, II, III, and IV in the hot gas section 10, wherein stages I, II, III, and IV are composed of non-stationary turbine blades B1, B2, B3, and B4 and stationary turbine blades V1, V2, V3, and V4, wherein only the turbine blades of the first few stages I, II, and III have ceramic coatings, and more particularly, the first three stages I, II, and III have ceramic coatings.

[0024] An advantageous design for a gas turbine is that the ceramic coatings of the turbine blades I, II: V1, B1, V2, B2 of the first two stages and optionally the ceramic coatings of the turbine blades V3, B3 of the third stage III have the same ceramic coating, particularly having segmented, especially dense (<10%), zirconium oxide-based ceramic layers.

[0025] One advantageous design is that the gas turbine has a super-porous ceramic layer in the turbine blades V3, B3 of the third stage III, which has a porosity of greater than 12%, particularly greater than 18%, and is specifically based on zirconium oxide.

[0026] An advantageous design of the gas turbine is that an improved compressor blade 40 is present in at least one compressor stage of the compressor 4, wherein the improved compressor blade 40 is characterized in that the skimming edge 44 does not extend at the edge of the blade body profile of the blade body 43, but is preferably arranged at the middle of the blade tip 43, and particularly extends across the entire length of the blade tip 43.

[0027] An advantageous design for a gas turbine is that the compressor 4 has only one compressor stage, particularly the compressor stage between the fourth and eighth compressor stages of the compressor 4, having improved compressor blades 40 according to claim 18.

[0028] One advantageous design is that the gas turbine has a combustion chamber 50, which has a tiled section 8, which has all-ceramic tiles 52, 53, 54 and a metal thermal barrier element 55.

[0029] An advantageous design of the gas turbine is that, viewed along the flow direction of the hot gas in the combustion chamber 50, there is a first ceramic material K1 for the all-ceramic tile 52, then preferably a second ceramic material K2 for the all-ceramic tile 53, then a third ceramic material K3 for the all-ceramic tile 53, and then preferably a metal thermal barrier 55 is present at the end, i.e., only the first and third materials can be used.

[0030] An advantageous design for a gas turbine is wherein the first ceramic material K1 is a mixture of corundum and mullite in a ratio of 40:60 to 60:40, particularly a mixture of equal parts; wherein the optional second ceramic material K2 has a mullite content of 15%–25%; and wherein the third ceramic material K3 contains almost no mullite, particularly less than 5%.

[0031] One advantageous design is that the gas turbine has three exhaust pipes 30 that can guide compressed air from the compressor 4 into the exhaust casing 33.

[0032] An advantageous method for modifying a gas turbine so that the gas turbine has at least one feature according to any one or more of the preceding claims.

[0033] This method is further advantageous in that the existing gas turbine's installed, fully assembled rotor is replaced by a new, fully assembled rotor (2) having at least one of the aforementioned features.

[0034] This method is further advantageous in that it includes an additional discharge pipe 30 that directs air from the compressor 4 into the exhaust housing 33.

[0035] Furthermore, this method is advantageous in that the gas turbine 1 has an exhaust casing 33 with an annular exhaust channel 65 in which a support strip 63 is present, which is modified by removing the forward region of the support strip 63 and replacing it with a new, modified profile with a smaller curvature.

[0036] The gas turbine 1 has at least: a rotor 2, a compressor 4, a downstream combustion zone 7, and a downstream hot gas section 10.

[0037] The compressor 4 has compressor disks 13, 16 for each compressor stage, and these compressor stages have corresponding compressor blades. The gas turbine 1 has at least twelve compressor stages, and particularly fifteen compressor stages, for the compressor 4.

[0038] Each stage of compressor 4 consists of guide vanes and rotor vanes in the corresponding compressor disks 13 and 16.

[0039] The compressor 4 was designed and configured to increase the thermal and mechanical loads of at least the last compressor stage, and especially the last two compressor stages.

[0040] Therefore, the material of at least the last compressor disk 16 of compressor 4 is preferably different from the material of the other preceding compressor disks 13, or different from the material of at least four compressor disks 13 up to the penultimate compressor disk and / or the last compressor disk 16. Preferably, the material of the last two compressor disks 16 of compressor 4 is preferably different from the material of the other preceding compressor disks 13, or different from the material of at least four compressor disks 13 up to the penultimate compressor disk and / or the last compressor disk 16.

[0041] Strength can be increased, especially at higher temperatures, by using different materials. This is achieved through alloy modification and / or heat treatment.

[0042] For steel, or generally, “different” or “different from” in this application means that at least one alloying element is present in greater or lesser quantities, or that the proportion of the alloying element is higher or lower, at least 10%, and particularly at least 20%.

[0043] Steel is always an alloy consisting of at least iron and carbon.

[0044] Different materials are intentionally used for compressor discs 13 and 16, wherein the last compressor disc 16 or the last two compressor discs 16 have a higher tensile strength, which is preferably 50 MPa higher or more preferably 100 MPa higher than the tensile strength of the preceding compressor disc 13.

[0045] Examples of compressor 4 having fifteen stages (a), b), or examples of compressor 4 having thirteen stages (c), d): a) Compressor disc 15: Second steel material; Compressor discs 11 to 14: First steel material; Compressor discs 1 to 10: First steel material or steel material different from the first and second steel materials. b) Compressor discs fourteen and fifteen: second type of steel; compressor discs ten to thirteen: first type of steel; compressor discs one to nine: first type of steel or different from first and second type of steel. c) Thirteenth compressor disc: second steel material; Ninth to twelfth compressor discs: first steel material; First to eighth compressor discs: first steel material or different from first and second steel materials. d) Compressor discs of the twelfth and thirteenth: second steel material; compressor discs of the ninth to eleventh: first steel material; compressor discs of the first to eighth: first steel material or different from the first and second steel materials.

