JP3977976B2 - Gas turbine stationary blade repair method - Google Patents

Description

[0001]
BACKGROUND OF THE INVENTION
  BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to a turbine vane in a gas turbine, and more particularly to a gas turbine vane that is located in a high-temperature combustion gas flow path and is susceptible to damage due to thermal fatigue and creep and requires regular maintenance and inspection. How to repairTo the lawRelated.
[0002]
[Prior art]
  Gas turbines have the advantage of superior operability compared to other prime movers for thermal power generation. For this reason, depending on the gas turbine power plant, DSS (Daily Start-up and Shutdown) While WSS (Weekly Start-up and Shutdown) is possible, the turbine components are in strict conditions.
[0003]
  This is because gas turbines have many parts that are located in the flow path of combustion gas and exposed to high temperatures, such as the moving blades and stationary blades of combustors and turbine impellers. This is due to repeated thermal strain and creep deformation in a high temperature environment during steady operation.
[0004]
  Therefore, for these gas turbine components exposed to high temperatures, alloys based on nickel or cobalt having excellent heat resistance, so-called heat-resistant superalloys, have been used.
[0005]
  However, even with such a heat-resistant superalloy, in the case of a gas turbine, it is used at a high temperature that is almost close to the limit operating temperature. Further, under the above-mentioned operation, it is considerably earlier than other parts. May be damaged.
[0006]
  In particular, the turbine vanes at the first stage are susceptible to damage due to thermal fatigue, and there have been reports of cases where cracks occur early and progress. Therefore, repair and replacement of new parts is indispensable. For this reason, the reduction of operation costs largely depends on the control of inspection and maintenance costs.
[0007]
  In particular, because the turbine vane is made of an expensive heat-resistant superalloy, the cost required for its replacement is large, and the proportion of the total operating cost is significant. Therefore, the thermal stress applied to it is reduced and damage is suppressed. Is an important issue.
[0008]
  Therefore, for example, in Japanese Patent Application Laid-Open No. 8-135402, a stiffener can be secured even if the thickness of the blade is reduced by disposing a reinforcing material in the height direction of the blade on the cooling side of the stationary blade. JP-A-8-135403 proposes an invention that can prevent damage caused by swaying, and the shroud is divided into a shroud having wings and a shroud that does not have such a constraint, and the constraint is given structurally. This invention proposes an invention that suppresses thermal stress and alleviates thermal stress.
[0009]
  In JP-A-11-93609, in a spelled stationary blade in which several single blades are connected, the blades are connected via pins, and the pins are linearly expanded from the stationary blade material. A method for reducing the rigidity of the shroud by making it from a material having a large coefficient is proposed. Further, JP-A-10-61406 proposes a method for reducing the stress concentration around the cooling hole by making the cooling hole into a slit shape. is doing.
[0010]
[Problems to be solved by the invention]
  In the above prior art, consideration is not given to stress concentration at the boundary between the stationary blade and the shroud, and there is a problem in damage to the stationary blade due to stress concentration.
  That is, in the prior art, consideration is given to stress relaxation for each of the blade and the shroud, but since it is the boundary between the stationary blade and the shroud that actually cracks, in the prior art, this boundary It is impossible to suppress the occurrence of cracks in the film, and thus the above problem occurs.
[0011]
  Although the stationary blades of a gas turbine are exposed to high-temperature gas, the thermal boundary conditions at this time are not necessarily uniform for all the blades, and there are different thermal boundary conditions for each blade. Therefore, it is considered that the part where the crack is generated is often different for each blade, and it appears that each blade is concentrated especially on the part exposed to high temperature. In order to extend the service life, it is necessary to find and adopt a configuration effective for reducing thermal stress at a site where a crack occurs in each stationary blade, that is, stress due to thermal deformation.
