18.104.22.168 Turbine Damage
Typical damages to high-pressure turbines
Problems with the blading are of special importance in the high-pressure turbine (HPT). Overheating of the turbine stator vanes has already been covered in Chapter 22.214.171.124. These parts are especially endangered by hot streaks in the gas flow (Fig. "Hot gas streaks as combustion chamber problem"). However, the turbine rotor blades can also be overheated (Fig. "High pressure turbine rotor blades overhaul intervals"), even though their rotation causes them to even out uneven temperature distribution around their circumference (Fig. "Temperature variation at the combustion chamber outlet"). Scorching and melting are typical damage symptoms (Fig. "Overheating signs on torbine rotor blades", Chapter 12.4). In extreme cases, entire sections of blading are missing (burned up). Signs of overheating are delaminated or rolled-up anti-diffusion coatings, orange peel effects and cracking on the inlet edge, and heavy localized oxidation of the surface ( ).
Typical causes of overheating of the turbine blading are a poor temperature profile of the hot gas flow following the combustion chamber (Fig. "Temperature variation at the combustion chamber outlet"), insufficient cooling air flow due to blockage or deformation of cooling air ducts (Fig. "High pressure turbine rotor blades overhaul intervals"), hindered heat removal due to oxidation layers on the duct walls, and increased gas temperatures due to the malfunction of other components (e.g. regulators, fuel injection system). Rust has been found to have an especially strong blocking effect. For this reason, it must be ensured that the inlet duct and all aggregates are rust-free in testing rigs. The deformation of cooling ducts can be caused by internal objects (OOD) that strike the inlet edge of the rotor blades. A potential threat is spalling thermal barrier coatings (Volume 1, Chapter 126.96.36.199). Typical foreign objects are coke particles from the combustion chamber (carbon impact, Chapter 188.8.131.52). Oxidation and hot gas corrosion in cooling air ducts can be caused by remnants from cleaning processes (acids), but also by materials used in repair or production processes (e.g. low-melting cast alloys).
The temperature levels in the engine parts are not the only factor causally relevant to damage. The temperature distribution (Fig. "Parameters influencing thermal fatigue") also influences part life spans. Thermal fatigue is the primary damage mechanism of cooled turbine blades (Chapter 12.6.2). Cracking at the inlet edges of the blades and inside (!) the parts, i.e. in the walls of the cooling air ducts, is typical of thermal fatigue. The phenomenon of increased cracking in the relatively cool walls of cooling air ducts is not explainable merely by creep damage due to high temperatures. In this cyclical fatigue process, the loads are also created by thermal strain differences (Fig. "Damaging colder part zones by thermal fatigue"). The cooler areas are subject to tensile stress, which reaches an equilibrium with the compressive stress in the hotter zones with greater heat expansion. In guide vanes, thermal strain between the shroud and vane leads to thermal fatigue cracking in the transition radius (Fig. "Temperature caused damages at high pressure turbine vanes"). Up to a certain size, these cracks are usually tolerated by the OEM. This size depends on the gas bending loads on the blades and/or the crack-dependent leakage air flow rate. Compressive stress in the blade can be powerful enough to cause it to fold over. In blading segments that consist of several blades, damage can be concentrated in a specific blade position. This is the case if the deformation of the crack-sensitive blades in the segment is prevented by the neighboring blades. Overly high part temperatures (into the range of material softening) promote localized bulging of the blade when the wall creeps under the pressure of the cooling air.
Thermal barrier coatings have reached a level of development that makes typical problems such as spalling and erosion controllable. This has increased their use on turbine blades. It must be remembered that the coating materials react with melted dust deposits and other air contaminants, causing them to fail prematurely. Due to the ion conductivity of hot ZrO2 coatings, there is a danger of oxygen being transported to the bond coating, where it can cause long-term oxidation and a corresponding loss of bond strength (Volume 1, Chapter 5.4.5).