[0046] The material of at least four compressor disks 13, up to the first compressor disk 13 or the penultimate compressor disk and / or the last compressor disk 16, preferably has a first steel (NiCrMoV), which has (in weight %): Carbon (C), especially 0.20% < C < 0.35%, Chromium (Cr), especially 1.0% < Cr < 2.5%, Molybdenum (Mo), especially 0.20% < Mo < 0.55%, and Vanadium (V), especially 0.03% < V < 0.20%, and Nickel (Ni), especially 3.0% < Ni < 4.4%, The proportions of silicon (Si), manganese (Mn), phosphorus (P), sulfur (S), arsenic (As), antimony (Sb), and tin (Sn) are kept as small as possible. Specifically, the first type of steel is composed of the above components. The tensile strength is 810 MPa–960 MPa.

[0047] The second steel differs from the first steel and, compared to the first steel, preferably has a lower proportion of carbon (C), a significantly higher proportion, particularly at least 100% higher, of chromium (Cr), a higher proportion of molybdenum (Mo), and intentionally introduced niobium (Nb), tungsten (W), nitrogen (N), and manganese (Mn). Increased strength, especially at higher temperatures, is achieved through the different materials. This is accomplished through alloy modification.

[0048] At least the last compressor disc 16 and torque disc 19 have a second steel (CrMoWVNbN) (in weight %): Carbon (C), especially 0.08% < C < 0.19%, Chromium (Cr), especially 9.5% < Cr < 11.0%, Molybdenum (Mo), especially 0.80%<Mo<1.5%, Vanadium (V), especially 0.10% < V < 0.30%, Nickel (Ni), especially 0.4% < Ni < 1.2%, Tungsten (W), especially 0.7% < W < 1.3%, Manganese (Mn), especially 0.2% < Mn < 0.7%, Nitrogen (N), especially 0.035% < N < 0.07%, Niobium (Nb), especially 0.02% < Nb < 0.08%. Specifically, the second steel is composed of the above-mentioned components.

[0049] For steel, or generally, “different” or “different from” in this application means that at least one alloying element is present in greater or lesser quantities, or that the proportion of the alloying element is higher or lower, at least 10%, and particularly at least 20%.

[0050] In the combustion zone 7, that is, between the compressor 4 and the hot gas section 10, there is a torque disc 19, which connects the portion of the rotor 2 of the compressor 4 to the hot gas section 10.

[0051] Preferably, three torque discs 19 are used.

[0052] Preferably, the torque disks, especially all torque disks 19, use the same material as the last compressor disk 16 or the last two compressor disks 16.

[0053] Therefore, at least the last compressor disc 16 and the torque disc 19 preferably also have the second steel (CrMoWVNbN).

[0054] The hot air section 10 has at least three levels, preferably four levels: I, II, III, and IV.

[0055] Each stage I, II, III, and IV of the hot gas section 10 consists of turbine blades 25, stationary guide vanes V, and non-stationary rotor blades B. The turbine blades 25 are arranged in the forward turbine disk 22 and the rearward turbine disk 23.

[0056] The turbine disks 22 used for at least the first two turbine stages, and especially for the first three turbine stages, are made of the same material, and preferably of a different material than the turbine disks of the fourth stage, in order to withstand the increased thermal loads.

[0057] The material used for torque disc 19 is preferably also used for the first, first two, or first three turbine discs 22. The material used for the first, first two, or first three turbine discs 22 is preferably different from the material used for the later stages III and IV.

[0058] The material used for the rear turbine disk 23 of stages III and IV, or only for the rear turbine disk 23 of stage IV, preferably has a first steel, but with a higher strength grade due to different heat treatment.

[0059] For steel, nickel-based superalloys, or generally, "different from" or "different from" in this application means that at least one alloying element is present in greater or lesser quantities, or that the proportion of the alloying element is higher or lower, at least 10%, and particularly at least 20%.

[0060] Also preferably, the foremost turbine disk 22 may be made of the same material as one or more rearmost turbine disks 23.

[0061] Therefore, the same materials, especially the second steel, can be used for the compressor disk 16, the torque disk, and all the turbine disks 22 and 23.

[0062] In the hot gas section 10, the gas turbine has at least three stages, particularly four stages I, II, III, IV, wherein each stage has stationary guide vanes V1, V2, V3, V4 and non-stationary rotating vanes B1, B2, B3, B4, wherein the turbine blades 25 of the first two stages V1, B1, V2, B2 and optionally the stationary turbine blade V3 or the stationary turbine blade V3 and non-stationary turbine blade B3 of the third stage III have different casting materials, which are different from the casting materials of the turbine blades of the third stage III and the fourth stage IV, or different from the casting materials of the non-stationary turbine blade B3 of the third stage and the casting materials of the turbine blades V4 and B4 of the fourth stage IV.

[0063] The example is: a) V1, B1, V2, and B2 are made of different materials than V3, B3, V4, and B4, wherein V1, B1, V2, and B2 are made of the same material, and V3, B3, V4, and B4 preferably are made of the same material. b) V1, B1, V2, B2, and V3 are made of different materials than B3, V4, and B4, wherein V1, B1, V2, B2, and V3 are made of the same material, and B3, V4, and B4 preferably are made of the same material. c) V1, B1, V2, B2, V3, B3 have materials different from V4, B4, wherein V1, B1, V2, B2, V3, B3 have the same material, and V4, B4 preferably have the same material.

[0064] The design depends primarily, but not solely, on the application conditions of the gas turbine 1 or the design of the compressor 4.

