Aeroengine Safety

Figure "High pressure turbine rotor blades overhaul intervals": Typical appearances of temperature-related damage to turbine rotor blades. This information is useful for evaluating boroscope data in the framework of a hot part appraisal. Further information regarding the specific damage mechanisms can be obtained from the indicated literature.

“A”: The orange peel effect occurs in the case of heavy oxidation with thermal fatigue cracking in zones with high operating temperatures (e.g. blade inlet edges, Fig. "Turbine guide vane thermal damages").

“B” Hot spots: Increased oxidation leads to roughening and rippling in the diffusion coating (indicated diagram Fig. "Coatings sensitive to thermal fatigue"). Rippling is a pronounced oriented roughness on the surface of a hot part. It is created by plastic deformation and wave formation in protective coatings due to thermal fatigue. In extreme cases, the blade wall may creep and bulge due to the cooling air pressure (Ref. 11.2.3.2-16, Example "Bulging turbine blades"). This type of bulge indicates a weakness in the protective cooling air film and/or poor inner cooling.

“C”: An especially critical spot is the blade tips (Ref. 3.3.2-2) of shroudless high-pressure turbine blades (see detail). Scorching and break-outs in the seal area of the blade tips are typical (Fig. "Overheating signs on torbine rotor blades"). The clearance gap-minimizing rubbing function prevents the use of anti-oxidation coatings. This means that unimpeded oxidation attack is possible in this especially hot engine zone. In these parts, especially, clearance gap increases have a considerable influence on efficiency (loss of power, SFC increases; Volume 2, Chapter 7.1.2) and are highly undesirable.

“D”: Burned-off upper blade inlet edge (Ref. 11.2.3.1-11). This type of damage can be caused by foreign objects (Volume 1, Chapter 5.2.1.1) or be the result of a blockage in the cooling air duct in the blade tip area (Fig. "Blockages in turbine blade cooling system").

“E”: Cracking in the middle of the pressure side of the blade. Local discoloration can also indicate this type of damage. These are probably thermal fatigue cracks that originate in the inner cooling air bores (Fig. "Damaging colder part zones by thermal fatigue", Example "Airflow leak causing fatigue failure").

“F”: Particle erosion is a damaging influence that is frequently underestimated. The erosive particles can be created in the engine itself. Typical examples include particles from thermal barrier coatings on upstream components (e.g. combustion chambers, Fig. "Typical combustion chamber damages") and soot or coke particles (Fig. "Carbon erosion at high pressure Turbine blades and vanes"). The particles may also be ingested sand. In this case, the erosive effect can be increased by an additional chemical attack (Volume 1, Ill. 5.3.2-12.1).

Findings such as deposits and coatings near the bores for the cooling air film are not depicted (Example "Dust ingestion"; Ills. 11,2.3.2-2, 11.2.3.1-13 and 11.2.3.1-14). They can be seen as indicators of considerable cooling air contamination. These deposits are made up of melted dusts (e.g. ingested or created internally by processes such as labyrinth rubbing; Volume 1, Chapter 5.3.2).

A special problem with boroscope data is misinterpretation of linear readings. These can be dangerous cracks, but may also simply be harmless deposits. In order to be certain about one`s conclusions, similar appearances on other blades and/or other part areas in which the probability of a specific damage type is known can be helpful.

Figure "Blockages in turbine blade cooling system" (Ref. 11.2.3.2-6): Film-cooled turbine rotor blades in the intermediate turbine of a triple-shaft fighter aircraft type showed a greenish glass-like coating after a few hundred hours of operation. This appeared in the direction of flow after the cooling air bores on the suction side of the blade (left diagram). Greenish deposits also formed below the root platform. However, they had a rough structure of the type expected from dust deposits. Analysis of these coatings revealed a high nickel content. This composition corresponded to wear products from the compressor housing coatings. The high temperature of the blade surface evidently melted the relatively low-melting particle mixture of dust and wear products (Fig. "Blocking as 'disease' of hot parts cooling systems"). A metallographic cross-section (far left detail) showed a partial blockage of the bores for the cooling air film The probable result would be overheating over longer operating periods.
The right diagram depicts a high-pressure turbine rotor blade which showed signs of considerable overheating on the inlet edge (“A”) after a testing run of a few hours (orange peel effect, Fig. "Turbine guide vane thermal damages"). A metallographic cross-section (far right diagram) revealed that the dust removal bores “B” and the radial cooling air bores at the tip behind the inlet edge were filled with foreign material (“C”). The following radial bore also had deposits.
Analysis of the radial bores revealed a high iron content. It was interesting that the structure of the deposits indicated ductile behavior when they were created. This behavior would increase the blocking effect. Because the engine was run in a testing rig with rusty metal sheeting on the walls of the inlet duct, it was suspected that vibrations caused rust to flake off. The rust was then ingested by the engine and entered the HP cooling air system.

