The components of a turbine, and especially blades and disks, are subjected to a wide variety of damaging influences (Ills. 188.8.131.52-2 and 184.108.40.206-6, Chapter 220.127.116.11), which act together in an exceptionally complex damage scenario. These influences are covered more extensively in separate chapters, and when possible, in connection with other topics. Therefore, this chapter does not completely cover all problems, but rather offers an overview. The blades of the high pressure turbine and low pressure turbine are dealt with separately using examples. Some of the diagrams and descriptions are based on Ref. 18.104.22.168-1.
HPT rotor blades
Due to the demands for high heat- and creep resistance, the rotor blades of modern gas turbines are produced by investment casting (Chapter 12.5). This allows the realization of complex inner cooling air bores. The air for the protective cooling air film is directed to the surface through many openings (Fig. "Cooling systems of turbine blades and vanes"). The typical blade materials are nickel-based alloys. A precipitation phase (Ni3Al) hardens these materials and ensures sufficient creep strength. These alloys can be polycrystalline, single crystal, or directionally solidified, depending on the casting process used (Fig. "Grain structure influencing metal properties"). Alignment or avoidance of grain boundaries is done because these are special weak points for creep stress (Chapter 12.5) and thermal fatigue (Chapter 12.6.2).
The blade tips and shroud (especially in older engine types) can be braced against neighboring blades (minimizes vibration). Rubbing systems (function similar to the compressor; Volume 2, Chapter 7.1.2) are used to seal the space between the tips of shroudless blades and the housing. To this end, the tips of some blades are armored with soldered particles of hard material (Volume 2, Ill. 7.1.4-14). Across from the blade tip, the housing has so-called turbine segments. In older engines (Volume 2, Chapter 7.1.3), filled or unfilled honeycomb structures form the rubbing surface. In modern engines, ceramic rub coatings are finding increasing use (thermally sprayed zirconium oxide). For the past several years, rotor blades have also been outfitted with ceramic thermal barrier coatings. Originally, these coatings served to minimize localized overheating and ensure that the required life span was reached. Today, the thermal insulating effect is already considered when the parts are designed, especially in the case of rotor blades. This philosophy can result in a coating failing (e.g. spalling) and quickly cause uncontrollable (e.g. through boroscope inspections) failure of the engine part (see Chapters 13 and 14). Physical vapor deposition (PVD) coatings with columnal structures (Fig. "Thermal barrier coatings of turbine rotor blades") especially resistant to thermal cycles have become the preferred type of blade coating, especially for rotor blades, taking the place of plasma spray coatings. Thermal barrier coatings require layered application with an oxidation-resistant bond coating and/or an anti-oxidation coating on the substrate.
HPT blades usually have anti-oxidation coatings on their outer surface. These coatings are made by indiffusion of aluminum or thermal spraying (MCrAlY coatings). During operation, a protective Al2O3 coating forms. In order to prevent unwanted diffusion of the coating components into the substrate, additional diffusion barriers with suitable enriched elements (e.g. planinum) are used. In order to prevent inner oxidation and/or sulfidation (Volume 1, Chapter 5.4.5), the inner cooling structure of the blades is also often given a diffusion protection coating. It must be remembered that many coatings behave brittly at low temperatures and can influence the thermal fatigue of the blades (Chapter 12.6.2).
Experience has shown that the roughness of the turbine blading (HPT and LPT) increases with run time due to deposits from the fuel, sintered dust deposits, hot gas corrosion pittings, erosion, and possibly crumbling ceramic thermal barrier coatings (Ref. 22.214.171.124-7). This can considerably increase thermal conductivity (up to 60%) and increase the temperatures in cooled blades.
HPT guide vanes
The guide vanes (Fig. "Cooling systems of turbine blades and vanes") at the combustion chamber exit (Fig. "Temperature variation at the combustion chamber outlet") are subject to the greatest thermal stress in the gas turbine. Localized temperature peaks in the gas flow are especially damaging. Even though these parts are stationary, they are subject to high mechanical loads. These are created primarily through restricted heat strain caused by large changing temperature gradients and gas bending loads. Today, high heat-resistant Ni-based alloys are used for these parts. These are more sensitive to overheating than the less heat-resistant Co-based alloys, which do not precipitation harden, and therefore do not suffer permanent strength losses. Guide vanes are generally fixed to shrouds at both ends and cast in two- or three-piece segments, or connected with high-temperature soldering.
The thermal loads necessitate intensive cooling (about 5-10% of the total airflow). For this reason, stator vanes in the high pressure turbine have more bores for the cooling air film (Fig. "Cooling systems of turbine blades and vanes") than rotor blades.
It is important for the axial gaps at the circumference between individual vanes or the vane segments to be well sealed. Overheating of the housing or rotor components through hot gas incursion must be prevented. Suitable seals are generally made of thin strips of metal that are pushed into slots in the blade platforms.
Because of their especially high temperature levels, guide vanes are protected with thermal barrier coatings and/or anti-oxidation coatings. Hot spots are especially pronounced at the guide vanes. Temperatures can exceed the melting point of the diffusion coatings. The high Al content lowers the melting point (coating rolls up, Fig. "Temperature caused damages at high pressure turbine vanes"). Guide vanes were outfitted with ceramic thermal barrier coatings on their gas-struck surfaces earlier than rotor blades. Spontaneous, uncontrollable failures resulting from coating failures are less likely in guide vanes than in rotor vanes, due to the lower mechanical stress on the former.
