12:126:1262:12622:12622 Damages due to Thermal Fatigue

 Damage due to thermal fatigue

Rising power concentrations and low fuel consumption force an increase in gas temperatures (Fig. "Development curve of thermal strength"). Because the thermal resistance of the available materials does not increase as quickly as the rising hot gas temperatures would require, the cooling of the hot parts must be intensified accordingly. This leads to greater temperature gradients in the engine parts (especially the combustion chamber and turbine) and, in the case of restricted thermal strain, increased thermal stress. Temperature cycles of the parts follow the gas temperature (i.e. engine power output). This also makes thermal fatigue damages more likely. The largest gradients are to be expected during startup and shutdown (Figs. "Loading of a turbine disk in the startup phase" and "Turbine disk loads during operation cycles").
Thermal fatigue damages are explained and discussed in the chapters dealing with individual engine components such as combustion chambers (Chapter 11.2.2) and turbines (Chapter 11.2.3). For this reason, the following text only treats a few exemplary damages and explains the influence of coatings on thermal fatigue.

 Cracks in integral turbine wheels

Figure "Cracks in integral turbine wheels" (Ref. 12.6.2-4): Integral turbine disks tend to annulus crack initiation. These cracks are allowable in some turbines within certain limits. This is a rare exception for cracks in highly-stressed rotating parts. These cracks are thermal fatigue cracks that form during cooling in the compressed annulus areas between the cracks. The compression can be easily recognized by certain crack characteristics (raised crack edges, gaping cracks, see detail). The crack growth rate slows because the annulus zone initially has large, radially-oriented stress gradients during operation. In order to ensure safe controllability of the cracks, they are not allowed to spread from the near-surface pressure zone into a zone with increasing tensile stress during the approved life span (Fig. "Loading of a turbine disk in the startup phase"). There is a danger of uncontrollable crack growth in the latter zone (Ref. 12.6.2-5). If the cracks spread into the blades, fractures due to high-frequency vibrations are likely (Fig. "Typical case of thermal fatigue", Example "Unequal thermal load leading to stress ").
Therefore, the cracks can only be tolerated for life spans that have been shown analytically and in tests to have controllable crack growth rates (i.e. slowing crack growth; Ref. 12.6.2-7). Generally, turbine disks with annulus cracks will not be approved for reinstallation. However, this should not be interpreted as meaning that annulus cracks are generally not tolerated, since the engine evidently “lived with” the cracks until the inspection.
The verification of benign annulus cracks requires that there are no disk vibrations that could cause accelerated crack growth, and that are known to have caused cracked disks to burst in several cases.
Optimized annulus cooling can be a successful remedy for annulus cracks. In the context of design, dangerous circumferential stresses in the annulus can be prevented through the use of a “drained annulus” (see Fig. "Preventing thermal fatigue cracks by design").

 Grain boundaries influencing thermal fatigue

Figure "Grain boundaries influencing thermal fatigue" (Ref. 12.6-4): Conventional cast turbine blades have a typical, axially-oriented columnar grain development (stem crystals) in the areas of the front and rear edges (right diagram). These grains are oriented perpendicular to the dominant loads of centrifugal force and thermal strain. In order to prevent creep cracks, directional solidification (DS) was introduced to orient the grain boundaries radially. Another important factor is the low modulus of elasticity in a lengthwise direction along the crystal orientation (Ref. 12.6.2-21). This leads to lower stress levels in strain-controlled processes like thermal fatigue. Unfortunately, these materials also have typical damage mechanisms that limit their use considerably.
The first DS generation (e.g. MAR M 200) corresponded to the conventional alloys. A small amount of hafnium was merely added to harden the grain boundaries. Due to problems with the casting technology, these alloys were only used in small blades. Because the longer solidification time and unique temperature gradients probably overstressed the form material, casting errors occurred. Diagonally-oriented grain boundaries and considerable micro-porosity additionally reduced strength.
The second DS generation (e.g. based on MAR M 247) contained about 3% rhenium. This increased the creep strength and prevents coarsening of the g' phase during operation. The creep strength approached that of the first single-crystals.
A characteristic weakness of the DS materials is the grain boundaries. Although these are oriented parallel to the main tensile stress direction, they are stressed by thermal stress that acts crossways. The high E-modulus that acts crossways also increases stress levels. This leads to thermal fatigue cracks along these grain boundaries during operation (right diagram). The cracking is promoted by oxidation of the grain boundaries. Due to the rubbing function of the blade tips, this oxidation cannot be prevented with a diffusion coating (Ref. 12.6.2-9). The damage to the blade tips leads to material removal and leakage losses. For this reason, these blades must be replaced even though there is no immediate safety risk.
An additional problem is the development on the grain boundaries of a phase that reduces the amount of g' phase in the grain, to the detriment of the creep strength.

Example "Unequal thermal load leading to stress " (Ref. 12.6.2-6):