[0065] Different materials are defined as materials that are significantly different in the content of carbon (C), chromium (Cr), molybdenum (Mo), and titanium (Ti), with a difference of at least 20%, and significantly different in the content of tungsten (W) and aluminum (Al), with a difference of at least 50%, and the first-grade materials additionally contain tantalum (Ta) and hafnium (Hf).

[0066] Turbine blades V4, B4, and possibly B3, V3, are made of nickel-based alloys (in weight percent): Cobalt (Co), especially 9.0%–10%, Chromium (Cr), especially 13.0%–15.0%, Tungsten (W), especially 3.5%–4.5%, Molybdenum (Mo), especially 3.7%–4.3%, Aluminum (Al), especially 2.8%–3.2%, Titanium (Ti), especially 4.8%–5.5%, Boron (B), especially 0.01%–0.2%, Zirconium (Zr), especially 0.02%–0.10%, Carbon (C), especially 0.15%–0.40%, Preferably, it does not contain tantalum (Ta), preferably does not contain hafnium (Hf), and preferably does not contain niobium (Nb). In particular, the nickel-based alloy is composed of the above-mentioned components.

[0067] However, unlike the materials of B4, V4, B3, and V3, at least the earlier grades V1, B1, V2, and B2 contain tantalum (Ta) and hafnium (Hf).

[0068] Turbine blades V1, B1, V2, B2, and possibly B3, V3 are made of a nickel-based alloy, said nickel-based alloy having (in weight %): Cobalt (Co), especially 9.0%–10.5%, Chromium (Cr), especially 8.0%–9.5%, Tungsten (W), especially 9.0%–10.5%, Molybdenum (Mo), especially 0.2%–1.0%, Aluminum (Al), especially 5.0%–6.0%, Titanium (Ti), especially 0.5%–1.5%, Tantalum (Ta), especially 2.5%–3.5%, Boron (B), especially 0.01%–0.25%, Zirconium (Zr), especially 0.004%–0.06%, Carbon (C), especially 0.05%–0.16%, Hafnium (Hf), especially 1.0%–2.0%. Specifically, the nickel-based alloy is composed of the above-mentioned components.

[0069] The rotating blade B1 preferably has an oriented structure. The rotating blade is not cast, but rather oriented and solidified solely by being guided by a known temperature. Therefore, it has a single-crystal (SX) or columnar (säulenartig) structure. This can be achieved by using the same material as the turbine blades V1, V2, or B2.

[0070] Here, it is preferred that all existing turbine blades of stages I, II, III, and IV 25 V1, B1, V2, B2, V3, B3, V4, and B4 have an oxide protective layer. This is preferably a metallic NiCoCrAlYRe layer, NiCoCrAlY layer, NiCoCrAlYTaSi layer, NiCoCrAlYTaHf layer, NiCoCrAlYTaSiHf layer, or NiCoCrAlYTa (=MCrAlY) layer.

[0071] At least the first two stages I and II of the hot gas section 10 have ceramic coatings on the turbine blades 25, namely turbine blades V3, B3, V4, and B4 do not have ceramic coatings (except for NiCoCrAlYRe, NiCoCrAlY, NiCoCrAlYTaSi, NiCoCrAlYTaHf, NiCoCrAlYTaSiHf, or NiCoCrAlYTaTGO).

[0072] Preferably, the turbine blades 25 of the first two stages I and II, as well as at least the non-stationary turbine blade V3, have a ceramic coating, i.e., turbine blades V4, B4, and optionally B3 do not have a ceramic coating (except for the TGO of MCrAlY).

[0073] The example is: V1, B1, V2, B2, and V3 have a ceramic coating, while B3, V4, and B4 do not. V1, B1, V2, and B2 have a ceramic coating, while V3, B3, V4, and B4 do not. V1, B1, V2, B2, V3, B3, and V4 have a ceramic coating, while B4 does not. V1, B1, V2, B2, V3, and B3 have a ceramic coating, while V4 and B4 do not.

[0074] Also preferably, the first ceramic coatings of the first two stages V1, B1, V2, and B2, and optionally the turbine blades V3 of the third stage III, may have the same ceramic coating. This is preferably a segmented, particularly dense (<10%) segmented ceramic layer, preferably based on zirconium oxide.

[0075] The non-stationary turbine blade B3, and optionally the stationary turbine blade V3 of the third stage, preferably have a second ceramic coating that differs from that of the turbine blades of the preceding stages V1, B1, V2, B2, and optionally V3. This is preferably a highly porous ceramic layer with a porosity of >12%, particularly >18%. Preferably, this porous layer has zirconium oxide, particularly in the form of partially stabilized zirconium oxide (PSZ).

[0076] according to Figure 1 The gas turbine 1 has an additional blow-off line 30. This blow-off line 30 draws compressed air from the compressor 4, particularly from the fifth or sixth stage of the compressor 4, and is preferably only needed when starting the gas turbine 1. It connects the compressor 4 to the exhaust casing 33. The gas turbine 1 preferably already has two blow-off lines, and may have three blow-off lines 30 after modification or as a new component.

[0077] Figure 2 The improved compressor blade 40 of the compressor blade 4 is shown. The compressor blade 40 has a flat profile 41 in a known manner. A blade body 42 extends from the flat profile 41 of the compressor blade 40 to the blade tip 43. The blade tip 43 is preferably planar and has a preferably single frontal edge 44 extending from the blade tip 43. The improved compressor blade 40 has a frontal edge 44 at the blade tip 43 that does not extend along the edge of the blade body profile, but rather preferably extends centrally along the blade tip 43, and particularly across the entire length of the blade tip 43. This is also derived from the improved design of the compressor 4.

[0078] Preferably, only one compressor stage of compressor 4, particularly the compressor stage between the fourth and eighth compressor stages of compressor 4, has improved compressor blades 40.

