3.3.2 Typical problems of the high pressure turbine - causes and remedies
In the high pressure turbine (HPT), the problems of the blades are of primary significance. As described in Chapter 3.3.1.2, the turbine nozzle guide vanes are especially endangered because of over heating through hot streaks in the gas flow ( "Ill. 3.3-9"). But turbine rotor blades can also be clearly overheated ( "Ill. 3.3-10"), even if temperature irregularities at the periphery are averaged through their rotation. Burns and meltdowns are typical. These are revealed, in extreme cases, through the missing of an entire blade section. Signs for overheating are detached diffusion coatings, the orange peel effect ( "Ill. 3.3-10" and "Ill. 3.3-12") and cracks at the leading edge, as well as strong local oxidation (‘burning’) and surface-fuse.
Typical causes for the overheating of the turbine blades are:
- The unfavorable temperature profile of the hot gas flow behind the combustor ( "Ill. 3.2.3-2").
- Lack of cooling air supply through blockage or deformation of the cooling air passages ( "Ill. 3.3-11").
- Bad heat removal because of oxidation layers at the passage walls ( "Ill. 3.3-12").
- High gas temperatures, on the grounds of malfunctioning of other components ( "Ill. 3.3-11").
Rust is especially blockage effective. Thus, it is imperative to take care that the inlet duct and all mountings are free from rust. The deformation of cooling air passages can originate through internal foreign objects (OOD, "Ill. 3.3-10"). Spalling of thermal barrier coatings is a typical cause (Chapter 3.2.3). Coke particles from the combustor are, indeed, not really a theme for the natural gas user. If a soot forming fuel is used for the start or, partly, for operation, the danger of carbon impact occurs. Oxidation and hot gas corrosion in cooling air passages can be facilitated through the residue of cleaning processes (etch product), but also through imported foreign material in repair and production procedures (e.g., low melting phases, Lit. 3.3-12). Especially dangerous are low melting positioning metals in the production process (alloys of lead, antimon, bismuth) that can react with the base material and trigger cracks. Melting drops from laser or electron beam processing are also able to block cooling holes and channels.
When there is suspicion of overheating, the combustion chamber is basically also to be examined for a possible unfavorable temperature profile ( "Ill. 3.2.3-2"). Characteristic are distortion, crack formation and break-outs as well as unusual coke formation, damaged or blocked fuel nozzles and so on ( "Ill. 3.2.3-1").
To initiate directed remedies also overheating without distinctive external failures must be identified. Hot parts like turbine blades consist of Ni-based materials that persipitation harden with the so called gamma prime phase ( ’), a special structure component. The experienced specialist can with the scanning electron microscope (SEM) from amount, size, form, orientation and alignment of this phase conclude on the temperature effect. Problematic is that longer operation times at normal temperatures, followng the over heating, change the assessable picture. A further possibility of later temperature determination comprises the metallographic examination of protective coatings (diffusion coatings, deposited coatings) on particular changes ( "Ill. 3.3-7").
In turbine rotor blades that are subject to high centrifugal forces, so called creep voids (pores) in the structure or on the fracture surface can give hints at the effective failure mechanism. ( "Ill. 3.3-13"). The usual casting alloys today do, other than the wrought alloys of elder engines, show unfortunately for the failure investigation no intensive creep void formation. The lack of creep pores does not imply the opposite conclusion: creep was not the cause of failure.
Yet, not only is the height of the temperature of importance, the temperature distribution itself has great effectiveness ( "Ill. 3.3-1"4). The failure mechanism pertaining to it is thermal fatigue ( "Ill. 3.3-1"). Crack initiation at the leading edge of the blade and in the interior (!) of the components, i.e., in the walls of the cooling air passages, is typical ( "Ill. 3.3-1"3, "Ill. 3.3-1"4 and "Ill. 3.3-1"5). The phenomenon of increased crack initiation in the relatively cold wall of the cooling air passages is, indeed, at first, not to be explained with creep damage. It relates here to another type of failure mechanism, a cyclical fatigue process by which the stress originates through hampered thermal expansion ( "Ill. 3.3-1"6) thermal fatigue. Here the cold areas find themselves under tensile stresses, which in turn balance with the compression stresses of the hot area with higher thermal expansion. In nozzle guide vanes, thermal stresses between the platforms and the blade airfoil lead to thermal fatigue cracks in the transition radius ( "Ill. 3.3-9" and "Ill. 3.3-1"7). Such cracks are frequently tolerated by the manufacturer (OEM) up to a certain size, depending on the gas bending load of the blades and /or the leakage air rate (as a consequence of the cracks). Among other things, the compression stresses in the airfoil are so high, they buckle out.