The disks of the high pressure turbine (Fig. "Operation loads on rotor disks") are exposed to a combination of loads that have a strong influence on their life span. They are affected by the centrifugal forces of the high-RPM rotor. They are also affected by cyclical thermal strain, which is temporally changing temperatures with gradients that are the highest of all disks in the engine. The following text uses these disks representatively for all other disks (compressor and LPT) to discuss the creation of the above loads, their typical distribution in the disk, and the effects they have on disk life.
Centrifugal forces cause the disk to expand. Unless they are relieved by constructive measures, the inner disk groups are subject to especially high tensile loads from tangental stress. These loads decrease considerably towards the annulus. At the annulus, the radial and tangental stress from the centrifugal force are at about the same level. Cursory observation does not result in an expectation that the centrifugal forces put more strain on the hub than on the annulus. A plausible explanation is that while the masses at the circumference are subject to greater centrifugal forces, these forces must be absorbed by the neighboring inner cross-sections. In this way, the centrifugal forces from the outer zones ultimately act on the hub and put extremely high strain on this area, even though the hub does not contribute greatly to the centrifugal force itself. In order to break down the stress, rotor blades are made considerably thicker at the hub than at the annulus. Despite this, the hub regions of modern engines show cyclical plastic deformations. These are LCF loads that determine the life span of these parts (Chapter 12.6.1).
Restricted thermal expansion caused by temperature gradients contributes greatly to the loads on the blade. The thermal strain overlays with the stress from the centrifugal forces ( ). The thermal strain changes with the operating states and temperature gradients.
In order to cool the turbine blades, the cooling air of the disk surface is directed along the blade root. In order to prevent the high pressures in the gas flow around the turbine blades from blowing the cooling air away, it is often additionally compressed by a type of radial compressor (“cover plate”) ahead of the disk. The air feed system to the HPT rotor blades can affect the mechanical behavior of the parts. Depending on the condition of the seals, gas vibrations in this system can influence the rotor. For example, this may be the cause of unexplained “rough” running of an engine in certain phases of operation (Fig. "Vibrations excited in ring shaped spaces 1" and Ill.126.96.36.199-10).
The low pressure turbine
Despite their relatively low load levels (low RPM, low temperature levels, even temperature distribution in the hot gases), low pressure turbines can experience cracking or even fractures due to damage-relevant mechanical loads caused by thermal fatigue and vibrations.
The long, slender rotor blades are generally outfitted with shrouds, which are braced against one another to stiffen them and prevent vibrations. These shrouds are subject to damage mechanisms such as fretting (Volume 2, Chapter 6.2), creep deformations (shroud flexure, Chapter 12.5.2), and erosion. Tension in the blades can be reduced by creep deformations (twisting open, Fig. "Typical damages to rotor blade shrouds").
Materials such as high-alloyed steels (CrNi type), hardened iron-based alloys, and also Ni alloys with typical limited thermal stability can experience structural changes during the long life spans typical for LPT blades. This results in decreases in the original static and dynamic strength values, as well as a reduction in toughness. One example of this type of change is the formation of brittle phases (such as sigma phases). Embrittlement in blades can promote the failure of <U>all blades</U> in the case of FOD (haircut). In materials that are altered in this way, the increased tendency to cracking makes repair welding difficult to use as part of an overhaul. Higher temperatures in the front turbine area cause the grain boundary carbides to grow, and also lead to coarsening or raftening of the hardening phase (also see Chapter 3.3. HPT).
As mentioned before, the low pressure turbine has specific damage types. These include special types of high temperature corrosion such as sulfidation (Volume 1, Chapter 188.8.131.52) or damage related to wet corrosion while standing (Chapter 184.108.40.206). In the case of a shaft failure, the core engine continues to supply hot gas that powers the LPT. This then reaches uncontrolled overspeed (runaway turbine) and the blading fails and is flung off (Volume 1, Ill. 4.5-4). For this reason, special measures are taken in the low pressure turbine to decelerate the rotor (Volume 1, Chapter 4.5). In contrast, a shaft failure in the HPT causes the core engine (gas generator) to fail, disrupting the flow of hot gases that power the turbine.
As with compressor housings, turbine housings are required to catch and contain all fragments resulting from a blade failure. However, containment (Volume 2, Chapter 8.2) cannot generally be guaranteed in the case of disk failures. There are special measures (e.g. burst protection rings) that can contain disk fragments (Volume 2, Chapter 8.2). Low pressure turbines are especially prone to dangerous overspeed if the undamaged core engine continues to supply hot gas to drive them. For weight reasons, the housing walls of the typical large-diameter low pressure turbines must be filigreed. In contrast, the high pressure turbines are outfitted with internals in the housing that fasten the seal segments and carry cooling air, and these, combined with the thicker housing walls, offer greater penetration resistance to blade fragments.