[0012]
  The present inventionPurpose ofIs to provide a gas turbine stationary blade repair method that is effective in reducing thermal stress.The
[0013]
[Means for Solving the Problems]
  Cracks in gas turbine stationary blades are generated mainly by repeated thermal stresses that accompany start and stop, and progress, but thermal stresses at this time are structurally affected by temperature changes or thermal deformation due to temperature distribution within the member. This is a so-called displacement control type load that is generated by being restrained by the load.
[0014]
  Therefore, in order to reduce the thermal stress, it is effective to make the member easily deformed, that is, to have a structure with low rigidity. For this reason, the simplest structural change is the plate of the member. This is to reduce the thickness, and this method is also effective in improving the cooling performance.
  However, reducing the rigidity of the member is accompanied by a decrease in reliability with respect to permanent deformation such as creep deformation of the member.
[0015]
  The above purpose isWhen a crack occurs in the gas turbine stationary blade, the back side of the cracked surface is cut, a groove is formed on both sides of the cracked portion, and is filled with weld metal, and a cooling hole that covers the cut portion The crack is repaired by joining the flat plate-shaped member havingAchieved.
[0017]
  This removes the cracked area, so there is no risk of the crack developing after repair.ThinAs a result of the reduction in rigidity and improvement in cooling efficiency due to the fleshing, thermal stress is reduced, andFlatBy attaching the reinforcing material, the resistance to local plastic deformation can be improved without impairing the cooling performance.
[0018]
  At this time, a part where the frequency of occurrence of cracks in the gas turbine stationary blade is high is obtained in advance from structural analysis and past actual machine damage data, and a member having the same shape as that part is prepared as a spare part. As a reference example, it is possible to consider a method in which a portion where a crack is generated is cut out completely and a spare part is joined in a manner to replace the cut portion.
[0019]
  In the case of this reference example,Since the part where the crack occurred is removed, there is no risk of the crack developing after repair,With replacement spare parts,As a result of low rigidity and improved cooling efficiency due to thinning, thermal stress is reduced, and the installation of a flat reinforcing material with cooling holes enables local plasticity without impairing cooling performance. Improved resistance to deformation is obtained.
[0020]
  At this time,The part where the frequency of occurrence of cracks in the gas turbine stationary blade is determined in advance from structural analysis and past actual machine damage data, a member having the same shape as that part is prepared, and this member is thinned from the cooling side. Cover the thinned part with a flat plate reinforcing material with cooling holes to make a spare part, cut out the part where the target vane crack occurred, and replace it with this part. When joining and repairing spare parts in the form, create damage data for each gas turbine stationary blade to be managed, create a database in advance, and manage the number of spare partsThe method can also be considered as a reference example.
[0021]
  In the case of this reference example,When a gas turbine stationary blade is damaged, spare parts suitable for the damage can always be prepared, and loss time due to repair can be minimized.
[0022]
  Furthermore, at this time,The part where the thermal stress of the gas turbine stationary blade is concentrated is thinned in advance, and reinforcement is performed with another member.The method can also be considered as a reference example,About the part where the thermal stress of the gas turbine stationary blade is concentrated in advance, it is divided into a plurality of members and then fitted to each other so that the stress is shared.The method of doing so can also be considered as a reference example.
[0023]
DETAILED DESCRIPTION OF THE INVENTION
  Hereinafter, repair of a gas turbine stationary blade according to the present inventionHow toThe embodiment will be described in detail.
[0024]
  First, an embodiment of a gas turbine stationary blade repair method according to the present invention will be described.The gas turbine stationary blade repair method considered as an example for reference is described in FIG. (a) , (b) According toIn these drawings, 1 is a shroud on the stationary blade casing side, 2 is a blade portion of the stationary blade, 3 is a portion where a crack K due to thermal fatigue is generated, and 4 is a spare part for repair. It is. The shroud 1 on the stationary blade casing side (outer peripheral side) is sometimes called an end wall.
[0025]
  Here, the crack K generated in the gas turbine stationary blade is often generated in a portion where the shroud 1 is joined to the blade portion 2 as shown in FIG.
