TURBINE BLADE AND AIRCRAFT ENGINE COMPRISING SAME

Abstract

The invention relates to a blade for use in a turbine of an aircraft engine. The blade is made of (a) a Mo-based alloy strengthened by intermetallic silicides or (b) a Ni-based single crystal superalloy. An aircraft engine and in particular, a turbofan aircraft engine including a corresponding turbine blade is also disclosed.

Claims

1. A blade for a turbine of an aircraft engine, wherein the blade is made of (a) a Mo-based alloy strengthened by intermetallic silicides or (b) a Ni-based single crystal superalloy.

2. The blade of claim 1, wherein the blade is made of (a).

3. The blade of claim 2, wherein (a) comprises molybdenum, silicon, boron and titanium as main constituents and further comprises one or both of iron and yttrium as minor alloying elements.

4. The blade of claim 2, wherein the alloy further comprises one or more of zirconium, niobium and tungsten as additional minor alloying elements.

5. The blade of claim 2, wherein the alloy is formed exclusively by molybdenum, silicon, boron, titanium, iron, yttrium, niobium, tungsten, zirconium or is formed exclusively by molybdenum, silicon, boron, titanium, iron, yttrium or is formed exclusively by molybdenum, silicon, boron, titanium, iron.

6. The blade of claim 2, wherein the alloy comprises a matrix of a molybdenum mixed crystal and one or more silicide phases.

7. The blade of claim 6, wherein the one or more silicide phases comprise (Mo,Ti).sub.5Si.sub.3 and/or (Mo,Ti).sub.5SiB.sub.2.

8. The blade of claim 1, wherein the blade is made of (b).

9. The blade of claim 8, wherein (b) comprises, in % by weight, from 3.7 to 7.0 Al, from 10 to 20 Co, from 2.1 to 7.2 Cr, from 1.1 to 3.0 Mo, from 5.7 to 9.2 Re, from 3.1 to 8.5 Ru, from 4.1 to 11.9 Ta, from 2.1 to 4.9 W, from 0 to 3.3 Ti, from 0 to 0.05 C, from 0 to 0.1 Si, from 0 to 0.05 Mn, from 0 to 0.015 P, from 0 to 0.001 S, from 0 to 0.003 B, from 0 to 0.05 Cu, from 0 to 0.15 Fe, from 0 to 0.15 Hf, from 0 to 0.015 Zr, from 0 to 0.001 Y, remainder Ni and unavoidable impurities, a weight ratio Ta:Al being from 1:1 to 2:1, and a weight ratio Co:W being from 2:1 to 5:1.

10. A turbine for an aircraft engine, wherein the turbine comprises at least one blade according to claim 1.

11. An aircraft engine, wherein the engine comprises a turbine according to claim 10.

12. An aircraft engine, wherein the engine comprises (i) a first turbine and (ii) a second turbine disposed downstream of (i) and having a plurality of turbine stages, at least a first stage of the plurality of turbine stages comprising at least one blade according to claim 1.

13. The aircraft engine of claim 12, wherein the engine is a turbofan aircraft engine.

14. The aircraft engine of claim 13, wherein the engine comprises a primary duct including a combustion chamber; the first turbine (i) disposed downstream of the combustion chamber; a compressor disposed upstream of the combustion chamber and coupled to the first turbine; and the second turbine (ii) disposed downstream of the first turbine (i) and coupled to a fan for feeding a secondary duct of the aircraft engine.

15. The aircraft engine of claim 14, wherein the blades of the first stage of (ii) are not cooled.

16. The aircraft engine of claim 14, wherein a square of a ratio of a maximum blade diameter of the fan to a maximum blade diameter of the second turbine is at least 3.5.

17. The aircraft engine of claim 16, wherein the square of the ratio of the maximum blade diameter of the fan to the maximum blade diameter of the second turbine is at least equal to the sum of one and a quotient of a bypass area ratio of an inlet area of the secondary duct to an inlet area of the primary duct divided by 3.6.

18. The aircraft engine of claim 14, wherein the second turbine has a total stage count (n.sub.St) of all turbine stages, a total blade count (N.sub.BV) of all rotor blades and stator blades of all turbine stages, a stage pressure ratio () of a pressure at an inlet to a pressure at an outlet at each turbine stage, and a total pressure ratio (p.sub.1/p.sub.2) of a pressure at an inlet of a first turbine stage to a pressure at an exit of a last turbine stage of the second turbine at a design point, and wherein the quotient (N.sub.BV/110) is less than a difference [(p.sub.1/p.sub.2)1] with the total pressure ratio being greater than 4.5; at least one stage pressure ratio is at least 1.5; and the turbine has from two to five turbine stages; and/or a quotient ((p.sub.1/p.sub.2)/n.sub.St) is greater than 1.6.

19. The aircraft engine of claim 14, wherein the second turbine has a total stage count (n.sub.St) of all turbine stages, a total blade count (N.sub.BV) of all rotor blades and stator blades of all turbine stages, a stage pressure ratio () of a pressure at an inlet to a pressure at an outlet at each turbine stage, and a total pressure ratio (p.sub.1/p.sub.2) of the pressure at an inlet of a first turbine stage to a pressure at an exit of a last turbine stage of the second turbine at a design point, and wherein a product (An.sup.2) of an exit area of the second turbine and a square of a rotational speed of the second turbine at the design point is at least 4.5.Math.10.sup.10 [in.sup.2.Math.rpm.sup.2], at least one stage pressure ratio is at least 1.5, and a blade tip velocity (u.sub.TIP) of at least one turbine stage of the second turbine at the design point is at least 400 meters per second.