Figure "Operation loads and damages on turbine rotor blades": Turbine rotor blades are subjected to loads and damage mechanisms that can influence one another. In order to give an overview of the part-specific influences, the different blade zones will be dealt with separately:
The blade tip has a deciding effect on the tip clearance gap and thus also on deterioration of engine efficiency over the operating life (Volume 2, Chapter 7.0). The hot gas leakage flow causes localized overheating with extreme oxidation (burns, Fig. "Turbine rotor blade tip erosion and oxidation"). When the blade tips rub against the hard (erosion resistant) and high temperature resistant seal segments in the housing (ceramic coatings), it causes heat creation and wear. (Shroudless) blade tips coated with hard material particles must undergo the necessary initial wear (i.e. run-in) within the first few hours of operation. The particles can lose their chipping properties after only a few hours of operation due to oxidation and/or reactions with the substrate (Volume 2, Ill. 7.1.4-14). If creep causes the shrouds to bend open, then the areas where they lay on neighboring blades are especially at risk of wear (Fig. "Typical damages to rotor blade shrouds"). Creep deformation can cause distorted shrouds to loosen, which makes them lose their vibration-damping effect ( ).
The blade body requires intensive cooling to ensure sufficient strength to absorb the centrifugal forces. The temperature gradients induce high cyclical thermal strain (thermal fatigue, Chapter 12.6.2). Creep changes (Chapter 12.5.2) residual stresses and/or increases them. Disturbance of the flow through a damaged upstream stator or pressure fluctuations typical for combustion chambers (Ill. 220.127.116.11-4) can cause considerable high frequency loads. Localized weakening of the walls occurs in hot blade areas due to oxidation and coating distortion. Depending on the temperatures and contaminants involved, hot gas corrosion (sulfidation) can be expected on the inside and outside of the blade. If the erosion of the soot and coke particles from the combustion chamber acts in combination with oxidation, then the blades can be damaged unforseeably quickly (Fig. "Carbon erosion at high pressure Turbine blades and vanes"). Larger coke particles are capable of plastically deforming the blades` inlet edge area constricting the following cooling air duct. Foreign particles (dust, wear products) in the cooling air can block the inner cooling air ducts, causing overheating (Volume 1, Chapter 18.104.22.168). If dust is melted in the combustion chamber, it can block the bores for the cooling air film from the hot gas side and/or react with the surfaces (Volume 1, Chapter 5.3.2). This accelerates erosive wear (coke, thermal barrier particles). Extreme localized overheating can melt the surface, causing permanent damage.
The blade root is the coolest part of the blade and subject to high mechanical loads from centrifugal forces and vibrations. Fir tree roots with their formed notches deserve special mention. They are affected by fretting, in addition to the contact forces (Volume 2, Chapter 6). For this reason, the surface of the root area is hardened (shot peening). The protective effect of the shot peening (pressure residual stress, hardening) is broken down by creep deformations.
Figure "Cooling systems of turbine blades and vanes": The stator vanes of the high pressure turbine that are directly behind the combustion chamber (Fig. "Temperature variation at the combustion chamber outlet") are the most intensively cooled parts of a gas turbine. In order to form a cooling air film despite the high pressure levels, the cooling air is taken from the compressor exit area. The cooling air consumption of the high pressure turbine stator vanes alone is several percent of the total air throughflow of an engine. The energy of the cooling air is not lost, and it releases energy upon expansion in the turbine. If the gas temperature is increased (e.g. in order to increase power and efficiency) with the same hot part technology, then the cooling air consumption must also increase. Minimizing cooling air consumption is a continuous endeavor. This can be done through the use of ceramic materials in the form of thermally sprayed thermal barrier coatings ( ). Thermal insulation coatings minimize the flow of heat into the blade wall, and thus also decrease the necessary intensity of the cooling.
Especially intensive and effective cooling of the blade surfaces and shrouds of metallic blades requires sieve-like perforation of the walls (air film cooling, bottom diagrams).
In order to reduce the unavoidable sintering effects of extremely heated coatings (e.g. within an hour at 1250°C) to acceptable levels, combinations of ceramic thermal barrier coatings and perforated walls (transpiration cooling, Ref. 22.214.171.124-12) are being developed. The creation of holes in coated blade walls seems to be a solvable problem.
The rotor blades of the high pressure turbine stages must absorb the high centrifugal forces and have correspondingly high heat resistance. This means that temperatures in the supporting cross-sections must be reduced considerably. The required cooling air acts as convection cooling (heat conductivity) inside the blade. The cooling of the blade must be especially intensive near the inlet edge. The necessary good heat transfer is ensured through impingement cooling (see detail). The air is blown directly against the surface of the inlet edge through the perforated wall of a neighboring cooling air duct. In order to maintain the film cooling on the surface of the blade inlet edge even against the flow, especially high cooling air pressure is required. This pressure can be created by an additional compressor in the form of a structured cover plate on the front of the turbine disk.
Effusion cooling with the aid of porous blade walls has not found serial use due to problems such as cracking, blocking of the walls, and sensitivity to foreign object impacts.
Inside the blades, the cooling air is distributed by metal inserts (bottom right diagram). The temperature of a turbine rotor depends on several factors. Welded or soldered connections between the porous “shirt” and the supporting inner (cast) structure have been determined to be especially susceptible to thermal fatigue cracks.
Figure "Thermal barrier coatings of turbine rotor blades": Thermally sprayed ceramic thermal barrier coatings have a lamellar construction. They can only react to the difference in thermal strain with the metallic substrate through island-like cracking, called segmenting (top left diagram). The coating “breathes”. This segmenting should occur during the introduction of heat in the production process or be affected in a way that it occurs in the desired manner during the early operating cycles. During the heating phase, the insulating ceramic coating expands more than the substrate due to its faster heating-up, even though it has a relatively small coefficient of thermal expansion. This closes cracks and creates controllable compressive stress. During stationary operation, a residual stress state is created that can be tolerated for long periods, depending on the temperature gradients in the coating and substrate. In cooled hot parts, the described effect can lead to especially high compressive stress levels during the heating phase. During the cooling phase, the process is reversed, and the cracks open again.
