EP2614219A2

EP2614219A2 - Method for component testing operating of a machine.

Info

Publication number: EP2614219A2
Application number: EP11771141.6A
Authority: EP
Inventors: Andreas Böttcher; Tobias Krieger
Original assignee: Siemens AG
Current assignee: Siemens AG
Priority date: 2010-11-02
Filing date: 2011-10-19
Publication date: 2013-07-17
Also published as: WO2012059324A2; EP2447470A1; WO2012059324A3; US20130213122A1

Abstract

Core drilling of a used component enables simultaneous refurbishment of the component and the state of the component to be analysed.

Description

Component verification and method of operating a machine

The invention relates to the inspection of a component after use in a machine and to a method for operating a machine with refurbished components.

Components of a gas turbine are designed for a specific life ^¬ duration and examined at various inspection intervals for damage. It is an objective, as far as possible to estimate the remaining ^¬ life of a component to failure to replace the affected component as optimally as possible and pinpoint shortly before the failure limit or repair. Such a prediction reduces the cost of unnecessary replacement of affected components that are still far from their failure limit.

So far, at various inspection intervals, components are visually inspected for cracks or similar signs of operational fatigue. This investigation is very inaccurate and can not say anything about the nature of the construction ^¬ partly inside to. Furthermore, components are destroyed in order to obtain information about their operating stress and mechanical properties inside. This leads to the fact that components can be examined in detail, but these are not used again after the examination. It is also possible that components be processing ^¬ tet are again (re uvenation) and are used again.

It is therefore an object of the invention o. G. Solve a problem.

The object is achieved by a method according to claim 1 and a method according to claim 14, 15.

The subclaims list further advantageous measures which can be combined with one another as desired in order to achieve further advantages. It is proposed to take samples of the component from one or more targeted core drillings and to deduce the operating load from these samples. With this core, all known investigation methods can be carried out, so that with the help of these results a much more accurate statement can be made about the remaining life of the component. As a result, components can be brought closer to the actual load limit and unjustified component changes are avoided. The extracted samples to the component geometry can preferably with a pin close again, so that the inspected component as ^¬ can be used for the operation of the.

Show it

Figure 1 - 4 schematically steps of the invention

Method,

FIG. 5 shows a turbine blade,

FIG. 6 shows a combustion chamber,

Figure 7 is a gas turbine and

Figure 8 is a list of superalloys

The figures and the descriptions represent only embodiments of the invention.

FIG. 1 shows a component 1 which was exposed to high temperatures (HT).

The component 1 is preferably a turbine blade 120, 130 which was in use and preferably consists of or comprises a superalloy, according to FIG. 8. Such components 120, 130, 155, which were in high temperature use, have layers which are then removed for reuse. Optionally, re uvenation measures (for example solution heat treatment and re-precipitation of precipitable hardenings) with the component 1, 120, 130, 155 are also carried out. FIG. 2 shows that at least one core 4, in particular only one core 4, is removed from the component 120, 130 at at least one point.

The core 4 can be removed from the component 1, 120, 130 before or after its re-uvenation measures.

Likewise, several cores 4, in particular taken at different locations. This provides an investigatable core 4.

The core 4 can have any shape.

The core 4 can also be further divided for the investigations (chemical, mechanical, optical) and / or used in succession.

Also preferably, the core 4 'have the same Wärmebe ^¬ actions of the rejuvenation as the component 1' is traversed, and are then tested mechanically and / or microscopically. Preferably, only mechanical tests are performed.

This core 4 can, as shown schematically in Figure 3 as a sample 4 ', to determine mechanical strength values under application of a force F (LCF, HCF, strength in general) are examined.

The component 1, 120, 130, 155 or the substrate to which a core was taken out 4 is filled with a filling material 7 again be ^¬.

The filling material 7 can be preferably soldering material, welding material ^¬ be a soldered or welded pin.

The process has the advantage that the component 1, 120, 130, 155 can be installed and used again and off ^¬ say can be made about the actual possible load limit or time later, so no time delay in replacement for used components (120 , 130) arise when the examinations are time-consuming, such as "low cycle fatigue". 5 shows a perspective view of a rotor blade 120 or guide vane show ^¬ 130 of a turbomachine, which extends along a longitudinal axis of the 121st The turbomachine may be a gas turbine of an aircraft or a power plant for power generation, a steam turbine or a compressor.

