FR3140130A1 - Gas turbine engine with improved heat dissipation - Google Patents

Abstract

A gas turbine engine for an aircraft includes: an engine core including a compressor, a combustion chamber, 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; turbomachine bearings; a power reducer adapted to drive the fan at a lower rotational speed than the turbine; and a heat management system configured to provide lubrication and cooling to the gearbox and the turbomachine bearings, and comprising a pipe assembly adapted to provide a flow of lubricant to the gearbox and the turbomachine bearings, at least one air heat exchanger -lubricant for dissipating a first amount of heat to a first heat sink, and at least one fuel-lubricant heat exchanger for dissipating a second amount of heat to a second heat sink, wherein the first heat sink is air and the second heat sink is fuel.

Description

Gas turbine engine with improved heat dissipation TECHNICAL FIELD OF THE INVENTION

The present description relates to a gas turbine engine, and in particular to a gas turbine engine with an improved heat management system.

TECHNICAL BACKGROUND

Gas turbine engines are typically used to power aircraft, and the like. A gas turbine engine generally includes an air intake, a fan, one or more compressors, a combustion chamber, one or more turbines, and an exhaust nozzle. Air entering the air intake is compressed by the compressor, mixed with fuel and then introduced into the combustion chamber, where combustion of the air/fuel mixture occurs. The high temperature, high energy exhaust fluids are then fed to the turbine, where the fluid energy is converted into mechanical energy to drive the compressor and blower rotating by one or more suitable interconnecting shafts to provide a propulsion thrust.

Gas turbine engines include turbomachine bearings provided between rotating and stationary parts of the engine, for example at either end of the interconnecting shaft(s). Such turbomachine bearings require adequate lubrication and cooling under all foreseeable operating conditions for optimal performance, minimize any wear, and therefore increase operating life. For this purpose, an oil system is provided. The heat released from the turbomachine bearings by the oil system is then dissipated into the air and/or fuel to obtain benefits in terms of specific fuel consumption (SFC).

A general purpose of a gas turbine engine is to improve efficiency and therefore reduce fuel consumption. As it is generally recognized that moving more air at a slower speed is an efficient way to achieve a given thrust and therefore improve SFC geared architectures, in which a fan of increased diameter is driven through a power reducer with a lower rotation speed than the compressor has been developed. The power reducer, in addition to the turbomachine bearings, generates heat, which must be removed to ensure correct and efficient operation. However, the additional amount of heat generated by the power reducer, if dissipated into the fuel, contributes to a risk of thermal degradation of the fuel at specific operating conditions.

This risk is even more serious in geared gas turbine engines with lean-burn combustion chambers. Lean burn is a combustion technology that aims to reduce nitrogen oxides (NOx), which begin to form at high temperatures and increase exponentially with increasing temperature. In lean burn combustion, the air/fuel ratio (AFR) is greater than a stoichiometric ratio, which helps keep the combustion temperature within known limits to reduce NOx production. However, lean-burn gas turbine engines pose severe constraints in terms of the amount of heat that can be dissipated into the fuel. For example, in fuel spray nozzles with pilot and main streams, when the main stream is stepped (turned off), the fuel in the main stream is generally stagnant and therefore picks up heat which is undesirable due to degradation thermal fuel.

Simply increasing the amount of heat dissipated into the air can neither be technically feasible due to limited oil-air heat exchange capacity available, nor energetically advantageous, as this would increase the amount of heat dissipated into the air. wasted energy, and at the same time would not guarantee proper operation of the engine in all operating conditions.

In essence, geared gas turbine engines, and in particular geared gas turbine engines with lean burn combustion chambers, provide new challenges in terms of managing heat generated by engine components.

There is therefore a need to provide a gas turbine engine with an improved heat management system which makes it possible to minimize energy waste, improve specific fuel consumption, and provide efficient cooling to the engine components, including the power reducer and the turbomachine bearings, under all foreseeable operating conditions.

In one aspect, a gas turbine engine for an aircraft is provided, comprising: an engine core including a compressor, a combustion chamber, a turbine, and a core shaft connecting the turbine to the compressor, wherein the shaft core shaft has a maximum core shaft takeoff thrust speed in the range of 5500 rpm to 9500 rpm, preferably in the range of 5500 rpm to 8500 rpm; a fan comprising a plurality of fan blades and arranged upstream of the engine core; turbomachine bearings; a power reducer adapted to drive the fan at a rotation speed lower than the turbine; and a heat management system configured to provide lubrication and cooling to the gearbox and the turbomachine bearings, and comprising a pipe assembly adapted to provide a flow of lubricant to the gearbox and the turbomachine bearings, at least one air heat exchanger -lubricant for dissipating a first quantity of heat towards a first heat sink, and at least one fuel-lubricant heat exchanger for dissipating a second quantity of heat towards a second heat sink; wherein the first heat sink is air and the second heat sink is fuel, wherein the heat management system is configured to provide the first amount of heat and the second amount of heat such that a first proportion of heat generated by the reducer and the turbomachine and dissipated in the air defined as

_85%PMD

at 85% of maximum core shaft takeoff thrust speed is in the range 0.25 to 0.70.

As used here, proportions correspond to percentages and vice versa, and so, for example, a proportion of 0.5 corresponds to 50% and a proportion of 1 corresponds to 100%.

In other words, at 85% of a maximum thrust speed at core shaft takeoff 25% to 70% of the heat generated by the power reducer and the turbomachine is dissipated in the first heat sink (the remainder of the heat generated by the power reducer and the turbomachine being dissipated in the second heat sink).

As used here, a maximum takeoff thrust (PMD) condition has the conventional meaning. Maximum takeoff thrust conditions are defined as engine operation under International Standard Atmosphere (ISA) sea level pressure and temperature conditions +15°C at maximum takeoff thrust at the end of the runway. , which is typically set at an aircraft speed of about 0.25 Mn, or between about 0.24 and 0.27 Mn. Maximum takeoff thrust conditions for the engine are therefore defined as operating the engine at maximum takeoff thrust (e.g., maximum acceleration) for the engine at sea level pressure and temperature of international standard atmosphere (ISA) +15°C with a fan inlet speed of 0.25 Mn. A core shaft maximum takeoff thrust (PMD) speed is the rotational speed of the core shaft under PMD conditions, and is measured in rpm (revolutions per minute). The maximum core shaft takeoff thrust speed is typically identified among the engine performance data and/or in the engine type certificate data sheet.

In this description, upstream and downstream are defined in relation to the air flow through the compressor; and the front and rear are defined relative to the gas turbine engine, i.e. the fan being at the front and the turbine at the rear of the engine.

As used herein, the term "turbomachine bearings" includes any component of the gas turbine engine other than the power reduction gear that generates heat and is cooled by the heat management system.

The heat generated by the power reducer and the turbomachine bearings and removed by the lubricant flow is the sum of the first heat quantity and the second heat quantity.

The present inventor understands that providing a heat management system configured to provide specific proportions of heat dissipated into the air at 85% of the maximum core shaft takeoff thrust speed, i.e. a speed at , or close to, cruise conditions, and a core shaft with a maximum takeoff thrust speed within a specific range helps provide adequate lubrication and cooling to the power reducer and turbomachine bearings, minimize the size and therefore the weight of the heat exchangers, and maximize the beneficial effects of SFC for the longest flight phase.

To configure the heat management system to achieve specific proportions of heat dissipated to air several design and/or operating parameters can be used. For example, one or more of the following parameters may be used: type (e.g. parallel or counterflow), efficiency, number, and area of heat exchange surface of air-lubricant and fuel exchangers -lubricant, the flow conditions of the lubricant, for example the mass flow rate of lubricant passing through each of the air-lubricant heat exchanger(s) and the fuel-lubricant heat exchanger(s), the ratio between the mass flow rate of lubricant passing through the air-lubricant heat exchanger(s) and the mass flow rate of lubricant passing through the fuel-lubricant heat exchanger(s), and the flow state of the cooling air, for example the mass flow of cooling air. For example, increasing (or decreasing) the total heat exchange surface of the at least one air-lubricant heat exchanger between the lubricant and the first heat sink would make it possible to increase (or decrease) the first amount of heat dissipated to the first heat sink; increasing (or decreasing) the mass flow of lubricant to the air-lubricant heat exchanger(s), and/or increasing (or decreasing) the mass flow of cooling air, would increase (or decrease) the first quantity of heat dissipated to the first heat sink. Similarly increasing (or decreasing) the total heat exchange surface area of the at least one fuel-lubricant heat exchanger between the lubricant and the second heat sink would increase (or decrease) the second amount of heat dissipated to the second heat sink. ; increasing (or decreasing) the mass flow of lubricant to the fuel-lubricant heat exchanger(s) would increase (or decrease) the second amount of heat dissipated to the second heat sink.

The pipe assembly may include a lubricant bypass to either or both of the at least one air-lubricant heat exchanger and the at least one fuel-lubricant heat exchanger. By adjusting the mass flow of lubricant in the bypass and through the heat exchangers, the amount of heat dissipated to the first and second heat sinks can be further adjusted. In embodiments, the lubricant bypass may be integrated into, and part of, the at least one air-lubricant heat exchanger and the at least one fuel-lubricant heat exchanger.

The maximum core shaft takeoff thrust speed can be in the range of 5500 rpm to 7500 rpm, preferably in the range of 5500 rpm to 6500 rpm.

The first proportion may be greater than 0.25, or greater than 0.30, or greater than 0.35, or greater than 0.40, or greater than 0.45, or greater than 0.50, or greater than 0.55, and less than 0.70, or less than 0.65, for example in the range from 0.25 to 0.70, or in the range from 0.35 to 0.70, or in the range range of 0.45 to 0.70, or in the range of 0.50 to 0.70, or in the range of 0.55 to 0.70, or in the range of 0.55 to 0 .65.

A second proportion of heat generated by the reducer and the turbomachine and dissipated in the air is defined as

_65%PMD

at 65% of the maximum core shaft takeoff thrust speed, and the heat management system can be configured to provide the first amount of heat and the second amount of heat such that the second proportion of heat generated by the reducer and the turbomachine and dissipated in the air is in the range from 0.60 to 1, preferably in the range from 0.70 to 1, more preferably in the range from 0.75 to 1 .

The second proportion may be greater than 0.60, or greater than 0.65, or greater than 0.70, or greater than 0.75 and less than 1, or less than 0.95, for example in the range from 0.60 to 1, or in the range of 0.65 to 1, or in the range of 0.70 to 1, or in the range of 0.75 to 1, or in the range of 0, 75 to 0.95.

The first heat sink may be bypass air flowing through a bypass line of the gas turbine engine.

The heat management system may be adapted to provide a heat ratio of the first proportion to the second proportion in the range of 0.45 to 0.65, preferably 0.45 to 0.60, more preferably 0.47 to 0.58.

In addition, the environmental conditions, and in particular the environmental temperature, may have an impact on the capacity of the first heat sink and the second heat sink to dissipate the heat generated by the power reducer and the turbomachine bearings. The inventor found that such a capacity does not vary with the environmental temperature in the same way for the first and second heat sinks. In other words, the relative amounts of heat that (external or bypass) the air and fuel can reject can vary with temperature.

For this reason, the heat management system may be configured to vary the first amount of heat and the second amount of heat at different environmental temperatures so as to provide specific first and second proportions that maximize the second amount of heat. heat, and therefore maximize the SFC, minimize the total heat exchange surface of the heat exchangers, without causing fuel degradation.

As motors are typically certified to operate in an environmental temperature range between ISA (International Standard Atmosphere) -69°C and ISA +40°C, the heat management system can be configured to vary the first amount of heat and the second quantity of heat, and therefore the first and second proportions, to maximize the SFC, minimize the dimensions (and therefore the weight) of the heat exchangers, without causing fuel degradation, depending on the environmental temperature .

The heat management system can be configured to provide the first proportion and the second proportion in the ranges described above for environmental temperatures in the range from ISA -69°C to ISA 40°C.

Furthermore, the heat management system can be configured to provide the first proportion at an environment temperature of ISA 40 °C in the range of 0.55 to 0.70, preferably in the range of 0 .60 to 0.70, preferably in the range of 0.62 to 0.68.

As the environment temperature decreases, the air and fuel temperatures decrease, and the amount of heat that can be rejected to the air and fuel increases, although not proportionally to each other. . Thus, to maximize the SFC, minimize the dimensions (and therefore the weight) of the heat exchangers without inducing fuel degradation, the heat management can also be configured to provide the first proportion in specific ranges at temperatures of different environment.

At an environment temperature of ISA -69°C the heat management can be configured to provide the first amount of heat and the second amount of heat so as to provide the first proportion in the range of 0.20 to 0 .40, preferably in the range of 0.20 to 0.35, more preferably in the range of 0.20 to 0.30.

At an environment temperature of ISA +10°C the heat management system can be configured to provide the first amount of heat and the second amount of heat so as to provide the first proportion in the range of 0.35 to 0.65, preferably in the range of 0.40 to 0.60, more preferably in the range of 0.45 to 0.55.

As the first proportion is set at a core shaft speed corresponding to, or close to, cruise conditions, providing the first and second quantities of heat so as to provide the first proportion within the above ranges, the SFC can be maximized for the entire temperature range for which a motor is certified.

The heat management system can be configured to provide the second proportion within specific ranges proven to be safe in terms of fuel degradation, and maximize the SFC and minimize the dimensions (and therefore weight) of the heat exchangers, at different environmental temperatures. In particular, at an environmental temperature of ISA +40°C the heat management system can be configured to provide the second proportion in the range from 0.85 to 1, or from 0.90 to 1, or from 0.92 to 1, or from 0.92 to 0.98. At an environmental temperature of ISA -69°C the heat management system can be configured to provide the second proportion in the range of 0.40 to 0.75, or in the range of 0.50 to 0.75, or in the range 0.55 to 0.70. At an environmental temperature of ISA +10°C the heat management system can be configured to provide the second proportion in the range of 0.60 to 0.95, or in the range of 0.70 to 0.95, or in the range of 0.75 to 0.95, or in the range of 0.80 to 0.92.

At an environment temperature of ISA +40°C the heat management system can be configured to provide a heat ratio of the first proportion to the second proportion in the range of 0.50 to 0.80, or in the range of 0.55 to 0.75, or in the range of 0.60 to 0.70, or in the range of 0.60 to 0.67.

At an environment temperature of ISA +10°C the heat management system can be configured to provide a heat ratio of the first proportion to the second proportion in the range of 0.45 to 0.65, or in the range of 0.45 to 0.60, or in the range of 0.50 to 0.60, or in the range of 0.50 to 0.58, or in the range of 0.47 at 0.58.

At an environment temperature of ISA -69°C the heat management system can be configured to provide a heat ratio of the first proportion to the second proportion in the range of 0.30 to 0.55, from preferably in the range of 0.30 to 0.50, more preferably in the range of 0.35 to 0.45.

Further, the heat management system may be configured to provide a ratio of the first proportion at an environment temperature of ISA +40°C to the first proportion at an environment temperature of ISA -69°C in the range of 1.5 to 4.5, preferably in the range of 2.0 to 4.0, more preferably in the range of 2.0 to 3.5.

The heat management system may be configured to provide a ratio of the second proportion at an environment temperature of ISA +40°C to the second proportion at an environment temperature of ISA -69°C in the range ranging from 1.0 to 2.1, preferably in the range of 1.2 to 2.1, more preferably in the range of 1.4 to 2.0.

Further, the heat management system may be configured to provide a ratio of the first proportion at an environment temperature of ISA +40°C to the first proportion at an environment temperature of ISA +10°C in the range of 1.20 to 1.42, or in the range of 1.22 to 1.41, or in the range of 1.25 to 1.40.

Further, the heat management system may be configured to provide the first amount of heat and the second amount of heat such that a ratio of the second proportion at an environmental temperature of ISA +40°C to the second proportion at an environment temperature of ISA +10°C in the range of 1.10 to 1.25, or in the range of 1.10 to 1.22, or in the range of 1 .11 to 1.20.

The heat management system may be configured to provide the first amount of heat and the second amount of heat such that a proportion of heat generated by the power reducer and the turbomachine and dissipated to the air is greater than A NH + B, and less than 1, wherein A is -1.15, B is 1.48 or greater, and NH is a core shaft speed expressed as a proportion of the thrust speed maximum at core shaft takeoff and is in the range 0.65 to 1.

B can be equal to or greater than 1.5, or 1.52, or 1.54, or 1.56.

In other words, when NH is 0.65 the core shaft speed is 65% of the maximum core shaft takeoff thrust speed, and when NH is 1 the core shaft speed is of 100%, or equal to, the maximum thrust speed at core shaft takeoff.

B can be equal to or greater than 1.5, or 1.52, or 1.54, or 1.56.

The heat management system may be configured to provide the first amount of heat and the second amount of heat such that the proportion of heat generated by the power reducer and the turbomachine and dissipated to the air is

less than the smallest among 1 and C NH + D,

where C is -1.84, D is in the range of 2.10 to 2.30, and NH is in the range of 0.65 to 1.

D may be in the range of 2.18 to 2.30, or in the range of 2.18 to 2.25, or in the range of 2.20 to 2.25.

The heat management system may be configured to provide the first amount of heat and the second amount of heat such that the proportion of heat generated by the power reducer and the turbomachine and dissipated to the air is

greater than A NH + B, and less than the smallest of 1 and E (NH - 1) + F,

wherein A is -1.15, B is 1.48 or greater, E is in the range of -1.16 to -3, F is 0.37 or greater, and NH is the core shaft speed expressed as a proportion of the maximum core shaft takeoff thrust speed and is in the range 0.65 to 1.

B can be equal to or greater than 1.5, or 1.52, or 1.54, or 1.56.

E can be in the range of -1.16 to -2.5, or in the range of -1.16 to -1.95.

In the above embodiments, the core shaft speed NH may be in the range of 0.65 to 0.95, or in the range of 0.65 to 0.90, or in the range ranging from 0.65 to 0.85.

The blower can have a blower diameter in the range of 210 cm to 380 cm.

The power reducer can have a gear ratio in the range of 2.9 to 4.0, or 3.0 to 3.8.

According to one aspect, a method of operating a gas turbine engine for an aircraft is provided, the method comprising providing a gas turbine engine comprising: an engine core comprising a compressor, a combustion chamber, a turbine, and a core shaft connecting the turbine to the compressor, wherein the core shaft has a maximum thrust speed at core shaft takeoff in the range of 5500 rpm to 9500 rpm, of preferably in the range from 5500 rpm to 8500 rpm; a fan comprising a plurality of fan blades and arranged upstream of the engine core; turbomachine bearings; a power reducer adapted to drive the fan at a rotation speed lower than the turbine; and a heat management system configured to provide lubrication and cooling to the gearbox and the turbomachine bearings, and comprising a pipe assembly adapted to provide a flow of lubricant to the gearbox and the turbomachine bearings, at least one air heat exchanger -lubricant for dissipating a first quantity of heat towards a first heat sink, and at least one fuel-lubricant heat exchanger for dissipating a second quantity of heat towards a second heat sink; wherein the first heat sink is air and the second heat sink is fuel; and wherein the method comprises the step of operating the heat management system to provide the first amount of heat and the second amount of heat such that a first proportion of heat generated by the gearbox and the turbomachine and dissipated in air defined as

_85%PMD

at 85% of maximum core shaft takeoff thrust speed is in the range 0.25 to 0.70.

The first proportion of heat generated by the reducer and the turbomachine and dissipated into the air may be in the range of 0.35 to 0.70, preferably in the range of 0.45 to 0.70, more preferably in the range from 0.50 to 0.70.

The method may include the step of operating the heat management system to provide at an environment temperature of ISA +40°C the first amount of heat and the second amount of heat such that the first proportion of heat generated by the reducer and the turbomachine and dissipated into the air at 85% of the maximum thrust speed at core shaft takeoff is in the range from 0.55 to 0.75, preferably in the range of 0.55 to 0.70, more preferably in the range of 0.60 to 0.70.

The method may include the step of operating the heat management system to provide the first amount of heat and the second amount of heat such that a ratio of the first proportion at an environment temperature of ISA +40°C to the first proportion at an environment temperature of ISA -69°C is in the range of 1.5 to 4.5, preferably in the range of 2.0 to 4.0, more preferably in the range of 2, 0 to 3.5

The method may include the step of operating the heat management system to provide the first amount of heat and the second amount of heat such that a ratio of the first proportion at an environment temperature of ISA +40°C to the first proportion at an environment temperature of ISA +10°C is in the range of 1.20 to 1.42, preferably in the range of 1.22 to 1.41, more preferably in the range of 1, 25 to 1.40.

In one aspect, there is provided a gas turbine engine for an aircraft comprising: an engine core including a compressor, a combustion chamber, 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; turbomachine bearings; a power reducer adapted to drive the fan at a rotation speed lower than the turbine; and a heat management system configured to provide lubrication and cooling to the power reducer and the turbomachine bearings, and including a pipe assembly adapted to provide a flow of lubricant to the power reducer and the turbomachine bearings to remove heat generated by the power reducer and the turbomachine bearings, at least one air-lubricant heat exchanger for dissipating a first quantity of heat to a first heat sink, and at least one fuel-lubricant heat exchanger for dissipating a second quantity of heat heat to a second heat sink, wherein the first heat sink is air and the second heat sink is fuel. A first proportion of the heat generated by the power reducer and the turbomachine and dissipated in the air is defined as

_85%PMD

at 85% of maximum thrust speed at core shaft takeoff, and

a second proportion of the heat generated by the power reducer and the turbomachine and dissipated in the air is defined as

_65%PMD

at 65% of maximum thrust speed at core shaft takeoff. The heat management system is configured to provide the first amount of heat and the second amount of heat such that the first proportion is in the range of 0.25 to 0.70, and the second proportion is within the range is 0.60 to 1.

As used here, proportions correspond to percentages and vice versa, and so, for example, a proportion of 0.5 corresponds to 50% and a proportion of 1 corresponds to 100%.

In other words, at 85% of a maximum thrust speed at core shaft takeoff 25% to 70% of the heat generated by the power reducer and the turbomachine is dissipated in the first heat sink (the remainder of the heat generated by the power reducer and the turbomachine being dissipated in the second heat sink), and at 65% of the maximum thrust speed at core shaft takeoff from 60% to 100% of the heat generated by the power reducer and the turbomachine is dissipated in the first heat sink (the remainder of the heat generated by the power reducer and the turbomachine being dissipated in the second heat sink).

The present inventor understands that providing a heat management system configured to provide specific proportions of heat dissipated into the air at 65% of the maximum core shaft takeoff thrust speed, i.e. a speed at , or close to idle in flight, and at 85% of the maximum thrust speed at core shaft takeoff, i.e. a speed at, or close to, cruise conditions, ensures adequate lubrication and cooling at power reducer and turbomachine bearings, minimize the size and therefore the weight of the heat exchangers, maximize the beneficial effects of SFC and at the same time avoid thermal degradation of fuel in all operating conditions.