The structure of the ceramic thermal barrier coatings (top right diagram) depends on the production process. It determines different coating behaviors, and can be optimized for various operating demands.
The typical lamellar structure of a thermally sprayed thermal barrier coating (APS coating = argon-plasma sprayed) is depicted in the bottom left diagram. Despite their lamellar structure, which is directed parallel to the surface, these coatings are considerably more sensitive to erosion (7-10 times) than vapor deposition coatings (Ref. 126.96.36.199-15). The tests were done with Al2O3 particles at room temperature and 930°C. The angle of impact was very important (bottom left diagram and Volume 1, Ill. 5.3.1-3). The erosion mechanism occurs through the spalling of relatively large sections, which are evidently the “static” of the spraying process. Due to the presence of microcracking and weakened grain boundaries, this process requires relatively little impact energy. This results in high material removal rates and high roughness (bottom right diagram). Experience has shown that a consequence of damage to the coating structure during operation (thermal fatigue) is spalling of the particles even without noticeable erosive loads (Volume 1, Ill. 188.8.131.52-4)
Physical vapor deposition coatings (EB-PVD coatings = electron physical vapor deposition) with a columnar structure are considerably less erosion-sensitive than thermal spray coatings. In PVD coatings, the erosion process must first create microcracks in the column crystals, which requires relatively large amounts of energy. The crystals that break out are very small particles and result in a minor loss of mass. PVD coatings are also finding increased use due to their good thermal fatigue behavior. They have been shown to be especially effective on the leaves of turbine rotor blades.
Ref. 184.108.40.206-11 gives extensive information about operating experience in civilian aircraft operation. Typical damage mechanisms were observed (volume 1, Ills. 220.127.116.11-4 and 18.104.22.168-5).
Silicate deposits from ingested dusts on the HPT blades. The deposits form especially on the pressure side of the blades. The special danger is acceleration of the sintering effect of the ZrO2 coatings. This worsens the thermal fatigue properties. Si and Ca can penetrate into the ceramic coating and react intensively. If the gaps of the columnar structure are filled, then cooling results in high compressive stress levels with tension stress in the bond coating. This, combined with oxidation of the bond coating, causes spalling.
The erosion of the thermal barrier coating greatly depends on the individual operating influences on the engine in question. This erosion can be detected before unrepairable damage occurs through suitable inspections, at least in turbine stator vanes. For repair to be feasible, it is necessary that the substrate is not unallowably damaged. One indicator is the remaining bond coating.
Figure "Thermal barrier coatings spalling limiting life span" (Ref. 22.214.171.124-13): The decisive phenomenon that determines the life span of a thermal barrier coating (TBC) is usually the oxidation of the bond zone. After longer run times, the bond strength of the TBC is often unallowably weakened. If the bond strength is overloaded by unavoidable thermal strain between the TBC and the substrate, then spalling can occur (top
diagram). This can cause overheating of the substrate beneath and/or coating particle impact damage to hot parts located downstream in the gas flow (OOD, see Volume 1, Ill. 126.96.36.199-12).
Zirconium oxide becomes an ion conductor at high temperatures and carries oxygen directly to the bond coating. In addition, the function-specific segmentation cracks and porousness allow the oxygen from the hot gases to reach the bond surface directly (Fig. "Thermal barrier coatings of turbine rotor blades"). This creates an oxide layer (TGO) between the TBC and the bond coating (BC). In order to suppress this oxidation, highly oxidation resistant bond coatings are applied to the substrate before the TBC. These are usually thermal spray coatings. Sintered slip coatings of type MCrAlY are also suitable.
As early as the production process (left detail), a thin Al2O3 layer forms on the bond coating. Further oxidation takes place in the hot gas (TGO, middle detail). This causes depletion of the bond coating`s oxidation-preventing aluminum content. If the brittle, relatively low strength oxide coating reaches a certain thickness, it will crack (right detail). The crack location can indicate its causal influences. Therefore, different types have come to be recognized. Cracks at the transition to the TBC (type “A”), in the oxide layer (type “B”) or at the transition to the bond coating. Type “C” refers to a compact oxide layer. Type “D” cracks run along the transition to the compact structure in oxide layers with “frayed” boundaries. In PVD-TBCs, the formation of the oxide layer seems to be related to the characteristics of the typical columnal structure. In thermal spray coatings (air plasma spray = APS coatings) and PVD coatings, cracking tends to occur in the oxide layer.
The bottom left diagram shows the 0.3 mm oxide growth of an APS coating during the operating time. One can see that increasing the operating temperature by about 50°C doubled the growth of the oxide layer. Similar growth can be expected even over several 1000 hours, whereby the “frayed” boundary especially contributes to oxide growth.
The middle diagram shows the trends for damage to APS coatings in relation to cracking in oxide layers. Cracking evidently slows considerably in thermal sprayed TBCs after several thousand hours of operation. By this point, an oxide layer thickness of about 0.06 mm is reached. The sudden flattening of the flaw curve indicates that this is a limit value for APS coatings. Above this value, spalling can be expected. Other damage, such as pore formation, evidently does not play a significant role in the specified test conditions.
