Turbofan aircraft engine

Patent metadata

Field	Value
Publication number	US-2016017797-A1
Application number	US-201414335107-A
Country	US
Kind code	A1
Filing date	Jul 18, 2014
Priority date	Jul 18, 2014
Publication date	Jan 21, 2016
Grant date	—

Abstract

A turbofan aircraft engine having a primary duct (C), including a combustion chamber (BK), a first turbine (HT) disposed downstream of the combustion chamber, a compressor (HC) disposed upstream of the combustion chamber and coupled (W 1 ) to the first turbine, and a second turbine (L) disposed downstream of the first turbine and coupled (W 2 ) via a speed reduction mechanism (G) to a fan (F) for feeding a secondary duct (B) of the turbofan aircraft engine. A square of a ratio of a maximum blade diameter (D F ) of the fan to a maximum blade diameter (D L ) of the second turbine is at least 3.5, in particular at least 4.

First claim

What is claimed is:

1 . A turbofan aircraft engine comprising:
a primary duct including a combustion chamber; a first turbine disposed downstream of the combustion chamber; a compressor disposed upstream of the combustion chamber and coupled to the first turbine; and a second turbine disposed downstream of the first turbine and coupled via a speed reduction mechanism to a fan for feeding a secondary duct of the turbofan aircraft engine, a square of a ratio of a maximum blade diameter of the fan to a maximum blade diameter of the second turbine being at least 3.5.

2 . The turbofan aircraft engine as recited in claim 1 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 4.

3 . The turbofan aircraft engine as recited in claim 1 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.

4 . The turbofan aircraft engine as recited in claim 3 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 3.2.

5 . The turbofan aircraft engine as recited in claim 1 wherein the square of the ratio of the maximum blade diameter of the fan to the maximum blade diameter of the second turbine is no greater than 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 2.6.

6 . The turbofan aircraft engine as recited in claim 5 wherein the square of the ratio of the maximum blade diameter of the fan to the maximum blade diameter of the second turbine is no greater than 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 2.8.

7 . The turbofan aircraft engine as recited in claim 1 wherein a sum of a product of a square of the maximum blade diameter of the fan in [m 2 ] and at least 0.1 m, a product of the maximum blade diameter of the fan in [m] and at most −0.1 m 2 , and at most 0.5 m 3 , is, in absolute value, at least equal to a volume of an outer wall of the second turbine between the entrance cross section and the exit cross section in [m 3 ] thereof.

8 . The turbofan aircraft engine as recited in claim 7 wherein a sum of a product of a square of the maximum blade diameter of the fan in [m 2 ] and 0.15 m, a product of the maximum blade diameter of fan in [m] and −0.28 m 2 , and 0.2 m 3 is, in absolute value, at least equal to a volume of an outer wall of the second turbine between the entrance cross section and the exit cross section in [m 3 ] thereof.

9 . The turbofan aircraft engine as recited in claim 1 wherein the maximum blade diameter of the fan is at least 1.2 m.

10 . The turbofan aircraft engine as recited in claim 1 wherein the second turbine has no more than 5 stages.

11 . The turbofan aircraft engine as recited in claim 10 wherein the second turbine has no more than 4 stages.

12 . The turbofan aircraft engine as recited in claim 1 wherein a product of an exit area of the second turbine in square inches and a square of a maximum allowable operating speed of the second turbine in rpms is at least 8000 m 2 /s 2 .

13 . The turbine as recited in claim 12 wherein a product of an exit area of the second turbine in square inches and a square of a maximum allowable operating speed of the second turbine in rpms is at least 9000 m 2 /s 2 .

14 . A passenger jet for at least ten passengers comprising the turbofan aircraft engine as recited in claim 1 .

15 . The passenger jet as recited in claim 13 wherein the passenger jet has a cruising altitude of at least 1200 m and/or no more than 15000 m and/or a cruising speed of at least 0.5 Ma and/or no more than 0.9 Ma.

16 . A method for designing a turbofan aircraft engine as recited in claim 1 wherein the maximum blade diameter of the fan and the maximum blade diameter of the second turbine are selected such that 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 3.3.

17 . The method as recited in 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 4.

18 . A method for manufacturing a turbofan aircraft engine as recited in claim 1 comprising: placing a first turbine downstream of the combustion chamber; placing a compressor upstream of the combustion chamber and coupling the compressor to the first turbine; placing a second turbine downstream of the first turbine and coupling the second turbine via a speed reduction mechanism to a fan for feeding a secondary duct of the turbofan aircraft engine, a square of a ratio of a maximum blade diameter of the fan to a maximum blade diameter of the second turbine being selected to be at least 3.5.

Assignees

• MTU Aero Engines AG

Inventors

• Rued Klaus Peter
• Humhauser Werner
• Klingels Hermann
• Stanka Rudolf
• Heinrich Eckart
• Hackenberg Hans-Peter
• Weber Stefan
• Rieger Claus
• Steinhardt Erich
• Gier Jochen
• Feldmann Manfred
• Huebner Norbert
• Maar Karl

Classifications

• F02C3/107

  with two or more rotors connected by power transmission · CPC title

• F02K3/025 Primary

  the by-pass flow being at least partly used to create an independent thrust component · CPC title

• F02C3/10 Primary

  with another turbine driving an output shaft but not driving the compressor · CPC title