Excerpt: “…The failure…was the fourth such failure experienced by the Coast Guard in 14 months…
…the maximum gross weight of the …aircraft was reduced from 8,700 lb. to 8,200 lb., and maximum continuous cruise engine operating temperatures have been reduced from an interim 749 °F (398°C) to 700 °F (371°C)…The directive also instructs…pilots to avoid cycling the engines on the ground and to minimize the time spent conducting operations in which the loss of a single engine could be critical, such as hovering near a ship…
The new restrictions are in effect `for the short term', possibly until newly designed replacement power turbine wheels become available. For now, however, …(OEM) personnel will conduct a one-time dye penetrant and eddy current inspection of the fleet-about 82 operational …(helicopters) and 280…engines…
Blisk failures first surfaced in the …(engine type) about six years ago,' but we weren't able to determine the exact cause of the failures until a little over two years ago,'..the program director, said. At that time the …(OEM) engineers discovered that the blades and the disk portion of the …(engine's) integrally bladed power turbine were heating and cooling at different rates, which caused stresses and eventually cracks in the blisk's blades. The Coast Guard has recorded cracks about 1/10 of an inch (0.25 mm) and has noted that the trailing edge of the blade, where it attaches to the disk, appears to be the area of greatest susceptibility. Engineers also discovered that cracks had not appeared in wheels with less than 350 operating hours… Following the release of the FAA directive, the Coast Guard instituted a 60-hr. interval between inspections…Since …(the last) failure, the Coast Guard has reduced this inspection interval to 30 hr.
Alternatives considered:
Despite the inspections, however, cracks have gone undetected. A Coast Guard preliminary investigation concluded that the blisk that failed…was cracked at the time of its last eddy current inspection. The in-flight failure occurred less than 1 flight hours after the inspection was completed…The crack was undetected because the inspection apparatus, a sleeve that fits over the blade, was not large enough to cover the area. In addition, the crack was in an area where it could not be seen by the inspection personnel…
According to the …(OEM), the long-term fix to the…blisk problem is development of a new power turbine wheel.”

Comments: The cited text does not indicate why it took years to realize that the thermal stress between the blades and annulus were causing the cracks. These were clearly thermal fatigue cracks in the shape of slightly modified annulus cracks. This damage mechanism has been known for years and is based on the aforementioned temperature differences (Ref. 12.6.2-7). The thermal fatigue cracks should also have been identifiable in a fracture surface inspection. The fact that the relatively short cracks, which are tolerated by other comparable turbines over long operating periods, led to such rapid and spontaneous fractures, could be related to an overlaying high-frequency vibration. This type of consequential damage, which originates in a thermal fatigue crack, has already been made known in other engines.

 Typical case of thermal fatigue

Figure "Typical case of thermal fatigue" (Example "Unequal thermal load leading to stress "): In this case, thermal fatigue cracks (Ills. and 12.2.21) evidently occurred at the blade root. If high-frequency vibrations, which always act on blades, cause the crack to grow rapidly, then blade failure cannot be prevented. Axially-oriented thermal fatigue cracks in the annulus (annulus cracks) are tolerated over long periods in many engines, even if cracked disks found during overhauls are not reused (Ref. 12.6.2-7).

 Coatings sensitive to thermal fatigue

Figure "Coatings sensitive to thermal fatigue": The damage symptoms of coated hot parts with thermal fatigue can be useful when evaluating causal influences and damage mechanisms.

Hot parts are coated for various reasons:

  • Oxidation protection (lengthening part life)
  • Protection against hot gas corrosion (e.g. against sulfidation).
  • Thermal insulation (lowering temperatures to guarantee life spans and/or prevent overheating, improving TF behavior, minimizing cooling air use).
  • Reflection (temperature stabilizing, in oil lines, etc.).

A single coating is generally not enough to meet these operating demands. Therefore, coating combinations are usually selected. A common combination is a metallic bond layer with a ceramic thermal barrier coating over it.
Depending on their properties, coatings have specific weak points and damage mechanisms under thermal cycles. This creates typical damage symptoms (top diagram).
Areas with poor bonds (with the base material or bond layer) may be caused by production, but may also have been created during the operating time (e.g. due to under-oxidation in thermal barrier coatings). These flaws lead to spalling. Brittle coatings (in the case of diffusion coatings, often temperature-dependent) tear (Ref. 12.6.2-9) and/or break out of the coating itself (e.g. thermal barrier coatings). In oxidation-resistant diffusion coatings, a crackle laquer-like cracking pattern may be created in the “brittle temperature range.” This reaches to the base material and only becomes clearly apparent after oxidation (bottom left diagram, Ref. 12.6.2-17). If the oxidation is promoted by hot gas corrosion, it can cause the protective coating to balloon up as in the bottom right diagram.
After a large number of load cycles, ductile coatings may show rumpling that is oriented across the direction of the deformation.
See Fig. "Metallographic findings of thermally stressed parts" for metallographic damage symptoms in hot parts.

 Turbine disk fracture during take off

Figure "Turbine disk fracture during take off": The highest loads in the disk occur during the startup phase, when stresses from centrifugal forces and restricted heat strain combine (Fig. "Loading of a turbine disk in the startup phase"). On the rare occasions that turbine disk fractures occur, it is usually during the startup phase (Example "Low pressure turbine stage ruptured"). This diagram of an older fighter aircraft type from the 1960s is based on a photograph published in the news. An amateur photographer was apparently able to capture the moment of takeoff when turbine disk damage occurred. The separation of the fuselage near the engine is clearly visible, and is most likely the consequence of rotor failure. It is assumed that this case involved a disk that had already been damaged by cracking (also see Ref. 12.6.2-18).

Example "Low pressure turbine stage ruptured" (Ref. 12.6.2-11):

Excerpt: “The flight had an uncontained failure of the No. 4 engine during the takeoff roll. The takeoff was rejected and the aircraft stopped on the remaining runway….Examination of the engine indicated the low pressure turbine fifth stage hub had ruptured. About 180 degrees of the hub rim had separated along with the blade attachment slots and blades. The hub ruptured from an area of a well oxidized, intergranular fracture that originated at a tierod hole. There was no apparent defect observed along the edge of the tierod hole that would cause the crack to initiate and propagate to critical length. Probable cause: failure of the …fifth stage turbine hub due to cyclic stress rupture …“

Comments: In this case, the fracture did not originate in the hub bore, but in a bolt bore located farther out. Evidently, the highest startup stresses in this part occur in the annulus area.

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12/126/1262/12622/12622.txt · Last modified: 2020/08/18 20:34 by