PVD coatings (0.3 mm thick) have several times more pronounced cracking than APS coatings (right diagram), as can be seen by comparing this with the 1050°C curve in the left diagram. This could be related to the gas permeability of the continuously porous columnal structure (Fig. "Thermal barrier coatings of turbine rotor blades"). It is possible that this drawback is compensated for by the resilience of the columnal structure and its corresponding lower thermal strain.
Figure "Operation loads on rotor disks": Cyclical loads from centrifugal forces and thermo-mechanical fatigue (TMF, Chapter 12.5.2) act simultaneously on the rotor disks of turbines. This creates LCF dynamic loads that determine the life of the parts (Chapter 12.6, ). The disks must also be sufficiently resistant to bursting in case of high overspeed (contained in engine specifications). High frequency vibrations can be induced in various ways. For example, pulsating gas forces can affect the disk or shaft through the blades. In extreme cases, disk fractures are also possible (Fig. "Disk fracture by flow vibrations"). Creep can induce residual stress or change any residual stress that is already present (Fig. "Creep effects and part behavior"). This can break down the protective effect of shot peening (e.g. ahead of scratches from handling) over longer operating periods.
The encroachment of hot gases, due to seal damage or cooling air disruption, for example, can unallowably heat up the annulus area to plastification. The overheated zone can suffer temporary or permanent material changes (e.g. solution annealing of the hardening phase) and lose dangerous amounts of strength. Rubbing or oil fires (Volume 2, Chapter 9.2 and Fig. "Causes of turbine disk overheating") can even heat massive inner disk areas until they are damaged and fail. Intensive rubbing after the failure of static part fasteners or rotor deflection (bearing damage) can split massive cross-sections in seconds through a localized melting process (Volume 2, Chapter 8.1). If integral labyrinth rings rub with turbine disks as part of their function, it is possible that hot cracking may occur (Volume 2, Chapter 7.2).
The typical turbine disk materials are much less sensitive to fretting than titanium alloys (Volume 2, Chapter 6). However, damage can still occur. For example, while renewing blading and a poor connection at the blade roots in an engine, cracking was discovered in the disk (Volume 2, Ill. 6.2-12). Very high surface pressure promotes cold welding of contacting surfaces with relative movements. In extreme cases, forceful removal of the blades is the only solution.
Contrary to expectations, corrosion can occur on turbine disks made from nickel alloys. This attack is often connected to the presence of silver (Volume 1, Ill. 188.8.131.52-4). Experience has shown that aggressive condensation water can dissolve silver from nuts and bolts. Sulfidation is promoted in the areas where this solution collects and evaporates.
Figure "Loading of a turbine disk in the startup phase": The temperature of the gas flow increases during start-up. The rotor absorbs a large amount of heat in a short time through the large surface area of the blading which is exposed to the gas. This rapidly heats up the relatively thin disk cross-section of the annulus. The thick hub section has much greater thermal inertia and remains relatively cool (left diagram). This results in very large temperature gradients between the annulus and hub during the start-up phase. The hot annulus area wants to expand, but is restricted by the cold, massive disk hub area. This creates powerful compressive stress in the annulus (Fig. "Cracks protecting from thermal fatigue", top right diagram). The hub area which ensures load balances is subjected to correspondingly high tensile stress. If one adds the tensile centrifugal loads to the compressive thermal loads, the annulus area is relieved. However, the tensile stress in the hub area continues to increase. For this reason, start-up is the most critical load state for the disks, and determines their life span.
The disk reaches unchanging temperature gradients during steady states of operation. Depending on the constructive design and cooling, this steady state is only reached after several minutes of operation (Fig. "Turbine disk loads during operation cycles"). The moment with the highest gradients, and therefore the greatest loads, passes in the start phase after a few minutes. The operator should be aware that the starting procedure to full power is of great importance for the expected life span of the turbine disks and, therefore, the engine. The temporal starting process is usually specified by the OEM and is a prerequisite for guaranteeing hot part life spans.
The diagram shows results from temperature measurements on an integral case turbine disk of a helicopter engine during its start phase. It was shut down again even before it reached steady state operation. During shutdown of the engine, the gas temperature and centrifugal force both decrease rapidly, doubly relieving the hub. The cooling of the annulus area through the gas flow leads to a large temperature difference relative to the still-hot hub. The cooling annulus contracts. The previously plastically compressed zone is now subject to high tensile loads into the plastic range (thermal fatigue). Compressive loads that are at an equilibrium with the tensile loads relieve the hot hub area. This causes the hub to experience low-frequency LCF cyclical loads (Chapter 12.6.1) when the engine is started up and shut down (Fig. "Turbine disk loads during operation cycles"). These loads have a decisive effect on the fatigue life of the disk.
Figure "Turbine disk loads during operation cycles": This diagram schematically depicts the temperature and stress during a full start-up/shutdown cycle, including steady state operation. The temperature distribution and mechanical loads in the disk cross section change in characteristic ways (also see Fig. "Loading of a turbine disk in the startup phase"). An integral turbine disk was selected as an example, as the effects in this case were especially pronounced. Turbine disks with individual blades, as are common in all larger engines, tend to have the same behavior, albeit less pronounced. Rotor blades are fixed into the disk by a fir-tree connection on the end of a neck. The root and the annulus area it is set into are intensively cooled by the cooling air for the blade.
