Gas turbine engine with improved heat management
An aircraft gas turbine engine includes: an engine core including a compressor, combustor, turbine, and a core shaft connecting the turbine to the compressor; a fan including fan blades and arranged upstream of the engine core; turbomachinery bearings; a power gearbox to drive the fan at a lower rotation speed than the turbine; and a heat management system providing lubrication and cooling to the power gearbox and turbomachinery bearings, and including a pipe assembly adapted to provide a lubricant flow to the power gearbox and turbomachinery bearings to remove the heat generated by the power gearbox and turbomachinery bearings, at least one air-lubricant heat exchanger to dissipate a first amount of heat to a first heat sink, and at least one fuel-lubricant heat exchanger to dissipate a second amount of heat to a second heat sink, wherein the first heat sink is air and the second heat sink is fuel.
1. A gas turbine engine for an aircraft comprising:
an engine core comprising a compressor, a combustor, a turbine, and a core shaft connecting the turbine to the compressor;
a fan comprising a plurality of fan blades and arranged upstream of the engine core;
turbomachinery bearings;
a power gearbox adapted to drive the fan at a lower rotation speed than the turbine; and
a heat management system configured to provide lubrication and cooling to the power gearbox and turbomachinery bearings, and comprising a pipe assembly, which includes first lubricant circuit and a second lubricant circuit separate from the first lubricant circuit, adapted to provide a lubricant flow to the power gearbox and turbomachinery bearings, at least one first air-lubricant heat exchanger arranged in the first lubricant circuit and configured to dissipate a first amount of heat to a first heat sink, at least one fuel-lubricant heat exchanger arranged in the second lubricant circuit and configured to dissipate a second amount of heat to a second heat sink, and at least one second air-lubricant heat exchanger arranged in the second lubricant circuit and configured to dissipate heat to the first heat sink,
wherein the first heat sink is air and the second heat sink is fuel,
wherein the heat management system is configured to control flow through the at least one first air-lubricant heat exchanger and at least one fuel-lubricant heat exchanger to provide the first amount of heat and the second amount of heat at an environment temperature of ISA +40° C. such that:
a first proportion of the heat generated by the power gearbox and the turbomachinery bearings and dissipated to air defined as
(
first
amount
of
heat
111
first
amount
of
heat
111
+
se
cond
amount
of
heat
112
)
85
%
MTO
at 85% of a core shaft maximum take-off speed is from 0.55 to 0.70; and
a second proportion of heat generated by the power gearbox and the turbomachinery bearings and dissipated to air defined as
(
first
amount
of
heat
111
first
amount
of
heat
111
+
se
cond
amount
of
heat
112
)
65
%
MTO
at 65% of the core shaft maximum take-off (MTO) speed is from 0.85 to 1.
2. The gas turbine engine of claim 1 , wherein the first proportion of the heat generated by the power gearbox and the turbomachinery bearings and dissipated to air at 85% of the core shaft maximum take-off speed at the environment temperature of ISA +40° C. is from 0.60 to 0.70.
3. The gas turbine engine of claim 1 , wherein the second proportion of heat generated by the power gearbox and the turbomachinery bearings and dissipated to air at 65% of the core shaft maximum take-off (MTO) speed at the environment temperature of ISA +40° C. is in from 0.90 to 1.
4. The gas turbine engine of claim 1 , wherein the heat management system is configured to provide at the environment temperature of ISA +10° C. the first amount of heat and the second amount of heat such that the first proportion of the heat generated by the power gearbox and the turbomachinery bearings and dissipated to air at 85% of the core shaft maximum take-off (MTO) speed is from 0.40 to 0.60.
5. The gas turbine engine of claim 1 , wherein the heat management system is configured to provide at the environment temperature of ISA +10° C. the first amount of heat and the second amount of heat such that the second proportion of heat generated by the power gearbox and the turbomachinery bearings and dissipated to air at 65% of the core shaft maximum take-off (MTO) speed is from 0.80 to 0.92.
6. The gas turbine engine of claim 1 , wherein the heat management system is configured to provide the first amount of heat and the second amount of heat such that a ratio of the first proportion of the heat generated by the power gearbox and the turbomachinery bearings and dissipated to air at 85% of the core shaft maximum take-off speed at the environment temperature of ISA +40° C. to the first proportion of the heat generated by the power gearbox and the turbomachinery bearings and dissipated to air at 85% of the core shaft maximum take-off speed at the environment temperature of ISA +10° C. is from 1.20 to 1.42.
7. The gas turbine engine of claim 1 , wherein the heat management system is configured to provide the first amount of heat and the second amount of heat such that a ratio of the first proportion of the heat generated by the power gearbox and the turbomachinery bearings and dissipated to air at 85% of the core shaft maximum take-off speed at the environment temperature of ISA +40° C. to the first proportion of the heat generated by the power gearbox and the turbomachinery bearings and dissipated to air at 85% of the core shaft maximum take-off speed at the environment temperature of ISA −69° C. is from 1.5 to 4.5.