At nozzle guide vanes consisting of many vanes, the failures can concentrate on one special vane. This is the case when this airfoil is hindered in its deformation through the neighboring vane of the segment. Too high component temperatures (up to the area of the softening) favor a local inflation of the blade airfoil wall ( "Ill. 3.3-10") through the cooling air pressure.
Today, turbine nozzle guide vanes of modern engines often show ceramic thermal barrier coatings. Such coatings have attained a level of development through which typical problems in the coating like flaking and erosion ( "Ill. 3.2.3-8") can be overcome. Even so, one should be aware that such coatings react with melt downs from dust deposits and other air impurities and can fail early ( "Ill. 3.2.3-7" and "Ill. 3.2.3-8"). As already mentioned in Chapter 3.2.3 (Combustors), because of the ion conductivity of the hot ZrO2 coating, there is the danger of oxygen transport to the bond coating, producing long term oxidation of the bond coat and a corresponding adhesion decrease.
"Illustration 3.3-9": The turbine nozzle guide vane sees all temperature dissimilarities of the hot gas flow ( "Ill. 3.2.3-2"), radial and periphery, in contrast to the rotor blade. Here, especial attentiveness during hot part inspections is valid. This is one of the most repair intensive components.
Some typical failure appearances:
- „A“: Orange peel effect in the area of high component temperatures through increased oxidation and shallow thermal fatigue cracks.
- „B“: Crack initiation as a consequence of thermal fatigue. The cracks are usually not deep, but heavily oxidized and broadened.
- „C“: Through over temperatures, burned (extreme oxidized) and/or melted regions, airfoil portions miss noticeably .
- „D“: Rippling (‘Reeling’, corrugation) diffusion coating show component temperatures in the region of softening of the base material. The effect is produced through local melting of an Al rich coating zone to the base material ( "Ill. 3.3-7").
- „E“: Burned surface and hot gas erosion in the platform area. This shows a dearth of film cooling air locally.
- „F“: Coating cracks in the coating, applied against oxidation and hot gas corrosion ( "Ill. 3.3-7").
- „G“: Radial crack in the blade airfoil at the pressure or suction side along the grain boundaries (heat tears, hot cracks), especially in directionally solidified components ( "Ill. 3.3-4"). Normally cast vanes show cracks preferrential along the transition to the platforms (see “I”). Mainly a hint of high thermal stresses and especially high component temperatures.
- „H“: Buckling and deformations of the airfoil. This effect is to be traced back to big thermal expansion differences between airfoil and platforms. High compression stresses lead to creep deformation. A hint of short term, especially high local temperatures, e.g., at the start or the acceleration.
- „I“: Typical crack formation through high thermal stresses and thermal fatigue at the transition blade airfoil/platforms ( "Ill. 3.3-17", see also “H”). These cracks are permitted up to a certain length specified by the manufacturer, because of reduced crack propagation.
"Illustration 3.3-10": It does not harm, if the operator is acquainted with the terminology of the appearance of typical failures of the hot parts. This case emerges, e.g., during borescope inspections ( "Ill. 4.1-5" and "Ill. 4.1-7"), in the frame of a hot part inspection. We assume a borescope finding is mentioned to the operator of the engine and perhaps actually shown or examined from photos at hand. If the assessment of the finding demands complicated and costly action, an understanding of the technical interrelationships is definitely a valuable help to decisions for the operator.
Typical findings in high pressure turbine rotor blades:
- „A“: ‘Orange peel effect’ on the grounds of stronger oxidation, with thermal fatigue cracks in a zone of high operation temperatures. We find this failure mode preferential at the blade leading edge.
- „B“: ‘Hot spot’ is shown through intensive roughness (‘ripples’), as a consequence of oxidation and corrugated formation in the diffusion coating. The term corrugation, indicates a magnification of roughness of the surfaces of a hot part through plastic deformation and corrugation of coatings under thermal fatigue. In extreme circumstances, a bulging of the blade wall develops, hinting at a weakness of the protective film cooling air (e.g., blocking).
- „C“: An especially neuralgic point is the blade tip (Lit.3.3-6) of high pressure turbine blades without shrouds (see detail). Ruptures and burns in the thick area of the blade tips are typical. Clearance minimization (rubbing) signifies a lack of oxidation protection coating (diffusion coating) together with heavy oxidation attacks. The clearance increase leads to a noticeable influence of the efficiency of the engine with this component. Hence, special efforts in the form of development programs with the goal to attain an improved long term behavior are being made. To this belong the application of hard particles (cubic boron nitride or coated silicion carbide).