The first proportion may be greater than 0.25, or greater than 0.30, or greater than 0.35, or greater than 0.40, or greater than 0.45, or greater than 0.50, or greater than 0.55, and less than 0.70, or less than 0.65, for example in the range from 0.25 to 0.70, or in the range from 0.35 to 0.70, or in the range range of 0.45 to 0.70, or in the range of 0.50 to 0.70, or in the range of 0.55 to 0.70, or in the range of 0.55 to 0 .65.

The second proportion may be greater than 0.60, or greater than 0.65, or greater than 0.70, or greater than 0.75 and less than 1, or less than 0.95, for example in the range from 0.60 to 1, or in the range of 0.65 to 1, or in the range of 0.70 to 1, or in the range of 0.75 to 1, or in the range of 0, 75 to 0.95.

The first heat sink may be bypass air flowing through a bypass line of the gas turbine engine.

The heat management system may be adapted to provide a heat ratio of the first proportion to the second proportion in the range of 0.45 to 0.65, preferably 0.45 to 0.60, more preferably 0.47 to 0.58.

The pipe assembly may include a first lubricant circuit adapted to provide a first flow of lubricant and a second lubricant circuit adapted to provide a second flow of lubricant, the at least one air-lubricant heat exchanger being arranged in the first lubricant circuit and the at least one fuel-lubricant heat exchanger being arranged in the second lubricant circuit.

The heat management system may include a modulation device adapted to adjust a lubricant flow distribution between the power reducer and the turbomachine bearings.

The modulation device may include a first pump device arranged in the first lubricant circuit to adjust the first lubricant flow and a second pump device arranged in the second lubricant circuit to adjust the second lubricant flow.

The first lubricant circuit can provide lubrication and cooling to the power reducer.

The second lubricant circuit can provide lubrication and cooling to the turbomachine bearings.

The gas turbine engine may include at least two air-lubricant heat exchangers for dissipating the first amount of heat in the first heat sink, at least one of which is arranged in the first lubricant circuit and at least one of which is arranged in the first heat sink. second lubricant circuit.

The heat management system may include a lubricant reservoir in fluid communication with, and supplying lubricant to, the first and second lubricant circuits.

The heat management system may include a first lubricant reservoir in fluid communication with, and supplying lubricant to, the first lubricant circuit and a second lubricant reservoir in fluid communication with, and supplying lubricant to, the second lubricant circuit. lubricant.

The blower may have a blower diameter in the range of 210 cm to 380 cm, or 210 cm to 370 cm, or 220 cm to 370 cm, for example 340 to 370.

The power reducer may have a gear ratio in the range of 2.9 to 4.0, or 3.0 to 3.8, or 3.1 to 3.7.

The at least one air-lubricant heat exchanger may be a matrix air-cooled oil cooler (MACOC).

The combustion chamber may be a lean burn combustion chamber. The lean burn combustion chamber may include a plurality of lean burn fuel spray nozzles, each fuel spray nozzle including a pilot fuel injector and a main fuel injector.

According to one aspect, a method of operating a gas turbine engine for an aircraft is provided, the method comprising providing a gas turbine engine comprising: an engine core comprising a compressor, a combustion chamber, 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; turbomachine bearings; a power reducer adapted to drive the fan at a rotation speed lower than the turbine; and a heat management system configured to provide lubrication and cooling to the power reducer and the turbomachine bearings, and including a pipe assembly adapted to provide a flow of lubricant to the power reducer and the turbomachine bearings, at least one an air-lubricant heat exchanger for dissipating a first amount of heat to a first heat sink, and at least one fuel-lubricant heat exchanger for dissipating a second amount of heat to a second heat sink, wherein the first heat sink is air and the second heat sink is fuel. A first proportion of heat generated by the power reducer and the turbomachine and dissipated in the air is defined as

_85%PMD

at 85% of maximum thrust speed at core shaft takeoff, and

a second proportion of heat generated by the power reducer and the turbomachine and dissipated in the air is defined as

_65%PMD

at 65% of the maximum thrust speed at core shaft takeoff,

wherein the method comprises the step of operating the heat management system to provide the first amount of heat and the second amount of heat such that the first proportion is in the range of 0.25 to 0 .70 and the second proportion is in the range 0.60 to 1.

The first proportion may be greater than 0.25, or greater than 0.30, or greater than 0.35, or greater than 0.40, or greater than 0.45, or greater than 0.50, or greater than 0.55, and less than 0.70, or less than 0.65, for example in the range from 0.25 to 0.70, or in the range from 0.35 to 0.70, or in the range range of 0.45 to 0.70, or in the range of 0.50 to 0.70, or in the range of 0.55 to 0.70, or in the range of 0.55 to 0 .65.

The second proportion may be greater than 0.60, or greater than 0.65, or greater than 0.70, or greater than 0.75 and less than 1, or less than 0.95, for example in the range from 0.60 to 1, or in the range of 0.65 to 1, or in the range of 0.70 to 1, or in the range of 0.75 to 1, or in the range of 0, 75 to 0.95.

The method may include the step of operating the heat management system to provide the first amount of heat and the second amount of heat such that a heat ratio of the first proportion to the second proportion is within the range of 0.45 to 0.65, or in the range of 0.45 to 0.60, or in the range of 0.47 to 0.58.

In one aspect, there is provided a gas turbine engine for an aircraft comprising: an engine core including a compressor, a combustion chamber, 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; turbomachine bearings; a power reducer adapted to drive the fan at a rotation speed lower than the turbine; and a heat management system configured to provide lubrication and cooling to the gearbox and the turbomachine bearings, and including a pipe assembly adapted to provide a flow of lubricant to the gearbox and the turbomachine bearings to remove heat generated by the gearbox and the turbomachine bearings, at least one air-lubricant heat exchanger for dissipating a first quantity of heat to a first heat sink, and at least one fuel-lubricant heat exchanger for dissipating a second quantity of heat in a second heat sink ; wherein the first heat sink is air and the second heat sink is fuel;

in which a first proportion of heat generated by the reducer and the turbomachine and dissipated in the air is defined as

_85%PMD

at 85% of maximum thrust speed at core shaft takeoff,

in which a second proportion of heat generated by the reducer and the turbomachine and dissipated in the air is defined as

_65%PMD

at 65% of maximum thrust speed at core shaft takeoff; And

wherein the heat management system is configured to provide at an environment temperature of ISA +10°C the first quantity of heat and the second quantity of heat such that a ratio of the first proportion to the second proportion is within the range is 0.45 to 0.65.

As used here, proportions correspond to percentages and vice versa, and so, for example, a proportion of 0.5 corresponds to 50% and a proportion of 1 corresponds to 100%.

The present inventor understood that the provision of a heat management system configured to provide, at a temperature of ISA +10 °C, i.e. during relatively hot days, specific proportions of heat dissipated into the air at 65 % of the maximum thrust speed at core shaft takeoff, i.e. a speed at or close to idle in flight, and at 85% of the maximum thrust speed at core shaft takeoff, i.e. a speed at, or close to, cruising conditions, makes it possible to ensure adequate lubrication and cooling to the power reducer and to the turbomachine bearings, minimize the size and therefore the weight of the heat exchangers, maximize the beneficial effects of SFC and at the same time avoid time thermal degradation of fuel under all operating conditions.

At an environmental temperature of ISA +10 °C, the ratio of the first proportion of heat generated by the reducer and the turbomachine and dissipated in the air to the second proportion of heat generated by the reducer and the turbomachine and dissipated in air may be in the range of 0.45 to 0.60, preferably in the range of 0.47 to 0.58.

The heat management system may be configured to provide the first amount of heat and the second amount of heat such that the first proportion of heat generated by the reducer and the turbomachine and dissipated to the air may be in the range from 0.25 to 0.70, preferably in the range of 0.35 to 0.70, more preferably in the range of 0.45 to 0.70, even more preferably in the range of 0.50 at 0.70.

The heat management system may be configured to provide the first amount of heat and the second amount of heat such that the second proportion of heat generated by the reducer and the turbomachine and dissipated to the air is in the range from 0.60 to 1, preferably in the range of 0.70 to 1, more preferably in the range of 0.75 to 1.

The pipe assembly may include a first lubricant circuit adapted to provide a first flow of lubricant and a second lubricant circuit adapted to provide a second flow of lubricant, the at least one air-lubricant heat exchanger being arranged in the first lubricant circuit and the at least one fuel-lubricant heat exchanger being arranged in the second lubricant circuit.

The maximum core shaft takeoff thrust speed can be in the range of 5500 rpm to 9500 rpm, preferably in the range of 5500 rpm to 8500 rpm.

The fan may have a fan rotation speed at PMD conditions in the range of 1500 rpm to 2800 rpm, preferably in the range of 1600 rpm to 2500 rpm, more preferably in the range of 1600 rpm to 2500 rpm. range from 1600 rpm to 2200 rpm

The heat management system can be configured to provide at an environment temperature of ISA +40°C the first amount of heat and the second amount of heat such that the first proportion of heat generated by the reducer and the turbomachine and dissipated in the air at 85% of the maximum core shaft takeoff thrust speed is in the range of 0.55 to 0.70, preferably in the range of 0.60 to 0 .70.

The heat management system can be configured to provide at an environment temperature of ISA +40°C the first amount of heat and the second amount of heat such that the second proportion of heat generated by the reducer and the turbomachine and dissipated in the air at 65% of the maximum core shaft takeoff thrust speed is in the range of 0.85 to 1, preferably in the range of 0.90 to 1.

The heat management system may be configured to provide at an environment temperature of ISA -69°C the first amount of heat and the second amount of heat such that the ratio of the first proportion to the second proportion is found in the range of 0.30 to 0.55, preferably in the range of 0.35 to 0.45.

The heat management system may be configured to provide the first amount of heat and the second amount of heat such that the first proportion is in the range of 0.40 to 0.60 at an environment temperature d 'ISA +10°C, and the second proportion at an environment temperature of ISA +10°C is in the range from 0.80 to 0.92.

The heat management system may be configured to provide the first amount of heat and the second amount of heat such that a ratio of the first proportion at an environment temperature of ISA +40°C to the first proportion at an environmental temperature of ISA -69°C is in the range of 1.5 to 4.5, preferably in the range of 2.0 to 4.0, more preferably in the range of 2.0 to 3.5.

The heat management system may be configured to provide the first amount of heat and the second amount of heat such that a ratio of the second proportion at an environment temperature of ISA +40°C to the second proportion at an environment temperature of ISA -69°C is in the range of 1.0 to 2.1, preferably in the range of 1.2 to 2.1, more preferably in the range of 1.4 to 2.0.

The heat management system may be configured to provide the first amount of heat and the second amount of heat such that a ratio of the first proportion at an environment temperature of ISA +40°C to the first proportion at an environment temperature of ISA +10°C is in the range of 1.20 to 1.42, preferably in the range of 1.22 to 1.41, more preferably in the range of 1 .25 to 1.40.

The heat management system may be configured to provide the first amount of heat and the second amount of heat such that a ratio of the second proportion at an environment temperature of ISA +40°C to the second proportion at an environment temperature of ISA +10°C is in the range of 1.10 to 1.25, preferably in the range of 1.10 to 1.22, more preferably in the range of 1.11 to 1.20.

According to one aspect, a method of operating a gas turbine engine for an aircraft is provided, the method comprising providing a gas turbine engine comprising: an engine core comprising a compressor, a combustion chamber, 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; turbomachine bearings; a power reducer adapted to drive the fan at a rotation speed lower than the turbine; and a heat management system configured to provide lubrication and cooling to the gearbox and the turbomachine bearings, and comprising a pipe assembly adapted to provide a flow of lubricant to the gearbox and the turbomachine bearings, at least one air heat exchanger -lubricant for dissipating a first quantity of heat towards a first heat sink, and at least one fuel-lubricant heat exchanger for dissipating a second quantity of heat towards a second heat sink; wherein the first heat sink is air and the second heat sink is fuel; And

in which a first proportion of heat generated by the reducer and the turbomachine and dissipated in the air is defined as

_85%PMD

at 85% of maximum thrust speed at core shaft takeoff, and

a second proportion of heat generated by the reducer and the turbomachine and dissipated in the air is defined as

_65%PMD

at 65% of the maximum thrust speed at core shaft takeoff,

wherein the method comprises the step of operating the heat management system to provide at an 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 in the range 0.45 to 0.65.

The ratio may be in the range of 0.45 to 0.60, preferably in the range of 0.47 to 58.

The method may include the step of operating the heat management system to provide the first amount of heat and the second amount of heat such that the first proportion of heat generated by the reducer and the turbomachine and dissipated in the air is in the range of 0.25 to 0.70, preferably in the range of 0.35 to 0.70, more preferably in the range of 0.45 to 0.70, even more preferably in the range from 0.50 to 0.70.

The method may include the step of operating the heat management system to provide the first amount of heat and the second amount of heat such that the second proportion of heat generated by the reducer and the turbomachine and dissipated in the The air is in the range of 0.60 to 1, preferably in the range of 0.70 to 1, more preferably in the range of 0.75 to 1.

The method may include the step of operating the heat management system to provide at an environment temperature of ISA +40°C the first amount of heat and the second amount of heat such that the first proportion of heat generated by the reducer and the turbomachine and dissipated into the air at 85% of the maximum thrust speed at core shaft takeoff is in the range from 0.55 to 0.75, preferably in the range of 0.55 to 0.70, more preferably in the range of 0.60 to 0.70; and/or the step of operating the heat management system to provide at an environment temperature of ISA +40°C the first amount of heat and the second amount of heat such that the second proportion of heat generated by the reducer and the turbomachine and dissipated in the air at 65% of the maximum thrust speed at core shaft takeoff is in the range from 0.85 to 1, preferably in the range from 0.90 to 1.

According to one aspect, a method of operating a gas turbine engine for an aircraft is provided, the method comprising providing a gas turbine engine comprising: an engine core comprising a compressor, a combustion chamber, 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; turbomachine bearings; a power reducer adapted to drive the fan at a rotation speed lower than the turbine; and a heat management system configured to provide lubrication and cooling to the gearbox and the turbomachine bearings, and comprising a pipe assembly adapted to provide a flow of lubricant to the gearbox and the turbomachine bearings, at least one air heat exchanger -lubricant adapted to receive cooling air from the bypass line to dissipate a first quantity of heat to a first heat sink, and at least one fuel-lubricant heat exchanger to dissipate a second quantity of heat to a second heat sink; wherein the first heat sink is bypass air and the second heat sink is fuel, the method further comprising:

- the operation of the heat management system to provide the first quantity of heat and the second quantity of heat such that a first proportion of heat generated by the reducer and the turbomachine and dissipated in the air defined as

_85%PMD

at 85% of a maximum core shaft takeoff thrust speed is in the range of 0.25 to 0.70; And

- operation of the fan at cruise conditions to provide a fan pressure ratio in the range from 1.35 to 1.43.

As used here, proportions correspond to percentages and vice versa, and so, for example, a proportion of 0.5 corresponds to 50% and a proportion of 1 corresponds to 100%.

In other words, at 85% of a maximum thrust speed at core shaft takeoff 25% to 70% of the heat generated by the power reducer and the turbomachine is dissipated in the first heat sink (the remainder of the heat generated by the power reducer and the turbomachine being dissipated in the second heat sink.

The present inventor understands that providing a heat management system configured to provide specific proportions of heat dissipated into the air at 85% of the maximum core shaft takeoff thrust speed, i.e. a speed at , or close to, cruise conditions, and a fan configured to provide at cruise conditions a fan pressure ratio within a specific range helps ensure adequate lubrication and cooling to the power reducer and turbomachine bearings, minimize the size and therefore the weight of the heat exchangers, maximize the beneficial effects of SFC for the longest flight phase.

The method may include operating the heat management system to provide the first amount of heat and the second amount of heat such that the first proportion is in the range of 0.35 to 0.70, preferably within the range of 0.45 to 0.70, more preferably in the range of 0.50 to 0.70.

The method may include operating the heat management system to provide at an environment temperature of ISA +40°C the first amount of heat and the second amount of heat such that the first proportion is within the range ranging from 0.55 to 0.70, preferably in the range from 0.60 to 0.70.

The method may include operating the heat management system to provide at an environment temperature of ISA +10°C the first amount of heat and the second amount of heat such that the first proportion is within the range ranging from 0.35 to 0.65, preferably in the range of 0.40 to 0.60, more preferably in the range of 0.45 to 0.55.

The method may include operating the heat management system to provide the first amount of heat and the second amount of heat such that a ratio of the first proportion to an environmental temperature of ISA +40°C at the first proportion at an environment temperature of ISA -69°C is in the range of 1.5 to 4.5, preferably in the range of 2.0 to 4.0, more preferably in the range range from 2.0 to 3.5.

The method may include operating the heat management system to provide the first amount of heat and the second amount of heat such that a ratio of the first proportion at an environment temperature of ISA +40°C to the first proportion at an environmental temperature of ISA +10°C is in the range of 1.20 to 1.42, preferably in the range of 1.22 to 1.41, more preferably in the range of 1.25 to 1 .40.

The method may include operating the heat management system to provide the first quantity of heat and the second quantity of heat such that a second proportion of heat generated by the reducer and the turbomachine and dissipated to the air defined as

_65%PMD

at 65% of the maximum core shaft takeoff thrust speed is in the range of 0.60 to 1, preferably in the range of 0.70 to 1, more preferably in the range of 0 .75 to 1.

The method may include operating the heat management system to provide the first amount of heat and the second amount of heat such that a ratio of the second proportion at an environmental temperature of ISA +40°C to the second proportion at an environment temperature of ISA -69°C is in the range of 1.1 to 2.1, preferably in the range of 1.2 to 2.1, more preferably in the range of range from 1.4 to 2.0.

The method may include operating the heat management system to provide the first amount of heat and the second amount of heat such that a ratio of the second proportion at an environmental temperature of ISA +40°C to the second proportion at an environment temperature of ISA +10°C is in the range from 1.10 to 1.25, preferably in the range from 1.10 to 1.22, more preferably in the range range from 1.11 to 1.20.

The method may include operating the heat management system to provide at an environment temperature of ISA -69°C the first amount of heat and the second amount of heat such that a ratio of the first proportion to the second proportion in the range of 0.30 to 0.55, preferably in the range of 0.35 to 0.45.

The method may include operating the heat management system to provide at an 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 is in the range of 0.45 to 0.65, preferably in the range of 0.45 to 0.60, more preferably in the range of 0.47 to 0.58.

The method may include operating the heat management system to provide at an environment temperature of ISA +10°C the first amount of heat and the second amount of heat such that the second proportion is in the range from 0.60 to 0.95.

The method may include operating the heat management system to provide at an environment temperature of ISA -69°C the first amount of heat and the second amount of heat such that the second proportion is within the range ranging from 0.40 to 0.75.

The method may include operating the heat management system to provide the first amount of heat and the second amount of heat such that a proportion of heat generated by the reducer and the turbomachine and dissipated to the air is

greater than A NH + B, and less than the smallest of 1 and C NH + D,

in which A is equal to -1.15, B is equal to or greater than 1.48, C is equal to -1.84, D is in the range from 2.10 to 2.30, preferably in the range range of 2.18 to 2.30, more preferably in the range of 2.18 to 2.25, even more preferably in the range of 2.20 to 2.25; and NH is the core shaft speed expressed as a proportion of the maximum core shaft takeoff thrust speed and is in the range 0.65 to 1.

B can be equal to or greater than 1.5, or 1.52, or 1.54, or 1.56.

The method may include operating the heat management system to provide the first amount of heat and the second amount of heat such that a proportion of heat generated by the reducer and the turbomachine and dissipated to the air is

greater than A NH + B, and less than the smallest of 1 and E (NH - 1) + F,

in which A is equal to -1.15, B is equal to or greater than 1.48, E is in the range of -1.16 to -3, preferably in the range of -1.16 to - 2.5, more preferably in the range of -1.16 to -1.95; F is equal to or greater than 0.37, and NH is the core shaft speed expressed as a proportion of the maximum core shaft takeoff thrust speed and is in the range of 0.65 to 1.

B can be equal to or greater than 1.5, or 1.52, or 1.54, or 1.56.

NH can be in the range of 0.65 to 0.85.

The heat management system may include a flow restriction valve arranged downstream of the air-lubricant heat exchanger, and the method may include operating the flow restriction valve to vary a mass flow rate of cooling air through the air-lubricant heat exchanger, which allows the first amount of heat to be varied.

In one aspect, there is provided a gas turbine engine comprising: an engine core including a compressor, a combustion chamber, 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; turbomachine bearings; a power reducer adapted to drive the fan at a rotation speed lower than the turbine; and a heat management system configured to provide lubrication and cooling to the gearbox and the turbomachine bearings, and comprising a pipe assembly adapted to provide a flow of lubricant to the gearbox and the turbomachine bearings, at least one air heat exchanger -lubricant adapted to receive cooling air from the bypass line to dissipate a first quantity of heat to a first heat sink, and at least one fuel-lubricant heat exchanger to dissipate a second quantity of heat to a second heat sink; wherein the first heat sink is bypass air and the second heat sink is fuel, wherein the heat management system is configured to provide the first amount of heat and the second amount of heat such that a first proportion of heat generated by the reducer and the turbomachine and dissipated in the air defined as

_85%PMD

at 85% of maximum core shaft takeoff thrust speed is in the range of 0.25 to 0.70; And

wherein the fan is configured to provide under cruise conditions a fan pressure ratio in the range of 1.35 to 1.43.

The heat management system can be configured to provide at an environment temperature of ISA +40°C the first amount of heat and the second amount of heat such that the first proportion is in the range of 0 .55 to 0.70, preferably in the range 0.60 to 0.70

The heat management system can be configured to provide at an environment temperature of ISA +10°C the first amount of heat and the second amount of heat such that the first proportion is in the range of 0 .35 to 0.65, preferably in the range of 0.40 to 0.60, more preferably in the range of 0.45 to 0.55.

The heat management system may be configured to provide the first amount of heat and the second amount of heat such that a ratio of the first proportion at an environment temperature of ISA +40°C to the first proportion at an environment temperature of ISA +10°C is in the range of 1.20 to 1.42, preferably in the range of 1.22 to 1.41, more preferably in the range of 1 .25 to 1.40.

In one aspect, there is provided a gas turbine engine for an aircraft comprising: an engine core including a compressor, a combustion chamber, 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; turbomachine bearings; a power reducer adapted to drive the fan at a rotation speed lower than the turbine; and a heat management system configured to provide lubrication and cooling to the gearbox and the turbomachine bearings, and comprising a pipe assembly adapted to provide a flow of lubricant to the gearbox and the turbomachine bearings, at least one air heat exchanger -lubricant for dissipating a first amount of heat to a first heat sink, and at least one fuel-lubricant heat exchanger for dissipating a second amount of heat to a second heat sink, wherein the first heat sink is air and the second heat sink is fuel, wherein the heat management system is configured to provide the first amount of heat and the second amount of heat such that a proportion of heat generated by the reducer and the turbomachine and dissipated to air at 65% of a core shaft maximum takeoff thrust (PMD) speed defined as

_65%PMD

is in the range of 0.6 to 1, and wherein the fan is configured to have a fan rotation speed at PMD conditions in the range of 1500 rpm to 2800 rpm.