[0079] Figure 3 Combustion chamber 50 of combustion zone 7 is shown. In combustion chamber 50, highly compressed air from compressor 4 mixes with fuel and burns. For this purpose, a burner system 51 is provided. The resulting hot gas then flows into hot gas section 10.

[0080] The combustion chamber 50 has a tiled portion 8, and is therefore preferably provided with a ceramic heat barrier 52, which is preferably made entirely of ceramic. For the all-ceramic design, preferably at least two different, preferably three different materials are used.

[0081] In the foreground region, at the beginning of the burner system 51, the ceramic thermal barrier 52 has a first material K1, while near the outflow region, the ceramic thermal barrier 54 has a third material K3 that is corrosion-resistant to all ceramic components. In between, for the intermediate ceramic thermal barrier 53, a second ceramic material is preferably used for the tile section 8.

[0082] For all-ceramic thermal barriers, "different" means that the microstructure is clearly distinguishable, and / or the chemical composition differs by at least 10%, particularly by at least 20%, in at least one component, or that the ceramic composition contains more or less of at least one chemical element.

[0083] The first ceramic material K1 used for the ceramic thermal barrier 52 in the foreground area is preferably a mixture of corundum and mullite in a ratio of 40:60 to 60:40, especially in a 50 / 50 ratio.

[0084] The second ceramic material K2 used in the ceramic thermal barrier 53 in the intermediate region preferably has a low mullite content. The mullite content is preferably 15%–25%.

[0085] The third ceramic material K3 preferably contains almost no mullite, or more precisely, the mullite content is <5%. The third ceramic material used for the ceramic thermal barrier near the outflow area preferably contains only corundum, and has a spinel content of 1% < spinel < 8%.

[0086] Viewed along the flow direction, the tile section 8 is constructed as multiple rows of thermal barrier elements.

[0087] Examples of material selection for ceramics K1, K2, and K3 with different numbers of rows in the combustion chamber. a) First line: K1 Second line: K2 Third line: K3, b) First row: K1 Second line: K3 Third line: K3, c) First line: K1 Second line: K2 Third line: K3 Fourth line: K3, d) First line: K1 Second line: K2 Third line: K2 Fourth line: K3, e) First row: K1 Second line: K1 Third line: K2 Fourth line: K3, f) First row: K1 Second line: K1 Third line: K2 Fourth line: K3 Fifth line: K3, g) First line: K1 Second line: K2 Third line: K2 Fourth line: K3 Fifth line: K3.

[0088] The end of the combustion chamber 50, i.e. the inlet shell, has a tiled portion 8 with a metal thermal barrier 55. The metal thermal barrier 55 has a metal substrate and a ceramic insulation layer system. The ceramic insulation layer system consists of a metal bonding layer (MCrAlY) and a ceramic protective layer, as is known from turbine blades.

[0089] The combustion chamber 50 is also adjusted for the preferred improved conditions of the compressor 4, and has at least two different types / materials for the tile section 8.

[0090] Figure 4 The exhaust housing 33 is shown, which has a support bar 63 for an external annular channel. Because the exhaust gas flowing through the preferred improved compressor exhibits different outflow characteristics, the support bar 63 was adjusted during the update. Figure 5 The method involves cutting off the front part of the support bar 63, especially the front half of the hollow structure, and replacing it with a new material with reduced curvature.

[0091] Figure 6 Rotor 2 is shown, which is fully assembled and has at least according to Figure 1 , Figure 2 , Figure 3 , Figure 4 and / or Figure 5 Improved or modified components. Thus, the rotor 2 already has, in particular, an improved compressor disc 16 with rotating blades of the compressor 4, or improved compressor blades 40, or ceramic thermal barrier elements 53, 55, 56 as part of the paving portion 8 of the combustion chamber 50, or an improved turbine disc, or improved turbine blades, and optionally has an modified exhaust housing 33.

[0092] This type of rotor 2 is easy and quick to install, eliminating the need for existing rotors to be disassembled, re-stacked, and reinstalled on-site. This significantly reduces downtime for equipment operators.

[0093] Even without prior improvements to the compressor, the individual features can be used to improve the gas turbine because each feature individually offers advantages and can be used for retrofitting, namely:

[0094] 1. A gas turbine 1, which has at least: Rotor 2; Combustion zone 7 is part of rotor 2; The compressor 4, as part of the rotor 2, is located upstream of the combustion zone 7; and The hot gas section 10, which is part of the rotor 2, is located downstream of the combustion zone 7. Among them, compressor 4 has a compressor disc for compressor blades, The hot gas section 10 includes a turbine disk for the turbine blades 25. Optionally, at least one compressor disc 16 of compressor 4 has a different material from the remaining compressor discs 13, and / or Optionally, at least one turbine disk 23 of the hot gas section 10 has a different material than the other turbine disks 22.

[0095] 2. The gas turbine according to claim 1, The gas turbine has a compressor 4 with at least twelve compressor stages. The gas turbine specifically features a compressor 4 with fifteen compressor stages. Among them, at least the last compressor stage, especially the last two compressor stages, have compressor disks 16 made of a different material than the earlier compressor disks 13.

[0096] 3. The gas turbine according to any one or both of claims 1 or 2, Among them, the first compressor disk 13, or at least four compressor disks 13 immediately preceding the last compressor disk or immediately preceding the penultimate compressor disk and the last compressor disk 16, have a first steel material. The first type of steel has the following characteristics (in weight %): Carbon (C), especially 0.20% < C < 0.35%, Chromium (Cr), especially 1.0% < Cr < 2.5%, Molybdenum (Mo), especially 0.20%<Mo<0.55%, Vanadium (V), especially 0.03% < V < 0.20%, Nickel (Ni), especially 3.0% < Ni < 4.4%, The proportions of silicon (Si), manganese (Mn), phosphorus (P), sulfur (S), arsenic (As), antimony (Sb), and / or tin (Sn) should be kept as small as possible. Specifically, the first steel is composed of the above-mentioned components.