- „D“: Burned away, upper blade leading edge, e.g., after a FOD by spalled off particles of thermal barrier coatings ( "Ill. 4.1-7" „C“) or a blockage of the cooling air passages in the blade tip region.
- „E“: Crack initiation in the middle of the pressure side of the blade airfoil. Such a damage can also be distinguished through local discoloration. Thermal fatigue cracks that start from the inner cooling air holes are possible. ( "Ill. 3.3-15").
- „F”: Ripple formation (rumpling, "Ill. 3.3-7" “C”) at the pressure side in the middle of the blade. This is an increased, orientated roughness. Prone are protection coatings which show a wave structure under cyclic plastic deformation by thermal fatigue.
- „G”: Crack formation at the tip of directionally cast blades. This concerns a combination of oxidation and thermal fatigue.
Not represented are findings like deposits and residues in the region of film cooling holes. They can indicate a critical cooling air pollution induced through melted dust. To these belong also abrasion particles of rub in coatings ( labyrinths, casings, "Ill. 3.3-4").
A special problem with regard to borescope findings is wrong interpretations of line shaped indications that could signify a dangerous crack ( "Ill. 3.3-5" , "Ill. 3.3-15" and "Ill. 4.1-7" „G“) or a harmless deposit. Expert knowledge is necessary here.
"Illustration 3.3-11": The component temperature of a turbine rotor is dependent on several influences. Especially important is the temperature distribution in the hot gas and the cooling air supply.
„1,2,3“ show different temperature profiles in the hot gas flow. One should aim at a uniform gradient ( "Ill. 3.2.3-2") with lower maximum temperatures („1“). For this purpose, the combustor (distortion, crack formation) itself, but also the cooling air entrance into the hot gas from previous blades, labyrinths and static gas guiding parts, is decisive. Hot gas ingestion at the rim area („4“) of the disc, e.g., as a consequence of a seal failure. The detail at the right shows the result of an extreme overheating of the disk rim. It came to clearly visible plastic deformation of the fir trees and centrifuging of the blades. „5,6,7“ indicate the cooling air flow around the disc and towards the blade. Changes in this region can lead to an unusually long time change of the sealing systems and increased leakages. The manufacturer had already considered the normal deterioration.
"Illustration 3.3-12": The effectiveness of the cooling of a turbine blade can be impaired through different influences. In addition, there is the danger of overheating and noticeable life reduction ( "Ill. 2.3-2").
- „A“: Narrowing of the film cooling holes near the leading edge through foreign object impact. Typical objects in the turbine are coke particles from the combustion chamber ( carbon impact) and spalled off thermal barrier coatings ( "Ill. 4.1-7" “C”)
- „B“: High operation temperatures initiate an isolating oxide layer in the cooling air holes (see also “E”). The less well cooled , hotter blade wall oxidizes more quickly. This is a self accelerating process.
- „C“: Blockage of the cooling air passage through dusts in the cooling air. Typical are, e.g., abrasives from labyrinth and casing liners ( "Ill. 3.1.2.4-4").
- „D“: Blockage of dust holes by much dust in the cooling air.
- „E“: Deterioration of thermal conduction through an oxide layer in the region of impurities in the cooling air passage. This can be the consequence of insufficient rinsing of the blade after an aggressive cleaning during overhaul. Compare with „B“.
"Illustration 3.3-13": A creep damage ( "Ill. 2.3-2") goes, in many materials, (not materials without grain boundaries, like single crystals), together with a formation of pores (so called creep pores or creep voids, Lit.3.3-2) preferably on to the grain boundaries, transverse to the tensile stress. Fe-base and Ni- base wrought alloys show this effect pronounced. Not so castings with the typical coarse grain and, understandable, single crystals.
The creep pores grow together with increasing life and can finally lead to creep rupture with typical, irregular surfaces (Lit. 3.2-25).
An evaluation of creep damage to draw conclusions for further implementation, perhaps, (rest life assessment) of not yet visibly damaged components is difficult, demanding many assumptions, requiring wide experiences concerning the affected parts and the behaviour of the used material under special operation conditions.