  So thisReference exampleThen, based on structural analysis, past actual machine damage data, and the like, a part where cracks occur frequently in the gas turbine stationary blade, such as a part where the shroud 1 is joined to the blade part 2, for example, is obtained in advance. A member having the same shape as that part is prepared as a spare part 4.
[0026]
  And when a crack K occurs in the shroud 1 and it needs to be repaired,Reference exampleThen, as shown in FIG. 2B, first, a region surrounding the crack K of the shroud 1 is set as a cracked portion 3 and then cut out from the shroud 1 as indicated by an arrow in the figure. .
[0027]
  Next, the spare part 4 is fitted in the space formed after the part 3 is removed from the shroud 1, and the part 3 is welded between the shroud 1 and the wing part 2 to be repaired. Here, W represents a joint portion by welding.
  As shown in FIG. 2, the spare part 4 at this time is thinned by scraping a part 4A of the member having the same outer shape as the part 3 prepared in advance from the cooling side on the upper side in the drawing, A hollow portion is formed by attaching a flat plate-like member 6 having a cooling hole 5 so as to close the thinned portion 4B.
[0028]
  So thisReference exampleAccording to the above, when a crack K occurs in the shroud 1, the region including the crack K is completely removed, and the spare part 4 is newly attached after that, so that the possibility that the crack K further develops is reliably suppressed. In addition, in the spare part 4, the portion where cracks are likely to occur is thinned, thereby reducing the thermal stress due to the low rigidity and improving the cooling property, and the addition of the flat plate member 6 reduces the strength. The cooling property by the cooling hole 5 is improved while suppressing the above.
[0029]
  This results in thisReference exampleAccording to the above, since the repair work of the gas turbine stationary blade is the replacement work by the spare part 1, the repair work is easy, the variation in the repair result can be reduced, and the thermal stress is reduced and the strength is increased by the spare part 1. Therefore, there is an advantage that the improvement of the reliability can be obtained by the repair, and the cost can be sufficiently reduced by extending the service life.
[0030]
  By the way, in the repair method using the spare parts as described above, it is a requirement that a spare part corresponding to the repair of the gas turbine stationary blade is provided immediately.
  Therefore,As a reference exampleFirst of all, the maintenance part management method for gas turbine stationary blades will be explained.In one exampleAs described below, the number of spare parts is managed.
[0031]
  First, in this reference example, as shown in FIG. 3, the casing-side shroud 1 is divided into six parts A to F as a whole, and as a result, as shown in FIG. The data indicating the crack occurrence status for the above is obtained, this is converted into a database, and a parts management system is constructed.
[0032]
  Here, FIG. 4 shows, for example, a periodical inspection of a gas turbine stationary blade, and data indicating a crack occurrence state is obtained from the result. Here, the vertical axis indicates a default value that needs to be repaired. The length of the crack is set, and the segment is each part of the gas turbine stationary blade divided into a plurality of unit parts in the circumferential direction.
[0033]
  Normally, in such a gas turbine, a plurality of the stationary blades in the circumferential direction, for example, as shown in FIG. 3, two blade portions are formed as a part in which the outer blade and the inner shroud are combined. Are arranged in the circumferential direction and assembled, and this portion is called a segment.
  Then, when preparing and preparing spare parts, the processing shown in the flowchart of FIG. 5 is executed to set the number of spare parts to be prepared and prepared.
[0034]
  In the flowchart of FIG. 5, first, data obtained in the past from other gas turbines used under conditions close to the usage history such as the operation time and the number of start / stops of the gas turbine scheduled to be repaired by a certain spare part. (S1). Then, the database is searched (S2), and from the search result (S3), how often a crack is generated in each of the parts A to F is obtained (S4).
[0035]
  Then, after obtaining the ratio from this frequency, the number of spare parts to be evaluated at this time is multiplied (S5), and the number of spare parts to be prepared is set (S6).