[0097] 4. The gas turbine according to any one or more of claims 1, 2, or 3, Among them, the last compressor disc or the second to last compressor disc and the last compressor disc 16 have a second steel material, and Compared with the first steel of the other compressor disc 13, the second steel has the following characteristics: A lower proportion of carbon (C). A significantly higher share, especially at least 100% higher share of chromium (Cr). A higher proportion of molybdenum (Mo), and Intentionally introduced niobium (Nb), tungsten (W), nitrogen (N) and manganese (Mn).

[0098] 5. The gas turbine according to claim 4, The second type of steel has the following characteristics (unit: weight %): Carbon (C), especially 0.08% < C < 0.19%, Chromium (Cr), especially 9.5% < Cr < 11.0%, Molybdenum (Mo), especially 0.80%<Mo<1.5%, Vanadium (V), especially 0.10% < V < 0.30%, Nickel (Ni), especially 0.4% < Ni < 1.2%, Tungsten (W), especially 0.7% < W < 1.3%, Manganese (Mn), especially 0.02% < Mn < 0.70%, Nitrogen (N), especially 0.035% < N < 0.070%, Niobium (Nb), especially 0.02% < Nb < 0.08%, Specifically, the second steel is composed of the above-mentioned components.

[0099] 6. The gas turbine according to any one or more of claims 1, 2, 3, 4 or 5, The gas turbine has multiple torque discs 19 between the compressor 4 and the hot gas section 10. At least one torque disc 19, and in particular all torque discs 19, have the same material as the last compressor disc 16 of the compressor 4.

[0100] 7. The gas turbine according to any one or more of claims 1, 2, 3, 4, 5 or 6, The gas turbine has at least three stages I, II, III, and IV in the hot gas section 10, and particularly at least four stages I, II, III, and IV. Among them, at least the first two turbine disks 22 of stages I, II, III, and IV, and especially the first three turbine disks 22 of stages I, II, III, and IV, have the same material, and Or it may be made of a different material than the turbine disk 23 of the third and fourth stages (III and IV). Or it may be made of a different material than the turbine disk of the fourth stage IV.

[0101] 8. The gas turbine according to any one or more of claims 1, 2, 3, 4, 5 or 6, The gas turbine has at least three stages I, II, III, and IV in the hot gas section 10, and particularly at least four stages I, II, III, and IV. All turbine disks 22 and 23 of stages I, II, III and IV are made of the same material.

[0102] 9. The gas turbine according to any one or more of claims 6, 7 or 8, Among them, at least the foremost turbine disk 22 has the same material as the torque disk 19.

[0103] 10. The gas turbine according to any one or more of claims 1, 7, 8 or 9, The gas turbine has four stages I, II, III, and IV in the hot gas section 10. Among them, stages I, II, III, and IV consist of stationary turbine blades V1, V2, V3, and V4, and non-stationary turbine blades B1, B2, B3, and B4. Among them, the turbine blades V1, B1, V2, and B2 of the first two stages I and II, and especially the stationary turbine blade V3 of the third stage III, and even more especially the stationary turbine blade V3 and the non-stationary turbine blade B3 of the third stage III, have: Or a material different from the non-stationary turbine blade B3 of stage III. Or a material different from at least one turbine blade V4, B4 among the multiple turbine blades of the fourth stage IV. Optionally, the first non-stationary turbine blade B1 has an SX structure or a columnar structure.

[0104] 11. The gas turbine according to claim 10, Among them, the turbine blades V4 and B4 of the fourth stage IV, and optionally the non-stationary turbine blade B3 and / or the stationary turbine blade V3 of the third stage III, have a nickel-based alloy, which has a weight percentage of: Cobalt (Co), especially 9.0%–10%, Chromium (Cr), especially 13.0%–15.0%, Tungsten (W), especially 3.5%–4.5%, Molybdenum (Mo), especially 3.7%–4.3%, Aluminum (Al), especially 2.8%–3.2%, Titanium (Ti), especially 4.8%–5.5%, Boron (B), especially 0.01%–0.2%, Zirconium (Zr), especially 0.02%–0.10%, Carbon (C), especially 0.15%–0.40%, Preferably, it does not contain tantalum (Ta), preferably it does not contain hafnium (Hf), and preferably it does not contain niobium (Nb). Specifically, the nickel-based alloy is composed of the above-mentioned components.

[0105] 12. The gas turbine according to claim 11, Among them, the turbine blades V1, B1, V2, and B2 of at least the first stage I and II are made of materials different from those of turbine blades B4 and V4, and optionally also different from those of turbine blades B3 and V3 of the third stage III, but instead have tantalum (Ta) and / or hafnium (Hf).

[0106] 13. The gas turbine according to any one or more of claims 1, 10, 11 or 12, The gas turbine has at least three stages I, II, III, and IV in the hot gas section 10, and more particularly, at least four stages I, II, III, and IV in the hot gas section 10. Among them, stages I, II, III, and IV consist of stationary turbine blades V1, V2, V3, and V4, and non-stationary turbine blades B1, B2, B3, and B4. Among them, the turbine blades of the first stage V1, B1, V2, B2, and optionally the turbine blades of the third stage V3, B3, are made of nickel-based alloys, which have (in weight %): Cobalt (Co), especially 9.0%–10.5%, Chromium (Cr), especially 8.0%–9.5%, Tungsten (W), especially 9.0%–10.5%, Molybdenum (Mo), especially 0.2%–1.0%, Aluminum (Al), especially 5.0%–6.0%, Titanium (Ti), especially 0.5%–1.5%, Tantalum (Ta), especially 2.5%–3.5%, Boron (B), especially 0.01%–0.25%, Zirconium (Zr), especially 0.004%–0.06%, Carbon (C), especially 0.05%–0.16%, Hafnium (Hf), especially 1.0%–2.0%, Specifically, the nickel-based alloy is composed of the above-mentioned components.

