Jet Engine Heat Transfer Bus Pumps

Abstract

A jet engine heat transfer bus pump is disclosed. An aircraft is disclosed herein, the aircraft comprising: a gas turbine engine, the gas turbine engine being configured to generate engine power (P engine ) of the fuel flow rate to burn the fuel, the fuel is composed of the first specific heat capacity (c p_fuel ) and net heat of combustion (NHC fuel ) characterizes; and a thermal management system configured to transfer heat from a working fluid to a fuel, the working fluid being characterized by a second specific heat capacity (c p_pump ) and the first density (ρ pump ) characterizes that the thermal management system includes a pump configured to generate a pump power (P pump ) to pressurize the working fluid, and #imgabs0#0.008≤POW / FFR 5 / 3 ≤12, FFR is between 0.05 lb-mass / s and 16 lb-mass / s, and ρ water and c p_water are the density and specific heat of water respectively.

Description

Technical Field

[0001] The present disclosure relates generally to fluid pumps and, more particularly, to pumps for aircraft jet engines. Background Art

[0002] Aircraft typically include various accessory systems that support the operation of the aircraft and / or its gas turbine engines. For example, such accessory systems may include a lubrication system that lubricates components of the engine, an engine cooling system that provides cooling air to engine components, an environmental control system that provides cooled air to the cabin of the aircraft, etc. Accordingly, heat is added to or removed from these accessory systems via fluids (e.g., oil, air, etc.) that are transported throughout a thermal management system (such as a closed-loop advanced Brayton cycle). BRIEF DESCRIPTION OF THE DRAWINGS

[0003] Figure 1 is a side view of an example aircraft.

[0004] Figure 2 is a schematic cross-sectional view of an example gas turbine engine for an aircraft.

[0005] Figure 3 is a schematic diagram of an example thermal management system that transfers heat between fluids.

[0006] Figure 4 FIG. 2 is a graph showing the corrected pump power versus time for an example heat transfer bus pump according to the teachings disclosed herein. Figure 2 A graph of the relationship between the corrected fuel flow rate of an example gas turbine engine.

[0007] Figure 5 is a diagram showing the corrected rotor diameter and Figure 2 A graph of the relationship between the corrected fuel flow rate of an example gas turbine engine.

[0008] Figure 6 yes Figure 4 A graph illustrating an example heat transfer bus pump embodiment according to the teachings disclosed herein.

[0009] Figure 7 yes Figure 5 A graph illustrating an example heat transfer bus pump embodiment according to the teachings disclosed herein.

[0010] Figure 8 is a flow chart illustrating an example method of operating a gas turbine engine according to the teachings disclosed herein.

[0011] The drawings are not drawn to scale. In general, the same reference numbers will be used throughout the drawings and accompanying written description to refer to the same or like parts. DETAILED DESCRIPTION

[0012] "Include" and "comprising" (and all their forms and tenses) are used herein as open-ended terms. Thus, whenever a claim adopts any form of "include" or "comprising" (e.g., includes, comprises, has, etc.) as a preamble or in any type of claim statement, it should be understood that additional elements, terms, etc. may be present without exceeding the scope of the corresponding claim or statement. As used herein, when the phrase "at least" is used as a transitional term, such as in the preamble of a claim, it is open-ended in the same manner as the terms "include" and "comprising" are open-ended. The term "and / or" when used, for example, in a form such as A, B, and / or C, refers to any combination or subset of A, B, C, such as (1) A alone, (2) B alone, (3) C alone, (4) A and B, (5) A and C, (6) B and C, or (7) A and B and C. As used herein in the context of describing structures, components, items, objects, and / or things, the phrase "at least one of A and B" is intended to refer to embodiments that include any of (1) at least one A, (2) at least one B, or (3) at least one A and at least one B. Similarly, as used herein in the context of describing structures, components, items, objects, and / or things, the phrase "at least one of A or B" is intended to refer to embodiments that include any of (1) at least one A, (2) at least one B, or (3) at least one A and at least one B. As used herein in the context of describing the performance or execution of processes, instructions, actions, activities, and / or steps, the phrase "at least one of A and B" is intended to refer to embodiments that include any of (1) at least one A, (2) at least one B, or (3) at least one A and at least one B. Similarly, as used herein in the context of describing the performance or execution of processes, instructions, actions, activities, and / or steps, the phrase "at least one of A or B" is intended to refer to embodiments that include any of (1) at least one A, (2) at least one B, or (3) at least one A and at least one B.

[0013] As used herein, singular references (e.g., "a", "a", "first", "second", etc.) do not exclude the plural. As used herein, the term "a" or "an" object refers to one or more of the object. The terms "a" (or "a"), "one or more", and "at least one" are used interchangeably herein. In addition, although listed separately, multiple devices, elements, or method actions can be implemented by, for example, the same entity or object. In addition, although individual features may be included in different examples or claims, these may be combined, and the inclusion of different examples or claims does not mean that the combination of features is not feasible and / or disadvantageous.