Hot gas corrosion and sulfidation
One of the specific types of damage in LP turbines is hot gas corrosion (HGC). There is a preponderance of corrosion in the lower temperature range up to about 950°C, which is typical for the operating conditions in the low-pressure turbine. Above this temperature, oxidation is dominant. If aggressive deposits have built up on the affected surface, they will cause rapidly progressing damage to the substrate. This is called sulfidation (Volume 1, Chapter 5.4.5). This can cause continuous damage to entire blade sections from one side to the other. This is especially dangerous for hollow profiles with poor air flow, which are commonly found in thick blade sections in order to reduce weight. Fragments breaking out of these walls create windows into the hollow spaces. Surfaces that tend to dust buildup are especially at risk. These include the front sides of stator shrouds, which create gaps that are struck by the airflow.
Oxidation and hot gas corrosion of honeycomb seals
Honeycomb seals are used as rub surfaces in housings opposite the rotor blade tips, and also on the inner stator shrouds opposite the turbine spacer labyrinths. They consist of a thin, honeycomb-shaped sheet metal structure and are designed to minimize the clearance gaps. These metal ridges, which are only a few tenths of a millimeter thick, are subjected to the long-term effects of the hot gases. The seal structures opposite the rotor blades tips, especially, can oxidize all the way through to the base material. Sheet metal honeycombs with this kind of extreme damage lose their strength and spall in large sections on the circumference. This leads to clearance gap increases and leakages (Volume 2, Ills. 7.1.3-12 and 7.1.3-14). If this type of damage is a concern, then suitable materials and/or coatings must be used for the honeycomb structure. Preoxidation of the new parts can be expected to improve their performance in this regard to a certain degree. However, whether or not this is sufficient is a question that must be answered in sufficiently realistic comparative tests for each specific application.
Wear and damage to labyrinth tips
The low-pressure turbine has an especially large number of labyrinth seals. The tips of these seals are altered and damaged even during normal operation. At the least, they are worn down. The material is often locally overheated. Armor can spall and cracking can occur (Volume 2, Ill. 7.2.2-3.1). Damaged seals are repaired during overhauls through welding, reworking, and even re-coating. Cracking must be taken seriously and prevented with proper measures. Some remedies are suitable tip materials and protective armor.
Wet corrosion in low-pressure turbines
Experience has shown that low-pressure turbines are subject to special corrosive loads when condensation water forms while standing. Increased HGC can then occur during subsequent operation. One known occurrence is location-specific damage (environmental factors) to nickel alloys in connection with silver coatings (e.g. on bolts and nuts, Volume 1, Chapter 184.108.40.206).
The fir-tree connections of martensite steel disks in older engine types are especially susceptible to condensation water corrosion while standing. Corrosion scars in highly-stressed disk zones such as the hub and the areas around bolt bores are points of origin for thermal fatigue cracks. For this reason, the bolt bores, which are usually especially difficult to reach, must be inspected very thoroughly (e.g. with special eddy current probes).
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. 220.127.116.11-16, 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. 18.104.22.168-11). This type of damage can be caused by foreign objects (Volume 1, Chapter 22.214.171.124) 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, 126.96.36.199-13 and 188.8.131.52-14). 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.
Example "Bulging turbine blades" (Ref. 184.108.40.206-16):
Excerpt: “….The problem with these turbine blades is, quite simply, consequential damage. It begins with a hot spot, a discoloration in the outer concave part of the blade, and is followed by a local deformation - we call it bulging (Fig. "High pressure turbine rotor blades overhaul intervals" “B”). This also damages the cooling air ducts…the outer part of the blades…burns off. A blade that is damaged in this way does not experience overly large vibrations, nor does it cause a noticeable decrease in power, and the engine practically continues to run normally… In the various models…(of the affected large fan engine type) there have been 22 turbine blade fractures to date (early 1994)…“
Comments: This evaluation of the effects of a partially fractured HPT rotor blade seems optimistic. The fact that the partially broken blade did not make itself known through vibrations or performance losses is probably due to the elastic and dampened bearings commonly used today. If the turbine rotor runs for an extended period with increased, unknown imbalance, there is a danger of the housing (e.g. struts), bearings, or rotor components being dynamically overloaded and suffering dynamic fatigue damage. These damages can cause the engine to fail.
Figure "Blockages in turbine blade cooling system" (Ref. 220.127.116.11-6): 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.
