Industrial Gas Turbines

"Illustration 3.3-1": The typical turbine disks of engines of higher performance classes (sketch above left) show fir tree serrations in the rim area, which accommodate the blade roots. Often, the retaining and sealing of the blades (e.g., through metal sheet segments) is planned in the rim. The disk diaphragm is frequently provided with labyrinth seals for cooling air management. The torque transmission is usually made through bolt connections. These can be in the outer disk area ( bolt holes in the disk), as well as through a flange near the hub. For the stress in the disks the flanges are more favorable because there is no notch effect in the highly loaded disk areas. Turbine disks have flanges near the hub, in compressors and LP turbines we find on the other hand flanges near the rim. In the hub area, the turbine disks, especially those of the high pressure stages are strongly thicker, in order to absorb the extreme thermal stresses and centrifugal forces ( "Ill. 2.2-3" and "Ill. 3.3-5"), at comparatively high component temperatures. The scetch above right shows a one piece turbine wheel, a so called integral turbine wheel (a blisk type). Those are normally cast and are found in small gas turbines, often derivates of helicopter engines. Integral turbine wheels have specific problems ( "Ill. 3.3-17" and "Ill. 3.3-18"). Integral turbine wheels made of wrought material by chipping are only used by some OEMs because of the lower hot strength and high production coasts. The blade of a blisk is directly ‘grown’ at the rim of the disk. In the region of the disk diaphragm normally labyrinth rings (seal rings) are fitted to prevent a hot gas invasion, especially into the cooling air stream. The disk often bears a balance collar at which material can be removed. In the high stressed, thickened hub region we find OEM-specific flanges for the torque transfer and centering purposes.

Turbine guide vanes are equipped with an inner and an outer platform (lower sketch). It is likely that two or three blades (in low pressure turbines =LPT, even more, sketch lower left) are connected into a segment, whereby a noticeable stiffening against high gas bend forces and vibration (long and slim blades of the LPT) is aimed at a certain „fail safe” behavior. An overloaded single vane does not lead to a spontaneous component failure during the overheating of single vanes. The vanes of the front stages, especially of the high pressure turbine, are provided with effective convection- and film cooling ( "Ill. 3.3-3" and "Ill. 3.3-6"). In gas turbines of small performance, integral guide vanes in the high pressure part, as well as in the low pressure turbine, come successfully into use (sketch low left). In such a design, all the guide vanes comprise a single part. This can be a casting and /or it is joined through welding and brazing. The disadvantage of this configuration are high thermal stresses and so a certain sensibility for thermal fatigue ( "Ill. 3.3-17" and "Ill. 3.3-18"). These guide vanes, however, do not show for casting technology reasons a comparatively complex cooling structure as do guide vane segments of bigger engines.

"Illustration 3.3-2": Turbine rotor blades can show different design principles. There are manufacturers who prefer blades with shrouds, as these have the advantage of the blade damping and supporting against vibration and are an effective sealing of the tip clearances with no special rubbing problems. Disadvantage is the additional centrifugal force of the blade as well as the wear problems on the contact surfaces and crack initiation at the shrouds ( "Ill. 3.4-1"). The blade airfoils of the high pressure turbine stages are usually intensively cooled ( "Ill. 3.3-3" and "Ill. 3.3-6"). Only in small gas turbines, under some hundred KW power output, with integral turbine wheels and without individual blades ( "Ill. 3.3-1"), are the rotor blades not cooled.

The airfoil is bordered by the root platform in the direction of the hub. Next to the gas guidance, the task of sealing is taken over also, in order to prevent gas ingestion to the blade root and/or to prevent a prohibited cooling air leakage. One of the tasks of the blade root shank is the favorable force transmission of the blade forces into the root. In some cases, through holes in the shank, a part of the cooling air is added to the blade. The blade root transports through the fir tree teeth, the blade forces into the disc. In contrast to the dove tail root of the compressor blades, the turbine blades have a fir tree connection, in order to control the high loads of the heavy blades at high operation temperatures. Manufacture specific is the size, respectively, the number of teeth of the fir tree. It effects, e.g., the over speed behavior, that means the tendency to loosen a blade during overspeeds by widening of the discs. So the burst of a disc can be avoided.