[0014] As used herein, stating that any part (e.g., a layer, film, region, area, or panel) is in any way on (e.g., positioned on, located on, disposed on, or formed on, etc.) another part indicates that the referenced part is in contact with the other part, or that the referenced part is located above the other part with one or more intermediate parts located therebetween.

[0015] As used herein, unless otherwise indicated, connection references (e.g., attached, coupled, connected, and joined) may include intermediate members between the elements referred to by the connection reference and / or relative movement between those elements. Thus, a connection reference does not necessarily infer that two elements are directly connected and / or are in fixed relation to each other. As used herein, stating that any part is "in contact with" another part is defined to mean that there is no intermediate part between the two parts.

[0016] Unless otherwise specifically stated, descriptors such as "first," "second," "third," etc., used herein do not confer or otherwise indicate any meaning of priority, physical order, arrangement in a list, and / or ordering in any manner, but are merely used as labels and / or arbitrary names to distinguish elements to facilitate understanding of the disclosed examples. In some examples, the descriptor "first" may be used to refer to an element in the detailed description, while a different descriptor (such as "second" or "third") may be used in the claims to refer to the same element. In such cases, it should be understood that such descriptors are only used to clearly identify elements that may, for example, otherwise share the same name.

[0017] The terms "front" and "rear" refer to relative positions within a gas turbine engine, pump, or vehicle, and refer to the normal operating attitude of the gas turbine engine, pump, or vehicle. For example, with respect to a gas turbine engine, the front refers to a position closer to the engine inlet, while the rear refers to a position closer to the engine nozzle or exhaust. Further, with respect to a pump, the front refers to a position closer to the pump inlet, while the rear refers to a position closer to the end of the pump opposite the inlet.

[0018] The terms "upstream" and "downstream" refer to locations along a fluid flow path relative to the direction of the fluid flow. For example, with respect to fluid flow, "upstream" refers to a location from which the fluid is flowing, and "downstream" refers to a location toward which the fluid is flowing. For example, with respect to a gas turbine engine, the engine inlet is considered upstream of the engine outlet, and the engine outlet is considered downstream of the engine inlet.

[0019] Various terms are used herein to describe the orientation of features. In general, some of the drawings are annotated with a set of axes, including the axis of rotation (e.g., axial axis) z and the radial axis r. In general, the drawings may be annotated with reference to the axial direction A, radial direction R, and / or circumferential direction C of the carrier associated with features, forces, and moments. The axial direction refers to the direction parallel to the axis of rotation z, about which the rotating parts of the turbine engine rotate. The radial direction refers to the direction perpendicular to the axis of rotation and pointing (radially inward) or away from (radially outward) the axis of rotation. The circumferential direction at a given point is a direction perpendicular to the local radial direction and perpendicular to the axial direction. With reference to the meridian plane, the meridian plane is a plane defined by a constant polar angle in a cylindrical coordinate system. The meridian plane refers to a plane formed by the axis of rotation and the radial axis.

[0020] The term "thrust class" refers to a group or class of gas turbine engines, or a type of gas turbine engine, specified in engine certification documents as being capable of generating similar amounts of thrust during various flight conditions (e.g., takeoff, cruise, landing, etc.). For example, a first thrust class includes one or more gas turbine engines generating a maximum cruise thrust of between 30,000 pounds-force (lbf) and 40,000 lbf.

[0021] The term "specific fuel consumption (SFC)" refers to the efficiency of a gas turbine engine in converting chemical energy into mechanical energy. Typically, the SFC of a gas turbine engine is based on the thrust class of the gas turbine engine, meaning that the SFC can vary significantly between different thrust classes, and engines within a thrust class share the same or similar SFC. In this case, the SFC (also called TSFC) can be defined as the mass flow rate of fuel injected into the combustion chamber per unit of thrust generated.

[0022] As used herein, "normalization" refers to creating shifted and scaled versions of the variables used in the relationships disclosed herein. Normalized values ​​allow comparison of variables, ratios, and / or relationships in a manner that eliminates or reduces the effects of certain overall effects (e.g., density, viscosity, specific heat capacity, etc.).

[0023] The term "pump power" refers to the power of the pump used to pressurize the working fluid to a mass flow rate. The term "corrected pump power" refers to the pump power normalized based on the properties of water to represent the pump power used to pressurize water to the same mass flow rate.