Excerpt 1 (Ref. 18.104.22.168-2): ”…(the OEM) has incorporated changes into the production…to correct problems with overheating and wear on reengined …aircraft…Tests have determined that ingested dirt was clogging the high-pressure turbine blade cooling holes. Changes now incorporated in production…include:
- Increased number of cooling holes
- Altered routing of cooling air passages to remove dirt before the cooling air reaches the high pressure turbine…”
Excerpt 2 (Ref. 22.214.171.124-3): “…(the airline) is considering redesign of the spinner and the addition of a vortex dissipator…to reduce particle ingestion that has been causing high levels of erosion and has necessitated a higher engine removal rate…The particle ingestion was discovered in routine boroscope examinations and did not cause in-flight shutdowns…“
Excerpt 3 (Ref. 126.96.36.199-4): ”…(The OEM) is developing two sets of fixes and improvements for its… powerplant based on operating experience…One set is designed to correct dust clogging in high pressure turbine blade cooling circuits….
The focus was overcoming dust clogging experienced on a number of engines. Fixes include the use of 13 holes in the high pressure turbine blades. The holes are located in the blade tipcap facing the channels of the blade's leading edge cooling circuit. The holes are drilled straight in the blade in contrast to the 10 angled holes used in the original blade design. The new holes are designed to allow dust to be ejected more freely from the cooling circuit, reducing chances for clogging. These fixes supersede an interim fix that involved drilling four new holes in the original blades (Excerpt 1). Also the arrangement of holes in the cooling path between the …combustor and high pressure turbine has been changed. This action was taken to stop the movement of dust before the dust reaches the high-pressure turbine. The modifications were evaluated in two ground test series…Approximately 350 lb. of dust were ingested by the engines in each of the test programs, officials said…“
Comments: In this case, the problems occurred in connection with the retrofitting of an older commercial aircraft type with fan engines with a large bypass ratio. The resulting low position of the engines over the ground (see Volume 1, Ill. 188.8.131.52-9) promoted unusually heavy dust ingestion.
The blockage of the cooling air film bores was not caused by particles from the outside that were melted in the hot gas. The dust was carried by the cooling air (Fig. "Blockages in turbine blade cooling system"). Evidently the cooling air extraction in the compressor was positioned in a way that heavily particle-laden air also reached the cooling air system (Fig. "Hot parts cooling structures blocked by dust"). It is interesting that the orientation of the bores helped to keep the dust blockages within acceptable limits (Fig. "Reduction of life span changes in cooling bores").
Example "Necessity of cleaning turbine blades" (Ref. 184.108.40.206-5):
Excerpt: ”…The problem centers on cracking of the first- and second-stage high-pressure turbine blades…The ultrasonic process removes foreign material that accumulates within the hollow blades during normal engine operation, but “over aggressive” cleaning made the internal surfaces of the blades vulnerable to high-cycle fatigue cracks…“
Comments: In relation to this chapter, cracking due to cleaning procedures is not relevant. The interesting fact is that cleaning was necessary to remove deposits from the cooling air structure. This shows that deterioration of the cooling system over the operating life is generally expected.
Example "Airflow leak causing fatigue failure" (Ref. 220.127.116.11-13):
Excerpt: “While engaged in cruise…several stage 1 high pressure turbine rotor blades within the No. 1 engine separated without warning from their platforms. The initial separation was found to have originated at a blade forward cooling hole and was determined to be fatigue induced. The fatigue was most likely brought on by a degradation of the cooling airflow pattern within the blade induced by an airflow leak not associated with the blade cooling holes. The blade separations resulted in disintegration of the high pressure turbine assembly and subsequent engine failure.”
Comments: This is a large first-generation fan engine. The fatigue crack in the inner cooling air duct is typical for thermal fatigue (
Example "Burnt turbine blades" (Ref. 18.104.22.168-14):
Excerpt: “During descent prior to the approach…the crew attempted to start the APU. The APU failed to start….Test results showed that the unit had turbine blades burned and eroded, fuel control cracking pressure set higher than specified, and the electrical harness wires to the fuel solenoid connector were reversed.”
Comments: This case involved extreme overheating due to failure of the fuel regulator. The blades of the turbine disk were evidently burned up (
Figure "Turbine rotor blade tip erosion and oxidation" (Ref. 22.214.171.124-7): The hot gas leakage flow places the tips of turbine rotor blades in increased danger of overheating (Ref. 126.96.36.199-9). 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. 188.8.131.52-10). 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. 184.108.40.206-8). 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. 220.127.116.11-9).