"Illustration 3.3-3": The rotor blades in the high pressure turbine must withstand the high centrifugal forces at high operation temperatures. To achieve this, the temperature of the supporting cross sections must be clearly lowered. The demanded cooling air acts as convection cooling (heat conduction) in the blade interior (right sketch). In the area of the leading edge, an especially intensive cooling is necessary. The required, good heat transportation is attained in that, through a perforated wall of a neighboring cooling air passage, the air is blown directly against the inner surface of the leading edge (impingement cooling).

In order to guarantee the film cooling (left sketch) of the blade surface on the leading edge against the gas flow, an especially higher cooling air pressure is necessary. In modern engines, this is achieved through a small additional compressor by a structured cover plate fitted onto the disc (“cover plate”). With this device the cooling air is additionally compressed in front of the inlet into the blade root.

"Illustration 3.3-4": The development trend goes in the direction of increasing the performance and efficiency together with even longer operation life. This trend assumes an increase of the hot gas temperature at the turbine entrance (combustion chamber exit). The increase of the high temperature strength, especially the creep and creep rupture ( "Ill. 2.3-1" and "Ill. 2.3-2") of the turbine blade materials is, next to the resistance against thermal fatigue ( "Ill. 3.3-16").

Because of their higher heat resistance, (coarse grain, dendritic structure and high portion of precipitation phases), the casting alloys have replaced the forging alloys of the sixties.

The components are produced in precision cast process (lost wax process). A true wax core is manufactured. Hollow (cooled) structures need inside the wax ceramic cores in the shape of the cooling system. This is covered with the ceramic layers in a complex immersion and drying process. The wax core is then melted out, the form burned. The casting process mostly under carried out in inert gas or vacuum.

The metal solidifies relatively quickly from the form surface into the inner and a polycrystalline structure originates (left sketch). The size of the grain is irregular and the direction/orientation is influenced by the local temperature gradients and solidification speed (conventionally cast, multi-axed).

Because the grain boundaries have materially typical weak points prone to creep and thermal fatigue, ( "Ill. 3.3-13"), it is attempted to orientate structures along the main load direction, in turbine rotor blades radially (zentrifugal forces) and in turbine vanes correspondent to the thermal stresses. This happens through suitably aimed temperature gradients during the solidification of the casting. The crystals grow in the direction of these gradients. An equi axed structure (directional solidified = DS), is created, (sketch in the middle). It was recognized, however, that near the main loading direction transverse loading takes place, especially in the form of thermal stresses. These can tear the longitudinal grain boundaries of the directionally solidified structures in a manner similar to wood.

In order to avoid all grain boundaries, single crystal (= SC) components (right sketch) were developed (Lit.3.3-1) and introduced in the nineties. Those “technical single crystals” consist of an alloy and are not high-purity and free of micro inhomogeneities like crystals in the chip industry. The temperature guidance during solidification of the cast is here even more exactly controlled than by directional solidification. Specific measures are necessary to guarantee the growth of one crystal at the beginning of solidification of the blade/part itself. These single crystal materials are also alloy technically optimized, to further increase high temperature strength and oxidation resistance.

"Illustration 3.3-5": During a start /shut down cycle, the temperature distribution and the mechanical load in the disk cross section changes in a characteristic manner. To exemplify, an integral turbine wheel ( "Ill. 3.3-1") is chosen by which the effects are especially distinctive. Turbine discs with single blades show the same, but less distinctive tendency, as is normal in all bigger engines. This is because the blades in the high pressure turbine region have a shank above the fir tree root where they are intensively cooled.

Assuming the hot gas temperature and rotational speed show the represented simplified trend in the rim region, „A“, a quick heating up through the hot gases follows at the start via the blades. The result is high compression stresses in the plastic region due to restriction of thermal expansion by the cool hub sections. This stress reduces in the disk with the balancing of temperature. During shut down of the engine, the rim cools quicker than the hub and there originate high tensile stresses in „A“. This LCF load leads to a thermal fatigue ( "Ill. 3.3-16"), occasionally, with crack initiation in the rim (‘rim cracking’, Lit. 3.3-2). These cracks are primarily axial (engine) orientated between the blades ( "Ill. 3.3-17" and "Ill. 3.3-18").

In the hub region „B“, high tensile loads initiate, reducing in time because of the temperature balance. After the shut down there are still existing thermal stresses (slow cooling of the massive hub). Therefore, without superimposed centrifugal force the hub region can come into compression.