[0024] The term "rotor diameter" refers to the outlet diameter of the most upstream rotor in a pump. The most upstream rotor can be the only rotor in a single-stage pump or it can be the most upstream rotor in a multi-stage pump. For example, in a multi-stage centrifugal pump (or compressor), the most upstream rotor is the impeller that the working fluid first contacts when flowing into the inlet of the pump. The term "corrected rotor diameter" refers to the outlet diameter of the most upstream rotor in a pump, which has been normalized based on the properties of water. For example, if a first pump includes a first impeller having a rotor diameter to pressurize the working fluid to a specific mass flow rate, a second pump may include a second impeller having a corrected rotor diameter to pressurize water to the same mass flow rate.

[0025] The term "outlet diameter" refers to the diameter of a rotor (e.g., an impeller of a centrifugal / radial pump / compressor or a rotor of an axial pump / compressor) at the most downstream point of the rotor. In some examples, when referring to an impeller of a centrifugal pump, the outlet diameter corresponds to the diameter at the outermost tips of the impeller blades. In some examples, when referring to a rotor of an axial pump, the outlet diameter corresponds to the diameter at the trailing edge of the rotor blades.

[0026] The term "fuel flow rate" refers to the mass flow rate of fuel injected into the combustor of the gas turbine engine to generate thrust for cruise conditions. The term "corrected fuel flow rate" refers to the fuel flow rate that is normalized based on the properties (e.g., specific heat capacity, etc.) of the working fluid (described below) in the heat transfer bus and the fuel being burned.

[0027] The term "net heat of combustion (NHC)" refers to the amount of energy (heat) released per unit mass of fuel burned, assuming the fuel is burned at a constant pressure and the products of combustion are gaseous. NHC is a property of the fuel that varies between different types of fuels (e.g., Jet-A, biodiesel, etc.).

[0028] The term "specific heat capacity (c p ) refers to the amount of heat that needs to be added to a unit mass of a substance to raise its temperature by one unit. For example, the amount of heat required to raise the temperature of one kilogram (kg) of water by one Kelvin (K) is 4184 joules (J), so the specific heat capacity (c p_water ) is 4184J / kg*K.

[0029] In the following detailed description, reference is made to the accompanying drawings forming a part thereof, wherein specific examples that can be put into practice are shown by way of example. These examples are described in sufficient detail to enable those skilled in the art to practice this theme, and it should be understood that other examples may also be used. Therefore, the following detailed description is provided to describe exemplary embodiments, rather than being considered as a limitation on the scope of the subject matter described in this disclosure. Certain features from the different aspects described below may be combined to form yet another new aspect of the subject matter discussed below.

[0030] A turbine engine (also referred to herein as a gas turbine engine) is an internal combustion engine that uses the atmosphere as a moving fluid. A gas turbine engine is a turbofan engine that includes a fan section and a bypass airflow passage upstream of a low-pressure compressor section. During operation, a certain amount of air enters the inlet of the engine and enters the fan section. A first portion of the air is directed or directed into the bypass airflow passage, and a second portion of the air is directed or directed into the low-pressure compressor section, where the pressure of the air increases. When the second portion of the air is directed through the high-pressure compressor section and into the combustion chamber, the pressure of the second portion of the air is further increased, where the pressurized air is mixed with fuel and burned to provide combustion gases. Subsequently, the combustion gases are directed through the high-pressure turbine section and the low-pressure turbine section, where a portion of the heat energy and / or kinetic energy from the combustion gases is extracted.

[0031] The combustion gases are then directed through the jet exhaust nozzle section of the gas turbine engine to provide propulsive thrust. At the same time, as the first portion of the air is directed through the bypass airflow passage before being discharged from the fan nozzle exhaust section of the turbofan engine, the pressure of the first portion of the air is significantly increased, also providing propulsive thrust. The combination of the propulsive thrust from the first and second portions of air determines the total thrust generated by the turbofan engine for propelling an aircraft in flight. In this sense, the power of the gas turbine engine can be defined as the product of the total thrust of the aircraft and the cruising speed. Generally, as the size of the gas turbine engine increases, the engine power, thrust rating and its SFC also increase.

[0032] Pumps (e.g., centrifugal pump / compressors, axial pump / compressors, etc.) are used in a thermal management system (TMS) to pressurize (drive) a working fluid (e.g., water, oil, supercritical carbon dioxide (sCO2), liquid helium, helium-xenon, etc.) through a heat transfer bus loop. Such a TMS can heat or cool accessory systems, sections, and / or components in an engine to improve their power, efficiency, and / or structural integrity. The TMS includes a heat transfer bus (TTB) to transfer a working fluid (or heat exchange fluid) between elements (e.g., accessory systems, sections, components, etc.) of a gas turbine engine so that heat can be transferred from the elements to the working fluid / from the working fluid to the elements.