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 18.104.22.168), 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:
Untwisting is a permanent straightening of the blade due to centrifugal forces (Ref. 22.214.171.124-15). 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. 126.96.36.199-8 and 188.8.131.52-11). 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. 184.108.40.206-12).
“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. 220.127.116.11-12, 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. 18.104.22.168-11). 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 (
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. 22.214.171.124-2).
Figure "Creep damages at turbine inlet guide vanes " (Ref. 126.96.36.199-12): 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.
188.8.131.52-1 P.König, A.Rossmann, “Ratgeber für Gasturbinenbetreiber”, ISBN 3-8027-2545-X, ASUE-Schriftreihe, Vulkan - Verlag Essen, 1999.
184.108.40.206-2 “CFM56-2 Changed to Correct Problems”, periodical “Aviation Week & Space Technology”, November 1, 1982, page 35.
220.127.116.11-3 “Delta Weighs Changes to CFM56-2”, periodical “Aviation Week & Space Technology”, December 20, 1982, page 31.
18.104.22.168-4 J.M. Lenorowitz, “CFM56 Powerplant Fixes Based on DC-8 Operations”, periodical “Aviation Week & Space Technology”, February 14, 1983, page 32.
22.214.171.124-5 E.H. Phillips, “Pratt & Whitney Recalls Defective Turbine Blades”, periodical “Aviation Week & Space Technology”,April 13, 1998, page 63.
126.96.36.199-6 P.König, T.Miller, A.Rossmann, “Damage of High Temperature Components by Dust Laden Air”, Proceedings AGARD-CP-558 of the conference “Erosion, Corrosion and Foreign Object Damage Effects in Gas Turbines”, Rotterdam 25-28 April 1994, pages 25-1 to 25-12.
188.8.131.52-7 J.P. Bindon, “Pressure Distributions in the Tip Clearance Region of an Unshrouded Axial Turbine Affecting the Problem of Tip Burnout”, Proceedings ASME 87-GT-230 of the “Gas Turbine Conference and Exhibition”, Anaheim, California, May31-June 4, 1987, pages 1-7.
184.108.40.206-8 J.D. Denton, “Loss Mechanisms in Turbomachines”, Proceedings ASME 93-GT-435 of the “Gas Turbine and Aero Engine Congress and Exhibition”, Cincinnati, Ohio, May 24-27, 1993.
220.127.116.11-9 Y.W. Kim, D.E. Metzger, “Heat Transfer and Effectiveness on Film Cooled Turbine Blade Tip Models”, Proceedings ASME 93-GT-208 of the “Gas Turbine and Aero Engine Congress and Exhibition”, Cincinnati, Ohio, May 24-27, 1993.
18.104.22.168-10 J.P. Bindon, “The Measurement of Formation of Tip Clearance Loss”,Proceedings ASME 88-GT-203 of the “Gas Turbine and Aero Engine Congress and Exhibition”, Amsterdam, June 5-9, 1988.
22.214.171.124-11 K.G. Rummel, “Investigation and Analysis of Reliability and Maintainability Problems Associated with Army Aircraft Engines”, NTIS-Report AD-772 950, August 1973, pages 59-63.
126.96.36.199-12 M.V. Averbeke, “Gammagraphy in Airline Maintenance”, AGARD-AG-201-Vol.1, “Non-Destructive Inspection Practices” (Editor E.Bolis), pages 295-325.
188.8.131.52-13 NTSB Identification SEA871A068, microfiche number 32820A, Incident occurred March 26, 1987.
184.108.40.206-14 NTSB Identification NYC87IA087, microfiche number 34387A, Incident occurred January 3, 1987.
220.127.116.11-15 “Metals Handbook, Ninth Edition”, Volume 11, “Failure Analysis and Prevention”, American Society for Metals”, 1986, page 283.
18.104.22.168-16 “Problem im Griff”, Interview mit Heinz Bart, technischer Direktor der Swissair, zu den Triebwerksausfällen der MD-11, periodical “Maintenance Special”, 1/1994, pages 50-53.