"Illustration 3.3-14": Typical temperature distribution of an intensive cooled turbine blade with large temperature gradients. Around the cooling air holes inside the blade, the temperatures are relatively low. These zones absorb a big portion of the centrifugal force related tensile stresses that are here especially high. With the tensile heat stresses cracks can form in the walls of the cooling holes ( "Ill. 3.3-15"). The cooling is so arranged that the blade temperature towards the root lessens, in order to justify the increasing centrifugal load and to keep the disk temperatures (on the disc serrations) low. The edges are, despite intensive cooling, especially hot and relieved through the greater thermal expansion of the centrifugal force stress. Blade edges are especially affected by oxidation and microstructure changes because of their high temperatures.
The irregular temperature distribution and thermal load in the cross section of the blade is also the reason why a simple growth measurement of the blade length towards the determination of creep elongation for the purpose of evaluation of the remaining life is usually unsuitable. Creep deformations lead rather to distortion of the airfoil and not to assessable elongation. The latest measurement methods with laser could promise success. They collect the entire blade geometry and compare the new blade with the one in operation.
"Illustration 3.3-15": In thermal stresses, the deformation hindered colder regions of a component are usually highly tensile stressed. The warmer regions with greater thermal expansion get under compression and produce in colder regions tensile stresses through deformation hindrances till a balance is reached. An example is the hub of turbine discs ( "Ill. 3.3-5"). A further typical example are cooled hot parts such as turbine blades ( "Ill. 3.3-14"). Around the cooling air holes, tensile stress zones are formed, from which thermal fatigue cracks could initiate ( "Ill. 3.3-14"). This can be first distinguished from the outside visually, if they appear through the airfoil surface ( "Ill. 3.3-10") and have noticeably weakened the component. This form of failure can be supported through oxidation or hot gas corrosion in the cooling air holes.
"Illustration 3.3-16": Thermal fatigue is a component damage with crack initiation through cyclic thermal stresses. An understandable example is the crack formation in turbine guide vanes ( "Ill. 3.3-17"). The sketch shows a model of the load procedures in the upper part.
A metallic rod is adjustable limited by two walls. This rod will expand on being heated up. The thermal expansion „lt“ occurs, without hindrance, through the walls. Through the expansion hindrance of the walls, a plastic compression „lp“ follows when overstepping the yield strength. After cooling, the rod around this compression is shorter. There is a gap between the wall and the rod. The rod is now relieved.
We imagine the rod would now be firmly connected to the walls at the beginning of the trial. In this case, there follows again the plastic compression (buckling) through high compression stress during heating. During cooling, there are now tensile stresses in the shortened rod. A temperature cycle (T) corresponds to a compression/ tensile stress cycle leading to mechanical fatigue ( "Ill. 3.1.2.1-0") in the LCF region, the so called thermal fatigue. We see then that a gap in the form of a crack between rod and wall leads to a relaxation of stress. This also signifies a slowing down of crack propagation during crack growth. This stress relief enables the OEM to permit component specific crack lengths in manuals and specifications ( "Ill. 3.3-17"). Without this, the safe repair with high temperature brazing frequently used at eroded and cracked hot parts would not be possible.
"Illustration 3.3-17": Thermal fatigue cracks can lead to local relief of the component ( "Ill. 3.3-16"), if they progress in an area of very low tensile or compression stress. Thus, its progress is slower, at least for some time, respectively, load cycles. A reduced crack propagation is only then to be expected, if, in the area of crack initiation, no high mechanical stresses, such as centrifugal forces, are superimposed.
The typical rim crack „A“ in an integral turbine wheel (characteristic for engines of smaller performance) can be controlled and is permitted by the OEM for certain periods of life respectively start-stop cycles.
The hub crack „B“ in contrast, finds itself in a zone of very high tensile stresses that hardly reduce across the cross section. Here also, through small faults and cracks, the danger of a component failure after short crack propagation arises. Such a situation is prohibited!
In the case of turbine nozzle guide vanes, typical thermal fatigue cracks in the transition platform/ airfoil „C“ are not seldom. These cracks are (specified length by the manufacturer) permitted up to a certain point, if they run in a compression stress zone.
"Illustration 3.3-18": Integral turbine wheel as is usual for small gas turbines:
- (1) Cracks through rub.
- (2) Cracks through blade vibrations in the HCF range.
- (3) Disc rim cracks through thermal fatigue.
- (4) Labyrinth cracks through rub and cycli- cal stress in LCF and / or HCF range.
- (5) and (6) fatigue fractures through LCF, on grounds of centrifugal force and temperature changes, especially at the start.
- (7) and (8) Crack formation through disc vibrations in the HCF area and/or LCF cracks through thermal change and centrifugal force alterations.
- (9) Crack formation through thermal fatigue and /or overheating and oxidation ( "Ill. 3.3-10").