[0033] In some examples, the TMS uses sCO2 as a working fluid because it has low viscosity and high specific heat, enabling the heat exchanger to effectively transfer heat to and / or from sCO2. In addition, sCO2 has chemical stability, reliability, availability, and non-flammability, making sCO2 more advantageous than some other heat exchange fluids (e.g., water, air, etc.). In addition, unlike other heat exchange fluids (such as water) that freeze or evaporate, sCO2 is not prone to phase change. The TMS includes a TTB pump (or sCO2 pump) to pressurize the sCO2 in the TTB. The TTB pump can be a centrifugal pump or axial pump (e.g., a single-stage centrifugal pump, a multi-stage centrifugal pump, a multi-stage axial pump, etc.) that uses an electric motor to rotate an axis connected to one or more rotors (e.g., an impeller, a rotor blade, etc.), which draws a working fluid into the pump inlet and accelerates the working fluid radially and / or axially to the pump outlet and / or other rotors in the pump. In some examples, the TTB is a rotary screw pump, as described below.

[0034] The motor converts the electrical energy of the armature into mechanical work of the shaft connected to it. The shaft generates torque and rotates at a certain rotational speed. Since the rotor is connected to the shaft, the torque is transferred to the rotor and the rotor rotates at the same angular velocity. The pump power (e.g., mechanical power, fluid power, etc.) corresponds to the product of the amount of torque generated and the rotational speed of the shaft. Therefore, as the pump power increases, the rotational speed also increases. When the shaft of the pump rotates, the rotor compresses or pressurizes the working fluid to a certain pressure head. The term "pressure head" is used in this article as a measurement of the kinetic energy generated by the pump. Generally, the pressure head is a measurement of the height of an incompressible fluid column, and the pump can create an incompressible fluid column based on the kinetic energy provided to the incompressible fluid. As the pressure head increases, the working fluid accelerates, and the mass flow rate of the working fluid (e.g., a compressible fluid, such as sCO2) also increases.

[0035] The angular velocity (or rotational speed) of a rotor (e.g., an impeller of a centrifugal pump, a rotor blade of an axial pump, etc.) remains constant along its diameter. However, the tangential velocity of a point on the rotor is equal to the product of the angular velocity and the distance from that point to the axis of rotation of the rotor. Therefore, the tangential velocity of the rotor is maximum at its tip and / or outer diameter. The pressure head generated by the rotor is proportional to the tangential velocity at the outlet diameter of the rotor. For example, when the pump is a single-stage centrifugal pump with an impeller as the rotor, the pressure head doubles as the impeller diameter doubles. In addition, the pressure head is proportional to the square of the angular velocity of the rotor. For example, when the angular velocity doubles, the pressure head becomes four times. Therefore, since the mass flow rate increases with the pressure head, and the pressure head increases with the angular velocity and the tangential velocity, both the pump power and the rotor diameter affect the mass flow rate that the pump can supply.

[0036] It should be understood that the temperature of the fuel can affect the SFC (or efficiency) of the engine. That is, increasing the temperature of the fuel before injection into the combustion chamber can increase the energy (heat) generated while reducing fuel consumption, which can improve (reduce) the SFC of the gas turbine engine. In addition, the mass flow rate of the working fluid within the TMS can affect the amount of heat transferred to the fuel. Since the pump power and rotor diameter of the TTB pump affect the mass flow rate, a certain pump power and / or rotor diameter can be determined based on a certain thrust level (or corrected fuel flow rate) of the gas turbine engine to improve the SFC of the gas turbine engine. In addition, the TTB pump can be designed to generate a pump power that can improve the SFC of the engine without introducing other unacceptable disadvantages, such as increased engine weight, increased pump volume, reduced engine efficiency, etc.

[0037] Designing a TTB pump to improve the SFC of an engine can be a challenging process involving many trade-offs between the size of the TTB pump and the SFC benefits provided by the TTB pump. To increase pump power, the pump can be adapted with a larger motor, drive shaft, compressor, etc., which ultimately increases the size and weight of the pump. Analysis and iteration are typically performed when designing a TTB pump that outputs the appropriate amount of power for a gas turbine engine producing a specific aircraft thrust without oversizing the TTB pump. Such analysis can be ambiguous, uncertain, time consuming, and is considered highly dependent on the specific engine architecture.

[0038] When designing a TTB pump for a specific thrust level, there is no traditional design standard that limits the size of the pump while also providing pump power to achieve the desired SFC benefit. Therefore, for a given thrust level, a traditional TTB pump can be too large or too small. Increasing the pump power improves the SFC, but also increases the size of the TTB pump. At a certain point, the benefits of SFC obtained from the increased pump power are offset by the increased size (e.g., weight and volume) of the pump. In other words, in an attempt to improve the SFC of the engine, the TTB pump design may be too large. On the other hand, in an attempt to limit the weight and volume of the pump, the power of the TTB pump may be reduced. However, this may result in the pump being too small and unable to provide SFC benefits for a given engine thrust level. There are significant challenges in designing a TTB pump for a thrust-generating jet engine for an aircraft that is neither too large nor too small and operates at a pump power that provides the desired SFC benefit to the engine.