[0036]
  Therefore, according to this reference example, accurate management of the number of spare parts is always obtained, and in the repair method using spare parts, a spare part corresponding to the repair is provided immediately, and the loss time due to repair is minimized. Therefore, the running cost can be sufficiently reduced.
[0037]
  Next, a gas turbine stationary blade repair method according to the present inventionOneThe embodiment will be described with reference to FIG.
  This embodiment relates to a method for dealing with repair of a gas turbine stationary blade without using a spare part. In FIG. 6, reference numeral 2 </ b> A denotes that the end of the blade portion 2 is joined to the shroud 1. In this case, as shown in FIG. 6 (a), a crack K occurred in a portion 1A in contact with the joint of the shroud 1 and the tip of the portion 2A. The case is shown.
[0038]
  In the case of the crack K in FIG. 6A, first, the crack K in the portion 1A is first filled and repaired by welding or brazing.
  Next, as shown in FIG. 6 (b), along the repaired crack K, the surface of the shroud 1 is cut so as to include the crack K, thereby forming the thinned portion 1B in the portion 1A. As shown in (c), the thinned portion 1B is closed, and the flat portion 7 having the cooling holes 8 is welded to the surface of the shroud 1 as a reinforcing material so as to include the entire portion 1A, thereby repairing the gas turbine stationary blade. To do.
[0039]
  Here, the part 1A where the crack K has occurred may be considered to correspond to the part exposed to high stress during the operation of the gas turbine. Therefore, the part 1A is shown in FIG. 6 (c). By performing repairs in this way, it is possible to achieve low thermal stress at sites under severe operating conditions during operation.
[0040]
  In addition, since the stationary vanes of the gas turbine are arranged in an arc shape along the circumference in the turbine, the shroud portion also has a curvature matching the arc, and is therefore on the casing side (outer peripheral side). In the shroud, the inner side becomes hot during operation, so it tries to deform in a direction where the curvature decreases due to heat, and in the shroud on the rotor side (inner side), the outer side becomes hot, so the deformation direction Is the opposite.
[0041]
  As a result of the thermal deformation appearing in the shroud member and the restraint due to the structural rigidity, thermal stress is generated. Therefore, when the locally thinned portion 1B exists in the embodiment shown in FIG. However, since the flat plate member 7 serving as a reinforcing material is provided in the above-described embodiment, the possibility of deformation can be sufficiently suppressed.
[0042]
  On the other hand, since the cooling structure is adopted in the gas turbine stationary blade, a strong temperature gradient appears inside the member, and a strength gradient is also generated in the stress, which is close to a bending stress.
  Therefore, in order to cope with this bending stress, the flat plate member 7 serving as a reinforcing material needs to be attached along the direction in which the main stress is generated.
  Here, since this main stress generally works in a direction perpendicular to the direction in which the crack K extends, in the embodiment shown in FIG. 6, the longitudinal direction of the flat plate member 7 extends in the crack K. It can be seen that the above requirements are satisfied.
[0043]
  Next, in the embodiment of FIG. 6, as a result of the provision of the thinned portion 1B, good results are easily obtained for the above-described repair by cracking welding or brazing. Hereinafter, this point will be described.
  First, as shown in FIG. 7 (a), the cross-sectional shape of such a crack K is a shape that enters the inside substantially perpendicularly from the surface of the member (in this case, the shroud 1). At this time, in order to facilitate the inflow of the weld metal into the crack, as shown in FIG. 7 (b), it is usual to perform groove processing 10 on the surface side of the member.
[0044]
  In this case, if the processing amount, especially the depth, of the groove processing 10 is small, a residual defect K ′ that remains unfilled by the weld metal 11 is generated as shown in FIG. Cannot be repaired.
  On the other hand, if the groove processing 10 is too large, thermal deformation due to welding appears greatly, resulting in problems.