As used here, proportions correspond to percentages and vice versa, and so, for example, a proportion of 0.5 corresponds to 50% and a proportion of 1 corresponds to 100%.

In other words, at 65% of the maximum core shaft takeoff thrust speed 60% to 100% of the heat generated by the power reducer and the turbomachine is dissipated in the first heat sink (the remainder of the heat generated by the power reducer and the turbomachine being dissipated in the second heat sink).

The present inventor understands that providing a heat management system configured to provide specific proportions of heat dissipated to the air at 65% of the maximum core shaft takeoff thrust speed, i.e. a speed at , or close to, idle in flight, and a fan configured to have a fan rotation speed at PMD conditions within a specific range helps ensure adequate lubrication and cooling of the power reducer and turbomachine bearings, minimize the size and therefore weight of the heat exchangers, maximize the beneficial effects of SFC and at the same time avoid thermal degradation of fuel when fuel flow is minimal.

The blower may be configured to have a fan rotation speed at PMD conditions in the range of 1600 rpm to 2500 rpm, preferably in the range of 1600 rpm to 2200 rpm, more preferably in the range from 1700 rpm to 1900 rpm.

The heat management system can be configured to provide the first amount of heat and the second amount of heat such that the proportion of heat generated by the reducer and the turbomachine and dissipated to the air at 65% of the speed of Maximum takeoff thrust (PMD) of core shaft is in the range of 0.70 to 1, preferably in the range of 0.75 to 1.

The heat management system can be configured to provide at an environment temperature of ISA +40°C the first amount of heat and the second amount of heat such that the proportion of heat generated by the reducer and the turbomachine and dissipated in air at 65% of the core shaft maximum takeoff thrust (PMD) speed in the range of 0.85 to 1, preferably in the range of 0.90 to 1.

The heat management system can be configured to provide at an environment temperature of ISA +10°C the first amount of heat and the second amount of heat such that the proportion of heat generated by the reducer and the turbomachine and dissipated in air at 65% of the core shaft maximum takeoff thrust (PMD) speed in the range 0.80 to 0.92.

The heat management system can be configured to provide the first amount of heat and the second amount of heat such that a ratio of the proportion of heat generated by the reducer and the turbomachine and dissipated to the air is 65% from the maximum take-off thrust speed (PMD) of the core shaft at an environmental temperature of ISA 40 °C to the proportion of heat generated by the gearbox and the turbomachine and dissipated in the air at 65% of the maximum takeoff thrust speed (PMD) of core shaft at an environment temperature of ISA -69 °C is in the range of 1.1 to 2.1, preferably in the range of 1.2 to 2.1, more preferably in the range of 1.4 to 2.0.

The heat management system can be configured to provide the first amount of heat and the second amount of heat such that a ratio of the proportion of heat generated by the reducer and the turbomachine and dissipated to the air is 65% from the maximum take-off thrust speed (PMD) of the core shaft at an environmental temperature of ISA +40 °C to the proportion of heat generated by the gearbox and the turbomachine and dissipated in the air at 65% of the maximum takeoff thrust speed (PMD) of core shaft at an environment temperature of ISA +10 °C is in the range of 1.10 to 1.25, preferably in the range of from 1.10 to 1.22, more preferably in the range of 1.11 to 1.20.

The heat management system can be configured to provide the first amount of heat and the second amount of heat such that a proportion of heat generated by the reducer and the turbomachine and dissipated into the air at 85% of a core shaft maximum takeoff thrust speed (PMD) defined as

_85%PMD

is in the range of 0.25 to 0.70, preferably in the range of 0.35 to 0.70, more preferably in the range of 0.45 to 0.70, even more preferably in the range range from 0.50 to 0.70.

The heat management system can be configured to provide at an environment temperature of ISA +40°C the first amount of heat and the second amount of heat such that the proportion of heat generated by the reducer and the turbomachine and dissipated in air at 85% of the core shaft maximum takeoff thrust (PMD) speed is in the range of 0.55 to 0.70, preferably in the range of 0.60 at 0.70.

The heat management system can be configured to provide at an environment temperature of ISA +10°C the first amount of heat and the second amount of heat such that the proportion of heat generated by the reducer and the turbomachine and dissipated in air at 85% of the core shaft maximum takeoff thrust (PMD) speed in the range of 0.40 to 0.60, preferably in the range of 0.45 to 0 .55.

The heat management system may be configured to provide the first amount of heat and the second amount of heat such that a ratio of the proportion of heat generated by the gearbox and the turbomachine and dissipated to air is 85%. from the maximum take-off thrust speed (PMD) of the core shaft at an environmental temperature of ISA +40 °C to the proportion of heat generated by the gearbox and the turbomachine and dissipated in the air at 85% of the maximum takeoff thrust speed (PMD) of core shaft at an environment temperature of ISA + 10 °C is in the range of 1.20 to 1.42, preferably in the range of from 1.22 to 1.41, more preferably in the range of 1.25 to 1.40.

The heat management system can be configured to provide at an environment temperature of ISA +10°C the first amount of heat and the second amount of heat such that a ratio of the proportion of heat generated by the speed of the reducer and the turbomachine and dissipated in the air at 85% of the maximum take-off thrust speed (PMD) of core shaft to the proportion of heat generated by the reducer and the turbomachine and dissipated in the air at 65% of the core shaft maximum takeoff thrust speed (PMD) is in the range from 0.45 to 0.65, preferably in the range from 0.45 to 0.60, more preferably in the range of 0.47 to 0.58.

The heat management system may be configured to provide the first amount of heat and the second amount of heat such that a proportion of heat generated by the reducer and the turbomachine and dissipated into the air is

greater than A NH + B, and less than the smallest of 1 and C NH + D,

in which A is equal to -1.15, B is equal to or greater than 1.48, C is equal to -1.84, D is in the range from 2.10 to 2.30, preferably in the range range of 2.18 to 2.30, more preferably in the range of 2.18 to 2.25; even more preferably in the range from 2.20 to 2.25; and NH is the core shaft speed expressed as a proportion of the maximum core shaft takeoff thrust speed and is in the range 0.65 to 1. B may be equal to or greater than 1 .5, or 1.52, or 1.54, or 1.56.

NH can be in the range of 0.65 to 0.85.

The pipe assembly may include a first lubricant circuit adapted to provide a first flow of lubricant and a second lubricant circuit adapted to provide a second flow of lubricant, the at least one air-lubricant heat exchanger being arranged in the first lubricant circuit and the at least one fuel-lubricant heat exchanger being arranged in the second lubricant circuit.

The first lubricant circuit can provide lubrication and cooling to the power reducer.

The second lubricant circuit can provide lubrication and cooling to the turbomachine bearings.

According to one aspect, a method of operating a gas turbine engine for an aircraft is provided, the method comprising providing a gas turbine engine comprising: an engine core comprising a compressor, a combustion chamber, 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, the fan being configured to have a fan rotation speed at PMD conditions in the range of 1,500 rpm to 2,800 rpm; turbomachine bearings; a power reducer adapted to drive the fan at a rotation speed lower than the turbine; and a heat management system configured to provide lubrication and cooling to the gearbox and the turbomachine bearings, and comprising a pipe assembly adapted to provide a flow of lubricant to the gearbox and the turbomachine bearings, at least one air heat exchanger -lubricant for dissipating a first quantity of heat towards a first heat sink, and at least one fuel-lubricant heat exchanger for dissipating a second quantity of heat towards a second heat sink; wherein the first heat sink is air and the second heat sink is fuel; in which a proportion of heat generated by the gearbox and the turbomachine and dissipated into the air at 65% of a core shaft maximum takeoff thrust speed (PMD) is defined as

_65%PMD

wherein the method comprises the step of operating the heat management system to provide the proportion of heat generated by the gearbox and the turbomachine and dissipated to the air at 65% of the maximum takeoff thrust speed (PMD ) of heart tree in the range 0.6 to 1.

The method may include the step of operating the heat management system to provide at an environment temperature of ISA +40°C the first amount of heat and the second amount of heat such that the proportion of heat generated by the gearbox and the turbomachine and dissipated into the air at 65% of the core shaft maximum takeoff thrust speed (PMD) is in the range from 0.85 to 1, preferably in the range from 0.90 to 1

The method may include the step of operating the heat management system to provide at an environment temperature of ISA +10°C the first amount of heat and the second amount of heat such that the proportion of heat generated by the gearbox and turbomachine and dissipated to air at 65% of the core shaft maximum takeoff thrust (PMD) speed in the range 0.80 to 0.92

The method may include the step of operating the heat management system to provide the first amount of heat and the second amount of heat such that a ratio of the proportion of heat generated by the gearbox and the turbomachine and dissipated in the air at 65% of the maximum take-off thrust speed (PMD) of the core shaft at an environmental temperature of ISA +40 °C to the proportion of heat generated by the gearbox and the turbomachine and dissipated in air at 65% of the core shaft maximum takeoff thrust (PMD) speed at an environmental temperature of ISA -69 °C is in the range 1.1 to 2, 1, preferably in the range of 1.2 to 2.1, more preferably in the range of 1.4 to 2.0.

In one aspect, there is provided a gas turbine engine for an aircraft comprising: an engine core including a compressor, a combustion chamber, 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; turbomachine bearings; a power reducer adapted to drive the fan at a rotation speed lower than the turbine; and a heat management system configured to provide lubrication and cooling to the gearbox and the turbomachine bearings, and comprising a pipe assembly adapted to provide a flow of lubricant to the gearbox and the turbomachine bearings, at least one air heat exchanger -lubricant for dissipating a first quantity of heat towards a first heat sink, and at least one fuel-lubricant heat exchanger for dissipating a second quantity of heat towards a second heat sink; wherein the first heat sink is air and the second heat sink is fuel, wherein the heat management system is configured to provide at an environment temperature of ISA +40°C the first amount of heat and the second quantity of heat such that:

a first proportion of the heat generated by the reducer and the turbomachine and dissipated towards the air defined as

_85%PMD

at 85% of maximum core shaft takeoff thrust speed is in the range 0.55 to 0.70; And

a second proportion of heat generated by the reducer and the turbomachine and dissipated in the air defined as

_65%PMD

at 65% of the core shaft maximum takeoff thrust (PMD) speed is in the range 0.85 to 1.

As used here, proportions correspond to percentages and vice versa, and so, for example, a proportion of 0.5 corresponds to 50% and a proportion of 1 corresponds to 100%.

In other words, at 85% of maximum thrust speed at core shaft takeoff and at an environment temperature of ISA +40°C 55% to 70% of the heat generated by the gearbox power and the turbomachine is dissipated in the first heat sink (the remainder of the heat generated by the power reducer and the turbomachine being dissipated in the second heat sink), and at 65% of the maximum thrust speed at takeoff core shaft and at an environment temperature of ISA +40 °C from 85% to 100% of the heat generated by the power reducer and the turbomachine is dissipated in the first heat sink (the rest of the heat generated by the power reducer and the turbomachine being dissipated in the second heat sink).

The present inventor understands that providing a heat management system configured to provide, at the highest temperature for which an engine is certified, specific proportions of heat dissipated to air at 65% of the engine speed maximum thrust at core shaft takeoff, i.e. a speed at, or close to, idle in flight, and at 85% of the maximum thrust speed at core shaft takeoff, i.e. a speed at, or close to, conditions cruise, ensures adequate lubrication and cooling to the power reducer and turbomachine bearings, minimizes the size and therefore the weight of the heat exchangers, maximizes the beneficial effects of SFC and at the same time avoids thermal degradation of fuel under all operating conditions.

The first proportion of heat generated by the gearbox and the turbomachine and dissipated in the air at 85% of a maximum thrust speed at core shaft takeoff at an environment temperature of ISA +40 °C can be in the range of 0.60 to 0.70, preferably in the range of 0.62 to 0.68.

The second proportion of heat generated by the gearbox and the turbomachine and dissipated into the air at 65% of the maximum core shaft takeoff thrust speed (PMD) at an environment temperature of ISA +40 °C may be in the range of 0.90 to 1, preferably in the range of 0.92 to 1.

The heat management system can be configured to provide at an environment temperature of ISA +10°C the first amount of heat and the second amount of heat such that the first proportion of heat generated by the reducer and the turbomachine and dissipated in the air at 85% of the maximum take-off thrust speed (PMD) of core shaft in the range of 0.40 to 0.60, preferably in the range of 0.45 to 0.55.

The heat management system can be configured to provide at an 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 reducer and the turbomachine and dissipated in the air at 65% of the maximum take-off thrust speed (PMD) of core shaft is in the range from 0.80 to 0.92.

The heat management system may be 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 gearbox and the turbomachine and dissipated into the air to 85% of the maximum thrust speed at core shaft takeoff at an environmental temperature of ISA +40 °C at the first proportion of the heat generated by the gearbox and the turbomachine and dissipated in the air at 85 % of a maximum core shaft takeoff thrust speed at an environment temperature of ISA +10°C is in the range of 1.20 to 1.42, preferably in the range of 1.22 to 1.41, more preferably in the range of 1.25 to 1.40.

The heat management system may be 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 gearbox and the turbomachine and dissipated into the air to 85% of the maximum thrust speed at core shaft takeoff at an environmental temperature of ISA +40 °C at the first proportion of the heat generated by the gearbox and the turbomachine and dissipated in the air at 85 % of a maximum core shaft takeoff thrust speed at an environment temperature of ISA -69°C is in the range of 1.5 to 4.5, preferably in the range of 2.0 to 4.0, more preferably in the range of 2.0 to 3.5.

The heat management system may be 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 gearbox and the turbomachine and dissipated into the air to 65% of the maximum thrust speed at core shaft takeoff at an environmental temperature of ISA +40 °C to the second proportion of the heat generated by the gearbox and the turbomachine and dissipated in the air at 65 % of the maximum core shaft takeoff thrust speed at an environment temperature of ISA +10 °C is in the range of 1.10 to 1.25, preferably in the range of 1 .10 to 1.22, more preferably in the range of 1.11 to 1.20.

The heat management system may be 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 gearbox and the turbomachine and dissipated into the air to 65% of the maximum thrust speed at core shaft takeoff at an environmental temperature of ISA +40 °C to the second proportion of the heat generated by the gearbox and the turbomachine and dissipated in the air at 65 % of the maximum core shaft takeoff thrust speed at an environment temperature of ISA -69 °C is in the range of 1.1 to 2.1, preferably in the range of 1 .2 to 2.1, more preferably in the range 1.4 to 2.0.

The heat management system may be configured to provide at an 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 in the range of 0.45 to 0.65, preferably in the range of 0.45 to 0.60, more preferably in the range of 0.47 to 0.58.

The heat management system can be configured to provide at an environment temperature of ISA -69°C a ratio of the first proportion of the heat generated by the reducer and the turbomachine and dissipate to air at 85% of the maximum thrust speed at core shaft takeoff to the second proportion of the heat generated by the gearbox and the turbomachine and dissipated in the air at 65% of the maximum thrust speed at core shaft takeoff in the range of 0.30 to 0.55, preferably in the range of 0.35 to 0.45.

The combustion chamber may be a lean burn combustion chamber.

The pipe assembly may include a first lubricant circuit adapted to provide a first flow of lubricant and a second lubricant circuit adapted to provide a second flow of lubricant, the at least one air-lubricant heat exchanger being arranged in the first lubricant circuit and the at least one fuel-lubricant heat exchanger being arranged in the second lubricant circuit.

The heat management system may include a modulation device adapted to adjust a lubricant flow distribution between the power reducer and the turbomachine bearings.

The blower can have a blower diameter in the range of 210 cm to 380 cm.

The power reducer can have a gear ratio in the range of 2.9 to 4.0, or 3.0 to 3.8.

According to one aspect, a method of operating a gas turbine engine for an aircraft is provided, the method comprising providing a gas turbine engine comprising: an engine core comprising a compressor, a combustion chamber, 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; turbomachine bearings; a power reducer adapted to drive the fan at a rotation speed lower than the turbine; and - a heat management system configured to provide lubrication and cooling to the gearbox and the turbomachine bearings, and comprising a pipe assembly adapted - to provide a flow of lubricant to the gearbox and the turbomachine bearings, at least one heat exchanger air-lubricant heat 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;

in which a first proportion of the heat generated by the reducer and the turbomachine and dissipated in the air is defined as

_85%PMD

at 85% of maximum thrust speed at core shaft takeoff; and a second proportion of heat generated by the reducer and the turbomachine and dissipated in the air is defined as

_65%PMD

at 65% of core shaft maximum takeoff thrust speed (PMD).

wherein the method comprises the step of operating the heat management system to provide the first amount of heat and the second amount of heat such that at an environmental temperature of ISA +40°C the The first proportion is in the range of 0.55 to 0.70, and the second proportion is in the range of 0.85 to 1.

At an environment temperature of ISA +40°C the first proportion may be in the range of 0.60 to 0.70, preferably in the range of 0.62 to 0.68.

At an environmental temperature of ISA +40°C the second proportion may be in the range from 0.90 to 1, preferably in the range from 0.92 to 1.

The method may include the step of operating the heat management system 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 reducer and the turbomachine and dissipated in the air at 85% of the maximum thrust speed at core shaft takeoff at an environment temperature of ISA +40 °C to the first proportion of the heat generated by the reducer and the turbomachine and dissipated in air at 85% of a maximum core shaft takeoff thrust velocity at an environmental temperature of ISA +10°C is in the range of 1.20 to 1.42 , preferably in the range of 1.22 to 1.41, more preferably in the range of 1.25 to 1.40.

In one aspect, there is provided a gas turbine engine for an aircraft comprising: an engine core including a compressor, a combustion chamber, 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; turbomachine bearings; a power reducer adapted to drive the fan at a rotation speed lower than the turbine; and a heat management system configured to provide lubrication and cooling to the gearbox and the turbomachine bearings, and comprising a pipe assembly adapted to provide a flow of lubricant to the gearbox and the turbomachine bearings, at least one air heat exchanger -lubricant for dissipating a first quantity of heat towards a first heat sink, and at least one fuel-lubricant heat exchanger for dissipating a second quantity of heat towards a second heat sink; wherein the first heat sink is air and the second heat sink is fuel; in which a first proportion of the heat generated by the reducer and the turbomachine and dissipated in the air is defined as

_85%PMD

at 85% of maximum thrust speed at core shaft takeoff,

in which a second proportion of the heat generated by the reducer and the turbomachine and dissipated in the air is defined as

_65%PMD

at 65% of maximum thrust speed at core shaft takeoff; And

in which the heat management system is configured to provide a ratio of the first proportion to the second proportion in the range from 0.30 to 0.55, preferably in the range from 0.35 to 0.45, at an environmental temperature of ISA -69 °C.

As used here, proportions correspond to percentages and vice versa, and so, for example, a proportion of 0.5 corresponds to 50% and a proportion of 1 corresponds to 100%.

The present inventor understands that providing a heat management system configured to provide, at the coldest temperature for which an engine is certified, specific proportions of heat dissipated to air at 65% of engine speed. maximum thrust at core shaft takeoff, i.e. a speed at, or close to, idle in flight, and at 85% of the maximum thrust speed at core shaft takeoff, i.e. a speed at, or close to, conditions cruise, ensures adequate lubrication and cooling to the power reducer and turbomachine bearings, minimizes the size and therefore the weight of the heat exchangers, maximizes the beneficial effects of SFC and at the same time avoids thermal degradation of fuel under all operating conditions.

The first proportion of the heat generated by the gearbox and the turbomachine and dissipated into the air at 85% of the maximum core shaft takeoff thrust speed at an environment temperature of ISA -69 °C can be in the range from 0.20 to 0.40.

The first proportion of the heat generated by the gearbox and the turbomachine and dissipated into the air at 85% of the maximum core shaft takeoff thrust speed at an environment temperature of ISA -69 °C can be in the range of 0.20 to 0.35, preferably in the range of 0.20 to 0.30.

The second proportion of the heat generated by the gearbox and the turbomachine and dissipated into the air at 65% of the maximum core shaft takeoff thrust speed at an environmental temperature of ISA -69 °C can be in the range from 0.50 to 0.70.

The second proportion of the heat generated by the gearbox and the turbomachine and dissipated into the air at 65% of the maximum core shaft takeoff thrust speed at an environmental temperature of ISA -69 °C can be in the range of 0.55 to 0.70, preferably in the range of 0.55 to 0.67.

The heat management system may be 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 gearbox and the turbomachine and dissipated into the air to 85% of the maximum thrust speed at core shaft takeoff at an environmental temperature of ISA +40 °C at the first proportion of the heat generated by the gearbox and the turbomachine and dissipated in the air at 85 % of the maximum core shaft takeoff thrust speed at an environment temperature of ISA -69°C is in the range 1.5 to 4.5.

The ratio of the first proportion of the heat generated by the gearbox and the turbomachine and dissipated into the air at 85% of the maximum thrust speed at core shaft takeoff at an environment temperature of ISA +40° C to the first proportion of the heat generated by the gearbox and the turbomachine and dissipated into the air at 85% of the maximum thrust speed at core shaft takeoff at an environmental temperature of ISA -69 °C may be in the range of 2.0 to 4.0, preferably in the range of 2.0 to 3.5.

The heat management system can be configured to provide the second proportion of the heat generated by the gearbox and the turbomachine and dissipated to the air at 65% of the maximum thrust speed at core shaft takeoff at a temperature of ISA environment +40 °C at the second proportion of the heat generated by the gearbox and the turbomachine and dissipated in the air at 65% of the maximum thrust speed at core shaft takeoff at a temperature of ISA environment -69°C is in the range of 1.1 to 2.1.

The heat management system can be configured to provide the second proportion of the heat generated by the gearbox and the turbomachine and dissipated to the air at 65% of the maximum thrust speed at core shaft takeoff at a temperature of ISA environment +40 °C at the second proportion of the heat generated by the gearbox and the turbomachine and dissipated in the air at 65% of the maximum thrust speed at core shaft takeoff at a temperature of ISA environment -69°C is in the range of 1.2 to 2.1, preferably in the range of 1.4 to 2.0.

The heat management system can be configured to provide at an environment temperature of ISA +40°C the first amount of heat and the second amount of heat such that the first proportion of heat generated by the reducer and the turbomachine and dissipated in the air at 85% of the maximum thrust speed at core shaft takeoff is in the range of 0.55 to 0.75, preferably in the range of 0.55 to 0 .70, more preferably in the range of 0.60 to 0.70.