Figure "Blocking as 'disease' of hot parts cooling systems" (Example "Dust ingestion"): In this case, ingested dust (“A”) blocked the high-pressure turbine rotor blades in a civilian fan engine (top diagram). Evidently the dust also had an erosive effect on the abradable coatings in the compressor (“B”). The dust was carried into the blades by the cooling air for the turbine (middle detail diagram). It is not clear whether the coating material removal also contributed to the inner blockage. The dust-removal bores (bottom left diagram) were evidently not sufficient to prevent the blockage.

Figure "Turbine rotor blade tip erosion and oxidation" (Ref. 11.2.3.2-7): The hot gas leakage flow places the tips of turbine rotor blades in increased danger of overheating (Ref. 11.2.3.2-9). This makes the blade tip a factor that limits the life span of the turbine and therefore also that of the entire engine. Cooled shroudless turbine rotor blades (top left diagram) are affected by typical localized overheating in the tip area that takes the form of a missing section of blade wall (burnout, right diagram). The heat transfer is evidently especially intensive in this area. In blades with smooth seal surfaces, the pressure-side edge overheats. The side walls of blades with a recess at the seal surface are subject to even greater thermal loads due to their poor heat removal (no cooling). Damage to the tips of these blades may indicate that the clearance gap flow in this configuration also dangerously overheats the suction side. Damage due to localized extreme overheating (oxidation, top left diagram, “B”) usually reaches from the seal surface (“A”) to the edge of the tip cap (“C”).
The cause, other than rubbing, for especially high blade tip edge temperatures is hot gas leakage flow from the pressure to the suction side (bottom diagram, Ref. 11.2.3.2-10). The leakage flow at the tip of a turbine blade, unlike that at the tips of compressor blades, is accelerated by the high pressure ratios (Ref. 11.2.3.2-8). The pressure decreases to a fraction of its original level over a very limited area of the blade tip edge. This pressure difference, combined with a very thin boundary layer (due to the high flow acceleration), is causally related to the burnout problem. The extreme low pressure at the exit side of the leakage flow is primarily a result of the sharp redirection at the clearance gap entrance. This redirection is especially tight because of the leakage flow along the edge. If the flow separated at the sharp edge, then the radius of the redirection would be considerably larger. The low pressure at the suction-side edge and a separation bubble in the gap contribute to the increase of the clearance gap flow. This means that the flow speeds in the clearance gap are greater than would be expected from the outer pressure distribution. This high hot gas speed with correspondingly high heat transfer coefficients promotes overheating and burnout.

In order to minimize blade tip damage due to localized overheating, the cooling in this area must be appropriately configured. The cooling air volume for the blade tip from the inside of the blade is often insufficient. In this case, the injection of a cooling air film along the pressure-side edge of the blade tip can prevent dangerously large heat transfers from the hot gas into the blade tip (Ref. 11.2.3.2-9).

Figure "Design of rotor blade tip shrouds": Shrouds on turbine blades improve the seal effect at the blade tips and brace the blade against high-frequency vibrations.
The long, slender bodies of the turbine blades in the rear stages (low-pressure turbine) require shrouds even in modern engine types. Unlike newer blade types, high-pressure turbine blades in older engine types are also outfitted with shrouds. Parallelogram-shaped shrouds are usually not braced against one another (right diagram). They only reduce Ährenfeld vibrations slightly (Chapter 12.6.3.2), create gaps under centrifugal loads, and shift against one another when the blades are twisted open (left diagram, “A” and “B”). In order to ensure the proper functioning of the shrouds, the blades are braced against one another in a way that places the blade under torsion stress. This gives the shrouds a Z-shape (interlocking shrouds, left diagram, “C”). The contact surfaces are protected from fretting through the application of armor (e.g. stellite welding; right diagram). These shrouds do not separate from one another under centrifugal expansion. However, the shape-specific notch promotes fatigue cracking (thermal fatigue and/or high-frequency vibrations), especially if the deposit welding unintentionally reaches the notch radius. Overly narrow contact surfaces (“b”) of neighboring shrouds (bottom right detail) cause partial radial overlapping with a positional shift in direction of the circumference (shingling) and cracking. Complete overlapping is called overriding. Radial overlapping causes the labyrinth fins to be angled, which results in more intensive localized rubbing with the danger of material being transferred between surfaces, and the beginning of a self-increasing damage process. Unlatching (middle diagram) is when the shrouds shift axially to the point that they lose contact. In this case, unallowable creep deformation of the blades is to be expected. The axially shifted labyrinth fins must now deal with all the material removal from rubbing on only one of their own circumferential tracks. Major axial shifting of the blades can cause serious rubbing damage to neighboring blade rows.