The blade 120, 130 has, along the longitudinal axis 121, a fastening area 400, an adjacent blade platform 403 and an airfoil 406 and a blade tip 415.

As a guide blade 130, the blade 130 may have at its blade tip ^¬ 415 another platform (not shown).

In the mounting region 400, a blade root 183 is formed, which serves for attachment of the blades 120, 130 to a shaft or a disc (not shown).

The blade root 183 is, for example, as a hammerhead out staltet ^¬. Other designs as Christmas tree or Schwalbenschwanzfuß are possible.

The blade 120, 130 has for a medium which flows past the scene ^¬ felblatt 406, a leading edge 409 and a trailing edge 412.

In conventional blades 120, 130, in all regions 400, 403, 406 of the blade 120, 130, for example, ^massive metallic materials, in particular superalloys, are used.

Such superalloys are known, for example, from EP 1 204 776 B1, EP 1 306 454, EP 1 319 729 A1, WO 99/67435 or WO 00/44949.

The blade 120, 130 can be made by a casting process, also by directional solidification, by a forging process, by a milling process or combinations thereof. Workpieces with a monocrystalline structure or structures are used as components for machines which are exposed to high mechanical, thermal and / or chemical stresses during operation.

The production of such monocrystalline workpieces takes place e.g. by directed solidification from the melt. These are casting processes in which the liquid metallic alloy is transformed into a monocrystalline structure, i. to the single-crystal workpiece, or directionally solidified.

Here, dendritic crystals are aligned along the heat flow and form either a columnar grain structure (columnar, ie grains that run the entire length of the workpiece and here, in common parlance, referred to as directionally solidified) or a monocrystalline structure, ie the whole workpiece be ^¬ is made of a single crystal. In these processes, it is necessary to avoid the transition to globulitic (polycrystalline) solidification, since undirected growth necessarily forms transverse and longitudinal grain boundaries, which negate the good properties of the directionally solidified or monocrystalline component.

The term generally refers to directionally solidified microstructures, which means both single crystals that have no grain boundaries or at most small angle grain boundaries, and stem crystal structures that have probably longitudinal grain boundaries but no transverse grain boundaries. These second-mentioned crystalline structures are also known as directionally solidified structures.

Such methods are known from US Pat. No. 6,024,792 and EP 0 892 090 A1.

Likewise, the blades 120, 130 may have coatings against corrosion or oxidation, e.g. B. (MCrAlX; M is at least one element of the group iron (Fe), cobalt (Co),

Nickel (Ni), X is an active element and stands for yttrium (Y) and / or silicon and / or at least one element of the rare earths, or hafnium (Hf)). Such alloys are known from EP 0 486 489 B1, EP 0 786 017 B1, EP 0 412 397 B1 or EP 1 306 454 A1.

The density is preferably 95% of the theoretical

Density.

A protective aluminum oxide layer (TGO = thermal grown oxide layer) is formed on the MCrAlX layer (as an intermediate layer or as the outermost layer).

Preferably, the layer composition comprises Co-30Ni-28Cr-8A1-0, 6Y-0, 7Si or Co-28Ni-24Cr-10Al-0, 6Y. Besides these cobalt-based protective coatings, nickel-based protective layers such as Ni-10Cr-12Al-0.6Y-3Re or Ni-12Co-21Cr-IIAl-O, 4Y-2Re or Ni-25Co-17Cr-10A1-0, 4Y-1 are also preferably used , 5Re.

On the MCrAlX may still be present a thermal barrier coating, which is preferably the outermost layer, and consists for example of ZrC> 2, Y2Ü3-Zr02, ie it is not, partially ^¬ or fully stabilized by yttria

and / or calcium oxide and / or magnesium oxide.

The thermal barrier coating covers the entire MCrAlX layer.

By suitable coating methods, e.g. Electron beam evaporation (EB-PVD) produces stalk-shaped grains in the thermal barrier coating.

Other coating methods are conceivable, for example atmospheric plasma spraying (APS), LPPS, VPS or CVD. The heat insulation layer may have ^¬ porous, micro- or macro-cracked ^compatible grains for better thermal shock resistance. The thermal barrier coating is therefore preferably more porous than the

MCrAlX layer.

Refurbishment means that components 120, 130 may have to be freed of protective layers after use (eg by sandblasting). This is followed by removal of the corrosion and / or oxidation layers or products. Optionally, even cracks in the component 120, 130 are repaired. Thereafter, a the coating of the component 120, 130 and a renewed use of the component 120, 130.