[0107] 14. The gas turbine according to any one or more of claims 1, 10, 11, 12 or 13, The gas turbine has at least three stages I, II, III, and IV in the hot gas section 10, and more particularly, at least four stages I, II, III, and IV in the hot gas section 10. Among them, stages I, II, III, and IV consist of stationary turbine blades V1, V2, V3, and V4, and non-stationary turbine blades B1, B2, B3, and B4. Among them, the turbine blades of all stages I, II, III, and IV, namely V1, B1, V2, B2, V3, B3, V4, and B4, have a metal oxide protective layer. Preferably, it has a metallic NiCoCrAlYRe layer, NiCoCrAlY layer, NiCoCrAlYTaSi layer, NiCoCrAlYTaHf layer, NiCoCrAlYTaSiHf layer, or NiCoCrAlYTa layer. Preferably, all turbine blades 25 have the same metal layer.

[0108] 15. The gas turbine according to any one or more of claims 1 to 14, It has at least three levels I, II, III, and IV in the hot gas section 10, and in particular, it has at least four levels I, II, III, and IV in the hot gas section 10. Among them, stages I, II, III, and IV consist of non-stationary turbine blades B1, B2, B3, and B4 and stationary turbine blades V1, V2, V3, and V4. Among them, only the turbine blades of the first few stages I, II, and III have ceramic coatings, especially the first three stages I, II, and III.

[0109] 16. The gas turbine according to claim 15, The ceramic coatings of the turbine blades I and II (V1, B1, V2, B2) in the first two stages, and optionally the ceramic coatings of the turbine blades V3 and B3 in the third stage III, have the following characteristics: The same ceramic coating, In particular, it has segmented, especially dense (<10%), zirconium oxide-based ceramic layers. And in the turbine blades V3 and B3 of the third stage III, there is a super-porous ceramic layer with a porosity greater than 12%, and particularly greater than 18%. Specifically, it uses zirconium oxide as the matrix.

[0110] 17. The gas turbine according to any one or more of the preceding claims, Specifically, improved compressor blades 40 are present in at least one compressor stage of compressor 4. The improved compressor blade 40 is characterized in that the skimming edge 44 does not extend at the edge of the blade body profile of the blade body 43, but is preferably arranged at the middle of the blade tip 43, and particularly extends across the entire length of the blade tip 43.

[0111] 18. The gas turbine according to claim 18, The compressor 4 has only one compressor stage, and in particular, the compressor stage between the fourth and eighth compressor stages of the compressor 4 has the improved compressor blades 40 according to claim 18.

[0112] 19. The gas turbine according to any one or more of the preceding claims, It has a combustion chamber 50, which has a tiled section 8, which has all-ceramic tiles 52, 53, and 54 and a metal thermal barrier element 55.

[0113] 20. The gas turbine according to claim 19, Viewed along the flow direction of the hot gas in the combustion chamber 50, the gas turbine has the following characteristics: K1, the first ceramic material used in all-ceramic tiles 52, The preferred second ceramic material K2 is used for the all-ceramic tile 53. Then, the third ceramic material K3 is preferred for use in the all-ceramic tile 53. And then preferably a metal thermal barrier 55 is present at the end.

[0114] 21. The gas turbine according to claim 20, The first ceramic material K1 is a mixture of corundum and mullite in a ratio of 40:60 to 60:40, particularly a mixture of equal parts. Among them, the optional second ceramic material K2 has a mullite content of 15%-25%. The third ceramic material K3 contains almost no mullite, and in particular, contains less than 5% mullite.

[0115] 22. The gas turbine according to one or more of the preceding claims, The gas turbine has three exhaust pipes 30 that can guide compressed air from the compressor 4 into the exhaust casing 33.

[0116] 23. A method for modifying a gas turbine. Specifically, the existing gas turbine is modified so that the gas turbine has at least one feature according to one or more of the preceding claims 1 to 22.

[0117] 24. The method according to claim 23, The existing gas turbine's installed, fully assembled rotor is replaced by a new, fully assembled rotor 2 having at least one feature of one or more of the aforementioned gas turbine claims 1 to 22.

[0118] 25. The method according to claim 23, The system includes another exhaust pipe 30 that directs air from the compressor 4 into the exhaust housing 33.

[0119] 26. The method according to any one or more of claims 23, 24 or 25, The gas turbine 1 has an exhaust casing 33 with an annular exhaust channel 65 in which a support bar 63 is present. The support bar is modified by removing the front part of the support bar 63 and replacing it with a new, modified profile with a smaller curvature.

Claims

1. A gas turbine (1), said gas turbine having at least: Rotor (2); Combustion zone (7); Compressor (4), which is located upstream of the combustion zone (7) as part of the rotor (2); and The hot gas section (10), which is part of the rotor (2) and located downstream of the combustion zone (7), in, The compressor (4) has compressor discs (13, 16) for compressor blades (40). The hot gas section (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 from the other compressor disks (13), and / or At least one turbine disk (23) of the hot gas section (10) has a different material from the other turbine disks (22).

2. The gas turbine according to claim 1, The gas turbine has a compressor (4) with at least twelve compressor stages. The gas turbine is particularly equipped with a compressor (4) with fifteen compressor stages. in, At least the last compressor stage, especially the last two compressor stages, have compressor disks (16) made of a different material than the earlier compressor disks (13).