[0045]
  On the other hand, in the embodiment of the present invention shown in FIG.FIG.When the crack K shown in (a) occurs, the thinned portion 1B is first formed as shown in FIG. 5 (b). Therefore, as shown in the figure, on the back side of the surface where the crack K is generated, The tip of the crack K can appear, and as a result,FIG.As shown in (c), the groove processing 10 'can be performed also from the back side.
[0046]
  Therefore,FIG.As shown in (d), the weld metals 11 and 11 ′ can be filled from both sides of the crack K, and the possibility of remaining defects remaining can be surely eliminated, and highly reliable repair can be easily performed. Will be obtained.
  At this time, since the thinned portion 1B is formed, it is possible to sufficiently suppress the possibility of thermal deformation when the groove processing 10, 10 'is filled with the weld metals 11, 11'.
[0047]
  Next, the present inventionIs targetedThe gas turbine stationary blade will be described.
  First, FIG., An example of a gas turbine stationary blade targetedIn the figure, the figure shows a state in which only one surface member 2A of the hollow wing 2 (see FIG. 1) appears as a cross section, and one end in the length direction thereof. 1 is shown integrated with a shroud on one side, for example, the shroud 1 on the casing side, and the other end is integrated with the shroud on the other side, that is, the shroud 1 'on the rotor side.
[0048]
  Accordingly, in FIG. 9, cooling air is to flow through the left side of the surface member 2 </ b> A on the side where the wing portion 2 is hollow, that is, inside, and in this embodiment, The reinforcing component 14 is attached along the surface member 2A from one shroud 1 to the other shroud 1 ′ so as to sandwich the corner portion 15 where stress is generated inside the surface member 2A, that is, on the cooling side. The reinforcing member for the surface member 2A is used.
[0049]
  ThisTargeted by the present inventionIn the case of the structure in which the moving blade 2 is hollow and both ends are integrated with the shrouds 1, 1 ′, as in the gas turbine stationary blade described above, the maximum stress appears in the corner portion 15 shown in the figure. FIG. 10 shows an example of a result obtained by analyzing the stress at this time by the voting element method.SupplementIn the stress characteristics in the case of a wing structure without a strong material, black dots are shown in FIG.PlaceStress characteristics.
[0050]
  This FIG. 10 analyzes the thermal conductivity and gas temperature of the member surface under a certain assumed operating condition of the gas turbine, and the change pattern associated with these start and stop as load conditions, and the result is analyzed for the blade 2. The temperature-stress hysteresis at the corner 15 is shown. At this time, in the embodiment of FIG. The total thickness is equal to the thickness of the wing part of the prior art.
[0051]
  As is apparent from FIG. 10, in the case of the gas turbine stationary blade of FIG.When there is no reinforcementBy comparison, it can be seen that the fluctuation range of the stress is reduced to about 20%, and the stress value in the steady state is reduced by about 50%.
  This is shown in FIG.Gas turbine stationary bladeIn this case, it is considered that as a result of reducing the plate thickness, the rigidity was reduced, and as a result of increasing the surface area, cooling was improved and thermal stress was reduced.
[0052]
  This FIG.Gas turbine stationary blade shown inFIG. 11 shows a specific example of FIG. 11, and the wing part 2 has a closed loop shape in structure. Therefore, as shown in FIG. 9, when the reinforcing member 14 has an integral structure, it cannot be attached.
  Therefore, as shown in FIG. 11, the reinforcing member 14 is divided into a plurality of portions 14 </ b> A and 14 </ b> B and these are joined and attached by a welded portion 16.
[0053]
  At this time, the reinforcing member 14 is provided with a plurality of cooling holes 8 as shown in the figure, so that the circulation of cooling air and the increase of the heat transfer area can be obtained.Of FIG.In this case, it is necessary to configure so that the rigidity of the entire stationary blade portion is not increased by attaching the reinforcing member 14.
  So thisGas turbine stationary blade of FIG.According to this, a sufficient reduction in thermal stress can be obtained, and a highly reliable gas turbine stationary blade can be easily obtained.