[0039] The inventors have developed an architecture for a TTB pump for a gas turbine engine and a TMS in an aircraft. In particular, the inventors proceeded in a manner to design a TTB pump having given performance characteristics of a pump motor (e.g., power, etc.) and dimensional characteristics of a pump rotor (e.g., outlet diameter, etc.). The inventors redesigned the TTB pump to achieve specific mass flow characteristics of the working fluid in the TMS and the heat transfer rate between the working fluid and the fuel in the engine, thereby increasing the temperature of the fuel, improving the SFC of the gas turbine engine, and reducing the overall fuel consumption of the engine. In designing several different types of gas turbine engines and TTB pumps (including the following combined Figure 6 and Figure 7 During the design of the TTB pump and engine described herein, the inventors re-examined the TTB pump power, outlet rotor diameter, size of the TTB pump, and the SFC of the engine produced by the redesigned TTB pump.

[0040] During the process of researching and evaluating various performance characteristics of TTB pumps that were considered feasible to best meet flight requirements, as well as the dimensional characteristics of TTB pumps (e.g., size, weight, etc.), the inventors unexpectedly discovered that there is a certain relationship between pump power and the fuel flow rate of the engine, and a certain relationship between the pump outlet rotor diameter and the fuel flow rate of the engine. These relationships provide a certain size TTB pump that improves the SFC of the engine while reducing the weight of the TTB pump and providing additional space for other components, systems and subsystems in the aircraft and / or engine. This discovery is described in more detail below.

[0041] For the drawings disclosed herein, the same numerals indicate the same elements throughout the drawings. Now referring to the drawings, Figure 1 is a side view of an example aircraft 10. Figure 1 As shown, the aircraft 10 includes a fuselage 12 and a pair of wings 14 (one shown) extending outwardly from the fuselage 12. In the example shown, a gas turbine engine 100 is supported on each wing 14 to propel the aircraft through the air during flight. In addition, the aircraft 10 includes a vertical stabilizer 16 and a pair of horizontal stabilizers 18 (one shown). However, in some examples, the aircraft 10 includes Figure 1 Illustrative examples of engines of different types and / or in different locations.

[0042] In addition, the aircraft 10 includes a thermal management system 102 (TMS 102) for transferring heat between fluids that support the operation of the aircraft 10. More specifically, the aircraft 10 includes one or more accessory systems that are configured to support the operation of the aircraft 10. For example, such accessory systems include a lubrication system that lubricates components of the engine 100, a cooling system that provides cooling air to components of the engine 100, an environmental control system that provides cooled air to a cabin of the aircraft 10, and the like. In such examples, the TMS 102 is configured to transfer heat from one or more fluids that support the operation of the aircraft 10 (e.g., oil of the lubrication system, air of the cooling system and / or environmental control system, and the like) to one or more other fluids that support the operation of the aircraft 10 (e.g., fuel supplied to the engine 100). However, in some other examples, the TMS 102 is configured to transfer heat directly to and / or from other components that support the operation of the aircraft 10 without an intermediate fluid.

[0043] Although the examples disclosed herein are reference Figure 1 The aircraft 10 is described in detail, but the examples disclosed herein may be applicable to aircraft 10 used with Figure 1-3 The present invention may also be applied to another type or configuration of aircraft having a thermal management system similar to TMS 102. Thus, the present subject matter may be readily applicable to another aircraft (e.g., a military jet aircraft, a cargo aircraft, etc.) having another engine (e.g., a turbojet engine, a turboprop engine, etc.). In addition, although Figure 1 The TMS 102 is shown as being located in the fuselage 12 of the aircraft 10 , but the TMS 102 (or a portion of the TMS 102 , such as a TTB, heat exchanger, TTB pump, etc.) may be located within the wing 14 , the engine 100 , and / or another location in the aircraft 10 .

[0044] Figure 2 is a schematic cross-sectional view of an example gas turbine engine 100. In the example shown, the engine 100 is configured as a high bypass turbofan engine. However, in some examples, the engine 100 is configured as a propfan engine, a turbojet engine, a turboprop engine, a turboshaft gas turbine engine, etc.