20. A method of reducing or eliminating the gap between a blade tip and a seal in a gas turbine of a turbomachine at an operating temperature which is at least 100 C. below a maximum operating temperature of the turbomachine of at least 1100 C., wherein the method comprises using the blade of claim 1.

Description

BRIEF DESCRIPTION OF THE DRAWING

[0099] The only FIG. shows, in partially schematic form, a turbofan aircraft engine of a passenger jet according to an embodiment of the present invention as set forth above.

DETAILED DESCRIPTION OF EMBODIMENTS OF THE INVENTION

[0100] The particulars shown herein are by way of example and for purposes of illustrative discussion of the embodiments of the present invention only and are presented in the cause of providing what is believed to be the most useful and readily understood description of the principles and conceptual aspects of the present invention. In this regard, no attempt is made to show details of the present invention in more detail than is necessary for the fundamental understanding of the present invention, the description in combination with the drawing making apparent to those of skill in the art how the several forms of the present invention may be embodied in practice.

[0101] FIG. 1 depicts a turbofan aircraft engine of a passenger jet in accordance with an embodiment of the present invention. The engine has a primary duct C containing a combustion chamber BK. The primary duct has a first turbine or high-pressure turbine HT, which is located immediately downstream (to the right in FIG. 1) of the combustion chamber and includes a plurality of turbine stages. The high-pressure turbine is fixedly coupled to a high-pressure compressor HC of the primary duct via a hollow shaft W1, and hence such that they rotate at the same speed, the high-pressure compressor being disposed immediately upstream of the combustion chamber. As used herein, a coupling providing for rotation at the same speed is understood to mean, in particular, a non-rotatable coupling having a constant gear ratio equal to one, such as is provided, for example, by a fixed connection.

[0102] The turbofan aircraft engine has a secondary duct B, which is arranged fluidically parallel to and concentric with the primary duct. A fan F is disposed immediately upstream of the primary and secondary ducts (to the left in FIG. 1) to draw in air and feed it into the primary and secondary ducts. An additional compressor or low-pressure compressor is disposed between the fan and the high-pressure compressor.

[0103] The fan is connected through a speed reduction mechanism including a transmission G and via a low-pressure shaft W2 to a second turbine or low-pressure turbine L of the primary duct. The low-pressure turbine includes a plurality of turbine stages and is disposed downstream of the high-pressure turbine (to the right in FIG. 1). The hollow shaft W1 is concentric with the low-pressure shaft W2.

[0104] The following are specific examples of alloys (a) which can be used from making the blade of the present invention (figures in each case represent at. %), and can also comprise small amounts of further elements as unavoidable impurities:

TABLE-US-00004 Mo Si B Ti Fe Y Zr Nb W 49.5 12.5 8.5 27.5 2.0 0 0 0 0 48.5 13.5 8.5 26.5 2.0 0 1.0 0 0 51.0 10.0 8.5 27.5 2.0 0 1.0 0 0 46.5 12.5 8.5 27.5 2.0 2.0 1.0 0 0 46.5 12.5 8.5 27.5 2.0 2.0 0 1.0 0 46.5 12.5 8.5 27.5 2.0 2.0 0 0 1.0 49.3 13.5 5.5 27.5 1.2 0 0 0 1.0

[0105] The following are specific examples of alloys (b) which can be used from making the blade of the present invention (figures in each case in % by weight). The balance to 100% by weight is constituted by Ni as main component and unavoidable impurities. Additionally, one or more of C, Si, Mn, P, S, B, Cu, Fe, Hf, Zr and Y may be present in a total concentration of less than 0.7% by weight.

TABLE-US-00005 Al Co Cr Mo Re Ru Ta Ti W 5.9 11.2 4.6 1.1 6.4 5.0 7.6 0 4 5.7 11.4 5.0 1.9 6.0 3.3 5.8 1.2 3.7 5.9 11.4 5.0 2.2 6.0 3.3 6.5 0.5 3.7 5.9 11.3 5.0 2.4 6.0 3.3 7.4 0 3.7

[0106] Although the present invention has been described herein with reference to particular means, materials and embodiments, the present invention is not intended to be limited to the particulars disclosed herein; rather, the present invention extends to all functionally equivalent structures, methods and uses, such as are within the scope of the appended claims.

LIST OF REFERENCE NUMERALS

[0107] A.sub.B inlet area of the secondary duct

[0108] A.sub.C inlet area of the primary duct

[0109] A.sub.L exit area of the low-pressure turbine

[0110] B secondary duct (bypass)

[0111] BK combustion chamber

[0112] C primary duct (core)

[0113] D.sub.Ba outer diameter of the secondary duct

[0114] D.sub.Bi, inner diameter of the secondary duct

[0115] D.sub.Ca outer diameter of the primary duct

[0116] D.sub.Ci inner diameter of the primary duct

[0117] D.sub.F maximum blade diameter of the fan

[0118] D.sub.L maximum blade diameter of the low-pressure turbine

[0119] F fan

[0120] G transmission (speed reduction mechanism)

[0121] HC (high-pressure) compressor

[0122] HT first turbine or high-pressure turbine

[0123] L second turbine or low-pressure turbine

[0124] V volume

[0125] W1 hollow shaft

[0126] W2 low-pressure shaft