Seal system for a gas turbine The present invention relates to an seal system for a gas turbine which seal system comprises a rotor blade tip with an abrasive coating system and a stationary heat shield ring coated with an abradable coating and facing opposite to the tip .

Shroudless gas turbine blades are increasingly used in gas turbines. However, it is very difficult to find a good seal ^¬ ing method between such blade tips and heat shield segments above the tips to minimize the hot gas leakage and ensure the engine efficiency, especially, on the tips of 1 ^st and 2 ^nd stage blades where the temperature is high, i.e. about 1000 °C or higher. It is often to observe that rubbing and oxidation damages on both the tips of the blades and the heat shield segments occur after a certain operating period. These dam- ages increase the gap between the blade tips and the heat shield segments continuously, leading to an increased hot gas leakage thereby reducing the efficiency of the gas turbine.

EP 0 707 091 Al describes a sealing system of gas turbines comprising a blade tip coated with a zirconium-based oxide having a plurality of vertical macrocracks . The zirconium- based oxide with the macrocracks shows good rub tolerance when contacting a seal ring of bare superalloy. The zirco ^¬ nium-based coating of EP 0 707 091 Al is a zirconia coating with macrocracks where the zirconia is partially stabilized, by an amount of 6.5 to 9 weight percent yttria.

With respect to this prior art it is objective of the present invention to provide an advantageous seal system for a gas turbine. It is a further objective of the present invention to provide an advantageous gas turbine. An inventive seal system for a gas turbine comprises a rotor blade tip with an abrasive coating system and a stationary heat shield ring located opposite to the blade tip. In par ^¬ ticular, the heat shield ring may be formed by a number of ring segments together forming the seal ring.

The abrasive coating system on the blade tip comprises an oxidation resistant and/or corrosion resistant bond coat and a thermal barrier coating atop of the bond coat. The thermal barrier coating is a Zr02 coating that is partially stabi ^¬ lized by Y2O ₃ (abbreviated YSZ) and is free of macrocracks, at least of vertical macrocracks. The heart shield ring is coated with an abradable coating. In particular, the abrad- able coating may be a ceramic coating, for example a YSZ coating i.e., a coating of yttria stabilized zirconium oxide. In the context of the present invention, a crack is consid ^¬ ered to be a macrocrack if its lengh is at least 0,1 mm.

When a gas turbine engine equipped with the invention seal system is started the thickness of both, the thermal barrier coating on the blade tip as well as the abradable coating will be abraded. Due to the rubbing of the blade tip a recess is formed in the abradable coating of the heat shield ring into which the tip projects so as to form some simple kind of labyrinth seal. Rubbing occurs in particular during start up because the blades extend their length due to thermal expan ^¬ sion. In full load operation when both the heat shield ring and the turbine blade will be at their maximum temperature, rubbing will eventually cease and the blade tip will project into the recess rubbed into the coating of the heat shield ring at a maximum depth. Due to a clearance that then exists between the abradable coating and the blade tip no rubbing will take place anymore. In this condition, the thermal bar ^¬ rier coating of the blade tip may be fully rubbed off. How- ever, the oxidation resistant and/or corrosion resistant bond coat is still present to protect the tip. The clearance be ^¬ tween the blade tip and the seal ring is small enough to pro ^¬ vide a good sealing action. In that sense, the thermal bar- rier coating can be regarded as sacrificial coating on top of the bond coat.

The thermal barrier coating without macrocracks of the seal system is preferably a Zr02 coating that is partially stabi ^¬ lized by 5 to 9 wt% Y2O ₃ , in particular by 7 wt% Y2O ₃ . More ^¬ over, the thermal barrier coating may be applied by air plasma spraying (APS) or by electron beam physical vapour deposition (EB-PVD) .

The abrasive coating system, i.e. the bond coat and the ther ^¬ mal barrier coating, may be particularly applied only on the tip region of the blade, or it may be applied as an extension from the coating system of the airfoil, or as an extension from the coating system of the airfoil and the platform.