[0045] In the illustrated example, the engine 100 includes a TMS 102. In some examples, the entire TMS 102 is located within the engine 100. In some examples, a portion of the TMS 102 is located in the engine 100, and other portions of the TMS 102 are located in other components of the aircraft 10 (e.g., the fuselage 12, the wing 14, etc.). The engine 100 extends along an axial centerline 103 and includes a fan 104, a low pressure (LP) spool 106, and a high pressure (HP) spool 108 that are at least partially surrounded by an annular casing or nacelle 110. More specifically, the fan 104 may include a fan rotor 112 and a plurality of fan blades 114 (one is shown) coupled to the fan rotor 112. In this regard, the fan blades 114 are circumferentially spaced apart and extend radially outward from the fan rotor 112. In addition, the LP and HP spools 106, 108 are positioned downstream of the fan 104 along the axial centerline 103. As shown, the LP spool 106 is rotatably coupled to a fan rotor 112, which allows the LP spool 106 to rotate fan blades 114. Additionally, a plurality of exit guide vanes or struts 116 are circumferentially spaced from one another and extend radially between a casing 118 and a nacelle 110 surrounding the LP and HP spools 106, 108. As such, the struts 116 support the nacelle 110 relative to the casing 118 such that the casing 118 and the nacelle 110 define a bypass airflow passage 120 positioned therebetween.

[0046] The casing 118 generally surrounds or encloses the compressor section 122, the combustion section 124, the turbine section 126, and the exhaust section 128 in a serial flow order. In some examples, the compressor section 122 includes a low pressure (LP) compressor 130 of the LP spool 106 and a high pressure (HP) compressor 132 of the HP spool 108 positioned downstream of the LP compressor 130 along the axial centerline 103. Each compressor 130, 132 may in turn include one or more rows of compressor stator vanes 134 interdigitated with one or more rows of compressor rotor blades 136. Thus, the compressors 130, 132 define a compressed air flow path 133 extending therethrough. Further, in some examples, the turbine section 126 includes a high pressure (HP) turbine 138 of the HP spool 108 and a low pressure (LP) turbine 140 of the LP spool 106 positioned downstream of the HP turbine 138 along the axial centerline 103. In turn, each turbine 138 , 140 may include one or more rows of turbine stator vanes 142 interdigitated with one or more rows of turbine rotor blades 144 .

[0047] In addition, the LP spool 106 includes a low pressure (LP) shaft 146, and the HP spool 108 includes a high pressure (HP) shaft 148 positioned concentrically about the LP shaft 146. In such an example, the HP shaft 148 rotatably couples the turbine rotor blades 144 of the HP turbine 138 and the compressor rotor blades 136 of the HP compressor 132, such that rotation of the turbine rotor blades 144 of the HP turbine 138 rotatably drives the compressor rotor blades 136 of the HP compressor 132. As shown, the LP shaft 146 is directly coupled to the turbine rotor blades 144 of the LP turbine 140 and the compressor rotor blades 136 of the LP compressor 130. In addition, the LP shaft 146 is coupled to the fan 104 via a gearbox 150. In this regard, rotation of the turbine rotor blades 144 of the LP turbine 140 rotatably drives the compressor rotor blades 136 of the LP compressor 130 and the fan blades 114.

[0048] In some examples, the engine 100 generates thrust to propel the aircraft. More specifically, during operation, air 152 enters an inlet portion 154 of the engine 100. The fan 104 supplies a first portion 156 of the air 152 to the bypass airflow passage 120 and a second portion 158 of the air 152 to the compressor section 122. The second portion 158 of the air 152 first flows through the LP compressor 130, where the compressor rotor blades 136 gradually compress the second portion 158 of the air 152. Next, the second portion 158 of the air 152 flows through the HP compressor 132, where the compressor rotor blades 136 continue to gradually compress the second portion 158 of the air 152. The compressed second portion 158 of the air 152 is then delivered to the combustion section 124. In the combustion section 124, the second portion 158 of the air 152 is mixed with fuel and combusted to generate high temperature and high pressure combustion gases 160. Thereafter, the combustion gases 160 flow through the HP turbine 138, wherein turbine rotor blades 144 of the HP turbine 138 extract a first portion of kinetic and / or thermal energy from the combustion gases 160. This energy extraction rotates the HP shaft 148, which drives the HP compressor 132. The combustion gases 160 then flow through the LP turbine 140, wherein turbine rotor blades 144 of the LP turbine 140 extract a second portion of kinetic and / or thermal energy from the combustion gases 160. This energy extraction rotates the LP shaft 146, which drives the LP compressor 130 and the fan 104 via the gearbox 150. The combustion gases 160 then exit the engine 100 through the exhaust section 128.

[0049] As described above, the aircraft 10 includes a TMS 102 for transferring heat between fluids that support the operation of the aircraft 10. In this regard, the TMS 102 is positioned within the engine 100. For example, Figure 2As shown, the TMS 102 is positioned within the casing 118 of the engine 100 . However, in some other examples, the TMS 102 (or a portion of the TMS 102 ) is positioned at another location within the engine 100 , such as the nacelle 110 .