3. The gas turbine according to any one or both of claims 1 or 2, in, The first compressor disk (13), or at least four compressor disks (13) immediately before the last compressor disk or immediately before the penultimate compressor disk and the last compressor disk (16), have a first steel. The first steel has the following characteristics (in weight %): Carbon (C), especially 0.20% < C < 0.35%, Chromium (Cr), especially 1.0% < Cr < 2.5%, Molybdenum (Mo), especially 0.20%<Mo<0.55%, Vanadium (V), especially 0.03% < V < 0.20%, Nickel (Ni), especially 3.0% < Ni < 4.4%, The proportions of silicon (Si), manganese (Mn), phosphorus (P), sulfur (S), arsenic (As), antimony (Sb), and / or tin (Sn) should be kept as small as possible. Specifically, the first steel is composed of the aforementioned components.

4. The gas turbine according to any one or more of claims 1, 2, or 3, in, The last compressor disc, or the penultimate compressor disc and the last compressor disc (16) have a second steel, and Compared with the first steel of other compressor discs (13), the second steel has the following characteristics: A lower share of carbon (C). A significantly higher share, especially at least 100% higher share of chromium (Cr). A higher proportion of molybdenum (Mo), and Intentionally introduced niobium (Nb), tungsten (W), nitrogen (N) and manganese (Mn).

5. The gas turbine according to claim 4, in, The second type of steel has the following characteristics (in weight %): Carbon (C), especially 0.08% < C < 0.19%, Chromium (Cr), especially 9.5% < Cr < 11.0%, Molybdenum (Mo), especially 0.80%<Mo<1.5%, Vanadium (V), especially 0.10% < V < 0.30%, Nickel (Ni), especially 0.4% < Ni < 1.2%, Tungsten (W), especially 0.7% < W < 1.3%, Manganese (Mn), especially 0.02% < Mn < 0.70%, Nitrogen (N), especially 0.035% < N < 0.070%, Niobium (Nb), especially 0.02% < Nb < 0.08%, Specifically, the second steel is composed of the aforementioned components.

6. The gas turbine according to any one or more of claims 1, 2, 3, 4 or 5, The gas turbine has multiple torque discs (19) between the compressor (4) and the hot gas section (10). in, At least one torque disc (19), and in particular all of the torque discs (19) have the same material as the last compressor disc (16) of the compressor (4).

7. The gas turbine according to any one or more of claims 1, 2, 3, 4, 5 or 6, The gas turbine has at least three stages (I, II, III) in the hot gas section (10). Specifically, the hot gas section (10) has at least four levels (I, II, III, IV). in, At least the first two turbine disks (22) of the stages (I, II, III, IV), and especially the first three turbine disks (22) of the stages (I, II, III, IV), are made of the same material, and Or it may be made of a different material than the turbine disks (23) of the third and fourth stages (III, IV). Or it may be made of a different material than the turbine disk of the fourth stage (IV).

8. The gas turbine according to any one or more of claims 1, 2, 3, 4, 5 or 6, The gas turbine has at least three stages (I, II, III) in the hot gas section (10). Specifically, the hot gas section (10) has at least four levels (I, II, III, IV). in, All turbine disks (22, 23) of the aforementioned stages (I, II, III, IV) are made of the same material.

9. The gas turbine according to any one or more of claims 6, 7 or 8, in, At least the foremost turbine disk (22) has the same material as the torque disk (19).

10. The gas turbine according to any one or more of claims 1, 7, 8 or 9, The gas turbine has at least three stages (I, II, III) in the hot gas section (10). Specifically, the hot gas section (10) has at least four levels (I, II, III, IV). in, Stages (I, II, III, IV) consist of stationary turbine blades (V1, V2, V3, V4) and non-stationary turbine blades (B1, B2, B3, B4). The turbine blades of the first two stages (I, II) (V1, B1, V2, B2), and especially the stationary turbine blade (V3) of the third stage (III), and particularly the stationary turbine blade (V3) and non-stationary turbine blade (B3) of the third stage (III), have the following characteristics: Or a material different from the non-stationary turbine blades (B3) of stage III, Or a material different from at least one turbine blade (V4, B4) of the multiple turbine blades in the fourth stage (IV). Optionally, the first non-stationary turbine blade (B1) has an SX structure or a columnar structure.

11. The gas turbine according to claim 10, in, The turbine blades of the fourth stage (IV) (V4, B4), and optionally the non-stationary turbine blades (B3) and / or stationary turbine blades (V3) of the third stage (III), are made of a nickel-based alloy having (in weight %): Cobalt (Co), especially 9.0%–10%, Chromium (Cr), especially 13.0%–15.0%, Tungsten (W), especially 3.5%–4.5%, Molybdenum (Mo), especially 3.7%–4.3%, Aluminum (Al), especially 2.8%–3.2%, Titanium (Ti), especially 4.8%–5.5%, Boron (B), especially 0.01%–0.2%, Zirconium (Zr), especially 0.02%–0.10%, Carbon (C), especially 0.15%–0.40%, Preferably, it does not contain tantalum (Ta), preferably it does not contain hafnium (Hf), and preferably it does not contain niobium (Nb). In particular, the nickel-based alloy is composed of the above-mentioned components.

12. The gas turbine according to claim 11, in, The turbine blades of at least the first stage (I, II) (V1, B1, V2, B2) are made of a material different from that of the turbine blades (B4, V4), and optionally also different from that of the turbine blades of the third stage (III) (B3, V3), but instead have tantalum (Ta) and / or hafnium (Hf).