The heat management system can be configured to provide at an environment temperature of ISA +40°C the first amount of heat and the second amount of heat such that the second proportion of heat generated by the reducer and the turbomachine and dissipated in the air at 65% of the maximum core shaft takeoff thrust speed is in the range of 0.85 to 1, preferably in the range of 0.90 to 1.

The heat management system may be configured to provide the first amount of heat and the second amount of heat such that a proportion of the heat generated by the reducer and the turbomachine and dissipated into the air is

greater than A NH + B, and less than the smallest of 1 and C NH + D,

in which A is equal to -1.15, B is equal to or greater than 1.48, C is equal to -1.84, D is in the range of 2.10 to 2.30, and NH is the core shaft speed expressed as a proportion of the maximum core shaft takeoff thrust velocity and is in the range of 0.65 to 1, preferably in the range of 0.65 to 0.85.

B can be equal to or greater than 1.5, or 1.52, or 1.54, or 1.56.

D may be in the range of 2.18 to 2.30, preferably in the range of 2.18 to 2.25, more preferably in the range of 2.20 to 2.25.

The maximum core shaft takeoff thrust speed may be in the range of 5500 rpm to 9500 rpm, preferably in the range of 5500 rpm to 8500 rpm, preferably in the range of 5500 rpm to 8500 rpm. range from 5500 rpm to 7500 rpm.

The combustion chamber may be a lean burn combustion chamber.

The power reducer can have a gear ratio in the range of 2.9 to 4.0, or 3.0 to 3.8.

According to one aspect, a method of operating a gas turbine engine for an aircraft is provided, the method comprising providing a gas turbine engine comprising: an engine core comprising a compressor, a combustion chamber, 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, the fan being configured to have a fan rotation speed at PMD conditions in the range of 1500 rpm to 2800 rpm; turbomachine bearings; a power reducer adapted to drive the fan at a rotation speed lower than the turbine; and a heat management system configured to provide lubrication and cooling to the gearbox and the turbomachine bearings, and comprising a pipe assembly adapted to provide a flow of lubricant to the gearbox and the turbomachine bearings, at least one air heat exchanger -lubricant for dissipating a first quantity of heat towards a first heat sink, and at least one fuel-lubricant heat exchanger for dissipating a second quantity of heat towards a second heat sink; wherein the first heat sink is air and the second heat sink is fuel;

in which a first proportion of the heat generated by the reducer and the turbomachine and dissipated in the air is defined as

_85%PMD

at 85% of maximum thrust speed at core shaft takeoff, and

a second proportion of the heat generated by the reducer and the turbomachine and dissipated in the air is defined as

_65%PMD

at 65% of the maximum thrust speed at core shaft takeoff,

wherein the method comprises the step of operating the heat management system to provide the first amount of heat and the second amount of heat such that at an environmental temperature of ISA -69°C a ratio of the first proportion to the second proportion is in the range of 0.30 to 0.55, preferably in the range of 0.35 to 0.45.

The method may include the step of operating the heat management system to provide the first amount of heat and the second amount of heat such that at an environmental temperature of ISA -69°C the first proportion of the heat generated by the gearbox and the turbomachine and dissipated into the air at 85% of the maximum thrust speed at core shaft takeoff is in the range of 0.20 to 0.40, preferably in the range of 0.20 to 0.35, more preferably in the range of 0.20 to 0.30.

The method may include the step of operating the heat management system to provide the first amount of heat and the second amount of heat such that at an environmental temperature of ISA -69°C the second proportion of the heat generated by the gearbox and the turbomachine and dissipated into the air at 65% of the maximum thrust speed at core shaft takeoff is in the range of 0.50 to 0.70, preferably in the range of 0.55 to 0.70, more preferably in the range of 0.55 to 0.67.

The method may include the step of operating the heat management system 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 reducer and the turbomachine and dissipated in the air at 85% of the maximum thrust speed at core shaft takeoff at an environment temperature of ISA +40 °C to the first proportion of the heat generated by the reducer and the turbomachine and dissipated in air at 85% of the maximum core shaft takeoff thrust speed at an environmental temperature of ISA -69 °C is in the range of 1.5 to 4.5, preferably in the range of 2.0 to 4.0, more preferably in the range of 2.0 to 3.5.

In one aspect, there is provided a gas turbine engine for an aircraft comprising: an engine core including a compressor, a combustion chamber, 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; turbomachine bearings; a power reducer adapted to drive the fan at a rotation speed lower than the turbine; and a heat management system configured to provide lubrication and cooling to the gearbox and the turbomachine bearings, and comprising a pipe assembly adapted to provide a flow of lubricant to the gearbox and the turbomachine bearings, at least one air heat exchanger -lubricant for dissipating a first quantity of heat towards a first heat sink, and at least one fuel-lubricant heat exchanger for dissipating a second quantity of heat towards a second heat sink; wherein the first heat sink is air and the second heat sink is fuel; in which a first proportion of heat generated by the reducer and the turbomachine and dissipated in the air is defined as

_85%PMD

at 85% of maximum thrust speed at core shaft takeoff; in which a second proportion of heat generated by the reducer and the turbomachine and dissipated in the air is defined as

_65%PMD

at 65% of maximum thrust speed at core shaft takeoff; and 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 at an environment temperature of ISA +40°C to the first proportion at an environment temperature of ISA -69 °C is in the range of 1.5 to 4.5, and a ratio of the second proportion at an environment temperature of ISA +40 °C to the second proportion at an environment temperature ISA -69°C is in the range 1.1 to 2.1.

As used here, proportions correspond to percentages and vice versa, and so, for example, a proportion of 0.5 corresponds to 50% and a proportion of 1 corresponds to 100%.

The present inventor understands that providing a heat management system configured to provide, at the highest and lowest temperatures for which an engine is certified, specific proportions of heat dissipated to air at 65% of the maximum thrust speed at core shaft takeoff, i.e. a speed at, or close to, idle in flight, and at 85% of the maximum thrust speed at core shaft takeoff, i.e. a speed at, or close to, of, cruising conditions, makes it possible to ensure adequate lubrication and cooling to the power reducer and to the turbomachine bearings, minimize the size and therefore the weight of the heat exchangers, maximize the beneficial effects of SFC and at the same time avoid the thermal degradation of fuel under all operating conditions.

The ratio of the first proportion at an environment temperature of ISA +40°C to the first proportion at an environment temperature of ISA -69°C can be in the range of 2.0 to 4.0 , preferably in the range 2.0 to 3.5.

The ratio of the second proportion at an environment temperature of ISA +40°C to the second proportion at an environment temperature of ISA -69°C may be in the range of 1.2 to 2.1 , preferably in the range 1.4 to 2.0.

The heat management system may be configured to provide the first amount of heat and the second amount of heat such that the first proportion at an environment temperature of ISA +40°C is in the range of 0 .55 to 0.70, preferably in the range of 0.60 to 0.70.

The heat management system may be configured to provide the first amount of heat and the second amount of heat such that the first proportion at an environment temperature of ISA -69°C is in the range of 0 .20 to 0.40, preferably in the range 0.20 to 0.35.

The heat management system may be configured to provide the first amount of heat and the second amount of heat such that the second proportion at an environment temperature of ISA +40°C in the range of 0.85 to 1, preferably in the range from 0.90 to 1.

The heat management system may be configured to provide the first amount of heat and the second amount of heat such that the second proportion at an environment temperature of ISA -69°C is in the range of 0 .50 to 0.70, preferably in the range of 0.55 to 0.70, more preferably in the range of 0.55 to 0.67.

The heat management system can be configured to provide at an environment temperature of ISA +10°C the first amount of heat and the second amount of heat such that the first proportion is in the range of 0 .40 to 0.60, preferably in the range of 0.45 to 0.55.

The heat management system can be configured to provide at an environment temperature of ISA +10°C the first amount of heat and the second amount of heat such that the second proportion is in the range of 0 .80 to .92.

The heat management system may be configured to provide the first amount of heat and the second amount of heat such that a ratio of the first proportion at an environment temperature of ISA +40°C to the first proportion at an environmental temperature of ISA +10 °C is in the range 1.20 to 1.42.

The ratio of the first proportion at an environment temperature of ISA +40°C to the first proportion at an environment temperature of ISA +10°C can be in the range of 1.22 to 1.41 , preferably in the range from 1.25 to 1.40.

The ratio of the second proportion at an environment temperature of ISA +40°C to the second proportion at an environment temperature of ISA +10°C may be in the range of 1.10 to 1.25 .

The ratio of the second proportion at an environment temperature of ISA +40°C to the second proportion at an environment temperature of ISA +10°C may be in the range of 1.10 to 1.22 , preferably in the range from 1.11 to 1.20.

The heat management system may include a modulation device adapted to adjust a lubricant flow distribution between the power reducer and the turbomachine bearings.

The pipe assembly may include a first lubricant circuit adapted to provide a first flow of lubricant and a second lubricant circuit adapted to provide a second flow of lubricant, the at least one air-lubricant heat exchanger being arranged in the first lubricant circuit and the at least one fuel-lubricant heat exchanger being arranged in the second lubricant circuit.

The heat management system may include a lubricant reservoir in fluid communication with, and supplying lubricant to, the first and second lubricant circuits.

According to one aspect, a method of operating a gas turbine engine for an aircraft is provided, the method comprising providing a gas turbine engine comprising: an engine core comprising a compressor, a combustion chamber, 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; turbomachine bearings; a power reducer adapted to drive the fan at a rotation speed lower than the turbine; and a heat management system configured to provide lubrication and cooling to the gearbox and the turbomachine bearings, and comprising a pipe assembly adapted to provide a flow of lubricant to the gearbox and the turbomachine bearings, at least one air heat exchanger -lubricant for dissipating a first quantity of heat towards a first heat sink, and at least one fuel-lubricant heat exchanger for dissipating a second quantity of heat towards a second heat sink; wherein the first heat sink is air and the second heat sink is fuel; in which a first proportion of heat generated by the reducer and the turbomachine and dissipated in the air is defined as

_85%PMD

at 85% of maximum thrust speed at core shaft takeoff; in which a second proportion of heat generated by the reducer and the turbomachine and dissipated in the air is defined as

_65%PMD

at 65% of maximum thrust speed at core shaft takeoff; and wherein the method comprises the step of operating the heat management system to provide the first amount of heat and the second amount of heat such that a ratio of the first proportion to an environmental temperature d 'ISA +40 °C at the first proportion to an environment temperature of ISA -69 °C is in the range of 1.5 to 4.5, preferably in the range of 2.0 to 4 .0, more preferably in the range of 2.0 to 3.5 and a ratio of the second proportion at an environment temperature of ISA +40°C to the second proportion at an environment temperature of ISA -69°C is in the range of 1.1 to 2.1, preferably in the range of 1.2 to 2.1, more preferably in the range of 1.4 to 2.0.

The method may include the step of operating the heat management system to provide the first amount of heat and the second amount of heat such that the first proportion at an environment temperature of ISA +40°C is in the range of 0.55 to 0.70, preferably in the range of 0.60 to 0.70.

The method may include the step of operating the heat management system to provide the first amount of heat and the second amount of heat such that the first proportion at an environment temperature of ISA -69°C is within the range from 0.20 to 0.40, preferably in the range from 0.20 to 0.35.

The method may include the step of operating the heat management system to provide the first amount of heat and the second amount of heat such that the second proportion at an environment temperature of ISA +40°C in the range of 0.85 to 1, preferably in the range of 0.90 to 1, and/or the second proportion at an environmental temperature of ISA -69 °C is in the range of 0.50 to 0.70, preferably in the range of 0.55 to 0.70, more preferably in the range of 0.55 to 0.67.

In one aspect, there is provided a gas turbine engine for an aircraft comprising: an engine core including a compressor, a combustion chamber, 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; turbomachine bearings; a power reducer adapted to drive the fan at a rotation speed lower than the turbine; and a heat management system configured to provide lubrication and cooling to the gearbox and the turbomachine bearings, and comprising a pipe assembly adapted to provide a flow of lubricant to the gearbox and the turbomachine bearings, at least one air heat exchanger -lubricant for dissipating a first amount of heat to a first heat sink, and at least one fuel-lubricant heat exchanger for dissipating a second amount of heat to a second heat sink, wherein the first heat sink is air and the second heat sink is fuel; wherein the heat management system is configured to provide the first quantity of heat and the second quantity of heat such that a proportion of heat generated by the reducer and the turbomachine and dissipated into the air is

greater than A NH + B and less than 1,

in which A is equal to -1.15, B is equal to or greater than 1.48, and NH is the core shaft speed expressed as a proportion of the maximum core shaft takeoff thrust speed and is in the range of 0.65 to 1, preferably in the range of 0.65 to 0.90.

As used here, proportions correspond to percentages and vice versa, and so, for example, a proportion of 0.5 corresponds to 50% and a proportion of 1 corresponds to 100%.

In other words, when NH is 0.65 the core shaft speed is 65% of the maximum core shaft takeoff thrust speed, and when NH is 1 the core shaft speed is of 100%, or equal to, the maximum thrust speed at core shaft takeoff.

The present inventor has understood that providing a heat management system configured to provide specific proportions of heat dissipated to air in the range of 65% to 100% of maximum thrust speed at shaft takeoff of heart, i.e. substantially over the entire operational range of the engine, ensures adequate lubrication and cooling to the power reducer and the turbomachine bearings, minimizes the size and therefore the weight of the heat exchangers, maximizes the beneficial effects of SFC and at the same time avoid thermal degradation of fuel under all operating conditions.

B can be equal to or greater than 1.5, or 1.52, or 1.54, or 1.56.

The heat management system may be configured to provide the first amount of heat and the second amount of heat such that the proportion of heat generated by the reducer and the turbomachine and dissipated to the air is

less than the smallest among 1 and C NH + D,

wherein C is -1.84, D is in the range of 2.10 to 2.30, and NH is in the range of 0.65 to 1, preferably in the range of 0.65 to 0.90.

D may be in the range of 2.18 to 2.30, preferably in the range of 2.18 to 2.25, more preferably in the range of 2.20 to 2.25.

The combustion chamber may be a lean burn combustion chamber. The lean burn combustion chamber may include a plurality of lean burn fuel spray nozzles, each fuel spray nozzle including a pilot fuel injector and a main fuel injector.

NH may be in the range of 0.65 to 0.85, preferably in the range of 0.65 to 0.80, more preferably in the range of 0.65 to 0.75.

The heat management system may be configured to provide the first amount of heat and the second amount of heat such that a first proportion of heat generated by the reducer and the turbomachine and dissipated into the air at 85% of the Maximum core shaft takeoff thrust speed is in the range of 0.50 to 0.70, preferably in the range of 0.50 to 0.65.

The heat management system may be configured to provide the first amount of heat and the second amount of heat such that a second proportion of heat generated by the reducer and the turbomachine and dissipated into the air at 65% of the Maximum core shaft takeoff thrust speed is in the range of 0.75 to 1, preferably in the range of 0.80 to 1.

The heat management system may be configured to provide the first amount of heat and the second amount of heat such that at an environmental temperature of ISA +40°C a first proportion of heat generated by the reducer and the turbomachine and dissipated in the air at 85% of the maximum thrust speed at core shaft takeoff is in the range from 0.55 to 0.70, preferably in the range from 0.60 to 0.70.

The heat management system may be configured to provide the first amount of heat and the second amount of heat such that at an environmental temperature of ISA +40°C a second proportion of heat generated by the reducer and the turbomachine and dissipated in the air at 65% of the maximum thrust speed at core shaft takeoff is in the range from 0.92 to 1.

The heat management system may be configured to provide the first amount of heat and the second amount of heat such that at an environmental temperature of ISA +10°C a second proportion of heat generated by the reducer and the turbomachine and dissipated in the air at 65% of the maximum thrust speed at core shaft takeoff is in the range from 0.80 to 0.92.

The heat management system may be configured to provide the first amount of heat and the second amount of heat such that a ratio of a second proportion of heat generated by the gearbox and the turbomachine and dissipated into the air at 65% of the maximum thrust speed at core shaft takeoff at an environmental temperature of ISA +40 °C at the second proportion of heat generated by the gearbox and the turbomachine and dissipated to air at 65% of the maximum core shaft takeoff thrust speed at an environment temperature of ISA +10 °C is in the range of 1.10 to 1.25, preferably in the range of 1, 10 to 1.22, more preferably in the range of 1.11 to 1.20.

The air from the first heat sink may be bypass air.

A flow restriction valve may be arranged downstream of the at least one air-lubricant heat exchanger to vary a mass flow rate of cooling air through the at least one air-lubricant heat exchanger, thereby which allows the first quantity of heat to be varied.

The pipe assembly may include a first lubricant circuit adapted to provide a first flow of lubricant and a second lubricant circuit adapted to provide a second flow of lubricant, the at least one air-lubricant heat exchanger being arranged in the first lubricant circuit and the at least one fuel-lubricant heat exchanger being arranged in the second lubricant circuit.

The blower can have a blower diameter in the range of 210 cm to 380 cm or 220 cm to 360 cm.

The power reducer can have a gear ratio in the range of 2.9 to 4.0, or 3.0 to 3.8.

According to one aspect, a method of operating a gas turbine engine for an aircraft is provided, the method comprising providing a gas turbine engine comprising: an engine core comprising a compressor, a combustion chamber, 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; turbomachine bearings; a power reducer adapted to drive the fan at a rotation speed lower than the turbine; and a heat management system configured to provide lubrication and cooling to the gearbox and the turbomachine bearings, and comprising a pipe assembly adapted to provide a flow of lubricant to the gearbox and the turbomachine bearings, at least one air heat exchanger -lubricant for dissipating a first quantity of heat towards a first heat sink, and at least one fuel-lubricant heat exchanger for dissipating a second quantity of heat towards a second heat sink; wherein the method comprises the step of operating the heat management system to provide the first amount of heat and the second amount of heat such that a proportion of heat generated by the gearbox and the turbomachine and dissipated in the air is

greater than A NH + B, and less than the smallest of 1 and C NH + D,

wherein A is -1.15, B is 1.48 or greater, C is -1.84, D is in the range of 2.10 to 2.30, preferably in the range ranging from 2.18 to 2.30, more preferably in the range of 2.18 to 2.25, more preferably in the range of 2.20 to 2.25, and NH is the core shaft speed expressed in as a proportion of the maximum core shaft takeoff thrust speed and is in the range 0.65 to 1.

B can be equal to or greater than 1.5, or 1.52, or 1.54, or 1.56.

The method may include the step of operating the heat management system to provide the first amount of heat and the second amount of heat such that a first proportion of heat generated by the reducer and the turbomachine and dissipated in air at 85% of the maximum core shaft takeoff thrust speed is in the range of 0.50 to 0.70, preferably in the range of 0.50 to 0.65.

The method may include the step of operating the heat management system to provide the first amount of heat and the second amount of heat such that a second proportion of heat generated by the reducer and the turbomachine and dissipated in air at 0.65 of the maximum core shaft takeoff thrust speed is in the range of 0.75 to 1, preferably in the range of 0.80 to 1.

The method may include the step of operating the heat management system to provide the first amount of heat and the second amount of heat such that a ratio of the first proportion to the second proportion is within the range ranging from 0.45 to 0.65.

The method may include the step of operating the heat management system to provide at an environment temperature of ISA +10°C the first amount of heat and the second amount of heat such that the ratio of the first proportion to the second proportion is in the range 0.47 to 0.58.

In one aspect, there is provided a gas turbine engine for an aircraft comprising: an engine core including a compressor, a combustion chamber, 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; turbomachine bearings; a power reducer adapted to drive the fan at a rotation speed lower than the turbine; and a heat management system configured to provide lubrication and cooling to the gearbox and the turbomachine bearings, and comprising a pipe assembly adapted to provide a flow of lubricant to the gearbox and the turbomachine bearings, at least one air heat exchanger -lubricant for dissipating a first amount of heat to a first heat sink, and at least one fuel-lubricant heat exchanger for dissipating a second amount of heat to a second heat sink, wherein the first heat sink is air and the second heat sink is fuel; wherein the heat management system is configured to provide the first quantity of heat and the second quantity of heat such that a proportion of heat generated by the reducer and the turbomachine and dissipated into the air is

greater than A NH + B, and less than the smallest of 1 and E (NH - 1) + F,

wherein A is -1.15, B is 1.48 or greater, E is in the range of -1.16 to -3, F is 0.37 or greater, and NH is the core shaft speed expressed as a proportion of the maximum core shaft takeoff thrust speed and is in the range of 0.65 to 1, preferably in the range of 0.65 to 0.95.

As used here, proportions correspond to percentages and vice versa, and so, for example, a proportion of 0.5 corresponds to 50% and a proportion of 1 corresponds to 100%.

In other words, when NH is 0.65 the core shaft speed is 65% of the maximum core shaft takeoff thrust speed, and when NH is 1 the core shaft speed is of 100%, or equal to, the maximum thrust speed at core shaft takeoff.

The present inventor has understood that providing a heat management system configured to provide specific proportions of heat dissipated to air in the range of 65% to 100% of maximum thrust speed at shaft takeoff of heart, i.e. substantially over the entire operational range of the engine, ensures adequate lubrication and cooling to the power reducer and the turbomachine bearings, minimizes the size and therefore the weight of the heat exchangers, maximizes the beneficial effects of SFC and at the same time avoid thermal degradation of fuel under all operating conditions.

B can be equal to or greater than 1.5, or 1.52, or 1.54, or 1.56.

E may be in the range of -1.16 to -2.5, preferably in the range of -1.16 to -1.95.

NH may be in the range of 0.65 to 0.90, preferably in the range of 0.65 to 0.85.

The heat management system may be configured to provide the first amount of heat and the second amount of heat such that a first proportion of heat generated by the reducer and the turbomachine and dissipated into the air at 85% of the Maximum thrust speed at heart shaft takeoff is in the range from 0.50 to 0.70.

The heat management system can be configured to provide at an environment temperature of ISA +40°C the first amount of heat and the second amount of heat such that the first proportion is in the range of 0 .55 to 0.70, preferably in the range of 0.60 to 0.70.

The heat management system may be configured to provide the first amount of heat and the second amount of heat such that a second proportion of heat generated by the reducer and the turbomachine and dissipated to the air at 0.65 of the maximum core shaft takeoff thrust speed is in the range of 0.75 to 1, preferably in the range of 0.80 to 1.

The heat management system can be configured to provide at an environment temperature of ISA +40°C the first amount of heat and the second amount of heat such that the second proportion is in the range of 0 .85 to 1, preferably in the range 0.90 to 1.

The heat management system may be configured to provide a ratio of the first proportion to the second proportion in the range of 0.45 to 0.65, preferably 0.47 to 0.58.

The heat management system can be configured to provide the first quantity of heat and the second quantity of heat such that under cruising conditions a proportion of heat generated by the gearbox and the turbomachine and dissipated in the air is found in the range of 0.56 to 0.75, preferably in the range of 0.56 to 0.70.