[0050] In the example shown, the TMS 102 transfers heat to the fuel flowing through the fuel supply flow line 170. In some examples, the fuel supply flow line 170 transfers fuel from a supply tank and / or a fuel pump to the combustion section 124. In some examples, the TMS 102 transfers heat from a working fluid (e.g., a heat exchange fluid, such as a supercritical fluid (e.g., supercritical carbon dioxide (sCO2)), etc.) within the TTB to the fuel within the fuel supply flow line 170 via a radiator exchanger (described below). The pressure and / or flow rate of the working fluid within the TMS 102 defines the rate at which heat energy is transferred between the working fluid and the fuel. In addition, the pump power and / or rotor diameter of the TTB pump of the TMS 102 affects the heat transfer rate between the working fluid and the fuel. In addition, the heat of the fuel in the fuel supply flow line 170 (upstream of the combustion section 124) can affect the SFC of the engine. It is therefore advantageous for the TMS 102 to utilize an effectively sized TTB pump to generate a pressure and / or flow rate such that the SFC of the engine is effectively improved without the TMS 102 adding significant weight to the engine 100 .

[0051] Although the examples disclosed herein are reference Figure 2 The gas turbine engine 100 is described in detail, but the examples disclosed herein may be applicable to use with Figure 1-3 The TMS 102 may be similar to a thermal management system for an engine of another type or configuration. Thus, the present subject matter may be readily adapted for use with another engine and / or another heat transfer application associated with another type of aircraft.

[0052] Figure 3 is a schematic diagram of an example implementation of a TMS 102 for transferring heat between fluids. In general, the TMS 102 is as described above and in Figure 1 and Figure 2 1 and 10. However, the TMS 102 may be implemented within another type of aircraft and / or another gas turbine engine of another configuration.

[0053] As shown, TMS102 includes a heat transfer bus 302 to transfer a working fluid (e.g., a heat exchange fluid) throughout TMS102. Specifically, heat transfer bus 302 (TTB 302) includes one or more fluid conduits through which the working fluid flows. As described below, the working fluid flows through various heat exchangers so that heat is added to and / or removed from the working fluid. In the example shown, the working fluid can be supercritical carbon dioxide (sCO2), oil, liquid helium, helium-xenon, etc. In addition, TMS102 includes a heat transfer bus pump 304 (TTB pump 304) to pump the working fluid through TTB 302.

[0054] The TTB pump 304 of the TMS102 includes a compressor 304a and a power source 304b connected to a shaft 304c. The compressor 304a is rotatably interlocked with the shaft 304c, and the power source 304b drives the rotation of the shaft 304c. Therefore, the power source 304b drives the rotation of the compressor 304a, and the rotation of the compressor 304a provides a pressure head to the working fluid in the TTB 302 downstream of the TTB pump 304. When the compressor 304a increases the pressure of the working fluid, the flow rate of the working fluid is accelerated downstream toward the low-pressure end (e.g., upstream of the TTB pump 304). Therefore, the TTB pump 304 drives the working fluid through the TMS102.

[0055] In some examples, the compressor 304a is a centrifugal compressor (or radial compressor) and includes one or more impellers. Therefore, the working fluid can enter along the axis of rotation of the compressor 304a and accelerate radially outward from the rotating impeller into the outlet port, which produces an increased pressure head. In some examples, the compressor 304a is a single-stage centrifugal compressor with a single impeller mounted on the end of the shaft 304c. In some examples, the compressor 304a is a multi-stage centrifugal compressor with multiple impellers installed serially along the shaft 304c. In some examples, the compressor 304a is an axial compressor with a multi-stage rotor and a stator that sequentially increase the pressure of the working fluid. Therefore, the working fluid can enter along the axis of rotation of the compressor 304a and pressurize / accelerate along a flow path parallel to the axis of rotation. In some examples, the compressor 304a is a rotary screw compressor, which includes two adjacent shafts with a helical thread meshed together. When the working fluid radially enters the rotary screw compressor, the two shafts rotate and force the fluid to flow axially along the shaft. Fluid pressure increases as the rotating threads drive flow into the outlet where the working fluid radially exits the rotary screw compressor.

[0056] Power source 304b generates torque from electrical power and / or mechanical power and transmits the torque to compressor 304a along shaft 304c. In some examples, power source 304b is an electric motor (e.g., a direct current (DC) brushless motor, etc.) including a field magnet that emits a magnetic field and an armature (or armature winding) that generates an alternating electromagnetic field. Based on the type of example motor implemented as power source 304b, the stator or rotor can be configured as an armature. In such an example, the motor includes a rotor and a stator, and the rotor is coupled to shaft 304c, which rotates based on the magnetic interaction between the field magnet and the armature.