Figure "Typical damages to rotor blade shrouds": Shrouds on turbine rotor blades are not unproblematic, and exhibit specific types of damage:

Creep deformations:
Untwisting is a permanent straightening of the blade due to centrifugal forces (Ref. 11.2.3.2-15). This breaks down the protective vibration-resistant effect.

Shroud flexure (Ill. 12.5.1-10), Example "Thermal impact by combustion chamber") is usually the result of a radial temperature profile that has been shifted outward (Ills. 11.2.2.1-8 and 11.2.3.1-11). In extreme cases, a corner of the shroud can break off.

Shingling is the overlapping of the shroud corners of neighboring blades (Fig. "Design of rotor blade tip shrouds"). Shingling is promoted by flexure of the shrouds and by one-sided fretting of the contact surfaces, which causes them to become very thin and slide over each other.

Fretting wear on neighboring blades promotes shingling. This is generally a combination of dynamic wear and hammering wear. This must be taken account when confirming the suitability of remedies and improvements: dynamic wear can break down the bracing of Z-shrouds and increase the gaps around parallelogram shrouds. This increases the probability of dangerous blade vibrations.

Fatigue cracking can occur in the corners of the shroud interlocking due to

• thermal fatigue
• load cycles under RPM changes
• high-frequency vibrations

as well as combinations of these factors. If alterations are made in an area where cracking can potentially occur, then these dynamic loads must be taken into account. Changes that could lower dynamic strength must be considered very carefully. This includes coatings, which can often result in “disimprovements” (Volume 1, Chapter 3).
Cracks due to thermal fatigue can form in the blade transition underneath the shrouds. Temperature gradients between the shrouds and blades in areas with notch effects (structural notch, stiffening notch, form notch) promote cracking.

Hot cracks and wear in seal fins caused by rubbing. The seal fins of the individual blades form a labyrinth around the circumference and are susceptible to the typical damage mechanisms of this seal type (see Volume 2, Chapter 7.2.2)

Figure "Temperature caused damages at high pressure turbine vanes": Unlike rotor blades, turbine stator vanes are subjected to all uneven temperature zones (radial and circumferential) in the hot gas flow. Stator vanes are the focal point of hot part inspections, and are one of the most repair-intensive engine parts.

Some typical damage symptoms:

“A”: Orange peel effect due to increased oxidation in zones with high part temperatures.

“B”: Cracking due to thermal fatigue. The cracks are usually shallow, but are heavily oxidized and gaping.

“C”: Burned (extremely oxidized) and/or melted zones due to overtemperature. Wall sections are missing (Ref. 11.2.3.2-12).

“D”: Delaminating diffusion coating indicates part temperatures in the softening range. The effect is caused by local melting of an Al-rich coating zone to the substrate.

“E”: Surface burning and hot gas erosion in the shroud area. This symptom indicates localized weaknesses in the cooling air film.

“F”: Cracks in the coating that is designed to protect against oxidation and hot gas corrosion (Fig. "Coatings sensitive to thermal fatigue"). The cracks occur primarily in temperature ranges in which the coating behaves brittly. These are usually low temperatures.

“G”: Radial cracks in the blade on the pressure or suction side along the grain boundaries of parts that are directionally solidified. This usually indicates high thermal strain and especially high part temperatures.

“H”: Folding and deformation of the blade (Ref. 11.2.3.2-12, Fig. "Creep damages at turbine inlet guide vanes "). This creep effect is caused by large thermal strain differences between the blade and platform. It is an indication of the short-term occurrence of especially high localized temperatures, e.g. during startup or acceleration (Fig. "Creep damages at turbine inlet guide vanes ").

“I”: High thermal strain causes this typical cracking in the blade/shroud transition (Fig. "Different behave of part zones under thermal fatigue", Ref. 11.2.3.2-11). These cracks are often tolerable up to a length specified by the OEM. The crack growth is relatively slow up to this length.

“K”: Creep overload and fretting wear occurs primarily on the lobes and bolts of the anti-twist device. The powerful circumferential force on the turbine stators is often underestimated at first glance. Failure of the fasteners has already led to serious consequential damages with uncontained fragments on several occasions (Volume 2, Example 6.2-4.1/2 and Chapter 8.1).

“L”: Erosion can decisively shorten the life span of the stator vanes. Ingested, often chemically aggressive, dust can cause astoundingly fast wear (Volume 1, Ill. 5.3.2-12.1). Coke and soot particles from the combustion chamber can have an extremely erosive effect (Figs. "Unstable combustion in 'low NOx' combustion chambers" and "Carbon erosion at high pressure Turbine blades and vanes"). The increased use of ceramic thermal barrier coatings results in increased erosion loads on the downstream blading due to spalling particles. The zirconium oxide particles wear the thin 50 m diffusion coatings down until the oxidation protection is gone. This can weaken the blade wall unallowably quickly and shorten the blade life to considerably less than any warranty period.