Coating also other regions of the blade than the tip region with the abrasive coating system allows to coat the turbine blade in a single coating process. The abrasive coating sys ^¬ tem is not subject to rubbing except for the tip region so that the thermal barrier coating will not be rubbed off in the other regions of the blade.

In a further development of the invention, the thermal bar ^¬ rier coating has a porosity of more than 10% by volume. This measure helps to prevent the formation of macrocracks in the thermal barrier coating during applying the coating.

A suitable bond coat for the abrasive coating system is a so called MCrAlY-coating where M stands for cobalt (Co) , nickel (Ni) , or both of them, Cr stands for chromium, Al stands for aluminium and Y stands for yttrium and/or silicon (Si) and/or Hafnium (Hf) and/or at least one rare earth element.

According to a further aspect of the invention, a gas turbine is provided. The gas turbine comprises a rotor with one or more stages of rotor blades, a heat shield encasing at least one stage of rotor blades and an inventive seal system. In the inventive gas turbine, the heat shield segments are coated with the abradable coating to form the seal ring. The tips of the rotor blades encased by the heat shield segments are coated with the abrasive coating system. Typically, at least the first stage of rotor blades would be fitted with the seal system. However, further stages, in particular the second stage, can also be fitted with the seal system. The inventive gas turbine has only small leakage so that high ef ^¬ ficiency of the gas turbine can be assured. Moreover, blade tip damages due to oxidation and/or corrosion can be reduced.

One preferred essential feature is that any ceramic coating being part of said abradable coating applied to said heat shield ring is free of vertical macrocracks . It is further preferred that any ceramic coating on top of said bond coat of said abrasive coating system of said blade tip is free of vertical macrocracks.

Further features, properties and advantages of the present invention will become clear from the following description of an embodiment in conjunction with the accompanying drawings.

Figure 1 schematically shows an inventive seal system.

Figure 2 shows the seal system of figure 1 after start of a gas turbine engine of which it is part.

Figure 3 shows the seal system of figure 1 after a while of full engine load operation.

In the following, the inventive seal system will be described with respect to figures 1 to 3. The seal system is part of a gas turbine with a rotor comprising a number of stages of turbine rotor blades arranged in axial direction of the rotor in alternating fashion with turbine nozzles. The stages of rotor blades are encased by annular heat shields composed of a number of heat shield segments. Typically, a stationary gas turbine has two to four stages of rotor blades and a corre ^¬ sponding number of vanes. However, a turbine may also have only one stage of turbine blades or even more than four stages. The gas turbine is equipped with an inventive seal system that comprises the tips of at least one of the stages of rotor blades and the respective heat shield.

Figure 1 schematically shows a heat shield segment 1 and a rotor blade 3. Other components of the gas turbine like, for example, the rotor to which the rotor blade 3 is fixed are not shown to keep the figure simple. The heat shield segment 1 is coated with an abradable coating 5 and represents a seal ring. In the present embodiment, the abradable coating is preferably a yttria stabilized zirconiumoxide coating (YSZ) . On the tip 7 of the rotor blade 3, an abrasive coating system is present that comprises an oxidation resistant and/or cor ^¬ rosion resistant bond coat 9 and an overlying thermal barrier coating 11 that forms the top coat of the abrasive coating system.

In the present embodiment, a MCrAlY-coating is used as a bond coat 9, where M stands for iron, cobalt and/or nickel, Cr stands for chromium, Al stands for aluminium and Y stands for yttrium, hafnium, silicon or at least one rare earth element, or a combination thereof. The top coat of the abrasive coat ^¬ ing system is, in present embodiment, zirconium oxide (Zr02) that is partially stabilized by 6 wt% yttrium oxide (Y ₂ O ₃ ) .