The maximum core shaft takeoff thrust speed may be in the range of 5500 rpm to 9500 rpm, preferably in the range of 5500 rpm to 8500 rpm, preferably in the range of 5500 rpm to 8500 rpm. range from 5500 rpm to 7500 rpm.

The fan may have a fan rotation speed at PMD conditions in the range of 1500 rpm to 2800 rpm, preferably in the range of 1600 rpm to 2500 rpm, more preferably in the range of 1600 rpm to 2500 rpm. range from 1600 rpm to 2200 rpm.

The first heat sink may be bypass air, and the at least one air-lubricant heat exchanger may be adapted to receive bypass air from the bypass line.

A flow restriction valve may be arranged downstream of the at least one air-lubricant heat exchanger to vary a mass flow rate of cooling air through the at least one air-lubricant heat exchanger, thereby which allows the first quantity of heat to be varied.

The pipe assembly may include a first lubricant circuit adapted to provide a first flow of lubricant and a second lubricant circuit adapted to provide a second flow of lubricant, the at least one air-lubricant heat exchanger being arranged in the first lubricant circuit and the at least one fuel-lubricant heat exchanger being arranged in the second lubricant circuit.

The first lubricant circuit can provide lubrication and cooling to the power reducer.

The second lubricant circuit can provide lubrication and cooling to the turbomachine bearings.

The gas turbine engine may include at least two air-lubricant heat exchangers for dissipating the first amount of heat in the first heat sink, at least one of which is arranged in the first lubricant circuit and at least one of which is arranged in the first heat sink. second lubricant circuit.

The combustion chamber may be a lean burn combustion chamber. The lean burn combustion chamber may include a plurality of lean burn fuel spray nozzles, each fuel spray nozzle including a pilot fuel injector and a main fuel injector.

According to one aspect, a method of operating a gas turbine engine for an aircraft is provided, the method comprising providing a gas turbine engine comprising: an engine core comprising a compressor, a combustion chamber, 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; turbomachine bearings; a power reducer adapted to drive the fan at a rotation speed lower than the turbine; and a heat management system configured to provide lubrication and cooling to the gearbox and the turbomachine bearings, and comprising a pipe assembly adapted to provide a flow of lubricant to the gearbox and the turbomachine bearings, at least one air heat exchanger -lubricant for dissipating a first amount of heat to a first heat sink, and at least one fuel-lubricant heat exchanger for dissipating a second amount of heat to a second heat sink, wherein the first heat sink is air and the second heat sink is fuel; and wherein the method comprises the step of operating the heat management system to provide the first amount of heat and the second amount of heat such that a proportion of heat generated by the gearbox and the turbomachine and dissipated in the air is

greater than A NH + B, and less than the smallest of 1 and E (NH - 1) + F,

in which A is equal to -1.15, B is equal to or greater than 1.48, E is in the range from -1.16 to -3, preferably in the range from -1.16 to - 2.5, more preferably in the range of -1.16 to -1.95, F is equal to or greater than 0.37, and NH is the core shaft speed expressed as a proportion of the maximum thrust speed at core shaft takeoff and is in the range of 0.65 to 1, preferably in the range of 0.65 to 0.95, more preferably in the range of 0.65 to 0.90.

B can be equal to or greater than 1.5, or 1.52, or 1.54, or 1.56.

The method may include the step of operating the heat management system to provide the first amount of heat and the second amount of heat such that a first proportion of heat generated by the reducer and the turbomachine and dissipated in air at 85% of the maximum thrust speed at core shaft takeoff is in the range 0.50 to 0.70.

The method may include the step of operating the heat management system to provide at an environment temperature of ISA +40°C the first amount of heat and the second amount of heat such that the first proportion is in the range of 0.55 to 0.70, preferably in the range of 0.60 to 0.70.

The method may include the step of operating the heat management system to provide the first amount of heat and the second amount of heat such that a second proportion of heat generated by the reducer and the turbomachine and dissipated in air at 0.65 of the maximum core shaft takeoff thrust speed is in the range of 0.75 to 1, preferably in the range of 0.80 to 1.

In one aspect, there is provided a gas turbine engine for an aircraft comprising: an engine core including a compressor, a combustion chamber, 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; turbomachine bearings; a power reducer adapted to drive the fan at a rotation speed lower than the turbine; and a heat management system configured to provide lubrication and cooling to the gearbox and the turbomachine bearings, and comprising a pipe assembly adapted to provide a flow of lubricant to the gearbox and the turbomachine bearings, at least one air heat exchanger -lubricant for dissipating a first amount of heat to a first heat sink, and at least one fuel-lubricant heat exchanger for dissipating a second amount of heat to a second heat sink, wherein the first heat sink is air and the second heat sink is fuel; in which the heat management system is configured to provide the first quantity of heat and the second quantity of heat such that under cruising conditions a proportion of heat generated by the gearbox and the turbomachine and dissipated in the air is found in the range from 0.35 to 0.80.

The present inventor understands that providing a heat management system configured to provide specific proportions of heat dissipated into the air at cruise conditions, which may be the longest phase of flight, provides lubrication and adequate cooling to the power reducer and turbomachine bearings, minimize the size and therefore the weight of the heat exchangers, maximize the benefits of SFC and at the same time avoid thermal degradation of fuel in all operating conditions.

The heat management system can be configured to provide the first amount of heat and the second amount of heat such that under cruising conditions the proportion of heat generated by the gearbox and the turbomachine and dissipated into the air is found in the range of 0.50 to 0.80, more preferably in the range of 0.57 to 0.80.

The heat management system may be configured to provide the first amount of heat and the second amount of heat such that under cruising conditions and at an environmental temperature of ISA +40°C a proportion of heat generated by the reducer and the turbomachine and dissipated in the air is in the range from 0.65 to 0.80, preferably in the range from 0.70 to 0.80.

The heat management system may be configured to provide the first amount of heat and the second amount of heat such that at cruise conditions and an environmental temperature of ISA +10°C a proportion of heat generated by the reducer and the turbomachine and dissipated in the air is in the range from 0.45 to 0.70, preferably in the range from 0.55 to 0.65.

The heat management system may be configured to provide the first amount of heat and the second amount of heat such that at cruise conditions and an environmental temperature in the range from ISA +10°C to ISA +40 °C a proportion of heat generated by the gearbox and the turbomachine and dissipated in the air is in the range from 0.58 to 0.75.

A first proportion of heat generated by the reducer and the turbomachine and dissipated in the air is defined as

_85%PMD

at 85% of a maximum core shaft takeoff thrust speed, and the heat management system may be configured to provide the first amount of heat and the second amount of heat such that the first proportion of heat generated by the reducer and the turbomachine and dissipated in the air is in the range from 0.35 to 0.70, preferably in the range from 0.45 to 0.70, more preferably in the range from 0.50 to 0.70.

The heat management system is configured to provide at an environment temperature of ISA +10°C the first amount of heat and the second amount of heat such that the first proportion of heat generated by the reducer and the turbomachine and dissipated in the air at 85% of the maximum core shaft takeoff thrust speed is in the range from 0.35 to 0.65, preferably in the range from 0.40 to 0, 60, more preferably in the range of 0.45 to 0.55.

A second proportion of heat generated by the reducer and the turbomachine and dissipated in the air is defined as

_0.65PMD

at 65% of the maximum core shaft takeoff thrust speed, and the heat management system can be configured to provide the first amount of heat and the second amount of heat such that the second proportion of heat generated by the reducer and the turbomachine and dissipated in the air is in the range from 0.60 to 1, preferably in the range from 0.70 to 1, more preferably in the range from 0.75 to 1 .

The heat management system can be configured to provide at an 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 reducer and the turbomachine and dissipated to the air at 0.65 of the maximum thrust speed at heart shaft takeoff is in the range from 0.80 to 0.92.

The heat management system may be configured to provide at an 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 lies in the range of 0.45 to 0.65.

The heat management system may be configured to provide the first amount of heat and the second amount of heat such that a proportion of heat generated by the reducer and the turbomachine and dissipated into the air is

greater than A NH + B, and less than the smallest of 1 and C NH + D,

in which A is equal to -1.15, B is equal to or greater than 1.48, C is equal to -1.84, D is in the range from 2.10 to 2.30, preferably in the range range of 2.18 to 2.30, more preferably in the range of 2.18 to 2.25; even more preferably in the range of 2.20 to 2.25, and NH is the core shaft speed expressed as a proportion of the maximum core shaft takeoff thrust speed and is within the range ranging from 0.65 to 1, preferably in the range from 0.65 to 0.85. B can be equal to or greater than 1.5, or 1.52, or 1.54, or 1.56.

The core shaft may have a rotational speed at cruising conditions in the range of 5000 rpm to 9000 rpm, or in the range of 5000 rpm to 7000 rpm, or in the range ranging from 5000 rpm to 6000 rpm.

The fan may have a fan rotation speed under cruising conditions in the range of 1400 rpm to 2600 rpm, preferably in the range of 1500 rpm to 2300 rpm, more preferably in the range of 1500 rpm to 2300 rpm. the range from 1600 rpm to 2000 rpm.

The pipe assembly may include a first lubricant circuit adapted to provide a first flow of lubricant and a second lubricant circuit adapted to provide a second flow of lubricant, the at least one air-lubricant heat exchanger being arranged in the first lubricant circuit and the at least one fuel-lubricant heat exchanger being arranged in the second lubricant circuit.

The heat management system may include a modulation device adapted to adjust a lubricant flow distribution between the power reducer and the turbomachine bearings.

The gas turbine engine may include at least two air-lubricant heat exchangers for dissipating the first amount of heat in the first heat sink, at least one of which is arranged in the first lubricant circuit and at least one of which is arranged in the first heat sink. second lubricant circuit.

The heat management system may include a lubricant reservoir in fluid communication with, and supplying lubricant to, the first and second lubricant circuits.

According to one aspect, a method of operating a gas turbine engine for an aircraft is provided, the method comprising providing a gas turbine engine comprising: an engine core comprising a compressor, a combustion chamber, 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; turbomachine bearings; a power reducer adapted to drive the fan at a rotation speed lower than the turbine; and a heat management system configured to provide lubrication and cooling to the gearbox and the turbomachine bearings, and comprising a pipe assembly adapted to provide a flow of lubricant to the gearbox and the turbomachine bearings, at least one air heat exchanger -lubricant for dissipating a first amount of heat to a first heat sink, and at least one fuel-lubricant heat exchanger for dissipating a second amount of heat to a second heat sink, wherein the first heat sink is air and the second heat sink is fuel; and wherein the method comprises the step of operating the heat management system to provide the first amount of heat and the second amount of heat such that under cruising conditions a proportion of heat generated by the gearbox and the turbomachine and dissipated in the air is in the range of 0.35 to 0.80, preferably in the range of 0.50 to 0.80, more preferably in the range of 0.57 to 0 .80.

The method may include the step of operating the heat management system to provide the first amount of heat and the second amount of heat such that at cruise conditions and an environment temperature of ISA + 40 °C a proportion of heat generated by the gearbox and the turbomachine and dissipated in the air is in the range from 0.65 to 0.80, preferably in the range from 0.70 to 0.80.

The method may include the step of operating the heat management system to provide the first amount of heat and the second amount of heat such that at cruise conditions and an environment temperature of ISA + 10 °C a proportion of heat generated by the gearbox and the turbomachine and dissipated in the air is in the range from 0.45 to 0.70, preferably in the range from 0.55 to 0.65.

The method may include the step of operating the heat management system to provide the first amount of heat and the second amount of heat such that under cruise conditions and at an environmental temperature in the range from ISA +10 °C to ISA +40 °C a proportion of heat generated by the gearbox and the turbomachine and dissipated in the air is in the range from 0.58 to 0.75.

In any of the aspects described above, one or more of the following characteristics may be present.

The maximum core shaft takeoff thrust speed may be in the range of 5500 rpm to 9500 rpm, preferably in the range of 5500 rpm to 8500 rpm, preferably in the range of 5500 rpm to 8500 rpm. range from 5500 rpm to 7500 rpm, preferably in the range from 5500 rpm to 6500 rpm, preferably in the range from 5800 rpm to 6200 rpm. Thus, 85% of the maximum core shaft takeoff thrust speed can be in the range of 4675 rpm to 8075 rpm, preferably in the range of 4675 rpm to 7225 rpm , preferably in the range from 4675 rpm to 6375 rpm, preferably in the range from 4675 rpm to 5525 rpm, preferably in the range from 4930 rpm to 5270 rpm min ; and 65% of the maximum core shaft takeoff thrust speed may be in the range from 3575 rpm to 6175 rpm, preferably in the range from 3575 rpm to 5525 rpm, preferably in the range from 3575 rpm to 4875 rpm, preferably in the range from 3575 rpm to 4225 rpm, preferably in the range from 3770 rpm to 4030 rpm .

The core shaft may have a rotational speed at cruising conditions in the range of 5000 rpm to 9000 rpm, or in the range of 5000 rpm to 7000 rpm, or in the range ranging from 5000 rpm to 6000 rpm, or in the range from 5200 rpm to 5800 rpm.

The fan may have a fan rotation speed at PMD conditions in the range of 1500 rpm to 2800 rpm, preferably in the range of 1600 rpm to 2500 rpm, more preferably in the range of 1600 rpm to 2500 rpm. range from 1600 rpm to 2200 rpm, even more preferably in the range from 1700 rpm to 1900 rpm.

The fan may have a fan rotation speed under cruising conditions in the range of 1400 rpm to 2600 rpm, preferably in the range of 1500 rpm to 2300 rpm, more preferably in the range of 1500 rpm to 2300 rpm. the range from 1600 rpm to 2000 rpm.

The first heat sink may be bypass air flowing through a bypass pipe of the gas turbine engine, and/or outside air.

A flow restriction valve may be arranged downstream of the at least one air-lubricant heat exchanger to vary a mass flow rate of cooling air through the at least one air-lubricant heat exchanger, thereby which allows the first quantity of heat to be varied.

The at least one air-lubricant heat exchanger can be arranged at, or in the immediate vicinity of, a bypass pipe of the gas turbine engine.

The pipe assembly may include a first lubricant circuit adapted to provide a first flow of lubricant and a second lubricant circuit adapted to provide a second flow of lubricant, the at least one air-lubricant heat exchanger being arranged in the first lubricant circuit and the at least one fuel-lubricant heat exchanger being arranged in the second lubricant circuit.

The first lubricant circuit may include a bypass of the at least one air-lubricant heat exchanger. By adjusting the first mass flow of lubricant in the bypass, the amount of heat dissipated to the first heat sink can be adjusted.

The second lubricant circuit may include a bypass of the at least one fuel-lubricant heat exchanger. By adjusting the second lubricant mass flow rate in the bypass, the amount of heat dissipated to the second heat sink can be adjusted.

The heat management system may include a modulation device adapted to adjust a lubricant flow distribution between the power reducer and the turbomachine bearings. The modulating device may include one or more pump devices, for example one or more hydraulic pumps, such as gear pumps, rotary vane pumps, and the like, and/or one or more metering orifices.

The modulation device may include a first pump device arranged in the first lubricant circuit to adjust the first lubricant flow and a second pump device arranged in the second lubricant circuit to adjust the second lubricant flow.

The heat management system may include a lubricant reservoir in fluid communication with, and supplying lubricant to, the first and second lubricant circuits.

The heat management system may include a first lubricant reservoir in fluid communication with, and supplying lubricant to, the first lubricant circuit and a second lubricant reservoir in fluid communication with, and supplying lubricant to, the second lubricant circuit. lubricant.

The first lubricant circuit can provide lubrication and cooling to the power reducer.

The second lubricant circuit can provide lubrication and cooling to the turbomachine bearings. The second lubricant circuit can provide cooling to the power electronics of the gas turbine engine.

The gas turbine engine may include at least two air-lubricant heat exchangers for dissipating the first amount of heat in the first heat sink, at least one of which is arranged in the first lubricant circuit and at least one of which is arranged in the first heat sink. second lubricant circuit.

The blower may have a blower diameter in the range of 210 cm to 380 cm, or 210 cm to 370 cm, or 220 cm to 370 cm, for example 340 to 370.

The power reducer may have a gear ratio in the range of 2.9 to 4.0, or 3.0 to 3.8, or 3.1 to 3.7.

The lubricant can be oil.

The at least one air-lubricant heat exchanger may be a matrix air-cooled oil cooler (MACOC).

The at least one fuel-lubricant heat exchanger may be a fuel-oil heat exchanger (FOHE).

The combustion chamber may be a lean burn combustion chamber. The lean burn combustion chamber may include a plurality of lean burn fuel spray nozzles, each fuel spray nozzle including a pilot fuel injector and a main fuel injector.

The inventor has found that even in gas turbine engines with a lean burn combustion chamber, in which the thermal requirement is more stringent, a heat management system as described makes it possible to maximize the SFC and to avoid a risk of thermal degradation of the fuel in all operating conditions.

As noted elsewhere herein, the present disclosure relates to a gas turbine engine. Such a gas turbine engine comprises an engine core comprising a turbine, a combustion chamber, a compressor, and a core shaft connecting the turbine to the compressor. Such a gas turbine engine includes a fan (having fan blades) located upstream of the engine core.

Arrangements of the present disclosure are particularly advantageous for blowers which are driven via a power reducer. Accordingly, the gas turbine engine includes a power reducer which receives input from the core shaft and provides drive to the fan so as to drive the fan at a rotational speed lower than that of the core shaft. heart. The input to the power reducer can be directly from the core shaft, or indirectly from the core shaft, for example via a spur shaft and/or gear. The core shaft can join the turbine and the compressor, such that the turbine and compressor rotate at the same speed (with the fan rotating at a lower speed).

The gas turbine engine as described herein may have any suitable general architecture. For example, the gas turbine engine can have any desired number of shafts that connect turbines and compressors, for example one, two or three shafts. By way of example only, the turbine connected to the core shaft may be a first turbine, the compressor connected to the core shaft may be a first compressor, and the core shaft may be a first core shaft . The engine core may further include a second turbine, a second compressor and a second core shaft connecting the second turbine to the second compressor. The second turbine, the second compressor and the second core shaft may be arranged to rotate at a higher rotational speed than the first core shaft.

In such an arrangement, the second compressor can be positioned axially downstream of the first compressor. The second compressor can be arranged to receive (for example receive directly, for example via a generally annular conduit) a flow from the first compressor.

The power reducer may be arranged to be driven by the core shaft which is configured to rotate (e.g. in use) at the lowest rotational speed (e.g. the first core shaft in the example above). For example, the power reducer may be arranged to be driven only by the core shaft which is configured to rotate (e.g. in use) at the lowest rotational speed (e.g. be only the first shaft of hearts, not the second heart tree, in the example above). Alternatively, the power reducer can be arranged to be driven by any shaft(s), for example the first and/or second shafts in the example above.

The power reducer can be a reduction box (in that the output to the fan has a lower rotation speed than the input from the core shaft). Any type of power reducer can be used. For example, the power reducer may be a "planetary" or "star" reducer, as described in more detail elsewhere in this document. The power reducer can have any desired reduction ratio (defined as the rotational speed of the input shaft divided by the rotational speed of the output shaft), for example greater than 2.5, for example in the range of 2.9 to 4.2, 3 to 4.2, 3 to 4, 3 to 3.8, or 3.2 to 3.8, for example of the order of or at minus 2.9, 3, 3.1, 3.2, 3.3, 3.4, 3.5, 3.6, 3.7, 3.8, 3.9, 4, 4.1 or 4 ,2. The gear ratio can be, for example, between any two of the values in the previous sentence. Strictly by way of example, the power reducer may be a "star" reducer having a ratio in the range from 3.1 or 3.2 to 3.8. In some arrangements, the gear ratio may be outside of these ranges.

In any gas turbine engine as described herein, a combustion chamber is provided axially downstream of the fan and compressor(s). For example, the combustion chamber may be directly downstream of (e.g. at the outlet of) the second compressor, when a second compressor is provided. As a further example, the flow at the outlet to the combustion chamber may be supplied to the inlet of the second turbine, when a second turbine is provided. The combustion chamber can be provided upstream of the turbine(s).

The or each compressor (e.g. the first compressor and the second compressor as described above) may include any number of stages, e.g. multiple stages. Each stage may include a row of rotor blades and a row of stator blades, which may be variable stator blades (in that their angle of incidence may be variable). The row of rotor blades and the row of stator blades can be axially offset from each other.

The or each turbine (e.g. the first turbine and the second turbine as described above) may comprise any number of stages, e.g. multiple stages. Each stage may include a row of rotor blades and a row of stator blades. The row of rotor blades and the row of stator blades can be axially offset from each other.

Each fan blade may be defined as having a radial span extending from one foot (or hub) at a gas-washed radially internal location, or 0% span position, to a tip at a 100% reach position. The ratio of the radius of the fan blade at the hub to the radius of the fan blade at the tip may be less than (or of the order of) any of: 0.4, 0 .39, 0.38, 0.37, 0.36, 0.35, 0.34, 0.33, 0.32, 0.31, 0.3, 0.29, 0.28, 0.27 , 0.26 or 0.25. The ratio of the radius of the fan blade at the hub to the radius of the fan blade at the tip can be included in a range bounded by any two of the values in the preceding sentence (i.e. say that the values can form upper or lower limits), for example in the range 0.28 to 0.32. These ratios may commonly be referred to as the hub to tip ratio. Both the hub radius and the tip radius can be measured at the leading edge (or axially most forward) portion of the blade. The hub-to-tip ratio refers, of course, to the gas-washed portion of the fan blade, that is, the portion radially outboard of any platform.

The fan radius can be measured from the engine centerline to the tip of a fan blade at its leading edge. The blower diameter (which may simply be twice the blower radius) may be greater than (or of the order of) any one of: 210 cm, 220 cm, 230 cm, 240 cm, 250 cm ( about 100 inches), 260 cm, 270 cm (about 105 inches), 280 cm (about 110 inches), 290 cm (about 115 inches), 300 cm (about 120 inches), 310 cm, 320 cm (about 125 inches) , 330 cm (approximately 130 inches), 340 cm (approximately 135 inches), 350 cm, 360 cm (approximately 140 inches), 370 cm (approximately 145 inches), 380 cm (approximately 150 inches), 390 cm (approximately 155 inches ), 400 cm, 410 cm (approximately 160 inches) or 420 cm (approximately 165 inches). The diameter of the blower may be within an inclusive range delimited by any two of the values in the preceding sentence (i.e. the values may form upper or lower limits), for example in the range 210 cm to 420 cm, or from 210 cm to 240 cm, or 250 cm to 280 cm, or 320 cm to 380 cm, or 340 cm to 370 cm.