8. The gas turbine engine of claim 1 , wherein the heat management system is configured to provide the first amount of heat and the second amount of heat such that a ratio of the second proportion of the heat generated by the power gearbox and the turbomachinery bearings and dissipated to air at 65% of the core shaft maximum take-off speed at the environment temperature of ISA +40° C. to the second proportion of the heat generated by the power gearbox and the turbomachinery bearings and dissipated to air at 65% of the core shaft maximum take-off speed at the environment temperature of ISA +10° C. is from 1.10 to 1.25.
9. The gas turbine engine of claim 1 , wherein the heat management system is configured to provide the first amount of heat and the second amount of heat such that a ratio of the second proportion of the heat generated by the power gearbox and the turbomachinery bearings and dissipated to air at 65% of the core shaft maximum take-off speed at the environment temperature of ISA +40° C. to the second proportion of the heat generated by the power gearbox and the turbomachinery bearings and dissipated to air at 65% of the core shaft maximum take-off speed at the environment temperature of ISA −69° C. is from 1.1 to 2.1.
10. The gas turbine engine of claim 1 , wherein
the heat management system is configured to provide at the environment temperature of ISA +10° C. the first amount of heat and the second amount of heat such that a ratio of the first proportion to the second proportion is from 0.45 to 0.65.
11. The gas turbine engine of claim 1 , wherein the heat management system is configured to provide at the environment temperature of ISA −69° C. a ratio of the first proportion of the heat generated by the power gearbox and the turbomachinery bearings and dissipated to air at 85% of the core shaft maximum take-off speed to the second proportion of the heat generated by the power gearbox and the turbomachinery bearings and dissipated to air at 65% of the core shaft maximum take-off speed in from 0.30 to 0.55.
12. The gas turbine engine of claim 1 , wherein the combustor is a lean burn combustor.
13. The gas turbine engine of claim 1 , wherein the heat management system further includes a modulation device adapted to adjust a lubricant flow distribution between the power gearbox and the turbomachinery bearings.
14. The gas turbine engine of claim 1 , wherein the fan has a fan diameter from 210 cm to 380 cm.
15. The gas turbine engine of claim 1 , wherein the power gearbox has a gear ratio from 2.9 to 4.0.
16. The gas turbine engine of claim 1 , wherein the fan has a fan diameter from 210 cm to 230 cm and the power gearbox has a gear ratio from 3.2 to 4.2.
17. A method of operating a gas turbine engine for an aircraft, the method comprising providing the gas turbine engine comprising:
an engine core comprising a compressor, a combustor, a turbine, and a core shaft connecting the turbine to the compressor;
a fan comprising a plurality of fan blades and arranged upstream of the engine core;
turbomachinery bearings;
a power gearbox adapted to drive the fan at a lower rotation speed than the turbine; and
a heat management system configured to provide lubrication and cooling to the power gearbox and turbomachinery bearings, and comprising a pipe assembly, which includes first lubricant circuit and a second lubricant circuit separate from the first lubricant circuit, adapted to provide a lubricant flow to the power gearbox and turbomachinery bearings, at least one first air-lubricant heat exchanger arranged in the first lubricant circuit and configured to dissipate a first amount of heat to a first heat sink, at least one fuel-lubricant heat exchanger arranged in the second lubricant circuit and configured to dissipate a second amount of heat to a second heat sink, and at least one second air-lubricant heat exchanger arranged in the second lubricant circuit and configured to dissipate heat to the first heat sink;
wherein the first heat sink is air and the second heat sink is fuel;
wherein a first proportion of the heat generated by the power gearbox and the turbomachinery bearings and dissipated to air is defined as
(
first
amount
of
heat
111
first
amount
of
heat
111
+
se
cond
amount
of
heat
112
)
85
%
MTO
at 85% of a core shaft maximum take-off speed; and
a second proportion of heat generated by the power gearbox and the turbomachinery bearings and dissipated to air is defined as
(
first
amount
of
heat
111
first
amount
of
heat
111
+
se
cond
amount
of
heat
112
)
65
%
MTO
at 65% of the core shaft maximum take-off (MTO) speed;
wherein the method comprises controlling flow through the at least one first air-lubricant heat exchanger and the at least one fuel-lubricant heat exchanger to provide the first amount of heat and the second amount of heat such that at an environment temperature of ISA +40° C. the first proportion is from 0.55 to 0.70, and the second proportion is from 0.85 to 1.
18. The method of claim 17 , wherein at the environment temperature of ISA +40° C. the first proportion is from 0.60 to 0.70.
19. The method of claim 17 , wherein at the environment temperature of ISA +40° C. the second proportion is from 0.90 to 1.
20. The method of claim 17 , comprising controlling the flow through the at least one first air-lubricant heat exchanger and the at least one fuel-lubricant heat exchanger to provide the first amount of heat and the second amount of heat such that a ratio of the first proportion of the heat generated by the power gearbox and the turbomachinery bearings and dissipated to air at 85% of the core shaft maximum take-off speed at the environment temperature of ISA +40° C. to the first proportion of the heat generated by the power gearbox and the turbomachinery bearings and dissipated to air at 85% of the core shaft maximum take-off speed at the environment temperature of ISA +10° C. is from 1.20 to 1.42.