Turbine blade shroud for gas turbine engine with power turbine and method of manufacturing same

BACKGROUND

0001This disclosure relates to turbomachinery, and more particularly, the disclosure relates to a shrouded turbine blade and method for making the same.

0002Gas turbine engines include a compressor that compresses air, a combustor that ignites the compressed air and a turbine across which the compressed air is expanded. The expansion of the combustion products drives the turbine to rotate, which in turn drives rotation of the compressor.

0003Gas turbine engines for applications such as helicopters incorporate a power turbine (PT) that is not mechanically coupled to the compressors in the gas generator portion of the gas turbine engine. The power turbine is rotationally driven by expanding gases from the gas generator portion to transmit power to a turboshaft. The turboshaft rotationally drives the helicopter propeller, typically at a constant speed, through a gearbox.

0004In order to increase efficiency, a clearance between the tips of the blades in the compressor, turbine and power turbine across the outer diameter of the flowpath is kept sufficiently small. This ensures that a minimum amount of air passes between the tips and the outer diameter. Turbine blades may incorporate a shroud to provide damping. Knife edges may be provided on the shroud to seal with respect to a blade outer air seal (BOAS) to maintain tight clearances.

0005The shroud adds mass to the end of the blade, which increases stress. One or more pockets may be provided on the shroud to reduce weight. For small diameter blades, traditional manufacturing approaches may prevent the pocket from being formed in the shroud.

0006A prior art blade for a gas turbine engine having the features of the preamble to claim 1 is disclosed in <patcit id="pcit0001" dnum="EP1013884A2"><text>EP 1,013,884 A2</text></patcit> or <patcit id="pcit0002" dnum="JPH0828303A"><text>JP H08 28303 A</text></patcit>.

SUMMARY

0007From one aspect, the present invention provides a blade for a gas turbine engine in accordance with claim 1.

0008In an embodiment of the above, a root supports a platform. The airfoil extends radially from the platform to the shroud. The internal passage extends from the root to the shroud.

0009In a further embodiment of any of the above, the airfoil extends from the platform to the shroud a span that is less than 66.0 mm (2.6 inches). In such an embodiment, the span extends in a radial direction from the platform to the shroud.

0010In a further embodiment of any of the above, the shroud includes axially spaced apart axial faces and circumferentially spaced apart circumferential faces. The axial and circumferential faces define the shroud perimeter.

0011In a further embodiment of any of the above, the knife edges extend to the circumferential faces.

0012In a further embodiment of any of the above, the shroud perimeter provides a first area. The knife edges include fillets. The second area is the area between and includes the fillets facing the pocket. The second area is in a range of 40% to 60% of the first area.

0013In a further embodiment of any of the above, the pocket has a third area defined by the perimeter edge. The third area is in a range of 80% to 90% of the second area.

0014In a further embodiment of any of the above, the outlet has a fourth area. The fourth area is in a range of 5% to 25% of the third area.

0015In a further embodiment of any of the above, the shroud has first and second thicknesses spaced in a direction perpendicular to the peak on at least one of the first and second surfaces. The first and second thicknesses extend in a radial direction. The first thickness is provided at a tangent point to a fillet that joins the shroud and the airfoil. The second thickness is provided at the perimeter edge. The ratio of the first thickness to the second thickness in a range of 1.1 to 2.1.

0016In a further embodiment of any of the above, the depth of the pocket corresponds to the location of the second thickness at the perimeter edge that is 1.52 mm (0.060 inch) or less.

0017From another aspect, the present invention provides a gas turbine engine in accordance with claim 11.

0018In an embodiment of the above, the shroud perimeter provides a first area. The knife edges include fillets. The second area is the area between the fillets facing the pocket. The pocket has a third area defined by the perimeter edge. The third area is in a range of 80% to 90% of the second area.

0019In a further embodiment of any of the above, the shroud has first and second thicknesses spaced in a direction perpendicular to the peak on at least one of the first and second surfaces. The first and second thicknesses extend in a radial direction. The first thickness are provided at a tangent point to a fillet that joins the shroud and the airfoil. The second thickness is provided at the perimeter edge. The ratio of the first thickness to the second thickness is in a range of 1.1 to 2.1.