[0054]
  by the way,The gas turbine stationary blade of FIG. 9 and the gas turbine stationary blade of FIG.In the above, the case where the reinforcing member 14 over the entire blade portion 2 is attached so as to reach from the one shroud 1 to the other shroud 1 ′ is shown, but the gas turbine stationary blade targeted by the present invention is shown in FIG. 12. And as shown in FIG. 13, you may implement in the form which attached the reinforcement member only to the corner part 15 which is a stress generation part.
[0055]
  Here, FIG. 12 shows a case where the reinforcing member 14C is attached to the corner portion 15 after the plate thickness of the wing portion 2 is reduced.In one exampleFIG. 13 shows an example in which the reinforcing member 14D is attached after the plate thickness of the shroud 1 or shroud 1 'is reduced.Take this caseUse a machine to reduce the plate thicknessprocessingIn this case, it is necessary to balance with the reduction amount of the plate thickness so that the rigidity of the entire stationary blade portion is not increased by attaching the reinforcing members 14C and 14D. .
[0056]
  Next, the present inventionReference examples of gas turbine stationary bladesWill be described. As already shown in FIG. 3, for example, in a gas turbine stationary blade, a plurality of blade portions are often configured as one segment, and a plurality of blade portions are formed in one segment. In this case, as shown in FIG. 3, the space between the wing part 2 and the wing part 2 of the shroud 1 is a high stress part, and the wing part which becomes a stress concentration part at this part. In many cases, the rear edge of the portion 2 and the corner portion 15 (see FIG. 9) of the shroud 1 are the starting points of cracks.
[0057]
  Therefore, the gas turbine stationary blade with such a spelling structure was targeted.Reference exampleWill be described with reference to FIG.
  FIG. 14 is a view of the rear side of the shroud, for example, the casing side shroud 1 as viewed from the outside of the circumference, and the rotor side shroud 1 ′ as viewed from the inside of the circumference. is there.
[0058]
  14Reference exampleThen, the shroud 1 is not made to have a uniform plate thickness, but the main portion 1A on the back side is reduced in thickness while leaving the peripheral portion 1B as shown in FIG. In addition, a portion 17 in which a portion of the portion 1A and the peripheral portion 1B is thinned so as to include a portion having a high probability of being a starting point of the crack generation described above is formed by, for example, a mechanization process, As shown in FIG. 2B, a flat reinforcing member 18 is welded so as to cover the portion 17.
[0059]
  At this time, for the portion where the peripheral portion 1B is lost due to the formation of this portion 17, the portion 18A of the reinforcing material 18 is fitted into the structure, thereby suppressing the leakage of cooling air, It is designed not to become an obstacle when the stationary blade is attached to the casing.
[0060]
  And this FIG. 14Reference exampleAlso in this case, the reinforcing member 18 is formed with cooling holes 19 so that the cooling air can be circulated to increase the heat transfer area.
  Therefore, according to the reference example of FIG. 14, the thermal stress can be sufficiently reduced, and a highly reliable gas turbine stationary blade can be easily obtained.
[0061]
  By the way, in such a segment type gas turbine stationary blade, when the stationary blade segment is attached to the casing, in order to be able to move smoothly along the casing to the attachment position, as shown in FIG. Generally, a guide portion 21 is formed on the back side.
[0062]
  However, in this case, since the plate thickness of the shroud 1 is increased at the portion where the guide portion 21 is formed, the rigidity of the shroud 1 is locally increased at this portion, which is desirable from the viewpoint of reducing thermal stress. None, but it is a necessary configuration in terms of segment installation and removal.
[0063]
  Therefore, a useful reference example for such a guided gas turbine stationary bladeIs described with reference to FIG.As shown in the drawing, instead of the guide portion 21 in the prior art, a guide member 22 having ears 23A and 23B formed on both side ends of the bottom portion is made separately from the shroud 1, and the shroud 1 Mounting portions 24A and 24B are formed in the periphery, and the guide members 22 are fixed to the shroud 1 by fitting the ear portions 23A and 23B of the guide members 22 to the mounting portions 24A and 24B of the shroud 1. Is.
