Gas turbine engine with acoustic spacing of the fan blades and outlet guide vanes
A gas turbine engine comprises a fan, a core turbine engine coupled to the fan, a fan case housing the fan and the core turbine engine, a plurality of outlet guide vanes extending between the core turbine engine and the fan case, and an acoustic spacing. The fan comprises a plurality of fan blades that define a fan diameter and a BEAL. The fan case comprises an inlet and an inlet length between the inlet and the fan. The acoustic spacing comprises a distance between the fan and the plurality of outlet guide vanes, and in combination with the BEAL determines an acoustic spacing ratio of the gas turbine engine. The combination of acoustic spacing and corrected specific thrust provide enhanced propulsive efficiency.
1. A gas turbine engine comprising:
a core engine comprising a compressor section, a combustion section, and a turbine section arranged in serial flow order, the turbine section including a low pressure turbine and the compressor section having a high pressure compressor defining a high pressure compressor exit area (A HPCExit ) in square inches;
a gearbox assembly coupled to the low pressure turbine;
a fan coupled to the gearbox assembly and having a fan diameter and a plurality of fan blades with a blade solidity that is greater than or equal to 0.8 and less than or equal to 2.0;
a blade effective acoustic length (BEAL) defined as:
BEAL
=
2
c
2
S
(
1
-
rr
)
N
b
cos
(
γ
)
wherein c is a chord length of a fan blade of the plurality of fan blades, S is a span of the fan blade, rr is a radius ratio of the fan, γ is a stagger angle of the fan blade, and N b is the number of the plurality of fan blades;
a nacelle that includes a fan case that surrounds the fan, the fan case comprising an inlet disposed forward of the fan and an inlet length, wherein the inlet length is an axial distance between a leading edge of one of the plurality of fan blades and the inlet, as measured at a 75% span position of the fan blade;
a plurality of outlet guide vanes disposed aft of the fan and extending radially between the core engine and the fan case;
an acoustic spacing from the fan blade trailing edge to an outlet guide vane leading edge;
an acoustic spacing ratio (ASR) defined as:
ASR
=
1
(
Nv
Nb
)
·
As
BEAL
wherein As is the acoustic spacing and Nv is the number of the plurality of outlet guide vanes; and
an inlet-to-nacelle (ITN) ratio defined as a ratio of the inlet length to a maximum diameter of the nacelle,
wherein the ASR of the gas turbine engine is 1.5 to 16.0, and the ITN ratio is 0.23 to 0.35, and
wherein the gas turbine engine defines a redline exhaust gas temperature (EGT) in degrees Celsius, a total sea level static thrust output (Fn Total ) in pounds, and a corrected specific thrust,
wherein the corrected specific thrust is greater than or equal to 42 and less than or equal to 90, the corrected specific thrust determined as follows: Fn Total ×EGT/(A HPCExit 2 ×1000).
2. The gas turbine engine of claim 1 , further comprising a disk-to-blade diametric (DBD) ratio defined as a ratio of a disk spacing length to the fan diameter, the disk spacing length being a distance between a forwardmost end of a fan disk and an intersection with the inlet taken along an engine centerline,
wherein the DBD ratio of the gas turbine engine is 0.09 to 0.59.
3. The gas turbine engine of claim 2 , wherein the DBD ratio of the gas turbine engine is 0.15 to 0.35.
4. The gas turbine engine of claim 2 , wherein the DBD ratio of the gas turbine engine is 0.19 to 0.27.
5. The gas turbine engine of claim 1 , wherein the EGT is greater than 1000 degrees Celsius and less than 1300 degrees Celsius.
6. The gas turbine engine of claim 1 , wherein the EGT is greater than 1100 degree Celsius and less than 1250 degrees Celsius.
7. The gas turbine engine of claim 1 , wherein the EGT is greater than 1150 degree Celsius and less than 1250 degrees Celsius.
8. The gas turbine engine of claim 1 , wherein the EGT is greater than 1000 degree Celsius and less than 1300 degrees Celsius, and wherein the corrected specific thrust is greater than or equal to 45.
9. The gas turbine engine of claim 1 , wherein the EGT is greater than 1000 degree Celsius and less than 1300 degrees Celsius, and wherein the corrected specific thrust is greater than or equal to 50.
10. The gas turbine engine of claim 1 , wherein the turbine section comprises a high pressure turbine having a first stage of high pressure turbine rotor blades, and wherein the gas turbine engine further comprises:
a cooled cooling air system in fluid communication with the first stage of high pressure turbine rotor blades.
11. The gas turbine engine of claim 1 , wherein the gas turbine engine defines a bypass passage, and wherein the gas turbine engine defines a third stream extending from a location downstream of a secondary fan to the bypass passage.
12. The gas turbine engine of claim 1 , further comprising a disk-to-inlet length (DIL) ratio defined as a ratio of a disk spacing length to the inlet length, the disk spacing length being a distance between a forwardmost end of a fan disk and an intersection with the inlet taken along an engine centerline,
wherein the DIL ratio of the gas turbine engine is 0.30 to 0.80.
13. The gas turbine engine of claim 12 , wherein the DIL ratio of the gas turbine engine is 0.30 to 0.70.
14. The gas turbine engine of claim 12 , wherein the DIL ratio of the gas turbine engine is 0.49 to 0.65.
15. The gas turbine engine of claim 1 , further comprising a fan pressure ratio from 1.25 to 1.45.
16. The gas turbine engine of claim 1 , wherein the ASR is 4.0 to 14.0.
17. The gas turbine engine of claim 1 , wherein the ASR is 6.6 to 13.5.
18. The gas turbine engine of claim 1 , wherein the fan case further comprises an acoustic treatment disposed on an interior surface of the fan case.
19. The gas turbine engine of claim 18 , wherein a length of the acoustic treatment is 50% to 90% of the inlet length.
20. The gas turbine engine of claim 1 , wherein the low pressure turbine comprises at least three low pressure turbine stages.