Blower rotation speed may vary during use. Generally, the rotation speed is lower for blowers with a larger diameter. Strictly by way of non-limiting example, the rotation speed of the fan at cruising conditions may be less than 2800 rpm, for example less than 2600 rpm, or less than 2500 rpm, or less than 2300 rpm, or less than 2200 rpm, or less than 2000 rpm. Strictly by way of non-limiting example, the rotation speed of the fan at cruising conditions may be greater than 1200 rpm, or greater than 1300 rpm, or greater than 1400 rpm, or greater than 1500 rpm. /min, or greater than 1600 rpm. By way of non-limiting example only, the rotation speed of the fan under cruising conditions may be in the range of 1400 rpm to 2800 rpm, or in the range of 1600 rpm to 2500 rpm. min, or in the range of 1600 rpm to 2200 rpm. Strictly by way of additional non-limiting example, the rotation speed of the fan at cruising conditions for an engine having a fan diameter in the range from 210 cm to 300 cm (for example from 240 cm to 280 cm or from 250 cm cm to 270 cm) can be in the range of 1700 rpm to 2600 rpm, for example in the range of 1800 rpm to 2300 rpm, for example in the range of 1900 rpm to 2100 rpm. By way of non-limiting example only, the rotation speed of the fan under cruising conditions for an engine having a fan diameter in the range of 330 cm to 380 cm may be in the range of 1200 rpm at 2000 rpm, for example in the range of 1300 rpm to 1800 rpm, for example in the range of 1400 rpm to 1800 rpm.

In use of the gas turbine engine, the fan (with associated fan blades) rotates about an axis of rotation. This rotation results in a movement of the tip of the fan blade with a speed U _tip . The work accomplished by the fan blades on the flow results in an increase in enthalpy dH of the flow. A fan peak load can be defined by dH/U _tip ² , where dH is the enthalpy increase (e.g. 1-D average enthalpy increase) across the fan and U _tip is the speed ( translation) of the fan tip, for example at the leading edge of the tip (which can be defined as the fan tip radius at the leading edge multiplied by the angular velocity). The peak fan load at cruise conditions may be greater than (or of the order of) any of: 0.28, 0.29, 0.30, 0.31, 0.32, 0, 33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or 0.4 (all values having no dimension). The peak fan load may be in an inclusive range delimited by any two of the values in the preceding sentence (i.e. the values may form upper or lower limits), for example in the range 0, 28 to 0.31 or 0.29 to 0.3.

Gas turbine engines in accordance with the present disclosure may have any desired bypass ratio, where the bypass ratio is defined as the ratio of the mass flow rate of the flow through the bypass duct to the mass flow rate of the flow through the heart to cruising conditions. In some arrangements the bypass ratio may be greater than (or of the order of) any of the following: 10, 10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5, 15, 15.5, 16, 16.5, 17, 17.5, 18, 18.5, 19, 19.5 or 20. The bypass ratio may be in an inclusive range delimited by any two of the values in the preceding sentence (i.e., the values can form upper or lower limits), for example in the range 12 to 16, 13 to 15, or 13 to 14. The bypass conduit can be substantially annular. The bypass conduit may be radially outboard of the core motor. The radially external surface of the bypass duct may be defined by a nacelle and/or a fan casing.

The overall pressure ratio of a gas turbine engine as described here can be defined as the ratio of the total pressure upstream of the fan to the total pressure at the outlet of the highest pressure compressor (before inlet). in the combustion chamber). By way of non-limiting example, the overall pressure ratio of a gas turbine engine as described here in cruise may be greater than (or of the order of) any of the following: 35, 40, 45, 50, 55, 60, 65, 70, 75. The overall pressure ratio may be within an inclusive range delimited by any two of the values in the preceding sentence (i.e., the values may form limits higher or lower), for example in the range of 50 to 70.

The specific thrust of an engine can be defined as the net thrust of the engine divided by the total mass flow through the engine. At cruise conditions, the specific thrust of an engine described here may be less than (or of the order of) any of the following: 110 Nkg ^-1 s, 105 Nkg ^-1 s, 100 Nkg ^-1 s , 95 Nkg ^-1 s, 90 Nkg ^-1 s, 85 Nkg ^-1 s or 80 Nkg ^-1 s. The specific thrust may be in an inclusive range delimited by any two of the values in the previous sentence (i.e. the values may form upper or lower limits), for example in the range 80 Nkg ^-1 s at 100 Nkg ^-1 s, or from 85 Nkg ^-1 s to 95 Nkg ^-1 s. Such engines can be particularly efficient compared to conventional gas turbine engines.

A gas turbine engine as described here can have any desired maximum thrust. Strictly by way of non-limiting example, a gas turbine as described here may be capable of producing a maximum thrust of at least (or of the order of) any of the following: 160 kN, 170 kN, 180 kN, 190 kN, 200 kN, 250 kN, 300 kN, 350 kN, 400 kN, 450 kN, 500 kN or 550 kN. The maximum thrust may be an inclusive range bounded by any two of the values in the preceding sentence (i.e., the values may form upper or lower limits). By way of example only, a gas turbine as described herein may be capable of producing maximum thrust in the range of 330 kN to 420 kN, for example 350 kN to 400 kN. The thrust mentioned above may be the maximum net thrust under standard atmospheric conditions at sea level plus 15 degrees C (ambient pressure 101.3 kPa, temperature 30 degrees C), with the engine static.

During use, the flow temperature at the high pressure turbine inlet can be particularly high. This temperature, called TET, can be measured at the outlet of the combustion chamber, for example immediately upstream of the first turbine blade, which itself can be called a nozzle guide vane. In cruise conditions, the TET can be at least (or of the order of) any of the following values: 1400 K, 1450 K, 1500 K, 1550 K, 1600 K or 1650 K The TET at cruise conditions may be within an inclusive range delimited by any two of the values in the preceding sentence (i.e., the values may form upper or lower limits). The maximum TET in use of the motor can be, for example, at least (or of the order of) any one of the following values: 1700 K, 1750 K, 1800 K, 1850 K, 1900 K , 1950 K or 2000 K. The maximum TET can be in an inclusive range delimited by any two of the values in the preceding sentence (i.e. the values can form upper or lower limits), e.g. in the range of 1800 K to 1950 K. The maximum TET can occur, for example, in a high thrust condition, for example in a maximum takeoff thrust (PMD) condition.

A fan blade and/or airfoil portion of a fan blade described herein may be made from any suitable material or combination of materials. For example at least part of the fan blade and/or the aerodynamic profile may be manufactured at least in part from a composite, for example a metal matrix composite and/or an organic matrix composite, such as 'a carbon fiber. As a further example at least part of the fan blade and/or the aerodynamic profile may be manufactured at least in part from a metal, such as a titanium-based metal or a titanium-based material. aluminum (such as an aluminum-lithium alloy) or a steel-based material. The fan blade may include at least two regions manufactured using different materials. For example, the fan blade may have a protective leading edge, which may be made using a material that is more able to withstand impact (e.g. by birds, ice or other material) than the rest of the dawn. Such a leading edge can, for example, be manufactured using titanium or a titanium-based alloy. So, strictly by way of example, the fan blade may have a carbon fiber or aluminum-based body (such as an aluminum-lithium alloy) with a titanium leading edge.

A fan as described here may comprise a central portion, from which the fan blades may extend, for example in a radial direction. The fan blades can be connected to the central part in any desired way. For example, each fan blade may include a fastener that may engage a corresponding notch in the hub (or disc). Strictly by way of example, such a fixing element may be in the form of a dovetail which can notch into and/or engage a corresponding notch in the hub/disc in order to secure the fan blade to hub/disc. As a further example, the fan blades may be formed integrally with a central part. Such an arrangement may be referred to as a bladed disk or a bladed crown. Any suitable method may be used to manufacture such a bladed disk or bladed crown. For example, at least a portion of the fan blades may be machined from a block and/or at least a portion of the fan blades may be connected to the hub/disc by welding, such as linear friction welding.

The gas turbine engines described here may or may not be equipped with a variable area nozzle (VAN). Such a nozzle with a variable section can make it possible to vary the outlet area of the bypass duct during use. The general principles of this description can apply to engines with or without VAN.

The fan of a gas turbine as described herein may have any desired number of fan blades, for example 14, 16, 18, 20, 22, 24 or 26 fan blades.

As used herein, cruise conditions have the conventional meaning and would be readily understood by those skilled in the art. Thus, for a given gas turbine engine for an aircraft, one skilled in the art would immediately recognize that cruise conditions mean the engine operating point at mid-cruise of a given mission (which may be designated in the industry as an “economic mission”) of an aircraft to which the gas turbine engine is designed to be attached. In this sense, mid-cruise is the point in an aircraft flight cycle at which 50% of the total fuel that is burned between the end of the climb and the start of the descent has been burned (which can be approximated by the midpoint – in terms of time and/or distance – between the end of the climb and the start of the descent). Cruise conditions thus define an operating point of the gas turbine engine which provides thrust which would ensure steady-state operation (i.e. maintaining a constant altitude and a Mach number, constant) at mid-cruise of an aircraft to which it is designed to be attached, taking into account the number of engines provided on that aircraft. For example when an engine is designed to be attached to an aircraft which has two engines of the same type, at cruise conditions the engine provides half the total thrust that would be required for steady state operation of that aircraft at mid-speed. cruise.

In other words, for a given gas turbine engine for an aircraft, cruise conditions are defined as the engine operating point that provides a specified thrust (required to provide – in combination with any other engines on the aircraft – a steady state operation of the aircraft to which it is designed to be fixed at a given mid-cruise Mach number) at mid-cruise atmospheric conditions (defined by the international standard atmosphere according to ISO 2533 at mid-cruise altitude). For any given gas turbine engine for an aircraft, the mid-cruise thrust, atmospheric conditions, and Mach number are known, and thus the engine's operating point at cruise conditions is clearly defined.

Strictly by way of example, the forward speed at cruise condition may be any point in the range from Mach 0.7 to 0.9, e.g. 0.75 to 0.85, e.g. 0, 76 to 0.84, for example 0.77 to 0.83, for example 0.78 to 0.82, for example 0.79 to 0.81, for example of the order of Mach 0.8, of l order of Mach 0.85 or in the range from 0.8 to 0.85. Any single speed within these ranges can be part of the cruise condition. For a certain aircraft, cruise conditions may be outside of these ranges, for example below Mach 0.7 or above Mach 0.9.

Strictly by way of example, cruise conditions may correspond to typical atmospheric conditions (according to the International Standard Atmosphere, ISA) at an altitude which is in the range from 10,000 m to 15,000 m, for example in the range from 10,000 m to 12,000 m, for example in the range from 10,400 m to 11,600 m (approximately 38,000 feet), for example in the range from 10,500 m to 11,500 m, for example example in the range from 10,600 m to 11,400 m, for example in the range from 10,700 m (approximately 35,000 feet) to 11,300 m, for example in the range from 10,800 m to 11,200 m, for example in the range from 10,900 m to 11,100 m, for example of the order of 11,000 m. Cruise conditions may correspond to typical atmospheric conditions at any given altitude within these ranges.

Strictly by way of example, cruise conditions may correspond to an engine operating point that provides a known required level of thrust (e.g. a value in the range 30 kN to 35 kN) at a forward Mach number of 0.8 and typical atmospheric conditions (according to the international standard atmosphere) at an altitude of 38,000 feet (11,582 m). Strictly as a further example, cruise conditions may correspond to an engine operating point that provides a known required level of thrust (e.g. a value in the range 50 kN to 65 kN) at a Mach number forward of 0.85 and typical atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 35,000 feet (10,668 m).

In use, a described gas turbine engine may operate at the cruise conditions defined elsewhere in this document. Such cruise conditions may be determined by the cruise conditions (e.g. mid-cruise conditions) of an aircraft to which at least one (e.g. 2 or 4) gas turbine engine may be fitted in order to provide propulsion thrust.

A gas turbine engine as described may be provided to an aircraft. The aircraft is the aircraft to which the gas turbine engine was designed to be attached. Thus, cruise conditions in this aspect correspond to mid-cruise of the aircraft, as defined elsewhere in this document.

A gas turbine engine described may be operated at cruise conditions as defined elsewhere herein (e.g. in terms of thrust, atmospheric conditions and Mach number). A gas turbine engine described herein may be operated at mid-cruise of the aircraft, as defined elsewhere in this document.

Unless mutually exclusive, any feature or setting described here may be combined with any other feature or setting described here. For example, any of the ranges of the proportion of heat generated by the power reducer and the turbomachine and dissipated into the air at any of the environmental temperatures can be applied to any of the aspect described here.

BRIEF DESCRIPTION OF THE FIGURES

Embodiments will now be described by way of example only, with reference to the Figures, in which:

is a side sectional view of a gas turbine engine;

is a side view in close section of an upstream part of a gas turbine engine;

is a partially cut away view of a power reducer for a gas turbine engine;

is a functional schematic arrangement of a heat management system according to the description;

is a schematic arrangement of the heat management system according to a first embodiment of the description;

is a schematic arrangement of the heat management system according to a second embodiment of the description;

is a schematic arrangement of the heat management system according to a third embodiment of the description; And

is a schematic arrangement of the heat management system according to a fourth embodiment of the description.

DETAILED DESCRIPTION OF THE INVENTION

There illustrates a gas turbine engine 10 having a main axis of rotation 9. The engine 10 includes an air intake 12 and a propulsion fan 23 which generates two air flows: a core air flow A and a bypass air flow B. The gas turbine engine 10 comprises a core 11 which receives the air flow from core A. The engine core 11 comprises, in axial flow series, a low pressure compressor 14, a high pressure compressor 15, combustion equipment 16, a high pressure turbine 17, a low pressure turbine 19 and a core exhaust nozzle 20. In one embodiment, the combustion equipment 16 may include a combustion chamber A lean burn nozzle comprising a plurality of fuel spray nozzles, each fuel spray nozzle comprising a pilot fuel injector and a main fuel injector. A nacelle 21 surrounds the gas turbine engine 10 and defines a bypass duct 22 and a bypass exhaust nozzle 18. The bypass air flow B flows through the bypass duct 22. The fan 23 is attached to and driven by the low pressure turbine 19 via a core shaft 26 and a power reducer 30 of the epicyclic type. The gas turbine engine 10 further includes a heat management system 100 which will be described in more detail hereinafter.

In use, core airflow A is accelerated and compressed by the low pressure compressor 14 and directed into the high pressure compressor 15 where further compression takes place. Compressed air exhausted from the high pressure compressor 15 is directed into the combustion equipment 16 where it is mixed with fuel and the mixture is burned. The resulting hot combustion products then expand, and thereby drive the high pressure and low pressure turbines 17, 19 before being discharged through the nozzle 20 to provide some propulsion thrust. The high pressure turbine 17 drives the high pressure compressor 15 through a suitable interconnecting shaft 27. The fan 23 generally provides the majority of the propulsion thrust. The epicyclic gearbox 30 is a reduction box.

An exemplary arrangement for a geared fan gas turbine engine 10 is shown in Fig. . The low pressure turbine 19 (see ) drives the core shaft 26, which is coupled to a sun gear, or sun gear, 28 of the epicyclic gear arrangement 30. Radially outwardly from and meshing with the sun gear 28 here, there are a plurality of planet gears 32 which are coupled together by a planet carrier 34. The planet carrier 34 forces the planet gears 32 to change orientation around the sun gear 28 in synchronism while allowing each satellite gear 32 to rotate around its own axis. The planet carrier 34 is coupled via links 36 to the fan 23 in order to cause its rotation around the motor axis 9. Radially towards the outside of the planet gears 32 and meshing with them , there is a ring or toothed ring 38 which is coupled, via links 40, to a stationary support structure 24.

It should be noted that the terms "low pressure turbine" and "low pressure compressor" as used herein may be taken to indicate the lowest pressure turbine stages and the lowest pressure compressor stages (i.e. i.e. not including the fan 23) respectively and/or the turbine and compressor stages which are connected together by the core shaft 26 with the lowest rotational speed in the engine (this is i.e. not including the gearbox output shaft which drives the blower 23). In some literature, the "low pressure turbine" and "low pressure compressor" referred to herein may alternatively be known as the "intermediate pressure turbine" and "intermediate pressure compressor". When such alternative nomenclature is used, the blower 23 may be designated the first compression stage or the lowest pressure compression stage.

The epicyclic gearbox 30 is shown by way of example in more detail on the . Each of the sun gear 28, the planet gears 32 and the ring gear 38 includes teeth around its periphery for meshing with the other gears. However, for clarity, only exemplary portions of the teeth are shown in the illustration. . There are four planet gears 32 illustrated, although it will be apparent to the skilled reader that more or fewer planet gears 32 may be provided within the scope of the invention. Practical applications of a planetary epicyclic gearbox 30 generally include at least three satellite gears 32.

The epicyclic gearbox 30 illustrated by way of example in Figures 2 and 3 is of the planetary type, in that the planet carrier 34 is coupled to an output shaft via connections 36, with the toothed ring 38 fixed . However, any other suitable type of epicyclic gearbox 30 may be used. As a further example, the epicyclic gearbox 30 may be a star arrangement, in which the planet carrier 34 is held fixed, with the toothed ring (or ring) 38 allowed to rotate. In such an arrangement, the fan 23 is driven by the ring gear 38. As another alternative example, the gearbox 30 can be a differential gearbox in which the ring gear 38 and the planet carrier 34 are one and the same. 'other allowed to turn.

It will be borne in mind that the arrangement shown in Figures 2 and 3 is by way of example only, and that various alternatives are within the scope of the present description. Strictly by way of example, any suitable arrangement may be used to position the gearbox 30 in the motor 10 and/or to connect the gearbox 30 to the motor 10. As a further example, connections (such as connections 36, 40 on the example of ) between the gearbox 30 and other parts of the motor 10 (such as the input shaft, the core shaft 26, the output shaft and the fixed structure 24) can have any desired degree of rigidity or flexibility. As a further example, any suitable arrangement of bearings between rotating and stationary parts of the motor (e.g. between input and output shafts from the gearbox and fixed structures, such as the gearbox housing) may be used, and the description is not limited to the exemplary arrangement of the . For example, when the gearbox 30 has a star arrangement (described above), those skilled in the art would easily understand that the arrangement of the output and support connections and the bearing locations would typically be different from that shown as example on the .

Thus, the present description extends to a gas turbine engine having any arrangement of gearbox styles (e.g. star or planetary), support structures, input shaft arrangement and exit, and landing locations.

Optionally, the reducer can drive additional and/or alternative components (e.g. the intermediate pressure compressor and/or a booster).

Other gas turbine engines to which the present description may be applied may have alternative configurations. For example, such engines may have another number of compressors and/or turbines and/or another number of interconnecting shafts. As an additional example, the gas turbine engine shown in has a split flow nozzle 18, 20 which means that the flow through the bypass conduit 22 has its own nozzle 18 which is independent of, and radially outside of, the core motor nozzle 20. However, this is not limiting, and any aspect of the present description may also apply to engines in which the flow through the bypass conduit 22 and the flow through the core 11 are mixed, or combined, before (or upstream of) a single nozzle, which may be called a mixed flow nozzle. One and/or the other of the nozzles (whether mixed or divided flow) can have a fixed or variable area. While the example described relates to a turbofan engine, the description can apply, for example, to any type of gas turbine engine, such as an open rotor (in which the stage fan is not surrounded by a nacelle) or a turboprop, for example. In some arrangements, the gas turbine engine 10 may not include a reduction gear 30.

The geometry of the gas turbine engine 10, and its components, is defined by a conventional axis system, comprising an axial direction (which is aligned with the axis of rotation 9), a radial direction (in the direction from bottom to top on the ) and a circumferential direction (perpendicular to the page on the view of the ). The axial, radial and circumferential directions are mutually perpendicular. As used herein, forward and reverse are relative to the gas turbine engine, i.e. the fan being in the front and the turbine being in the rear of the engine, and forward refers to the direction running from the rear to the front of the gas turbine engine.

There represents a functional schematic layout of the heat management system 100.

The heat management system 100 includes a lubricant circuit 113 with a hose assembly adapted to provide a flow of lubricant, e.g. oil, to all components of the gas turbine engine that require lubrication and cooling. The lubricant circuit 113 may have any suitable arrangement, as will be further illustrated with reference to Figures 5 to 8. The housing 101 represents heat generating gas turbine engine components, or in other words , the heat generated by such components and removed by the lubricant. Components requiring lubrication and cooling typically include the power gearbox and turbomachine bearings, and may also include one or more accessory gearboxes, power electronics, and electrical machines (if applicable). Turbomachine bearings designate bearings, such as plain bearings or rolling element bearings, arranged between rotating and stationary parts of the engine, such as the core shaft and any other interconnecting shafts, the compressor(s) and the turbine(s). Power electronics can be part of an electrical power management/generation system comprising one or more electrical machines, generators and/or batteries. For the sake of simplification, in the present description the term “turbomachine bearing” includes any component of the gas turbine engine other than the power reducer which generates heat and is cooled by the heat management system 100.

The heat 101 generated by the power reducer and the turbomachine bearings is dissipated to a first heat sink 102 and to a second heat sink 103. The first heat sink 102 is air. The second heat sink 103 is fuel. The heat management system 100 includes an air-oil heat exchanger (air-cooled oil heat exchanger) 104 and a fuel-oil heat exchanger (fuel-cooled oil heat exchanger) 105. The heat exchangers air-oil and fuel-oil 104, 105 may be a plurality of air-oil heat exchangers 104 and a plurality of fuel-oil heat exchangers 105.

The air-oil exchanger 104 is for example a matrix air-cooled oil cooler (MACOC). The air-oil heat exchanger 104 may be arranged in, or in close proximity to, the bypass line 22 such that a first quantity of heat 111 is rejected to the bypass air flow B. In a embodiment, the air-oil heat exchanger 104 can be arranged in the nacelle 21 and the first quantity of heat 111 is rejected to the outside air. In another embodiment, the air-oil heat exchanger 104 may be arranged in other parts of the engine. In another embodiment, the heat management system 100 may include one or more air-to-oil heat exchangers 104 arranged at, or in close proximity to, the bypass line 22 and/or one or more heat exchangers air-oil 104 arranged in the nacelle 21. In other words, the first heat sink 102 can be bypass air or outside air, or both. When more than one air-oil heat exchanger 104 is used, the first heat quantity 111 is the sum of the heat dissipated by each air-oil exchanger 104.

In more detail, an air circuit 106 supplies cooling air (either, one, or both of the bypass air and outside air) to the air-oil heat exchanger. 104. More than one air-oil heat exchanger 104 may be provided in the air circuit 106, either in series or in parallel. An air bypass 107 may be provided in the air circuit 106. The air bypass 107 makes it possible to vary the mass flow rate of air passing through the air-oil exchanger 104, for example to compensate for different environmental conditions. that the gas turbine engine can experience. Heated air is then discharged either into the bypass pipe 22 or channeled elsewhere, for example into the atmosphere. A modulation device 108 is provided to vary the first quantity of heat 111. In the illustrated embodiment, the modulation device 108 is a flow restriction valve arranged downstream of the air-oil heat exchanger 104 and adapted to vary the mass flow rate of cooling air through the air-oil heat exchanger 104. The air bypass 107 may be omitted. If air bypass 107 is present, the modulation device is arranged upstream of air bypass return 107 along air circuit 106, as illustrated. In embodiments, the modulation device 108 can be arranged elsewhere, for example upstream of the air-oil heat exchanger 104 as illustrated in dotted lines on the .

