COMPRESSOR AREA SPLITS FOR GEARED TURBOFAN
A gas turbine engine has an upstream compressor having an upstream entrance area leading into a first vane upstream of the upstream compressor. A downstream compressor has an exit area at a leading edge of an exit vane for the downstream compressor. The entrance area divided by the exit area is greater than or equal to 13.8 and less than or equal to 15.3.
1 . A gas turbine engine comprising:
an upstream compressor having an upstream entrance area leading into a first vane upstream of the upstream compressor, and a downstream compressor having an exit area at a leading edge of an exit vane for said downstream compressor, and wherein the following ratio applies:
said entrance area divided by said exit area being greater than or equal to 13.8 and less than or equal to 15.3.
2 . The gas turbine engine as set forth in claim 1 , wherein a ratio of a downstream entrance area defined at a first vane upstream of said downstream compressor rotor to said upstream entrance area being less than or equal to 0.500 and greater than or equal to 0.245.
3 . The gas turbine engine as set forth in claim 2 , wherein a ratio of said exit area to said downstream entrance area being greater than or equal to 0.145 and less than or equal to 0.270.
4 . The gas turbine engine as set forth in claim 3 , wherein a higher pressure turbine for driving said downstream compressor rotor has two stages.
5 . The gas turbine engine as set forth in claim 4 , wherein said downstream compressor rotor having more stages than said high pressure turbine rotor.
6 . The gas turbine engine as set forth in claim 3 , wherein a fan rotor having a fan hub of a first diameter at a leading edge of a fan blade, and said fan blade having a tip diameter at said leading edge, with a ratio of said fan hub diameter to said fan tip diameter being less than or equal to 0.40.
7 . The gas turbine engine as set forth in claim 6 , wherein said engine is utilized on a short range aircraft.
8 . The gas turbine engine as set forth in claim 7 , wherein said short range aircraft having a single aisle between passenger section areas.
9 . The gas turbine engine as set forth in claim 6 , wherein said gas turbine engine is utilized on a long range aircraft.
10 . The gas turbine engine as set forth in claim 9 , wherein said long range aircraft having at least two aisles between passenger seating areas.
11 . The gas turbine engine as set forth in claim 2 , wherein said first vane is a variable vane.
12 . The gas turbine engine as set forth in claim 1 , wherein a ratio of said exit area to a downstream entrance area, at a first vane upstream of said downstream compressor rotor being greater than or equal to 0.145 and less than or equal to 0.270.
13 . The gas turbine engine as set forth in claim 1 , wherein a higher pressure turbine for driving said downstream compressor rotor has two stages.
14 . The gas turbine engine as set forth in claim 13 , wherein said downstream compressor rotor having more stages than said high pressure turbine rotor.
15 . The gas turbine engine as set forth in claim 14 , wherein said downstream compressor rotor having eight or nine stages.
16 . The gas turbine engine as set forth in claim 1 , wherein a fan rotor having a fan hub of a first diameter at a leading edge of a fan blade, and said fan blade having a tip diameter at said leading edge, with a ratio of said fan hub diameter to said fan tip diameter being less than or equal to 0.40.
17 . The gas turbine engine as set forth in claim 1 , wherein said engine is utilized on a short range aircraft.
18 . The gas turbine engine as set forth in claim 17 , wherein said short range aircraft having a single aisle between passenger section areas.
19 . The gas turbine engine as set forth in claim 1 , wherein said gas turbine engine is utilized on a long range aircraft.
20 . The gas turbine engine as set forth in claim 19 , wherein said long range aircraft having at least two aisles between passenger seating areas.