Cooled turbine blade

The invention relates to a coolable turbine blade according to Claim first

Internally cooled turbine blades, which a meandering Interior (<patcit id="pcit0001" dnum="EP1022434A2"><text>EP 1022434 A2</text></patcit>) As well as impingement cooling having, are known.

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

The <patcit id="pcit0002" dnum="DE3234906A1"><text>DE 32 34 906 A1</text></patcit>. <patcit id="pcit0003" dnum="US5857837A"><text>US-PS 5,857,837</text></patcit>. <patcit id="pcit0004" dnum="US5873695A"><text>5,873,695</text></patcit>. <patcit id="pcit0005" dnum="US5902093A"><text>5,902,093</text></patcit>. <patcit id="pcit0006" dnum="US5462405A"><text>5,462,405</text></patcit>. <patcit id="pcit0007" dnum="US6139269A"><text>6,139,269</text></patcit> show a turbine blade whose leading edge is cooled by impingement cooling.

The <patcit id="pcit0008" dnum="DE19963716A1"><text>DE 199 63 716 A1</text></patcit>. <patcit id="pcit0009" dnum="US4474532A"><text>US-PS 4,474,532</text></patcit>. <patcit id="pcit0010" dnum="US4753575A"><text>4,753,575</text></patcit> and the <patcit id="pcit0011" dnum="US4767268A"><text>4,767,268</text></patcit> show a turbine blade having a cavity in the region of the blade tip.

The <patcit id="pcit0012" dnum="US6431832B"><text>US-PS 6,431,832</text></patcit> shows a turbine blade, which has a cavity in the region of the blade tip and a cavity in the region of its leading edge.

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

It is therefore an object of the invention to provide a turbine blade in which the cooling of the leading edge is improved.

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

In the subclaims, further advantageous embodiments of the turbine blade are listed. The measures listed in the dependent claims may be combined in an advantageous manner with each other.

Show it:<dl id="dl0001"><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 turbine blade according to the invention formed is used.</dd></dl>

<figref idrefs="f0001">figure 1</figref> shows a perspective view a blade 120, 130 (<figref idrefs="f0003">Fig. 3</figref>) Extending along a longitudinal axis 121 (radial direction). The blade 120, 130 has along the longitudinal axis 121, a securing region 400, an adjoining blade platform 403, an airfoil portion 406 and a blade tip 19. The airfoil section 406 is flown at a leading edge 13 of 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 (<figref idrefs="f0003">Fig. 3</figref>) Is used. The blade root 183 is configured as a hammerhead. Other configurations, such as a fir-tree or dovetail root are possible. In conventional blades 120, 130 are used in all regions 400, 403, 406 of the blade 120, 130 used solid metallic materials. The blade 120, 130 may this case be produced by a casting process, by a forging process, by a milling process or combinations.

The <figref idrefs="f0002">figure 2</figref> shows such a turbine blade or vane 120, 130 in longitudinal section. In the <figref idrefs="f0002">figure 2</figref> is a turbine blade or vane 120 130 a gas (<figref idrefs="f0003">Fig. 3</figref>But also aircraft turbine) or steam turbine ready.

The turbine blade 120, 130 is at least partially hollow and is cooled inside. The turbine blade 120, 130 has the blade root 400 to at least one first coolant inlet region 22, a second coolant inlet region 25 and a third coolant inlet region 28th The coolant inlet areas 22, 25, 28 or may be a constriction 31 (here, for example, the coolant inlet region 22) by means of a coolant flow can be set into the interior either closed (for example, 28 here). The constriction 31 can be controlled, ie increased during operation or reduced. In this embodiment, the third coolant inlet region 28 is completely closed and narrowed the first coolant inlet region 25th The coolant flows through the first and second coolant inlet region 22, 25 into the interior of the turbine blade 120, the 130th The turbine blade 120, 130 is in the axial direction 34 flows around a hot medium. The hot medium meets in the axial direction 34 first on the leading edge 13 (shovel nose) and then flows past the blade airfoil 406 passing to the trailing edge 10th The refrigerant in a nasal cavity 37, which follows in the axial direction 34 directly behind the leading edge 13, penetrates through the first coolant inlet region 22, 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, at least a first exit opening 40 is, for example, but not necessarily, provided for the coolant, so that there is a film cooling of the blade tip can be made. In any case, the refrigerant flows through at least one connecting opening 41 in a crown pocket 43 (cavity in the region of the blade tip 19) and cools there the blade tip 19 along 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 axially 34th In the area of ​​the trailing edge 10 of this portion of the refrigerant leaves the turbine blade 120, 130th

The second coolant inlet region 25 also always flows coolant in an internal cavity 46, which adjoins in the axial direction 34 to the nasal cavity 37th The inner cavity 46 is, for example, at least partially, in particular formed very meandering so that a portion of the coolant meandering inside the turbine blade 120, 130 back flows to the region of the trailing edge 10 and at least partially over the length of the trailing edge 10 of the turbine blade 120, 130 exits. The inner cavity 46 is bounded by the nasal cavity 37 in the radial direction 121 by a wall 49th In the wall 49 as many as small as possible impingement cooling apertures 52 are present. Through the impingement cooling openings 52, the coolant from the internal cavity 46 flows into the nose cavity 37, where it meets the inner wall surface 53 of the leading edge 13 and cools it by means of impingement cooling. On the inner wall surface 53 on which impinge the coolant is, for example, at least one, in particular, many small projections 55 is provided, which serve to turbulence of the cooling medium.