13. The gas turbine according to any one or more of claims 1, 10, 11 or 12, The gas turbine has at least three stages (I, II, III) in the hot gas section (10). Specifically, the hot gas section (10) has at least four levels (I, II, III, IV). in, Stages (I, II, III, IV) consist of stationary turbine blades (V1, V2, V3, V4) and non-stationary turbine blades (B1, B2, B3, B4). The turbine blades of the first stage (V1, B1, V2, B2) and optionally the turbine blades of the third stage (V3, B3) are made of nickel-based alloys, wherein the nickel-based alloys have (in weight %): Cobalt (Co), especially 9.0%–10.5%, Chromium (Cr), especially 8.0%–9.5%, Tungsten (W), especially 9.0%–10.5%, Molybdenum (Mo), especially 0.2%–1.0%, Aluminum (Al), especially 5.0%–6.0%, Titanium (Ti), especially 0.5%–1.5%, Tantalum (Ta), especially 2.5%–3.5%, Boron (B), especially 0.01%–0.25%, Zirconium (Zr), especially 0.004%–0.06%, Carbon (C), especially 0.05%–0.16%, Hafnium (Hf), especially 1.0%–2.0%, Specifically, the nickel-based alloy is composed of the above-mentioned components.

14. The gas turbine according to any one or more of claims 1, 10, 11, 12 or 13, The gas turbine has at least three stages (I, II, III) in the hot gas section (10). Specifically, the hot gas section (10) has at least four levels (I, II, III, IV). in, Stages (I, II, III, IV) consist of stationary turbine blades (V1, V2, V3, V4) and non-stationary turbine blades (B1, B2, B3, B4). Among them, the turbine blades of all stages (I, II, III, IV) (V1, B1, V2, B2, V3, B3, V4, B4) have a metal oxide protective layer. Preferably, it has a metallic NiCoCrAlYRe layer, NiCoCrAlY layer, NiCoCrAlYTaSi layer, NiCoCrAlYTaHf layer, NiCoCrAlYTaSiHf layer, or NiCoCrAlYTa layer. Preferably, all turbine blades (25) have the same metal layer.

15. The gas turbine according to any one or more of claims 1 to 14, The gas turbine has at least three stages (I, II, III) in the hot gas section (10). Specifically, the hot gas section (10) has at least four levels (I, II, III, IV). in, Stages (I, II, III, IV) consist of non-stationary turbine blades (B1, B2, B3, B4) and stationary turbine blades (V1, V2, V3, V4). Of these, only the turbine blades of the first three stages (I, II, III) have ceramic coatings. Specifically, only the turbine blades of the first three stages (I, II, III) have a ceramic coating.

16. The gas turbine according to claim 15, in, The ceramic coatings of the turbine blades of the first two stages (I, II: V1, B1, V2, B2) and optionally, the ceramic coatings of the turbine blades of the third stage (III) (V3, B3) have: The same ceramic coating, In particular, it has segmented, especially dense (<10%), zirconium oxide-based ceramic layers, and / or The turbine blades (V3, B3) of the third stage (III) have a super-porous ceramic layer with a porosity greater than 12%, particularly greater than 18%. Specifically, it uses zirconium oxide as the matrix.

17. The gas turbine according to any one or more of the preceding claims, in, Improved compressor blades (40) are present in at least one compressor stage of the compressor (4). The improved compressor blade (40) is characterized in that, The skimming edge (44) does not extend at the edge of the blade body profile of the blade body (43), but rather the skimming edge (44) is preferably arranged in the middle of the blade tip (43) and extends particularly across the entire length of the blade tip (43).

18. The gas turbine according to claim 18, in, The compressor (4) has only one compressor stage, particularly the compressor stage between the fourth and eighth compressor stages of the compressor (4), having improved compressor blades (40) according to claim 18.

19. The gas turbine according to one or more of the preceding claims, The gas turbine has a combustion chamber (50). The combustion chamber (50) has a tiled section (8) consisting of rows of tiles. The tiled section (8) has all-ceramic tiles (52, 53, 54) and a metal thermal barrier element (55).

20. The gas turbine according to claim 19, in, Viewed along the flow direction of the hot gas in the combustion chamber (50), the gas turbine has: First ceramic material (K1) for all-ceramic tiles (52). The preferred second ceramic material (K2) is used for all-ceramic tiles (53). Then the third ceramic material (K3) is used for all-ceramic tiles (53). And preferably, a metal thermal barrier (55) is present at the end.

21. The gas turbine according to claim 20, in, The first ceramic material (K1) is a mixture of corundum and mullite in a ratio of 40:60 to 60:40, particularly a mixture of equal parts. The optional second ceramic material (K2) contains 15%–25% mullite. The third ceramic material (K3) contains almost no mullite, and in particular, contains less than 5% mullite.

22. The gas turbine according to one or more of the preceding claims, The gas turbine has three exhaust pipes (30) that guide compressed air from the compressor (4) into the exhaust casing (33).

23. A method for modifying a gas turbine. in, An existing gas turbine is modified to have at least one feature according to any one or more of claims 1 to 22.

24. The method according to claim 23, in, The existing gas turbine's installed, fully assembled rotor is replaced by a new, fully assembled rotor (2) having at least one of the features described in any one of claims 1 to 22 of the aforementioned gas turbine.

25. The method according to claim 23 or 24, in, Another discharge pipe (30) is installed, which guides air from the compressor (4) into the exhaust housing (33).

26. The method according to any one or more of claims 23, 24 or 25, in, The gas turbine (1) has an exhaust casing (33) with an annular exhaust passage (65). A support strip (63) is present in the discharge channel. The support strip is modified by removing the front area of ​​the support strip (63) and replacing it with a new, modified profile with a smaller curvature.

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