Assuming that the hot gas temperatures have the depicted simple progress, the annulus area
“A” experiences rapid heating-up due to the hot gases during start-up. This creates high compressive stress due to restricted thermal expansion. This stress is then reduced by the thermal equilibrium in the disk. When the engine is shut down, the annulus cools more quickly than the hub and high tensile stress is created in “A”. This LCF stress causes thermal fatigue loads (Chapter 12.6.2), in some cases with annulus cracking (Refs. 184.108.40.206-3 and 220.127.116.11-4).
High tensile stresses are created in the hub area “B”, and decrease over time as thermal equilibrium is reached. After shutdown, the remaining thermal strain (slow cooling of the massive hub) can become compressive without the overlaying centrifugal forces.
Illustrations 18.104.22.168-9 and 22.214.171.124-10: Typical signs of overheating on the blading can indicate the temporal temperature changes even during external damage inspections (also see Chapter 12.4).
“Burning”: This extreme form of oxidation can be seen in missing blade sections. The edges bordering this area are rounded. The rough surface is usually coated with oxides (Example "Burnt turbine blades").
Creep fractures: If the temperatures are near the solidus line, at which the first melting in the structure (usually near grain boundaries, right diagram) occurs, the outer regions of the blades can be expected to be flung off. The damage symptoms are even around the entire circumference (middle diagram). The cracks/fractures run along the grain boundaries, and are usually perpendicular to the centrifugal force. These single cracks are connected by shear planes (Fig. "Pore formation as creep damage"). Some grain boundaries gape open (Fig. "Damage at overheated turbine disk").
At lower temperatures, plastic deformations (creep) occur. Typical symptoms include bent-open shrouds (Fig. "Typical damages to rotor blade shrouds" ) and constricted blades (Fig. "Creep deformations in turbine blades").
Melting: This only refers to cases in which the liquidus temperature was exceeded in a limited blade area (e.g. inlet edge). This is only possible if very high temperatures act for a very short time (e.g. darting flame). Longer times would affect larger areas. Causes of short-duration overheating include compressor stalls, damage to injection nozzles, or ignition problems in the combustion chamber. The typical damage symptom is an intact oxide skin that contains the melted material. This creates a skin-like structure that is characteristically deformed (often wrinkled) by the melt.
In diffusion-coated blades, the coating can melt beneath the surface near increased concentrations of low-melting components (e.g. aluminum). This causes the coating to delaminate, tear open, and/or roll up (Fig. "Temperature caused damages at high pressure turbine vanes").
Thermal shock: This is a one-time, short-duration occurrence with localized high temperatures below the solidus range (Fig. "Symptoms of 'thermal shock' cracks"). The result is plastic compression followed by laceration. A typical sign is crack fields in the overheated zone.
Orange peel effect: see Fig. "Turbine guide vane thermal damages".
Hot cracking: If softened or melted grain boundaries tear open, it is referred to as hot cracking (Volume 2, Chapter 7.2.2). This is typical for cracking during rubbing and welding.
Figure "Turbine guide vane thermal damages": These damage symptoms occur when engine part zones are overheated to temperatures slightly below the solidus temperature for extended periods. The appearance is described as an orange peel effect. This phenomenon can be observed especially on the inlet edges of high-pressure turbine blades. It comprises flat cracking fields in a zone with heavy oxidation and considerable material removal from the surface. This material removal, combined with the cracking, results in an oriented surface structure (rippling). The thermal fatigue cracks are enlarged by oxidation (“washed out”). This involves a noticeable temporal influence and, at least for a time, very slow crack growth.
Figure "Causes of excessively high part temperatures": The temperature of a turbine rotor is especially dependent on the temperature distribution in the hot gas and the cooling air inflow.
“1”, “2”, and “3” show various temperature profiles in the hot gas flow. The most desirable is a trapezoidal profile (Fig. "Temperature variation at the combustion chamber outlet") with a low maximum temperature (“1”). The combustion chamber itself, as well as the flow of cooling air into the hot gas from upstream blades, labyrinths, and static gas-carrying parts, influence temperature distribution.
Hot gas encroachments into the annulus area (“4”) of the disk are the result of events such as seal failures, rotor blade root fractures (Fig. "Blade root failure causing hot gas incursion to disk"), or a lack of cooling air. Depending on the intensity, this can lead to catastrophic damage in a fraction of a minute or after several hours. Typical damage symptoms are shown in the right diagram. The disk side flow always tends to allow hot gas into the ring duct (disk side space) in areas of low pressure. This can be controlled by designs that ensure appropriate pressure ratios, e.g. through air inflow (barrier air, cooling air; Fig. "Suck effect of flow in ducts with rotating walls").
“5”, “6”, and “7” show the cooling air around the disk to the blade. Changes in this area can occur in connection with an unusual long-term change in the seal systems and increased leakages, for example. The OEM should have taken normal wear into account.