When the top coat is applied by air plasma spraying, the spraying parameters are used to avoid the formation of macro- cracks, in particular the formation of vertical macrocracks. In other words, the parameters, when applying the YSZ-coating by air plasma spraying, are chosen such that only cracks oc- cure that are shorter than 0.1 mm. Typical parameters that can be varied to achieve a macrocrack-free coating are tem ^¬ perature and velocity of the particles used in the spray process. Avoiding macrocracks can, for example, be achieved by setting the parameters in the air plasma spray process such that the applied YSZ-coating has a porosity larger that 10 % by volume. In this case, the formation of macrocracks is typically prevented. Although the abrasive coating system is only shown on the tip of the rotor blade in figure 1 the coating system may be an extension of the coating system applied on the whole airfoil 13 and the platform 15 of the rotor blade 3 which are exposed to hot combustion gases during gas turbine operation.

Figure 1 shows the inventive seal system formed by the seal ring 1 with the abradable coating 5 and the blade tip 7 with the abrasive coating system composted of the MCrAlY-bond coat 9 and the YSZ-top coat 11 just after applying it.

Figure 2 shows the seal system after start of the gas tur ^¬ bine. In gas turbines, the rotor blades 3 are arranged with a tiny gap between the tips 7 of the rotor blades 3 and the en ^¬ casing seal rings 1. Please note that in the figures the gap between the surface of the top coat 11 and the surface of the abradable coating 5 is exaggerated for clarity reasons. Dur ^¬ ing start up operation of the gas turbine, an elongation of the turbine blades 3 occurs due to thermal expansion and cen ^¬ trifugal forces experienced by the blades 3 due to the rota- tion of the rotor. On the other hand, since the seal ring ex ^¬ periences no centrifugal force the diameter of the ring in ^¬ creases less than the length of the rotor blades. In addi ^¬ tion, during start up of the gas turbine the thermal expan ^¬ sion of the rotor blades 3 and the seal ring 1 typically dif- fer from each other. Both effects together lead to rubbing between the tips 7 and the seal ring 1. During this rubbing, a part of the abradable coating 5 is rubbed off by the abra ^¬ sive coating system on the blade tip 7, in particular by the ceramic top coat 11. However, also the ceramic top coat is rubbed off so that its thickness will be reduced over time. After start up and a while of full load operation of the gas turbine the ceramic top coat 11 may be fully disappeared due to rubbing and due to lifetime limiting effects by bond coat oxidation in combination with thermal fatigue leading to spallation. However, the sealing is still ensured and the bond coat 9 will not be rubbed off due to a space 17 that has been generated by the ceramic top coat 11 abrading into the abradable coating 5. Hence, the bond coat 9 will still pro ^¬ tect the blade tip 7 from high temperature oxidation and/or corrosion for the rest of the operation time.

The YSZ-coating on the blade tip 7 provides a better abrad- ability at the high temperatures together with an abradable coating than a bare blade tip does. Moreover, the coated blade tip withstands a high temperature exposure up to 1200°C, and is even protected from oxidation and/or corrosion damages. Although the YSZ-layer 11 of the abrasive coating system may be fully rubbed off by abrading to the abradable coating of the seal ring or by spallation when it reaches its lifetime during gas turbine operation within an overhaul in ^¬ terval, the YSZ has already rubbed into the abradable coating with a maximum depth after a short time of full-load opera ^¬ tion and, hence, fulfilled its function. In this sense, the YSZ-coating could be regarded as a sacrificial coating. The removal of the YSZ-coating has left a space between the bond coat surface and the rubbed surface of the abradable coating 5, so that the bond coat 9 will not or minimally be rubbed off in the later operation and can still protect the blade tip 7 from the oxidation.

The invention has been described with respect to an exemplary embodiment thereof as an illustrative example of the inven- tive seal system and the inventive gas turbine. However, please note that although a special embodiment has been de ^¬ scribed to explain the invention deviations from this embodi ^¬ ment are possible. For example, although the zirconium oxide was stabilized by 7 wt% yttrium oxide in the described em- bodiment the contend of yttrium oxide may vary between 5 wt% and 9 wt%. In addition, although an MCrAlY-coating has been described as bond coat other oxidation and/or corrosion resistant bond coats, in particular other alumina scale forming bond coats, may be used. Hence, the scope of the invention shall not be limited by the described exemplary embodiment but only by the appended claims.