Cooled turbine blade

The invention relates to a coolable turbine blade according to Claim. 1

Internally cooled turbine blades, the meandering a Interior (EP 1022434 A2) comprise as well as impingement cooling, are known.

The impingement cooling in the interior of turbine blades is also known.

The DE 32 34 906 A1, US Patent No. 5,857,837, 5,873,695, 5,902,093, 5,462,405, 6,139,269 show a turbine blade whose leading edge is cooled by impingement cooling.

The DE 199 63 716 A1, US Patent No. 4,474,532, 4,753,575 and 4,767,268 show a turbine blade in a cavity exhibit area of ​​their blade tip.

The US-PS 6,431,832 shows a turbine blade which has a Cavity has in the region of the blade tip and a Cavity in the region of its leading edge.

The effectiveness of the cooling in the region of the leading edge However, a turbine blade is not sufficient.

It is therefore an object of the invention to provide a turbine blade, wherein the cooling of the leading edge improves is.

The object is solved by a turbine blade according to Claim 1, wherein the leading edge is cooled by impingement cooling is and when a crown pocket (cavity in the region the blade tip) is present.

In the subclaims, further advantageous embodiments the turbine blade listed. The measures listed in the subclaims can in advantageous manner be combined.

Show it:<dl tsize="7"><dt>figure 1</dt><dd>a turbine blade,</dd><dt>figure 2</dt><dd>a cross section of a turbine blade and</dd><dt>figure 3</dt><dd>a gas turbine, in which the inventive design The turbine blade is used.</dd></dl>

Figure 1 shows a perspective view of a blade 120, 130 (Fig. 3) extending along a longitudinal axis 121 (radial Direction). The blade 120, 130 has successively along the longitudinal axis 121 a securing region 400, a fact adjoining blade platform 403, a blade region 406 and a blade tip 19. The airfoil section 406 is at a leading edge 13 is flowed through by a medium. In the fastening region 400, a blade root 183 is formed, used to secure the rotor blades 120, 130 on the shaft 103 (FIG. 3) is used. The blade root 183 is hammerhead configured. Other configurations, for example as Firtree or dovetail are possible. In conventional blades 120, 130 are in all areas 400, 403, 406 of the blade 120, 130, solid metallic Materials used. The blade 120, 130 may in this case by a casting process, by a forging process, by a milling process be made or by combinations thereof.

2 shows such a turbine blade 120, 130 in Longitudinal section. 2 shows a turbine blade or vane 120 130 of a gas (Fig. 3, but also aircraft turbine) or Steam turbine ready.

The turbine blade 120, 130 is at least partially hollow executed and is cooled inside. The turbine blade 120, 130 has the blade root 400 at least one first coolant inlet region 22, a second coolant inlet region 25 and a third coolant inlet region 28. The coolant inlet areas 22, 25, 28 can selectively closed (here, for example, 28) or a narrowing 31 (here, for example, the coolant inlet region 22) by means of a coolant stream into the interior can be adjusted. The constriction 31 can also be regulated , ie increased or decreased during operation will. In this embodiment, the third coolant inlet region 28 completely closed, and the first coolant inlet region 25 narrows. The coolant flows through the first and second coolant inlet region 22, 25 into the interior of the turbine blade 120, 130th The turbine blade 120, 130 is of axially 34 a hot medium flows around. The hot medium meets in the axial Direction 34 first on the leading edge 13 (shovel nose) and then flows to the blade surface 406 by up to Trailing 10th The coolant through the first coolant inlet region 22 in a nasal cavity 37, which in the axial direction 34 follows directly behind the leading edge 13, penetrates, flows in the radial direction 121 to the blade tip 19th The nasal cavity 34 extends for example parallel to the leading edge . 13

In the region of the blade tip 19 and the leading edge 13 for example, but not necessarily, at least one first outlet opening 40 is provided for the coolant, so that there is a film cooling of the blade tip can be done. In any case, the coolant flows through at least one Connection opening 41 in a crown pocket 43 (in the cavity Region of the blade tip 19) and cools there the blade tip 19 over its length in the axial direction 34th The crown pocket 43 extending from the nasal cavity 43 to the trailing edge 10. The expansion of the crown pocket 43 in the radial direction 121 widens, for example in axially 34th In the area of ​​the trailing edge 10 leaves this part of the coolant the turbine blades 120, 130th

Through the second coolant inlet region 25 also flows always coolant in an internal cavity 46, which adjoined in the axial direction 34 to the nasal cavity 37th The inner cavity 46 is, for example, at least partially, especially designed very meandering so a meander-shaped portion of the coolant inside the Turbine blade 120, 130 toward the area of ​​the trailing edge 10 flows and at least partially over the length of Trailing edge 10 exiting the turbine blade 120, 130th The inner cavity 46 is of the nasal cavity 37 in the radial Direction 121 delimited by a wall 49th In the wall 49 are as many as small as possible impingement cooling openings 52 available. Through the impingement cooling openings 52, the coolant flows from the interior cavity 46 in the nasal cavity 37, meets there the inner wall surface 53 of the leading edge 13 and cools this means impingement cooling. On the inner wall surface 53 on which impinge the coolant, For example, at least one, in particular, many small bumps 55 mounted that of the turbulence Cooling medium used.

The cooling means (air and / or steam) passing through the impingement cooling openings 52 flows into the nasal cavity 37, flows if necessary together with the coolant through the first coolant inlet region 22 is flowing through the at least a connection opening 41 in the crown pocket 43 and / or through the first outlet opening 40 to the outside. The first coolant inlet region 22 may also be completely be closed so that the coolant from the inner only Cavity 46 from entering the nasal cavity 37th

A transverse web 58 separates the inner cavity 46 in the axial Direction 34 of the crown pocket 43rd Also, by at least one, in particular two or more second outlet openings 56 in the crosspiece 58 refrigerant from the meandering portion 46 in the nose bag 43 penetrate (for example, as impingement cooling).