Figure "Suck effect of flow in ducts with rotating walls" (Ref. 126.96.36.199-14): Ring spaces (disk side spaces) are formed between the rotor disks and stators (top left diagram; Fig. "Flow conditions in ring-shaped chambers"). The wall formed by the rotor disk rotates and centrifuges the flow outward through friction with the gas (air). On the stator side, the flow is directed inward. Inside these gas flows along the wall, a core flow forms and flows around the circumference at roughly half the rotating speed of the rotor wall (bottom diagram). The friction forces between the walls and gas, and within the gas itself (agitation losses) cause considerable heating-up of the ring space if it is not cooled by inflowing air (barrier air, cooling air). The flow in the disk side spaces interacts with the hot gas flowing past outside. This creates pressure differences of the type that typically occur at the blade profiles (pressure and suction side). In areas of the gas flow with relatively high pressure levels, the gas flow tends to encroach into the ring space. In areas with lower pressure levels, the disk side flow can enter into the main flow (top right and bottom right diagrams).
Figure "Flow conditions in ring-shaped chambers" (Ref. 188.8.131.52-14): Various types of ring spaces are formed between rotor systems or rotors and stators. The selected example is the depicted turbine of a fighter aircraft engine with one HP stage and one LP stage. The different systems are rotor/rotor (both side walls rotate), rotor/stator (one side wall rotates), and rotating ducts (at least the outer wall rotates; e.g. in hollow shafts). The flows in these spaces are very different and influence effects such as heat development, erosion (Volume 1, Ill. 5.3.1-8), cooling air throughflow, and tendency for hot gas encroachment.
Figure "Hot parts cooling structures blocked by dust": Dust damages compressors not only through erosion. In modern engine types with complex cooling systems for the front turbine stages (Fig. "Cooling systems of turbine blades and vanes"), there is an increased danger of overheating damage due to a blockage in the inner cooling air duct and corrosion and/or oxidation.
High temperature corrosion is especially promoted by sulfuric dusts such as gypsum (Volume 1, Chapter 5.4.5). This danger depends on the area where the cooling air is extracted (zone “A”). If possible, cooling air should be extracted from an area of the compressor with a low dust concentration. Because of the centrifugal effect of the compressor (Volume 1, Chapter 5.3.1), the area near the rotor hub is suitable for air extraction (bottom diagram, Ref. 184.108.40.206-5). If this is not possible in every case in triple-shaft engines (top diagram), additional measures may be necessary to reduce the concentration of dust.
One possibility may be a redirection of the cooling air flow before it enters the blading. This would at least eliminate larger particles (Example220.127.116.11-1 and Fig. "Blocking as 'disease' of hot parts cooling systems").
Figure "Reduction of life span changes in cooling bores" (Ref. 18.104.22.168-8): In most cases, the deposits on and in blades consist of particles carried by the hot gas flow or the cooling air. In the literature, it is assumed that particles smaller than 0.001 mm do not strike the part surfaces on the hot gas side or create dangerous deposits. Evidently, sufficiently small particles will usually simply follow the gas flow. Particles between 0.001 and 0.01 mm will strike the surfaces and make deposits unavoidable. Even larger particles capable of unallowably shifting the surface section can enter into the cooling air flow. Typical examples include remaining melt drops from laser boring during production, shot peening particles, or spalling remnants of core material. Particles from spalled labyrinth armor in the cooling air flow and metal shavings from rubbing are also capable of creating blockages.
The trajectory of particles in the hot gas flow is influenced by the cooling air film. This affects their tendency to build up in layers around air bleed bores (top diagram, Ref. 22.214.171.124-10). Evidently, considerably fewer particles are deposited in front or behind the discrete air bleed jet than on the surrounding surface (Ref. 126.96.36.199-9).
Deposits can also be expected on the cooling air side (inside the blade) if the cooling air flow carries particles prone to depositing (Ref. 188.8.131.52-8). The location and size of deposits depend especially on the angle between the bores and the direction of flow (hot gas, cooling air), as well as the flow speed (middle diagrams). Local changes in the flow speed, which can occur in areas where the flow is constricted after a separation in the bore, can influence the location and size of deposits. Typical areas are bore zones with high flow speeds (“jetting”), and separation bubbles with low flow speeds. Depending on the consistency of particles and their tendency to stick, zones with either high impact speed or with low speeds, in which the particles can settle, are prone to depositing. Under consideration of the above factors, suitable shaping of the bore and optimization of the cooling air flow can reduce the tendency to blockages.
The interplay of influences (table) during the creation of deposits is extremely complex. When considering the development of a coating from individual particles, bonding is a matter of special concern. Typical bonding mechanisms act like thermal spray coatings (Fig. "Thermal barrier coatings spalling limiting life span"):
It is to be expected that problems with deposits are engine-specific, and depend on its operation and location. In order to prevent damaging deposits, influences must already be known and considered during design, construction, and testing. If this is not the case, subsequent improvements will become necessary for specific applications of the serial engine (e.g. operation in especially dusty environments).
Figure "Deposits influencing turbine blade operation behavior": The following are considered to be particles that create deposits on the hot parts:
Deposits form dependent on of the specific conditions (Fig. "Reduction of life span changes in cooling bores"). They influence the operating behavior of the components, and therefore also that of the entire engine (Ref. 184.108.40.206-8):
Thick coatings can considerably change the blade profiles, worsening the efficiency of the turbine. This situation can occur in air that carries large amounts of sand (Volume 1, Ill. 5.3.2-12.1). In extreme cases, such as flight through clouds of vulcanic ash (Volume 1, Ill. 5.3.2-14), the flow cross-section in the front turbine stages can be constricted to the point that the engines fail during flight.