In embodiments, the modulation device 108 may be any device or state that allows the air flow and/or lubricant flow to be varied, such as: one or more air flow control valves. air flow ; one or more oil flow control valves; one or more compressors/pumps (one or the other or both in the air circuit 106 and the lubricant circuit 113); air bypass 107; and/or variations in gas turbine engine states that vary the mass flow rate of air in the air circuit 106 and/or the mass flow rate of lubricant in the lubricant circuit 113 to vary the first quantity of heat 111 exchanged between the lubricant and the first heat sink 102.

The fuel-oil heat exchanger 105, or FOHE, is for example a shell and tube heat exchanger where fuel is piped into the tubes, or a finned plate heat exchanger. The fuel-oil heat exchanger 105 is arranged along a fuel circuit 109 which supplies fuel to the combustion equipment 16. More than one fuel-oil heat exchanger 105 may be provided in the fuel circuit 109. fuel 109, either in series or in parallel. The fuel may be aviation kerosene. The fuel circuit 109 may include a fuel bypass 110, which allows the mass flow rate of fuel passing through the fuel-oil heat exchanger 105 to be varied, for example to compensate for different environmental conditions than the gas turbine engine. can suffer. The fuel receives a second quantity of heat 112. The heat dissipated in the fuel is retained in the thermodynamic cycle of the engine and therefore the fuel is a convenient heat sink for the heat generated by the power reducer and turbomachine bearings.

A modulation device 118 is provided to vary the second quantity of heat 112. In the illustrated embodiment, the modulation device 118 is a flow restriction valve arranged in the fuel circuit 109 downstream of the exchanger fuel-oil heat 105 and upstream of the fuel bypass return 110. In embodiments, the modulation device 118 can be arranged elsewhere, for example upstream of the fuel-oil heat exchanger 105 as illustrated in dotted lines on the , and can be any device or state making it possible to vary the flow of fuel and/or the flow of lubricant, such as for example: one or more fuel flow control valves; one or more lubrication flow control valves; one or more pumps (one or the other or both in the fuel circuit 109 and the lubricant circuit 113); fuel bypass 110; and/or variations in gas turbine engine states that vary the mass flow rate of fuel in the fuel circuit 109 and/or the mass flow rate of lubricant in the lubricant circuit 113 to vary the second quantity of heat 112 exchanged between the lubricant and the second heat sink 103.

Ideally, the second quantity of heat 112 is maximized to obtain SFC benefits, without inducing fuel degradation. However, the ability of the second heat sink 103 to exchange heat and the amount of heat generated by the motor generally varies depending on motor conditions. Under low power conditions, for example at flight idle, the amount of heat generated by the engine is relatively low, but at the same time also the mass flow rate of fuel in the fuel circuit 109 is relatively low, therefore decreasing the second amount of heat 112 that can be safely transferred to the fuel; at high power conditions, for example at maximum takeoff thrust, the amount of heat generated by the engine is relatively high, but also the mass flow rate of fuel in the fuel circuit 109 is relatively high, which allows increase the second amount of heat 112 that can be safely transferred to the fuel.

The operating conditions of the motor also have an impact on the capacity of the first heat sink 102 to dissipate heat, that is to say on the first quantity of heat 111; for example, on the ground at idle the bypass flow B is minimal and generally increases with core shaft speed. However, the ability of the first heat sink 102 and the second heat sink 103 to dissipate heat generally does not vary proportionally with core shaft speed. To this end, the heat management system 100 is configured to vary the first heat quantity 111 and the second heat quantity 112 at different engine conditions, i.e. different core shaft speeds, such that their ratios are in specific ranges which make it possible to minimize the dimensions (and therefore the weight) of the heat exchangers 104, 105 and to maximize the second quantity of heat 112 for all engine operating conditions, particularly in cruise where the majority of the fuel is burned, and therefore maximize the SFC.

The heat management system 100 is configured to provide the first quantity of heat 111 and the second quantity of heat 112 such that a first proportion of the heat generated by the power reducer and the turbomachine and dissipated to the air defined as:

_85%PMD

at 85% of the maximum thrust speed at core shaft takeoff is in the range from 0.25 to 0.70, for example equal to 0.55; and a second proportion of the heat generated by the power reducer and the turbomachine and dissipated towards the air defined as:

_65%PMD

at 65% of the maximum thrust speed at core shaft takeoff is in the range from 0.60 to 1, for example equal to 0.85.

In embodiments, the first proportion may be greater than 0.25, or greater than 0.30, or greater than 0.35, or greater than 0.40, or greater than 0.45, or greater than 0 .50, or greater than 0.55, and less than 0.70, or less than 0.65, for example in the range from 0.25 to 0.70, or in the range from 0.35 to 0 .70, or in the range of 0.45 to 0.70, or in the range of 0.50 to 0.70, or in the range of 0.55 to 0.70, or in the range from 0.55 to 0.65.

In embodiments, the second proportion may be greater than 0.60, or greater than 0.65, or greater than 0.70, or greater than 0.75, and less than 1, or less than 0.95, for example in the range from 0.60 to 1, or in the range from 0.65 to 1, or in the range from 0.70 to 1, or in the range from 0.75 to 1, or in the range from 0.75 to 0.95.

The maximum thrust speed at core shaft takeoff can be in the range of 5500 to 9500 rpm. Thus, 85% of the maximum core shaft takeoff thrust speed can be in the range of 4675 to 8075 rpm, and 65% of the maximum core shaft takeoff thrust speed can be in the range from 3575 to 6175 rpm.

For example, for gas turbine engines with a fan diameter between 210 cm and 330 cm the maximum thrust speed at core shaft takeoff can be in the range from 6700 to 9500 rpm; For gas turbine engines with a fan diameter between 330 cm and 380 cm the maximum thrust speed at core shaft takeoff can be in the range from 5500 to 6700 rpm.

In one embodiment, the maximum core shaft takeoff thrust speed may be in the range from 5500 rpm to 6500 rpm, and therefore 85% of the maximum shaft takeoff thrust speed. core shaft may be in the range of 4675 to 5525 rpm, and 65% of the maximum core shaft takeoff thrust speed may be in the range of 3575 to 4225 rpm.

The fan may have a fan rotation speed at PMD conditions in the range of 1500 rpm to 2800 rpm, preferably in the range of 1600 rpm to 2500 rpm, more preferably in the range of 1600 rpm to 2500 rpm. range from 1600 rpm to 2200 rpm, even more preferably in the range from 1700 rpm to 1900 rpm.

Under cruising conditions, the core shaft may have a rotational speed in the range of 5000 rpm to 9000 rpm, or in the range of 5000 rpm to 7000 rpm, or in the range of 5000 rpm to 7000 rpm. the range from 5000 rpm to 6000 rpm, or in the range from 5200 rpm to 5800 rpm.

At cruise conditions, the fan may have a fan rotation speed in the range of 1400 rpm to 2600 rpm, preferably in the range of 1500 rpm to 2300 rpm, more preferably in the range of 1500 rpm to 2300 rpm. the range from 1600 rpm to 2000 rpm.

At cruise conditions, the heat management system 100 can be configured to provide the first quantity of heat 111 and the second quantity of heat 112 such that the proportion of heat generated by the power reducer and the turbomachine and dissipated in air is in the range of 0.35 to 0.80, or in the range of 0.50 to 0.80, or in the range of 0.57 to 0.80, or in the range ranging from 0.57 to 0.75, or in the range from 0.56 to 0.70, for example 0.65.

Additionally, the heat management system 100 may be configured to provide the first amount of heat 111 and the second amount of heat 112 such that a ratio of the first proportion to the second proportion is in the range of 0.45 to 0.65, preferably in the range of 0.45 to 0.60, more preferably in the range of 0.47 to 0.58, for example 0.57.

Furthermore, the environmental conditions, and in particular the environmental temperature, have an impact on the capacity of the first heat sink 102 and the second heat sink 103 to dissipate the heat generated by the power reducer and the turbomachine bearings. The inventor has found that such capacitance does not vary with the environmental temperature in the same way for the first and second heat sinks 102, 103. In other words, the quantities of heat (external or bypass) that air and fuel can release vary with temperature.

For this reason, the heat management system 100 may be configured to vary the first amount of heat 111 and the second amount of heat 112 depending on the temperature of the environment so as to provide specific first and second proportions of heat. heat generated by the power reducer and the turbomachine and dissipated in the air which makes it possible to maximize the second quantity of heat 112, maximize the SFC, minimize the dimensions (and therefore the weight) of the heat exchangers 104, 105, without causing fuel degradation.

For this purpose, the heat management system 100 can be configured to provide the first proportion of heat generated by the power reducer and the turbomachine and dissipated in the air at an environment temperature of ISA +40 ° C in the range of 0.55 to 0.70, or in the range of 0.60 to 0.70, or in the range of 0.62 to 0.68, for example 0.65.

Furthermore, the heat management system 100 can be configured to provide the second proportion of heat generated by the power reducer and the turbomachine and dissipated in the air at an environmental temperature of ISA +40 ° C in the range from 0.85 to 1, or in the range from 0.90 to 1, or from 0.92 to 1, or from 0.92 to 0.98, for example 0.95.

As the environment temperature decreases, the air and fuel temperatures decrease, and the amount of heat that can be rejected to the air and fuel increases, although not proportionally to each other. . Thus, to maximize the SFC, minimize the dimensions (and therefore the weight) of the heat exchangers 104, 105, without inducing fuel degradation, the heat management 100 can also be configured to ensure the first proportion in specific ranges at different environmental temperatures.

At an environmental temperature of ISA +10°C, the heat management system 100 may be configured to provide the first amount of heat 111 and the second amount of heat 112 so as to provide the first proportion of heat generated by the power reducer and the turbomachine and dissipated in the air in the range from 0.35 to 0.65, or in the range from 0.40 to 0.60, or in the range from 0.45 to 0.55, for example 0.50.

At an environmental temperature of ISA +10°C the heat management system 100 may be configured to provide the first amount of heat 111 and the second amount of heat 112 so as to provide the second proportion of heat generated by the power reducer and the turbomachine and dissipated in the air in the range from 0.60 to 0.95, or in the range from 0.70 to 0.95, or in the range from 0.75 to 0 .95, or in the range 0.80 to 0.92, for example, 0.86.

At an environmental temperature of ISA -69°C the heat management 100 may be configured to provide the first amount of heat 111 and the second amount of heat 112 so as to provide the first proportion of heat generated by the reducer d power supply and the turbomachine and dissipated in the air in the range from 0.20 to 0.40, or in the range from 0.20 to 0.35, or in the range from 0.20 to 0, 30, for example 0.25.

At an environmental temperature of ISA -69°C the heat management system 100 may be configured to provide the first amount of heat 111 and the second amount of heat 112 so as to provide the second proportion of heat generated by the power reducer and the turbomachine and dissipated in the air in the range from 0.40 to 0.75, or in the range from 0.50 to 0.75, or in the range from 0.55 to 0 .70, for example 0.63.

At an environmental temperature of ISA +40°C the heat management system 100 may also be configured to provide the first amount of heat 111 and the second amount of heat 112 so as to provide a ratio of the first proportion to the second proportion in the range from 0.50 to 0.80, or in the range from 0.55 to 0.75, or in the range from 0.60 to 0.70, for example 0.65.

At an environmental temperature of ISA +10°C the heat management system 100 may be configured to provide the first amount of heat 111 and the second amount of heat 112 so as to provide a ratio of the first proportion to the second proportion in the range of 0.40 to 0.65, or in the range of 0.45 to 0.65, or in the range of 0.50 to 0.60, for example, 0.57.

At an environmental temperature of ISA -69°C the heat management system 100 may be configured to provide the first amount of heat 111 and the second amount of heat 112 so as to provide a ratio of the first proportion to the second proportion in the range from 0.30 to 0.55, or in the range from 0.30 to 0.50, or in the range from 0.35 to 0.45, for example 0.40.

By providing the first amount of heat 111 and the second amount of heat 112 such that the ratios of the first proportion to the second proportion at different environmental temperatures are as provided above, the SFC can be maximized and Fuel degradation can be avoided under all environmental conditions and engine operating conditions.

Furthermore, the heat management system 100 may be configured to provide the first amount of heat 111 and the second amount of heat 112 such that the ratio of the first proportion to an environment temperature of ISA +40° C at the first proportion at an environment temperature of ISA -69 °C is in the range of 1.5 to 4.5, or in the range of 2 to 4, or in the range of 2 to 3.5, for example 2.6.

The heat management system 100 may be configured to provide the first amount of heat 111 and the second amount of heat 112 such that the ratio of the second proportion at an environmental temperature of ISA +40°C to the second proportion at an environment temperature of ISA -69 °C is in the range of 1.0 to 2.1, or in the range of 1.2 to 2.1, or in the range of 1.4 to 2.0, for example 1.6.

Further, the heat management system 100 may be configured to provide the first amount of heat 111 and the second amount of heat 112 such that the ratio of the first proportion to an environment temperature of ISA +40° C at the first proportion at an environment temperature of ISA +10 °C is in the range of 1.20 to 1.42, or in the range of 1.22 to 1.41, or in the range range from 1.25 to 1.40, for example 1.3.

Further, the heat management system 100 may be configured to provide the first amount of heat 111 and the second amount of heat 112 such that a ratio of the second proportion to an environment temperature of ISA +40 °C to the second proportion at an environment temperature of ISA +10 °C is in the range of 1.10 to 1.25, or in the range of 1.10 to 1.22, or in the range from 1.11 to 1.20, for example 1.16.

The heat management system 100 may be configured to provide the first amount of heat 111 and the second amount of heat 112 such that a proportion (expressed as a fraction) of heat generated by the power reducer and the turbomachine and dissipated in air is greater than A NH + B, and less than 1, in which A is equal to -1.15, B is equal to or greater than 1.48, and NH is a core shaft speed expressed as a proportion of the maximum core shaft takeoff thrust speed and is in the range 0.65 to 1.

B can be equal to or greater than 1.5, or 1.52, or 1.54, or 1.56.

The heat management system may be configured to provide the first amount of heat and the second amount of heat such that the proportion of heat generated by the power reducer and the turbomachine and dissipated to the air is

less than the smallest among 1 and C NH + D,

where C is -1.84, D is in the range of 2.10 to 2.30, and NH is in the range of 0.65 to 1.

D may be in the range of 2.18 to 2.30, or in the range of 2.18 to 2.25, or in the range of 2.20 to 2.25.

For example at 85% of the maximum thrust speed at core shaft takeoff (i.e. NH = 0.85), and being D equal to 2.20, the proportion of heat generated by the power reducer and the turbomachine and dissipated in the air can be greater than 50.25% (i.e. -1.15·0.85 + 1.48 = 0.5025, or 50.25%) and less than 63.6% (i.e. -1, 84·0.85 + 2.20 = 0.636, or 63.6%).

The heat management system may be configured to provide the first amount of heat and the second amount of heat such that the proportion of heat generated by the power reducer and the turbomachine and dissipated to the air is

greater than A NH + B, and less than the smallest of 1 and E (NH - 1) + F,

wherein A is -1.15, B is 1.48 or greater, E is in the range of -1.16 to -3, F is 0.37 or greater, and NH is the core shaft speed expressed as a proportion of the maximum core shaft takeoff thrust speed and is in the range 0.65 to 1.

B can be equal to or greater than 1.5, or 1.52, or 1.54, or 1.56.

E can be in the range of -1.16 to -2.5, or in the range of -1.16 to -1.95.

For example at 65% of the maximum thrust speed at core shaft takeoff (i.e. NH = 0.65), being E equal to -1.70 and F equal to 0.37, the proportion of heat generated by the power reducer and the turbomachine and dissipated in the air can be greater than 73.25% (i.e. -1.15·0.65 + 1.48 = 0.7325, or 73.25%) and less than 96, 5% (i.e. -1.70 · (0.65-1) + 0.37 = 0.965, or 96.5%).

For example at 100% of the maximum thrust speed at core shaft takeoff (i.e. NH = 1), being E equal to -1.70 and F equal to 0.37, the proportion of heat generated by the gear reducer power and the turbomachine and dissipated in the air can be greater than 33% (i.e. -1.15·1 + 1.48 = 0.33, or 33%) and less than 37% (i.e. -1.70· ( 1-1) + 0.37 = 0.37, or 37%).

In the above embodiments, the core shaft speed NH may be in the range of 0.65 to 0.95, or in the range of 0.65 to 0.90, or in the range ranging from 0.65 to 0.85.

Figures 5, 6, 7 and 8 show embodiments of the heat management system 100 as described. Similar characteristics between the and Figures 5 to 8 have received the same reference numbers, and will not be described in detail again in connection with Figures 5 to 8.

There shows a heat management system 200 where the air-oil heat exchanger 104 and the fuel-oil heat exchanger 105 are arranged in series along the lubricant circuit 113. In detail, the lubricant circuit 113 provides lubrication and cooling to the power reducer and to the turbomachine bearings. The housing 201 represents the power reducer and the turbomachine bearings lubricated and cooled by the lubricant in the lubricant circuit 113. In the illustrated embodiment, the air-oil heat exchanger 104 is arranged upstream of the exchanger fuel-oil heat exchanger 105 along the lubricant circuit 113. More than one air-oil heat exchanger 104 and more than one fuel-oil heat exchanger 105 may be provided along the lubricant circuit 113. In a embodiment not illustrated, the air-oil heat exchanger 104 can be arranged downstream of the fuel-oil heat exchanger 105 and downstream of the power reducer and the turbomachine bearings along the lubricant circuit 113.

The cooling air supplied to the air-oil heat exchanger 104 by means of the air circuit 106 represents the first heat sink 102 and may be bypass air or outside air, or both. The air circuit 106 includes an air bypass 107 to vary the mass flow of air passing through the first heat exchanger 104 and therefore the first quantity of heat 111. For this purpose, a modulation device 209 is arranged on along the air bypass 107. In embodiments, the air bypass 107 may be omitted and the mass flow rate of air passing through the first heat exchanger 104 may be varied by means of the modulation device 209 arranged downstream , or upstream of the first heat exchanger 104 along the air circuit 106.

The fuel supplied to the fuel-oil heat exchanger 105 by means of the fuel circuit 109 represents the second heat sink 103. Once the second quantity of heat 112 has been transferred from the lubricant to the fuel, the fuel is directed to the combustion equipment for combustion. The fuel circuit 109 comprises a fuel bypass 110 to vary the mass flow of fuel passing through the second heat exchanger 105 and therefore the second quantity of heat 112. For this purpose, a modulation device 210 is arranged along the bypass of fuel 110.

The heat management system 200 further comprises one or more reservoirs 120 for supplying the lubricant, for example oil, via one or more pumps.

Modulation devices 208, 209, 210 are provided in the lubricant circuit 113, the air circuit 106 and the fuel circuit 109 respectively, to vary the first quantity of heat 111 and the second quantity of heat 112. The modulation device 208 in the lubricant circuit 113 can be arranged upstream (as illustrated in the ), or downstream of the heat exchangers 104, 105 and may be one or more pumps, one or more flow control valves, or any other device adapted to vary the mass flow of lubricant through the heat exchangers 104, 105. The modulation device 209 in the air circuit 106 may be arranged in the air bypass 107 as illustrated, or downstream of the air-oil heat exchanger 104, and may be one or more compressors, a or several flow control valves, or any other suitable device adapted to vary the mass flow of air through the first heat exchanger 104. The modulation device 210 in the fuel circuit 109 can be arranged in the bypass of fuel 110 as illustrated, and may be one or more pumps, one or more flow control valves, or any other suitable device adapted to vary the mass flow of fuel through the second heat exchanger 105.

To further modulate the first heat quantity 111 and the second heat quantity 112, the lubricant circuit 113 may also include bypass circuits to adjust the mass flow rate of lubricant passing through the air-oil heat exchanger 104 and/or or the fuel-oil heat exchanger 105. In the illustrated embodiment of the , the lubricant circuit 113 comprises a first lubricant bypass circuit 122 adapted to divert a portion of lubricant away from the air-oil heat exchanger 104, and a second lubricant bypass circuit 124 adapted to divert a portion of lubricant away from the fuel-oil heat exchanger 105, thereby allowing the first amount of heat 111 and the second amount of heat 112 to be varied. In embodiments, one or more devices ) modulation 208, such as valves, can be arranged along the lubricant bypass circuits 122, 124 to adjust the portion of lubricant passing through the heat exchangers 104, 105, which allows the first quantity to be varied of heat 111 and the second amount of heat 112. In embodiments, either or both of the first lubricant bypass circuit 122 and the second lubricant bypass circuit 124 may be omitted.

There shows a heat management system 300 in which the air-oil heat exchanger 104 and the fuel-oil heat exchanger 105 are arranged in parallel in the lubricant circuit 313. In detail, the lubricant circuit 113 provides lubrication and cooling to the power reducer and to the turbomachine bearings. Housing 301 represents the power reducer and the turbomachine bearings lubricated and cooled by the lubricant. The lubricant circuit 313 comprises a first sub-circuit 314 and a second sub-circuit 315. The first and second sub-circuits form respective first and second loops of the lubricant circuit 313. The first loop and the second loop are arranged in parallel. The first sub-circuit 314 supplies the air-oil heat exchanger 104, and the second sub-circuit 315 supplies the fuel-oil heat exchanger 105.

The air-oil heat exchanger 104 rejects part of the heat generated by the power reducer and the turbomachine bearings corresponding to the first quantity of heat 111 towards the first heat sink 102. As described with reference to Figures 4 and 5 , the first heat exchanger 104 is an air-oil heat exchanger and receives cooling air from the air circuit 106 which represents the first heat sink 102. The air circuit 106 includes the air bypass 107 to adjust the mass flow rate of air passing through the air-oil heat exchanger 104 and therefore vary the first quantity of heat 111. The air circuit 106, in particular the air bypass 107, further comprises the device for modulation 209, for example one or more compressors and/or one or more air flow regulation valves to control the mass flow of air passing through the first heat exchanger 104 and therefore vary the first quantity of heat 111. or several additional compressors can be arranged upstream of the air-oil heat exchanger 104 in the air circuit 106 to increase the cooling air pressure.