The blade 120, 130 may be hollow or solid. If the blade 120, 130 is to be cooled, it is hollow and also has, if necessary, film cooling holes 418 ^(indicated by dashed lines) on.

FIG. 6 shows a combustion chamber 110 of a gas turbine.

The combustion chamber 110 is configured, for example, as so-called ^an annular ^combustion chamber, in which a plurality of in the circumferential direction about an axis of rotation 102 arranged burners 107 open into a common combustion chamber space 154 and generate flames 156th For this purpose, the combustion chamber 110 is configured in its entirety as an annular structure, which is positioned around the axis of rotation 102 around.

To achieve a comparatively high efficiency, the combustion chamber 110 is designed for a comparatively high temperature of the working medium M of about 1000 ° C to 1600 ° C. To even with these operating parameters unfavorable for the materials, a comparatively long operating time to be made ^¬ union, the combustion chamber wall 153 is provided on its side facing the medium M, Arbeitsme- facing side with a formed from heat shield elements 155. liner.

Each heat shield element 155 made of an alloy is equipped on the working fluid side with a particularly heat-resistant protective layer (MCrAlX layer and / or ceramic coating) or is made of high-temperature-resistant material (solid ceramic blocks).

These protective layers may be similar to the turbine blades, so for example MCrAlX means: M is at least one element of the group iron (Fe), cobalt (Co), nickel (Ni), X is an active element and stands for yttrium (Y) and / or

Silicon and / or at least one element of the rare earths, or hafnium (Hf). Such alloys are known from the EP 0 486 489 B1, EP 0 786 017 B1, EP 0 412 397 B1 or EP 1 306 454 A1.

On the MCrAlX can still be present, for example, a ceramic thermal insulation layer and consists for example of ZrC> 2, Y203-ZrC> 2, ie it is not, partially or fully ^¬ dig stabilized by yttrium and / or calcium oxide and / or magnesium oxide.

Other coating methods are conceivable, for example atmospheric plasma spraying (APS), LPPS, VPS or CVD. The heat insulating layer can ^¬ ner to have better thermal shock resistance porous, micro- or macro-cracked pERSonal.

Reprocessing (Refurbishment) means that ^heat shield elements may need to be removed 155 after use of protective layers (for example by sandblasting). This is followed by removal of the corrosion and / or oxidation layers or products. If necessary, cracks in the heat shield element 155 are also repaired.

This is followed by a recoating of the heat shield elements 155 and a renewed use of the heat shield elements 155.

Due to the high temperatures inside the combustion chamber 110 may also be provided for the heat shield elements 155 and for their holding elements, a cooling system. The heat shield elements 155 are then, for example, hollow and may still have cooling holes (not shown) which open into the combustion chamber space 154.

FIG. 7 shows by way of example a gas turbine 100 in a longitudinal partial section. The gas turbine 100 has a rotatably mounted about a rotational axis 102 ^¬ rotor 103 having a shaft 101, which is also referred to as the turbine rotor.

Along the rotor 103 successively follow an intake housing 104, a compressor 105, for example, a torus-like

Combustion chamber 110, in particular annular combustion chamber, with a plurality of coaxially arranged burners 107, a turbine 108 and the exhaust housing 109th

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

Each turbine stage 112 is formed, for example, from two blade ^rings . As seen in the direction of flow of a working medium 113, in the hot gas channel 111 of a row of guide vanes 115, a series 125 formed of rotor blades 120 follows.

The guide vanes 130 are fastened to an inner housing 138 of a stator 143, whereas the moving blades 120 of a row 125 are attached to the rotor 103 by means of a turbine disk 133, for example.

Coupled to the rotor 103 is a generator or work machine (not shown). During operation of the gas turbine 100 104 air 135 is sucked by the compressor 105 through the intake housing and ver ^¬ seals. The 105 ^¬ be compressed air provided at the turbine end of the compressor is ge ^¬ leads to the burners 107, where it is mixed with a fuel. The mixture is then burned to form the working fluid 113 in the combustion chamber 110. From there, the working medium 113 flows along the hot gas channel 111 past the guide vanes 130 and the rotor blades 120. On the rotor blades 120, the working medium 113 expands in a pulse-transmitting manner, so that the rotor blades 120 drive the rotor 103 and drive the machine coupled to it. The components exposed to the hot working medium 113 are subject to thermal loads during operation of the gas turbine 100. The guide vanes 130 and rotor blades 120 of the first turbine stage 112, viewed in the flow direction of the working medium 113, are subjected to the greatest thermal stress in addition to the heat shield elements lining the annular combustion chamber 110.