"Illustration 3.3-6": The nozzle guide vanes of the high pressure turbine (‘nozzle’), directly behind the combustor ( "Ill. 3.2.3-2"), are the most intensively cooled components of the gas turbine.Thus, a film cooling (left sketch) can form despite the high pressure niveau of the gas flow. The cooling air is removed in the area of the compressor exit. The cooling air usage of the high pressure turbine nozzle guide vanes alone lies within several percentages of the air flow rate of the engine. This energy of the cooling air is not lost. During expansion in the turbine, it improves the power output up to a certain grade. If the gas temperature increases, (e.g., in order to increase the performance and efficiency), the cooling air consumption must also increase at the same hot part technologies. Despite that, one makes the effort to minimize the cooling air use of this component. This is successful through the use of ceramic materials in the form of thermal barrier coatings ( "Ill. 3.2.3-3", "Ill. 3.2.3-4" and "Ill. 3.2.3-5"). These thermal barrier coatings decrease the heat flow in the vane wall and the necessary intensity of the cooling. There are already vanes that use monolithic ceramic components in development (‘hybrid blade’, "Ill. 5.2-1").

An especially intensive and effective cooling of the vane surfaces and the platforms of metallic vanes demands a sieve like perforation of the vane walls for the film cooling. This so called effusion cooling with the help of porous vanes walls is, however, not implemented in serial use, because of diverse problems such as crack formation and blockage.

In the inside of the vanes, the cooling air is distributed through sheet metal inserts (right sketch).

"Illustration 3.3-7": Diffusion coatings serve to avoid prohibited oxidation and hot gas corrosion of the hot parts. They are mostly produced through diffusion of aluminum, sometimes also with platinum at higher temperatures in a suitable atmosphere. The aluminium forms a dense protective aluminum oxide layer at operation temperature. The coating thickness lies normally clearly above 0.1 mm. These coatings are particularly suited to diminish oxidation. Where there are other demands with noticeable corrosion, e.g., sulfidation protection ( "Ill. 3.4-2"), other metals (e.g., chrome) are brought in for diffusion. In the diffusion process during production, a portion of the coating with especially high Al content is built up on the base material ( top layer), a further part is diffused in the base material (diffusion zone).The coatings can be altered through the operation. Those changes can give important hints about the operation to the expert. To this belong damages, remaining lifetime and temperatures. Here are some typical effects:

• „A“: Crack initiation (so called coating cracks) occurs by low temperatures, where the coating is relatively brittle.
• „B“: Formation of „oxide nails“ at coating cracks in the area of the substrate. Here, the relatively low Al content offers no sufficient oxidation protection any more.
• „C“ : Rippling through cyclical heat expansions and plastic coating deformations at high temperatures.
• „D“: Delamination of the coating through crack formation in the area of brittle phases.
• „E“: Formation of brittle phases through diffusion processes in the region. Affected are base materials and coatings dependent on temperature and time.
• „F“: Pore formation at the transition to the base material through the Kirkendall effect. This is a diffusion process which balances alloying constituents.
• „G“: Loss of coating through the combined effect of oxidation, erosion and cyclical heat expansions. It is the main criterion of the coating life.
• „H“: Micro structural changes in the coating can give the expert important hints as to the real emerging operation temperatures. Thus one can infer, e.g., from structure alterations, whether damages have taken place through overheating .
• „I“: Inter diffusion between base materials and coating change the structure and composition of both. Such diffusion processes can, even when slow, also arise at normal operation temperatures. They can be decreased through so called diffusion barriers. Intermediate layer made out of platinum is appropriate here.
• „K“: Beginning melting of the transition zone with coating delamination at temperatures in the region of 1250°C. Thereby they are an indication of extremely high surface temperatures.

"Illustration 3.3-8": The smallest cross section in the gas stream is in the area of the the HPT nozzle ( "Ill. 3.2.3-2") guide vanes at the combustor exit. This cross section is individually and very exakt adjustet to the gas turbine. Even seemingly small changes can serious affect the operation behavior. To this belongs, e.g., the surge margin ( "Ill. 3.1.1-2"). Typical problems can emerge during operation with buckling by creep under thermal pressure stresses. It may also concern erosion, oxidation, overheating (‘burning’) and breakouts. The OEM manual or directives should give informations about the limits (e.g., the borescope indications).

It is important that with a repair this cross section is not incorrect influenced. To this belong not exact sized repairs.

A further problem are flow disturbances on deformations, that provoke dangerous vibrations of the rotor blades.