The transverse web 58 runs in the axial direction 34 or in Figure 2 under a certain angle to the axial direction 34, since the crown pocket 43 is widened axially 34th

Due to the location and / or the thickness of the crossbar 58, the third own form of a blade vibration to higher frequencies be moved. Stresses on the fastening area near trailing edge 88 can be through bores (not shown) in the third coolant inlet region 28 easily influenced.

In the turbine blade 120, 130 is at least in the region the leading edge 13 does not use film cooling as it passes through the arrangement of nasal cavity 37, crown pocket 43 and internal cavity 46 inside according to the invention strongly is cooled improved. can in the area of ​​the trailing edge 10 film cooling are used, but need not, so that by the invention cooling air is saved and the Efficiency of the turbine is increased.

can be obtained by the second outlet openings 56 in the transverse web 58 and the crossbar 58 are further cooled.

Both in the crown pocket 43 and in the region of the trailing edge 10 are bars 61 present between the interior the two sides of the blade 406 extend. The webs 61 serve for swirling the refrigerant flowing. In the crown bag 43, the webs 61, for example, evenly distributed. In the area of ​​the leading edge 10 extending, for example, two staggered rows of ridges 61 in the radial Direction 121. This also allows for a distribution of the coolant, flowing out of a Mäanderende 79, via the Length trailing 10th The outflow 82 of the turbine blade 120, 130 in the area the trailing edge 10 is composed of the outflow 73 of the internal cavity 46 and the outflow 85 crown pocket 43rd

Through the third coolant inlet region 28 flows when a cover 91 is missing or suitable through holes having, for example, the coolant in a Fußhohlraum 76 (for example, extends to the blade platform 403, that in the blade root 400) and by a third Outlet opening 64 in the discharge area 73 of the inner Cavity 46 in the shielding edge 10th Likewise, the coolant through the third coolant inlet region 28 flows through a fourth outlet opening 67 flow out directly in the region of the trailing edge 10 again. Both in the area of ​​the outflow region 73 and in the region the trailing edge 10 of the crown pocket 43 can elongate Webs 70 be present (rather than circular ridges 61). Elongated means that they are in the radial cross-section in axial direction 34 extend.

The entire cavity of the turbine blade 120, 130 is here for example, through the nasal cavity 37, the crown pocket 43, the inner cavity 46 and the Fußhohlraum 76 formed.

Figure 3 shows an example of a gas turbine 100 in a Partial longitudinal section. The gas turbine 100 internally comprises about a rotation axis 102 rotatably mounted rotor 103, also known as the turbine rotor referred to as. Along the rotor 103 are an intake manifold 104, a compressor 105, an eg toroidal Combustion chamber 110, in particular annular combustion chamber 106, with several coaxially arranged burners 107, a turbine 108 and the exhaust manifold 109th The annular combustion chamber 106 communicates with an example annular hot gas channel 111. There, for example, make four successive turbine stages 112 form the turbine 108th Each turbine stage 112 is formed from two blade rings. In the direction of flow of a working medium 113, seen following one of the hot gas passage 111 a row of guide vanes 115 Rotor blades 120 formed series 125th

The guide vanes 130 (Fig. 2) are secured to an inner housing 138 of a stator 143, whereas the rotor blades 120 (Fig. 2) of a row 125 for example by means a turbine disk 133 are attached to the rotor 103rd At the rotor is coupled 103 A generator or a working machine (not shown).

During operation of the gas turbine 100 is from the compressor 105 104 135 air sucked through the intake and compressed. The provided at the turbine-side end of the compressor 105 compressed air is passed to the burners 107 and there mixed with a fuel. The mixture then, forming the working medium 113 in the combustion chamber 110 burned. From there, the working medium 113 flows along the hot gas duct 111 past the guide vanes 130 and the rotor blades 120. At the rotor blades 120 relaxes the Working medium 113, transferring its momentum, so that the rotor blades 120 drive the rotor 103 and the generator coupled to it Working machine.

subject to the hot working medium 113 components exposed thermal during operation of the gas turbine 100 Loads. The guide vanes 130 and rotor blades 120 of seen in the flow direction of the working medium 113, first Turbine stage 112, in addition to the annular combustion chamber 106 heat shield bricks most thermally stressed. In order to withstand the temperatures prevailing there, be this cooled by a coolant (see Figure 2). The substrates may likewise have a directional structure, ie they are monocrystalline (SX structure) or have only longitudinally directed grains (DS structure). As materials are iron, nickel or cobalt-base Superalloys. For example, superalloys are used as from EP 1204776, EP 1306454, EP 1319729, WO 99/67435 or WO 00/44949 are known; these documents are part of the Epiphany.

Similarly, the blades 120, 130 can coatings against Corrosion (MCrAlX; M is at least one element of group Iron (Fe), cobalt (Co), nickel (Ni), X stands for yttrium (Y) and / or at least one rare earth element) and heat exhibit by a thermal barrier coating. The thermal barrier coating consists, for example, ZrO₂, Y₂O₄-ZrO₂. ie unstabilized, partially stabilized or fully stabilized by yttrium oxide and / or calcium and / or magnesium oxide.

By suitable coating methods, for example electron (EB-PVD), columnar grains in the generates thermal barrier coating.

The guide vane 130 has a the inner housing 138 of Turbine 108 facing vane root (not shown here) opposite and a guide vane head Vane root. The guide vane rotor 103 facing and on a fastening ring 140 of the stator 143 set.