To withstand the prevailing temperatures, they can be cooled by means of a coolant.

Likewise, substrates of the components may have a directional structure, i. they are monocrystalline (SX structure) or have only longitudinal grains (DS structure).

As a material for the components, in particular for the ^turbine blade or vane 120, 130 and components of the combustion chamber 110. For example, iron-, nickel- or cobalt-based superalloys are used.

Also, the blades 120, 130 may be anti-corrosion coatings (MCrAlX; M is at least one element of the group iron (Fe), cobalt (Co), nickel (Ni), X is an active element and is yttrium (Y) and / or silicon , Scandium (Sc) and / or at least one element of the rare earth or hafnium). Such alloys are known from EP 0 486 489 B1, EP 0 786 017 B1, EP 0 412 397 B1 or EP 1 306 454 A1.

On the MCrAlX may still be present a thermal barrier coating, and consists for example of r02, Y203-Zr02, ie it is not, partially or completely stabilized by Ytt ^¬ riumoxid and / or calcium oxide and / or magnesium oxide.

The guide blade 130 has a guide blade foot facing the inner housing 138 of the turbine 108 (not shown here). ) and a vane head opposite the vane root. The vane head faces the rotor 103 and fixed to a mounting ring 140 of the stator 143.

Claims

claims

1. Method for checking a component (1, 120, 130, 155),

in particular for determining the maximum period of use for reuse of the component (1, 120, 130, 155), in which the used component (1, 120, 130, 155) before reuse (1, 120, 130, 155)

and especially before reprocessing

a core (4) is removed at at least one point, which is examined (4) and

its (14) test results for determining the duration of use and / or load data of the component (1, 120, 130, 155) during the re-use of the component (1,

120, 130, 155) are used for reuse.

2. The method according to claim 1,

in which the same heat treatments are carried out with the core (4) as with the component (1, 120, 130, 155).

3. The method according to claim 1 or 2,

in which mechanical examinations are carried out with the core (4).

4. The method according to claim 1, 2 or 3,

in which chemical investigations are carried out with the core (4).

5. The method according to claim 1, 2, 3 or 4,

in which metallographic investigations are carried out with the core (4).

6. The method according to claim 1, 2, 3, 4 or 5,

where the body,

from which the core (4) was taken,

replenished,

in particular by soldering or welding,

in particular by means of a pin.

7. The method of claim 1, 2, 3, 4, 5 or 6,

in which the core (4) is solid,

especially elongated.

8. The method of claim 1, 2, 3, 4, 5, 6 or 7,

in which at various points of the component (1, 120, 130, 155), a core (4) is removed.

9. The method of claim 1, 2, 3, 4, 5, 6, 7 or 8, wherein only a core (4) is removed.

10. The method of claim 1, 2, 3, 4, 5, 6, 7 or 8, wherein a core (4) for the investigations is divided into several parts.

11. The method according to one or more of the preceding claims,

in which the core (4) is removed before reprocessing.

12. The method according to one or more of the preceding claims 1 to 10,

in which the core (4) after reprocessing

is removed.

13. The method according to one or more of the preceding claims,

during reprocessing, stripping and one

Heat treatment means

in particular a solution annealing and a

Excretion heat treatment.

14. Method for operating a machine (120),

having a component (1, 120, 130, 155),

in which a core (4) is removed from the component (1, 120, 130, 155),

in particular according to claim 1, 2, 3, 4, 5, 6, 7, 8, 9, 10, 11, 12 or 13,

in which the component (1, 120, 130, 155) for operating the machine (100) is reprocessed and

is used again in a machine (100) and the examination results of the core (4) take place only after the renewed first use of the remanufactured component (1, 120, 130, 155).

15. A method for operating a machine (120),

having a component (1, 120, 1301, 155),

in which a core (4) is removed from the component (1, 120, 130, 155),

in particular according to claim 1, 2, 3, 4, 5, 6, 7, 8, 9, 10, 11, 12 or 13,

in which the component (1, 120, 130, 155) for operating the

Machine (100) is reprocessed and

in a machine (100) and the results of the examination of the core (4) before the first th use of the remanufactured component (1, 120, 130, 155).