Deposits in cooling air bores are influenced by the flow conditions of the hot gases and the cooling air (Fig. "Reduction of life span changes in cooling bores"). Deposits can form and constrict the flow cross-section considerably. One unique phenomenon is radially oriented deposits in the middle of the bore (“B”, “shark fin”, Ref. 220.127.116.11-8). On the hot gas side, there is a danger of bores for the cooling air film becoming partially (“C”) or, in extreme cases, completely blocked.
Left-over particles from the production process can have especially serious effects. At typical high operating temperatures, SiC grains from grinding and peening processes react with the substrate and damage it through diffusion or melting. The same is true for coating powders (for the production of anti-diffusion coatings) with high Al content and aggressive halogen components (Fluorine, Chlorine).
Even surfaces with ceramic thermal barriers are susceptible to reactions (Volume 1, Ills. 18.104.22.168-4, 22.214.171.124-5, and 126.96.36.199-4).
Figure "Causes of turbine disk overheating": Causes of the overheating of turbine disks (Ref. 188.8.131.52-5):
Hot gas encroachment in the annulus area: The failure of a seal (e.g. intermediate stage labyrinth) or the fracture of a blade at the root (Fig. "Blade root failure causing hot gas incursion to disk") can cause the annulus area to be struck by hot gas. Within a fraction of a minute, dangerous heat levels can plastically deform the fir-tree connections and cause blades to be released. The preliminary stage of this process is a permanent decrease in strength (can be detected by loss of hardness) that causes damage only after longer operating times.
Lack of cooling air: This occurs due to seal damage in the cooling air system and/or pressure changes in the air system. It can cause temperature increases both in the annulus (blading/disk) and in the hub.
In addition to hot gas encroachment and a lack of cooling air (Ill. 184.108.40.206-11), there are other causes of overheating:
Oil fires: If oil escapes from the turbine bearings, it will almost certainly ignite immediately in the high temperatures in this area (Volume 2, Chapter 9.2). Experience has shown that this situation even puts thick disk sections at risk of overheating. The hub near the leakage, especially, can dangerously overheat within minutes. This poses the immediate threat of the disk bursting.
Rubbing: Rubbing in the disk area can dangerously weaken supporting sections in seconds. Damage mechanisms include hot cracking, softening, and creep (Volume 2, Chapter 8.2). Excessive rubbing of integral labyrinth rings, as are used in turbine disks in smaller engines (e.g. helicopter engines), is especially likely to result in a catastrophic failure (Volume 2, Chapter 7.2.3).
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126.96.36.199-4 A.K.Koul, “Hot Section Materials for Small Turbines”, AGARD Proceedings CP-537, of the conference “Technology Requirements for Small Gas Turbines”, Montreal, Canada, 4th-8th October 1993, pages 40-1 to 40-9.
188.8.131.52-5 Rolls-Royce plc, “The Jet Engine”, ISBN 0 902121 2 35, Fifth Edition 1996, pages 85-93.
184.108.40.206-6 E.A. Witmer, T.R. Stagliano, J.J.A. Rodal, “Engine Rotor Burst Containment/Control Studies”, conference proceedings AGARD-CP-248, 1978, pages 15-3 to 15-29.
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18.104.22.168-8 V.H.M.Kuk, P.T.Ireland, T.V.Jones, M.C. Rose, “Particle Deposition in Gas Turbine Blade Film Cooling Holes”, conference on “Erosion, Corrosion and Foreign Object Damage Effects in Gas Turbines”, Rotterdam, The Netherlands, Proceedings AGARD-CP-558, 1994, pages 10-1 to 10-18.
22.214.171.124-9 D.G.Bogard, D.L.Schmidt, M.Tabbita, “Characterization and Laboratory Simulation of Turbine Airfoil Surface Roughness and Associated Heat Transfer”, Proceedings Paper ASME 96-GT-386 of the “International Gas Turbine and Aeroengine Congress & Exhibition”, Birmingham, UK, June 10-13, 1996, pages 1-7.
126.96.36.199-10 S.Baldauf, “Filmkühlung thermisch höchstbelasteter Oberflächen: Korrelation thermographischer Messungen”, research reports from the Institute of Thermal Flow Machines, University of Karlsruhe (TH), Volume 12/2001, Logos Verlag Berlin, ISBN 3-89722-639-1, pages 4-35.
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184.108.40.206-12 E.Lugscheider, K.Bobzin, A.Etzkorn, “Widening the usability of yttria stabilised zirconia by advanced cooling technology”, periodical “Mat.-Wiss. u. Werkstofftech”, 32, 2001, pages 660-664.
220.127.116.11-13 E. Berghof-Hesselbächer, H. Echsler, P. Gawenda, M. Schorr, M. Schütze, “Zeit- und temperaturabhängige Entwicklung von physikalischen Defekten in Wärmedämmschichtsystemen”, periodical (Carl Hanser Verlag) “Praktische Metallografie”, 40, 2003 (5), pages 219-231.
18.104.22.168-14 C. Lechner, J. Seume, “Stationäre Gasturbinen”, Springer Verlag Berlin Heidelberg New York, ISBN 3-540-42831-3, 2003 pages 592-595.
22.214.171.124-15 P. Morrell, D.S. Rickerby, “Advantages/Disadvantages of Various TBC Systems as Perceived by the Engine Manufacturer”, Proceedings Paper AGARD-R-823 of the 85th Meeting of Agard Structures and Materials Panel“, Aalborg, Denmark 15-16 October 1997, pages 20-1 to 20-9.