0020In a further embodiment of any of the above, the depth of the pocket corresponds to the location of the second thickness at the perimeter edge that is 1.52 mm (0.060 inch) or less.

0021In a further embodiment of any of the above, the gas turbine engine includes a root that supports a platform. The airfoil extends radially from the platform to the shroud. The internal passage extends from the root to the shroud. The airfoil extends from the platform to the shroud a span. The span is less than 66.0 mm (2.6 inches).

0022In a further embodiment of any of the above, the gas generator portion includes a compressor section. The power turbine is mechanically disconnected from the compressor section.

0023From yet another aspect, the present invention provides a method of manufacturing a blade for a gas turbine engine in accordance with claim 13.

0024In an embodiment of the above, the pocket is formed by a portion of the internal passage core.

0025In a further embodiment of any of the above, the pocket forming step includes electro-discharge that machines the pocket.

BRIEF DESCRIPTION OF THE DRAWINGS

0026The disclosure can be further understood by reference to the following detailed description when considered in connection with the accompanying drawings wherein: <ul id="ul0001" list-style="none" compact="compact"><li><figref idref="f0001">Figure 1</figref> is a schematic view of a gas turbine engine for use in a helicopter.</li><li><figref idref="f0002">Figure 2</figref> is a schematic view of a power turbine.</li><li><figref idref="f0002">Figure 3</figref> is a top elevational view of a turbine blade with a tip shroud.</li><li><figref idref="f0003">Figure 4</figref> is an enlarged cross-sectional view of the turbine blade taken along line 4-4 in <figref idref="f0002">Figure 3</figref>.</li><li><figref idref="f0003">Figure 5</figref> is a schematic view of the turbine blade with an example casting core.</li><li><figref idref="f0003">Figure 6</figref> is a flow chart depicting an example manufacturing method for the disclosed turbine blade.</li></ul>

DETAILED DESCRIPTION

0027<figref idref="f0001">Figure 1</figref> schematically illustrates a gas turbine engine 20. In this example, the engine 20 is a turboshaft engine, such as for a helicopter. The engine 20 includes an inlet duct 22, a compressor section 24, a combustor section 26, and a turbine section 28.

0028The compressor section 24 is an axial compressor and includes a plurality of circumferentially-spaced blades. Similarly, the turbine section 28 includes circumferentially-spaced turbine blades. The compressor section 24 and the turbine section 28 are mounted on a main shaft 29 for rotation about an engine central longitudinal axis A relative to an engine static structure 32 via several bearing systems (not shown).

0029During operation, the compressor section 24 draws air through the inlet duct 22. Although gas turbine engines ingest some amount of dust, such engines are typically not designed for highly dusty environments. Engines such as the engine 20 are subject to operating in highly dusty environments during takeoff and landing. In this example, the inlet duct 22 opens radially relative to the central longitudinal axis A. The compressor section 24 compresses the air, and the compressed air is then mixed with fuel and burned in the combustor section 26 to form a high pressure, hot gas stream. The hot gas stream is expanded in the turbine section 28, which may include first and second turbine 42, 44.

0030The first turbine 42 rotationally drives the compressor section 24 via a main shaft 29. Together these components provide a gas generator portion of the engine 20.

0031The second turbine 44, which is a power turbine in the example embodiment, rotationally drives a power shaft 30, gearbox 36, and output shaft 34. Although fluidly coupled to the gas generator portion, the power turbine 44 is mechanically disconnected from the gas generator portion. That is, the main shaft 29 and power shaft 30 are not connected to one another such that the shafts 29, 30 rotate separately and at different speeds. Moreover, there are no compressors mounted to the power shaft 30. The power turbine 44 may include a single or multiple stages of blades and vanes. The output shaft 34 rotationally drives the helicopter rotor blades 39 used to generate lift for the helicopter. The hot gas stream is expelled through an exhaust 38.

0032The engine 20 also includes a seal system in the turbine section 28 around the blades. Such a seal system may be referred to as a blade outer air seal (BOAS). The seal system serves to provide a minimum clearance around the tips of the blades, to limit the amount of air that escapes around the tips.

0033The power turbine 44 is shown in more detail in <figref idref="f0002">Figure 2</figref>. The power turbine 44 includes stages of stator vanes 48 axially spaced apart from one another and supported with respect to the turbine case structure 46, which is part of the engine static structure 32. Stages of rotor blades 50 are axially interspersed between the stages of stator vanes 48.