A successfully implemented solution for especially pronounced abrasive erosion loads in operating temperatures up to 980 °C is the use of thermal spray coatings made from tungsten carbide-cobalt and chromium carbide-nickel-chromium (Ref. 11.2.3.3-2).

Figure "Creep damages at turbine inlet guide vanes " (Ref. 11.2.3.2-12): This is a civilian engine type of the first serial engines with small bypass ratios (middle diagram). Various damage mechanisms caused typical turbine damage. The damages were described in the literature as follows:

Excerpt: “This problem has been world-wide since 1969 for all operators of ..(this engine type). The stationary first stage turbine vanes (also called nozzle guide vanes = NGV) were subject to creep in service due to high temperature of the combustion chamber exhaust gases. This creep occurred in the unpleasant form of “bowing” of the vane: this is a deflection of the vane in the aft direction due to the aerodynamic loads. This aft deflection increased with time up to the moment that it reached the original distance between the stationary and rotating first stage turbine blades. Then rubbing occurred between stationary and rotating blades (bottom middle diagram), very soon, complete first stage turbine “salad” and destruction of the whole turbine blading…“

Excerpt: ”…This failure had in fact nothing to do with NGV bowing. It was a case of mixed creep and fatigue failure (bottom left diagram). … (This) shows that the damage due to such a failure is not very important but the missing blade can cause a higher vibration pattern in the remaining turbine blading and cause further premature fatigue failures (bottom right diagram).”

Bowed NGVs can create dangerous flow disturbances. These can excite dangerous vibrations in turbine rotor blades and/or the entire rotor.
Integral turbine stators of the type used in smaller engines as one-piece cast parts are very sensitive to bowing. This construction principle cannot absorb thermal strain differences between the shroud rings and the blading due to high stiffness of the inner and outer shroud rings (Ill. 12.5.1-15).

Figure "Damage at overheated turbine disk": An integral cast, uncooled turbine disk (left) and a forged turbine disk with separate cast cooled blades were extremely overheated during startup. They had typical damage symptoms at the blade annulus.
The uncooled disk displayed gaping grain boundaries that were primarily oriented perpendicular to the direction of centrifugal force. Especially hot blade zones on the inlet and exit edges showed signs of burning (also see Example "Burnt turbine blades").
The cooled blades were destroyed to about the middle of the blade. The front edges showed signs of burning (Fig. "Overheating signs on torbine rotor blades"). Cooler areas had high-temperature violent fractures with crack surfaces that are created by hammering loads (Fig. "Annealed structure indicates overheating") and are typical for cast Ni alloys.

Figure "Blade root failure causing hot gas incursion to disk": If the annulus of a turbine disk is struck by hot gas, it can cause a catastrophic failure. This results in uncontainable fragments, since several blades and/or annulus fragments are released. Dangerous hot gas encroachment can occur in various ways:

• Failure of the seals at the root platforms
• Insufficient barrier/cooling air (pressure drop)
• Fracture of a rotor blade at the root. Hot gas can enter into the gap created by the released blade. This can dangerously overheat annulus sections that follow against the direction of rotation.

Figure "Overheating of turbine blades during stand stillS": Overheating during standing is also a possibility. A typical example is a failed attempted start of a helicopter engine. Possible causes are jammed power turbines or an attempted start before fuel has drained completely (Fig. "Overheating during startup with jammed turbine rotor"). Characteristic damage symptoms of this type of overheating permit later identification.
Only limited circumferential zones of the blade annulus showed signs of overheating such as tarnishing, deformation of blade edges, and burning (right diagram). Depending on the distribution of the fuel nozzles, this can be caused by a single circumferential area or by several sections.
An unmistakable characteristic is melted blading (usually the inlet edge). In this case, the higher-melting oxide skin wrapped around the melt like a wrinkled plastic bag (details). If this skin breaks, it will create shiny metallic melt droplets that spray and orient themselves on the downstream hot part surfaces.

Figure "Overheating during startup with jammed turbine rotor": Overheating process in the case of a jammed power turbine (Fig. "Overheating of turbine blades during stand stillS"):

An improper waiting period before a restart (Volume 2, Ill. 7.1.2-9) can lead to serious overheating damage. Differences in thermal strain between the turbine housing and the power turbine rotor bridge the tip clearance gap in this case (top diagram, Volume 2, Chapter 7.1.2). If the gas generator does not supply enough air for combustion of the injected fuel, then fuel can collect in the 6 o'clock position. The inflow of cooling air from the high-pressure turbine allows the fuel to ignite in the rear turbine stages. This results in burning and melting in a limited segment at the 6 o'clock position.