The cooling means (air and / or steam) that flows through the impingement cooling openings 52 in the nasal cavity 37, flows, optionally together with the refrigerant which has flowed through the first coolant inlet region 22, through the at least one connecting opening 41 in the crown pocket 43 and / or through the first outlet opening 40 to the outside. The first coolant inlet region 22 can also be completely closed so that the coolant enters only from the internal cavity 46 in 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 58 coolant from the meandering portion 46 penetrate the crossbar in the nose bag 43 (for example, as impingement cooling).

The crosspiece 58 is as in <figref idrefs="f0002">figure 2</figref> at a certain angle to the axial direction 34, as the crown pocket 43 is widened axially 34th

By the location and / or the thickness of the transverse web 58, the third natural mode of a blade vibration can be shifted to higher frequencies. Stresses on the fastening area near the trailing edge 88 can be (not shown) through holes in the third coolant inlet region 28 easily influenced.

In the turbine blade 120, 130 no film cooling is used at least in the area of ​​the leading edge 13, since it is 43 and inner cavity 46 is cooled according to the invention greatly improved by the arrangement of nasal cavity 37, crown pocket inside. In the area of ​​the trailing edge 10 may have a film-cooling are used, but need not, so that by the invention cooling air is saved, and the efficiency of the turbine is increased.

Through the second outlet openings 56 in the transverse web 58 and the transverse web 58 may be additionally cooled.

Both in the crown pocket 43 and in the region of the trailing edge 10 bridges 61 are present, which extend in the interior between the two sides of the airfoil 406th The webs 61 serve to swirl the flowing coolant. In the crown bag 43, the webs 61, for example, uniformly 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 distribution of the refrigerant flowing out of a Mäanderende 79, over the length of trailing 10th The outflow 82 of the turbine blade 120, 130 in the area of ​​the trailing edge 10 is composed of the outflow 73 of the inner cavity 46 and the outflow 85 of the crown pocket 43rd

By the third coolant inlet region 28 flows when a cover 91 is missing or in appropriate through holes, for example, the coolant in a Fußhohlraum 76 (for example, extends to the blade platform 403, ie in the blade 400) and through a third port 64 in the outflow region 73 of the inner cavity 46 in the shielding edge 10th Likewise, the coolant flowing through the third coolant inlet region 28, through a fourth outlet opening 67 directly in the area of ​​the trailing edge 10 again flow out. Both in the area of ​​the outflow region 73 and in the region of the trailing edge 10 of the crown pocket 43 elongated ridges may 70 (instead of circular ridges 61) exist. Elongated means that they extend in the axial direction 34 in radial cross section.

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

The <figref idrefs="f0003">figure 3</figref> exemplifies a gas turbine 100 in a longitudinal partial section. The gas turbine 100 internally comprises a about a rotation axis 102 mounted rotor 103, which is also referred to as the turbine rotor. 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 a plurality of coaxially arranged burners 107, a turbine 108 and the exhaust manifold 109th The annular combustion chamber 106 communicates with an eg annular hot gas channel 111. There, for example, four successive turbine stages 112 form the turbine 108th Each turbine stage 112 is formed from two blade rings. Viewed in the flow direction of a working medium 113, followed by 120 formed from rotor blades 125 series in the hot gas passage 111 a row of guide vanes 115th

The guide vanes 130 (<figref idrefs="f0002">FIG. 2</figref>) Are secured to an inner housing 138 of a stator 143 while the rotor blades 120 (<figref idrefs="f0002">FIG. 2</figref>) Are a row 125 are attached, for example by means of a turbine disk 133 rotor 103rd Coupled to the rotor 103 is a generator or a work engine (not shown).

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

To the hot working medium 113, the components exposed 100 thermal stresses during operation of the gas turbine. The guide vanes 130 and rotor blades 120 of the flow direction of the working medium 113, the first turbine stage 112, in addition to the annular combustion chamber 106 heat shield bricks which the highest thermal stresses. In order to withstand the temperatures prevailing there, they are cooled by a coolant (see <figref idrefs="f0002">figure 2</figref>). 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-based superalloys. For example, superalloys are used as from the <patcit id="pcit0013" dnum="EP1204776A"><text>EP 1204776</text></patcit>. <patcit id="pcit0014" dnum="EP1306454A"><text>EP 1306454</text></patcit>. <patcit id="pcit0015" dnum="EP1319729A"><text>EP 1319729</text></patcit>. <patcit id="pcit0016" dnum="WO9967435A"><text>WO 99/67435</text></patcit> or <patcit id="pcit0017" dnum="WO0044949A"><text>WO 00/44949</text></patcit> are known; these documents form part of the disclosure.

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

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

The guide vane 130 has a the inner housing 138 of the turbine 108 facing vane root (not shown here) and an opposing guide vane guide vane. The guide vane head faces the rotor 103 and is fixed to a securing ring 140 of the stator 143.