0034Axially spaced apart arrays of blades 50 are supported on a rotor 40 connected to the power shaft 30. BOAS 56 are supported by the turbine case structure 46 to provide a seal with respect to an end of the blade. In the example shown, the end of the blade includes a shroud 52 supporting radially extending knife edges 54, which cooperate with the BOAS 56 to provide a seal.

0035Referring to <figref idref="f0002 f0003">Figures 3-5</figref>, each blade 50 includes a root 62 supporting a platform 60. An airfoil 58 has a span extending in a radial direction R from the platform 60 to the shroud 52. The span is less than 66.0 mm (2.6 inches). For blades having such a short span, it is difficult to reliably manufacture complex geometries on the surface of the blade.

0036An internal passage 70 (<figref idref="f0003">Fig. 5</figref>) extends from the root 62 to the shroud 52 in the example shown. The internal passage 70 may be supplied with cooling fluid from, for example, the compressor section 24. Alternatively, the internal passage 70 may not be supplied with any cooling fluid and may be incorporated into the blade 50 to further reduce the overall weight of the blade 50.

0037Referring to <figref idref="f0002">Figure 3</figref>, the shroud 52 includes a shroud perimeter 64 defined by axially spaced apart faces 66 and circumferentially spaced apart faces 68, which are spaced apart from one another in a circumferential direction C. Circumferential faces of adjacent blades typically do not contact one another. The knife edges 54 extend to and connect the circumferential faces 68. The shroud perimeter 64 provides a first area A1.

0038The knife edges 54 include fillets 84. A second area A2 is the area between and including the fillets 84 facing a pocket 72 provided in the shroud 52 between the knife edges 54. The second area A2 also includes the area of the pocket 72 itself. The second area A2 is in a range of 40% to 60% of the first area A1.

0039The pocket 72 is circumscribed by a perimeter edge 74. A perimeter edge defines a third area A3, which is in a range of 80% to 90% of the second area A2. An outlet 76 fluidly connects the internal passage 70 and the pocket 72. The outlet 76 has a fourth area A4, and the fourth area A4 is in a range of 5% to 25% of the third area A3.

0040The pocket 72 includes first and second surfaces 78, 80, which are generally planar, joined at a peak 82. This configuration provides tapered walls that become thinner in opposing directions distant from peak 82. Fillets 53 join the airfoil 58 to the shroud 52. The thicker wall section near the airfoil 58 provides strength and resists bending moments, or curling stresses, under centrifugal loading. In accordance with the claims, the outlet 76 is arranged on at least one side of peak 82 (although on only one side in the illustrated example). First and second surfaces 78, 80 form an angle, for example, in a range of 4° to 8°, relative to an untapered equivalent surface.

0041The shroud 52 has first and second thicknesses T1, T2 spaced in a direction perpendicular to peak 82 on at least one of the first and second surfaces 78, 80. The first and second thicknesses T1, T2 extend in the radial direction R. The first thickness T1 is provided at a tangent point to the fillet 53. The second thickness T2 is provided at the perimeter edge 74. The ratio of the first thickness T1 to the second thickness T2 is in a range of 1.1 to 2.1. In one example, a depth D of the pocket 72 corresponding to the location of the second thickness T2 at the perimeter edge is 1.52 mm (0.060 inch) or less.

0042A method of manufacturing the blade 50 is shown generally at 94 in <figref idref="f0003">Figure 6</figref>. The method 94 includes providing an internal passage core 88 (<figref idref="f0003">Figure 5</figref>), as indicated at block 96. The blade 50 is cast about the internal passage core 88, as indicated at block 98. The pocket 72 is formed in the shroud 52, as indicated at block 100. In one example, the pocket 72 is formed by a second portion 92 of the core 88, for example. In such an example, the core 88 includes a first core portion 90 arranged within the airfoil 58 and radially beneath the shroud 52 to form the internal passage 70. The second core portion 92 may be a unitary, continuous structure with the first core portion 90. In another example, the pocket 72 is formed by electro-discharge machining (EDM). The outlet 76 connecting the pocket 72 to the internal passage 70 may also be formed by EDM.