The fuel-oil heat exchanger 105 rejects the remainder of the heat generated by the power reducer and the turbomachine bearings corresponding to the second heat quantity 112 to the second heat sink 103. Again, as described with reference to and 5, the fuel-oil heat exchanger 105 receives cooling fuel from the fuel circuit 109 which represents the second heat sink 103. The fuel circuit 109 includes the fuel bypass 110 to adjust the mass flow rate of fuel passing through it. fuel-oil heat exchanger 105 and therefore vary the second quantity of heat 112. The fuel circuit 109, in particular the bypass 110, further comprises the modulation device 210, for example one or more pumps and/or a or more fuel flow control valves to control the mass flow of fuel passing through the second heat exchanger 105 and therefore vary the second quantity of heat 112. One or more additional pumps may be provided in the fuel circuit, either in upstream or downstream of the fuel-oil heat exchanger 105.

As illustrated with reference to , either or both of the air bypass 107 and the fuel bypass 110 may be omitted and the modulating devices arranged accordingly along the air circuit 106 (either upstream or downstream of the air-oil heat exchanger 104) and the fuel circuit 109 (either upstream or downstream of the fuel-oil heat exchanger 105).

In addition, each sub-circuit 314, 315 includes a lubricant reservoir 120 and a modulation device 208. In embodiments not illustrated, a single lubricant reservoir 120 can supply both the first sub-circuit 314 and the second sub-circuit 315. Each modulation device 208 can be arranged either upstream (as illustrated in the ) or downstream of the respective heat exchangers 104, 105.

The lubricant circuit 313 may further include lubricant bypass circuits of either or both of the air-to-oil heat exchanger 104 and the fuel-to-oil heat exchanger 105. In the mode illustrated embodiment, the lubricant circuit 313 comprises a first lubricant bypass circuit 322 adapted to divert a portion of lubricant away from the air-oil heat exchanger 104, and a second lubricant bypass circuit 324 adapted to divert a portion of lubricant away from the fuel-oil heat exchanger 105. The first and second lubricant bypass circuits 322, 324 can make it possible to modulate the first and second quantities of heat 111, 112 exchanged at level of the air-oil heat exchanger 104 and at the level of the fuel-oil heat exchanger 105. For this purpose, one or more modulation devices 208, such as valves, can be provided in the bypass circuits of lubricant to adjust the portion of lubricant passing through the heat exchangers 104, 105.

There illustrates another embodiment of the heat management system 350 in which the air-oil heat exchanger 104 and the fuel-oil heat exchanger 105 are arranged in parallel along the lubricant circuit 363. The lubricant circuit 363 provides lubrication and cooling to the power reducer and turbomachine bearings, illustrated as a 301 housing on the . The lubricant circuit 363 comprises a first sub-circuit, or branch, 364 and a second sub-circuit, or branch, 365.

The air-oil heat exchanger 104 is arranged along the first branch 364 and the fuel-oil heat exchanger 105 is arranged along the second branch 365. More than one air-oil heat exchanger 104 and more than one fuel-oil heat exchanger 105 may be arranged in series along the respective branches.

The heat management system 350 further comprises at least one lubricant reservoir 120 for supplying lubricant, for example oil, via one or more pumps.

The air-oil heat exchanger 104 and the fuel-oil heat exchanger 105 can be arranged downstream of the tank 120 and upstream of the power reducer and the turbomachine bearings (as in the illustrated embodiment), or downstream of the power reducer and the turbomachine bearings.

The cooling air supplied to the air-oil heat exchanger 104 by means of the air circuit 106 represents the first heat sink 102 and may be bypass air or outside air, or both. The air circuit 106 includes an air bypass 107 to vary the mass flow rate of air passing through the air-oil heat exchanger 104 and therefore the first quantity of heat 111. In embodiments, the bypass d The air 107 can be omitted and the mass flow rate of air passing through the air-oil heat exchanger 104 can be varied by means of a modulation device arranged downstream, or upstream of the air-oil heat exchanger. 104.

The fuel supplied to the fuel-oil heat exchanger 105 by means of the fuel circuit 109 represents the second heat sink 103. Once the second quantity of heat 112 has been transferred from the lubricant to the fuel, the fuel is directed to the combustion equipment for combustion. The fuel circuit 109 includes a fuel bypass 110 to vary the mass flow of fuel passing through the fuel-oil heat exchanger 105 and therefore the second quantity of heat 112.

Lubricant bypass circuits may be provided at the air-oil heat exchanger 104 and/or the fuel-oil heat exchanger 105. In the illustrated embodiment, the lubricant circuit 363 includes a first circuit lubricant bypass circuit 372 adapted to divert a portion of lubricant away from the air-oil heat exchanger 104, and a second lubricant bypass circuit 374 adapted to divert a portion of lubricant away from the fuel-oil heat exchanger 105, which allows the first amount of heat 111 and the second amount of heat 112 to be varied. In embodiments, either or both of the first bypass circuit lubricant 372 and the second lubricant bypass circuit 374 can be omitted. Lubricant mass flow adjustment devices, such as valves, may be arranged in either or both of the first lubricant bypass circuit 372 and the second lubricant bypass circuit 374 to adjust the mass flow rate of lubricant in the respective lubricant bypass circuits and thereby adjust the mass flow rate of lubricant passing through the respective heat exchangers.

Modulation devices 208, 209, 210 are provided in the lubricant circuit 363, the air circuit 106 and the fuel circuit 109 respectively, to vary the first quantity of heat 111 and the second quantity of heat 112. The modulation device 208 in the lubricant circuit 363 may be one or more pumps, one or more flow control valves, or any other device adapted to vary the mass flow of lubricant through the heat exchangers. For example, the modulation device 208 may include a flow distribution valve to divide the flow of lubricant between the first branch 364 and the second branch 365; the modulation device 208 may also include modulation valves arranged along one and/or the other of the first lubricant bypass circuit 372 and the second lubricant bypass circuit 374. The modulation device 209 in the air circuit 106 may be disposed in the air bypass 107 upstream of the air-oil heat exchanger 104 as illustrated, or downstream of the air-oil heat exchanger 104, and may be one or more compressors, one or more flow control valves, or any other device adapted to vary the mass flow of air through the first heat exchanger 104. The modulation device 210 in the fuel circuit 109 can be arranged in the fuel circuit 109 upstream of the fuel-oil heat exchanger 105 as illustrated, or downstream of the fuel-oil heat exchanger 105, and may be one or more pumps, one or more flow control valves, or any other device adapted to vary the mass flow of fuel through the second heat exchanger 105.

There shows a heat management system 400 comprising a pipe assembly 401 adapted to provide lubricant and cooling to a power reducer 30 and turbomachine bearings 430. The pipe assembly 401 comprises a lubricant supply reservoir 120, for example example of oil, to a first lubricant circuit 414 and a second lubricant circuit 415. In other words, a single reservoir 120 supplies lubricant to both the first lubricant circuit 414 and the second lubricant circuit 415 In alternative embodiments, each of the first and second lubricant circuits 414, 415 may comprise a dedicated reservoir 120.

In detail, the first lubricant circuit 414 provides lubrication and cooling to the turbomachine bearings 430, while the second lubricant circuit 415 provides lubrication and cooling to the power reducer 30.

One or more pumps 408 are provided in the first lubricant circuit 414 to vary the mass flow rate of lubricant.

Furthermore, a first air-oil heat exchanger 104 is provided to reject the heat generated by the turbomachine bearings 430 towards the first heat sink 102. The air-oil heat exchanger 104 is for example a cooled oil cooler by air matrix (MACOC) of the type described with reference to the previous embodiments. As illustrated, a pump 408 can be arranged downstream of the tank 120 and upstream of the air-oil heat exchanger 104.

Cooling air is supplied to the first air-oil heat exchanger 104 by means of an air circuit 406. The air supplied to the air-oil heat exchanger 104 may be bypass air, for example from downstream of the fan outlet guide vanes (FOGV) arranged in the bypass pipe 22. In alternative embodiments, the air supplied to the air-oil heat exchanger 104 can be air outside, or a combination of outside air and bypass air. In other words, the first heat sink 102 can be bypass air or outside air, or both. A modulation device, such as a flow restriction valve, may be arranged along the air circuit 406 downstream of the first air-oil heat exchanger 104.

The first air-oil heat exchanger 104 is arranged downstream of the tank 120 and upstream of the turbomachine bearings 430. In embodiments not illustrated, the air-oil heat exchanger 104 can be arranged downstream of the bearings of the turbomachine. turbomachine 430 and upstream of the tank 120. The first air-oil heat exchanger 104 may include one or more first air-oil heat exchangers 104 arranged in series or in parallel along the first lubricant circuit 414 to reject the heat generated by the turbomachine bearings 430 towards the first heat sink 102.

An additional, or second, air/oil heat exchanger 104', for example an additional MACOC, is arranged in the second lubricant circuit 415 to reject part of the heat generated by the power reducer 30 towards the first heat sink 102 An additional air circuit 406' provides cooling air to the second air-oil heat exchanger 104'. The cooling air of the additional air circuit 406' can be bypass air, for example from downstream of the FOGVs arranged in the bypass pipe 22, or outside air, or a combination of outside air and bypass air, as described with reference to the first air-oil heat exchanger 104. A modulating device, such as a flow restriction valve, may be arranged along the supplemental air circuit 406' downstream of the second air-oil heat exchanger 104'. The second air-oil heat exchanger 104' may include one or more second air-oil heat exchangers 104 arranged in series or in parallel along the second lubricant circuit 415 to reject a portion of the heat generated by the power reducer 30 towards the first heat sink 102.

A lubricant bypass circuit 430 is arranged at the second air-oil heat exchanger 104' to divert part of the lubricant from the second air-oil heat exchanger 104'. For this purpose, an oil flow control valve 420 is arranged in the lubricant bypass circuit 430.

The first(s) air-oil heat exchanger(s) 104 and the second(s) air-oil heat exchanger(s) 104' together reject the first quantity of heat 111 towards the first sink thermal 102. It should be noted that in the embodiment of the , the first quantity of heat 111 is the sum of the heat generated by the turbomachine bearings 430 and part of the heat generated by the power reducer 30.

Downstream of the second air-oil heat exchanger 104' and of the lubricant bypass circuit 430 in the second lubricant circuit 415, there is arranged a fuel-oil heat exchanger 105, for example a fuel-oil heat exchanger ( FOHE) of the type described with reference to the previous embodiments, to reject the second quantity of heat 112 towards the second heat sink 103, namely the cooling fuel. The second quantity of heat 112 is the remainder of the heat generated by the power reducer 30 and not rejected by the second air-oil heat exchanger 104' to the first sink 102. The fuel-oil heat exchanger 105 may include one or more fuel-oil heat exchangers 105 arranged in series or in parallel along the second lubricant circuit 415 to reject the second quantity of heat 112 generated by the power reducer 30 towards the second heat sink 103.

A fuel circuit 409 supplies cooling fuel to the fuel-oil heat exchanger 105. The cooling fuel is supplied to the fuel-oil heat exchanger 105 from a fuel tank 432, e.g. fuel tank of the aircraft, on which the gas turbine engine is mounted, via a low pressure pump 436. The fuel leaving the fuel-oil heat exchanger 105 is then directed to nozzles spraying fuel from the combustion equipment 16 via a high pressure pump 438. The fuel flow regulation valve(s) (not shown) can be arranged along the fuel circuit 409, either in downstream, or upstream, of the fuel-oil heat exchanger 105.

The power reducer 30 is arranged downstream of the second heat exchanger 105 along the second lubricant circuit 415. In embodiments not illustrated, the second air-oil heat exchanger 104' may be arranged downstream of the fuel-oil heat exchanger 105 and upstream of the power reducer 30.

According to the illustrated embodiment, a pump 408 is arranged along the second lubricant circuit 415 downstream of the reservoir 120 and upstream of the second air-oil heat exchanger 104'. In embodiments not illustrated, the pump 408 can be arranged downstream of the second air-oil heat exchanger 104', or downstream of the fuel-oil heat exchanger 105, and upstream of the power reducer 30.

The heat management systems 200, 300, 350, 400 illustrated with reference to Figures 5, 6, 7 and 8 are configured to provide the first quantity of heat 111 and the second quantity of heat 112 such that the proportions of heat generated by the power reducer and the turbomachine and dissipated in the air are within the ranges described with reference to the .

It will be understood that the invention is not limited to the embodiments described above and that various modifications and improvements can be made without departing from the concepts described here. Unless mutually exclusive, any feature may be employed separately or in combination with other features and the description extends to and includes all combinations and sub-combinations of one or more features described herein.

Claims (14)

1 Gas turbine engine (10) for an aircraft comprising:
- an engine core (11) comprising a compressor (14), a combustion chamber (16), a turbine (19), and a core shaft (26) connecting the turbine (19) to the compressor (14);
- a fan (23) comprising a plurality of fan blades and arranged upstream of the engine core (11);
- turbomachine bearings;
- a power reducer (30) adapted to drive the fan (23) at a rotation speed lower than the turbine (19); And
- a heat management system (100, 200, 300, 350, 400) configured to provide lubrication and cooling to the gearbox (30) and to the turbomachine bearings which generate heat (101),
in which the heat management system (100, 200, 300, 350, 400) comprises:
- a pipe assembly adapted to provide a flow of lubricant to the reduction gear (30) and to the turbomachine bearings,
- at least one air-lubricant heat exchanger (104) for dissipating a first quantity (111) of the heat of the lubricant towards a first heat sink (102) which is air,
- at least one fuel-lubricant heat exchanger (105) for dissipating a second quantity (112) of the heat of the lubricant towards a second heat sink (103) which is fuel,
- a set for controlling the first and second quantities of heat comprising:
-- at least one modulation device (108, 208, 209) for varying the air flow and/or the flow of lubricant passing through said at least one air-lubricant heat exchanger (104);
-- at least one modulation device (118, 208, 210) for varying the flow of fuel and/or the flow of lubricant passing through said at least one fuel-lubricant heat exchanger (105);
in which the first quantity of heat (111) and the second quantity of heat (112) are controlled by the heat management system (100, 200, 300, 350, 400) so that the heat (101), generated by the reducer (30) and the turbomachine, is dissipated in the following air:
- a first proportion in operating conditions of the gas turbine engine (10) at 85% of a maximum thrust speed at core shaft takeoff (26), this first proportion being defined as
_85%PMD ,
- a second proportion in operating conditions of the gas turbine engine (10) at 65% of a maximum thrust speed at core shaft takeoff (26), this second proportion being defined as
_65%PMD
wherein the heat management system (100, 200, 300, 350, 400) is configured to provide a ratio of the first proportion to the second proportion in the range of 0.30 to 0.55, preferably in the range range from 0.35 to 0.45, at an environmental temperature of ISA -69 °C. 2 gas turbine engine (10) according to the preceding claim, in which the first proportion of the heat generated by the reduction gear (30) and the turbomachine and dissipated in the air at 85% of the maximum thrust speed at heart shaft takeoff at an environment temperature of ISA -69°C is in the range of 0.20 to 0.40, preferably in the range of 0.20 to 0.35, more preferably in the range of 0, 20 to 0.30. 3 Gas turbine engine (10) according to any one of the preceding claims, in which the second proportion of the heat generated by the reduction gear and the turbomachine and dissipated in the air at 65% of the maximum thrust speed at takeoff of core shaft at an environment temperature of ISA -69°C is in the range of 0.50 to 0.70, preferably in the range of 0.55 to 0.70, more preferably in the range from 0.55 to 0.67. 4 gas turbine engine (10) according to any one of the preceding claims, wherein the heat management system (100, 200, 300, 350, 400) is configured to control the first quantity of heat (111) and the second quantity of heat (112) such that a ratio of the first proportion of the heat generated by the gearbox (30) and the turbomachine and dissipated into the air at 85% of the maximum takeoff thrust speed d the core shaft (26) at an environmental temperature of ISA +40 °C at the first proportion of the heat generated by the reducer (30) and the turbomachine and dissipated in the air at 85% of the speed of maximum takeoff thrust of core shaft (26) at an environment temperature of ISA -69 °C is in the range of 1.5 to 4.5, preferably in the range of 2.0 to 4.0, more preferably in the range of 2.0 to 3.5. 5 Gas turbine engine (10) according to any one of the preceding claims, wherein the heat management system (100, 200, 300, 350, 400) is configured to provide the second proportion of the heat generated by the reducer (30) and the turbomachine and dissipated in the air at 65% of the maximum thrust speed at takeoff of core shaft (26) at an environmental temperature of ISA +40 ° C at the second proportion of the heat generated by the gearbox (30) and the turbomachine and dissipated into the air at 65% of the maximum thrust speed at core shaft takeoff (26) at an environmental temperature of ISA -69 °C is in the range of 1.1 to 2.1, preferably in the range of 1.2 to 2.1, more preferably in the range of 1.4 to 2.0. 6 Gas turbine engine (10) according to any one of the preceding claims, wherein the heat management system (100, 200, 300, 350, 400) is configured to drive at an environment temperature of ISA +40 °C the first heat quantity (111) and the second heat quantity (112) such that the first proportion of heat generated by the reducer (30) and the turbomachine and dissipated in the air at 85% of the maximum takeoff thrust speed of core shaft (26) is in the range of 0.55 to 0.75, preferably in the range of 0.55 to 0.70, more preferably in the range ranging from 0.60 to 0.70; and/or in which the heat management system (100, 200, 300, 350, 400) is configured to control at an environmental temperature of ISA +40 °C the first quantity of heat (111) and the second amount of heat (112) such that the second proportion of heat generated by the gearbox and the turbomachine and dissipated into the air at 65% of the maximum core shaft takeoff thrust speed is in the range from 0.85 to 1, preferably in the range from 0.90 to 1. 7 Gas turbine engine (10) according to any one of the preceding claims, wherein the heat management system (100, 200, 300, 350, 400) is configured to control the first quantity of heat (111) and the second quantity of heat (112) such that a proportion of heat generated by the reducer (30) and the turbomachine and dissipated in the air is
greater than A NH + B, and less than the smallest of 1 and C NH + D,
where A is -1.15, B is 1.48 or greater, C is -1.84, D is in the range of 2.18 to 2.30, and NH is the speed of core shaft (26) expressed as a proportion of the maximum takeoff thrust speed of core shaft (26) and is in the range of 0.65 to 1, preferably in the range of 0.65 to 0, 85. 8 Gas turbine engine (10) according to the preceding claim, wherein D is in the range from 2.18 to 2.25, preferably in the range from 2.20 to 2.25. 9 Gas turbine engine (10) according to any one of the preceding claims, wherein the maximum thrust speed at takeoff of core shaft (26) is in the range from 5500 rpm to 9500 rpm min, preferably in the range from 5500 rpm to 8500 rpm, more preferably in the range from 5500 rpm to 7500 rpm. 10 Gas turbine engine (10) according to any one of the preceding claims, wherein the combustion chamber (16) is a lean mixture combustion chamber; and/or wherein the power reducer (30) has a gear ratio in the range of 2.9 to 4.0, preferably 3.0 to 3.8. 11 Method of operating a gas turbine engine (10) for an aircraft, the method comprising providing a gas turbine engine (10) comprising:
- an engine core (11) comprising a compressor (14), a combustion chamber (16), a turbine (19), and a core shaft (26) connecting the turbine (19) to the compressor (14);
- a fan (23) comprising a plurality of fan blades and arranged upstream of the engine core (11), the fan (23) being configured to have a fan rotation speed at PMD conditions in the range from 1500 rpm to 2800 rpm;
- turbomachine bearings;
- a power reducer (30) adapted to drive the fan (23) at a rotation speed lower than the turbine (19); And
- a heat management system (100, 200, 300, 350, 400) configured to provide lubrication and cooling to the gearbox (30) and to the turbomachine bearings which generate heat (101),
in which the heat management system (100, 200, 300, 350, 400) comprises:
- a pipe assembly adapted to provide a flow of lubricant to the reduction gear (30) and to the turbomachine bearings,
- at least one air-lubricant heat exchanger (104) for dissipating a first quantity (111) of the heat of the lubricant towards a first heat sink (102) which is air,
- at least one fuel-lubricant heat exchanger (105) for dissipating a second quantity (112) of the heat of the lubricant towards a second heat sink (103) which is fuel,
- a set for controlling the first and second quantities of heat comprising:
-- at least one modulation device (108, 208, 209) for varying the air flow and/or the flow of lubricant passing through said at least one air-lubricant heat exchanger (104);
-- at least one modulation device (118, 208, 210) for varying the flow of fuel and/or the flow of lubricant passing through said at least one fuel-lubricant heat exchanger (105); in which the first quantity heat (111) and the second quantity of heat (112) are controlled by the heat management system (100, 200, 300, 350, 400) to provide:
- a first proportion of the heat (101), generated by the reduction gear (30) and the turbomachine, which is dissipated in the air under operating conditions of the gas turbine engine (10) at 85% of a speed maximum thrust at core shaft takeoff (26), this first proportion being defined as
_85%PMD , and
- a second proportion of the heat (101), generated by the reducer (30) and the turbomachine, which is dissipated in the air under operating conditions of the gas turbine engine (10) at 65% of the speed of maximum thrust at core shaft takeoff (26), this second proportion being defined as
_65%PMD ,
wherein the method comprises the step of operating the heat management system (100, 200, 300, 350, 400) to control the first amount of heat (111) and the second amount of heat (112) such so that at an environmental temperature of ISA -69 °C a ratio of the first proportion to the second proportion is in the range from 0.30 to 0.55, preferably in the range from 0, 35 to 0.45. 12 Method according to the preceding claim, comprising the step of operating the heat management system (100, 200, 300, 350, 400) to control the first quantity of heat (111) and the second quantity of heat (112 ) so that at an environmental temperature of ISA -69 °C the first proportion of the heat generated by the gearbox (30) and the turbomachine and dissipated in the air at 85% of the maximum thrust speed at core shaft takeoff (26) is in the range of 0.20 to 0.40, preferably in the range of 0.20 to 0.35, more preferably in the range of 0.20 at 0.30. 13 Method according to any one of claims 11 and 12, comprising the step of operating the heat management system (100, 200, 300, 350, 400) to control the first quantity of heat (111) and the second quantity of heat (112) such that at an environmental temperature of ISA -69 °C the second proportion of the heat generated by the reducer (30) and the turbomachine and dissipated in the air is 65% of the maximum core shaft takeoff thrust speed is in the range of 0.50 to 0.70, preferably in the range of 0.55 to 0.70, more preferably in the range of 0.55 to 0.67. 14 Method according to any one of claims 11 to 13, comprising the step of operating the heat management system (100, 200, 300, 350, 400) to control the first quantity of heat (111) and the second quantity of heat (112) such that a ratio of the first proportion of the heat generated by the gearbox (30) and the turbomachine and dissipated into the air at 85% of the maximum thrust speed at takeoff of core shaft (26) at an environmental temperature of ISA +40 °C at the first proportion of the heat generated by the reduction gear (30) and the turbomachine and dissipated in the air at 85% of the thrust speed maximum takeoff of core shaft (26) at an environment temperature of ISA -69 °C is in the range from 1.5 to 4.5, preferably in the range from 2.0 to 4.0, more preferably in the range 2.0 to 3.5.