[0064]
  At this time, one end portion of the guide member 22 and the mounting portion 24A are provided with pin holes 25 and 26, respectively, and after the guide member 22 is inserted and fitted into the shroud 1, the pin hole on the guide side is provided. A pin (not shown) is inserted and fixed from the pin 25 into the shroud side pin hole 26 so that the guide member 22 does not move relative to the shroud 1.
[0065]
  At this time, in the other mounting portion 24B, it is important that the guide member 22 is not fixed and is slidable, so that stress due to thermal deformation of the guide member 22 is not transmitted to the shroud 1 side. In addition, the rigidity by the guide member 22 can be prevented from being added to the shroud 1.
[0066]
  Therefore, in FIG.Reference exampleTherefore, a sufficient reduction in thermal stress can be obtained, and a highly reliable gas turbine stationary blade can be easily obtained.
  In FIG. 16, the description of the part where the inside (hollow part) of the wing part 2 is visible from the shroud 1 is omitted.
[0067]
【The invention's effect】
  According to the present invention, since the thermal stress can be reduced with a simple configuration, the life of the gas turbine stationary blade can be extended, and the reliability of the gas turbine system can be improved and the maintenance cost can be sufficiently reduced. be able to.
[Brief description of the drawings]
[Figure 1]A reference example is shown when spare parts are used to repair a gas turbine stationary blade.It is explanatory drawing.
[Figure 2]Other reference examples when spare parts are used for repair of gas turbine stationary bladesIt is explanatory drawing.
[Fig. 3]Crack initiation sites appearing on gas turbine vanesIt is explanatory drawing of.
[Fig. 4]Statistics of crack initiation sites appearing on gas turbine vanes.FIG.
[Figure 5]Flow for explaining repair parts management method of gas turbine stationary blade by reference exampleFIG.
FIG. 6 shows a gas turbine stationary blade according to the present invention.Repair methodIt is explanatory drawing of one Embodiment.
[Fig. 7]Welding repair of gas turbine vanesIt is explanatory drawing.
[Fig. 8]The welding repair in the embodiment of the gas turbine stationary blade repair method according to the present inventionIt is explanatory drawing.
FIG. 9First reference example of a gas turbine stationary bladeIt is explanatory drawing which shows.
FIG. 10Stress in the first reference example of a gas turbine stationary bladeFIG.
FIG. 11Second reference example of gas turbine stationary bladeIt is explanatory drawing which shows.
FIG.Third reference example of gas turbine stationary bladeIt is explanatory drawing which shows.
FIG. 13Fourth reference example of a gas turbine stationary bladeIt is explanatory drawing which shows.
FIG. 14A reference example of a gas turbine stationary blade to which a spelling blade is appliedIt is explanatory drawing which shows.
FIG. 15Conventional gas turbine stationary blades with guidesIt is explanatory drawing which shows an example.
FIG. 16A reference example of a gas turbine stationary blade with a guideIt is explanatory drawing which shows.
[Explanation of symbols]
  1 Shroud on the casing side
  1 ', 20 Rotor shroud
  2 Wings of stationary vanes
  3 Parts where cracks occurred due to thermal fatigue
  4 Spare parts applied at the time of repair
  5, 8, 19 Cooling hole
  6, 7 Flat member
  10 groove
  11, 16 Welded part
  14 Reinforcing member
  15 Corner
  21 Guide section
  22 Guide member
  23A, 23B Ear
  24A, 24B Mounting part
  25, 26 pin holes
  K crack
  K 'residual defect
  W welded joint

Claims (1)

When a crack occurs in the gas turbine stationary blade,
Cutting the back side of the cracked surface,
Groove processing from both sides of the cracked part and filling with weld metal,
A repairing method for a gas turbine stationary blade, wherein the crack is repaired by joining a flat plate member having a cooling hole covering the cut portion.

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