[0057] In some examples, the power source 304b is a turbine that extracts thermal energy of the working fluid in the TTB 302 to generate mechanical power. This configuration includes multiple sequential stages of rotating rotor blades and stationary stator blades that generate mechanical power based on the kinetic energy and thermal energy of the working fluid. In some examples, the TTB pump 304 includes a separate motor that is coupled to the shaft 304c in conjunction with the power source 304b (e.g., a turbine) to supplement the power available to the compressor 304a. In some examples, the power source 304b is configured in another way to provide power corresponding to a certain pressure output of the TTB pump 304 to the compressor 304a.

[0058] The TMS 102 includes one or more heat source exchangers 306 arranged along the heat transfer bus 302. More specifically, the heat source exchangers 306 are fluidly coupled to the heat transfer bus 302 so that the working fluid flows through the heat source exchangers 306. In some examples, the heat source exchangers 306 transfer heat from the fluid supporting the operation of the aircraft 10 to the working fluid, which cools the fluid supporting the operation of the aircraft 10 and heats the working fluid. Figure 3 Two heat source exchangers 306 are shown, but Figure 1-3 The TMS 102 may include a single heat source exchanger 306 or multiple (eg, two, three, five, etc.) heat source exchangers 306 .

[0059] Heat source exchangers 306 may correspond to heat exchangers for cooling fluids supporting the operation of aircraft 10. In some examples, at least one of heat source exchangers 306 is coupled to a lubrication system of engine 100. Thus, heat source exchangers 306 may transfer heat from oil lubricating engine 100 to the working fluid. In some examples, at least one of heat source exchangers 306 is coupled to a cooling system of engine 100. Thus, heat source exchangers 306 may transfer heat from cooling air exhausted from compressor section 122 (or compressor discharge plenum) of engine 100 to the working fluid. In some examples, heat source exchangers 306 correspond to another way of cooling fluids supporting the operation of aircraft 10.

[0060] In addition, TMS 102 includes a plurality of heat sink exchangers 308 arranged along heat transfer bus 302. More specifically, heat sink exchangers 308 are fluidly coupled to heat transfer bus 302 such that the working fluid flows through heat sink exchangers 308. In some examples, heat sink exchangers 308 transfer heat from the working fluid to other fluids supporting the operation of aircraft 10, which heats the other fluids supporting the operation of aircraft 10 and cools the working fluid. Figure 3 Two heat sink exchangers 308 are shown, but Figure 1-3 The TMS 102 may include a single heat sink exchanger 308 or multiple (eg, two, three, five, etc.) heat sink exchangers 308 .

[0061] Radiator exchangers 308 may correspond to heat exchangers used to heat fluid to support operation of aircraft 10. In some examples, at least one of radiator exchangers 308 is coupled to a fuel system of engine 100. Thus, radiator exchangers 308 may transfer heat from the working fluid to the fuel flowing through the working fluid. Figure 2 The fuel of the fuel supply flow line 170 of the engine 100. In some examples, such as Figure 2 and Figure 3 In the embodiment of the present invention, at least one of the heat exchangers 308 is in contact with the first portion 156 of the air 152 flowing through the bypass air flow passage 120 of the engine 100. Therefore, the heat exchanger 308 can transfer heat from the working fluid to the first portion 156 of the air 152 flowing through the bypass air flow passage 120.

[0062] Figure 3 The TMS 102 includes one or more bypass conduits 310 to allow the working fluid to bypass the heat source exchanger 306 or the radiator exchanger 308. As shown, each of the bypass conduits 310 is fluidly coupled to the heat transfer bus 302 so that at least a portion of the working fluid can bypass the heat source exchanger 306 or the radiator exchanger 308. In addition, the TMS 102 includes one or more heat source valves 312 to allow a portion of the working fluid to enter the bypass conduit 310 and bypass one or more heat source exchangers 306. The TMS 102 also includes one or more radiator valves 314 to allow a portion of the working fluid to enter the bypass conduit 310 and bypass one or more of the radiator heat exchangers 308. The heat source valve 312 and the radiator valve 314 are fluidly coupled to the heat transfer bus 302 and the corresponding bypass conduit 310. In addition, the heat source valve 312 and the radiator valve 314 can be moved between fully and / or partially open and / or closed positions to selectively block the heat exchange flow through the bypass conduit 310.

[0063] In some examples, heat source valve 312 is opened to allow a portion of the working fluid to enter bypass conduit 310, bypass heat source exchanger 306, and reduce the temperature and pressure of the working fluid in TTB 302 (e.g., to inhibit accelerated wear of TTB 302). In some examples, radiator valve 314 is opened to allow a portion of the working fluid to enter bypass conduit 310, bypass radiator exchanger 308, and increase the temperature and pressure of the working fluid in TTB 302 (e.g., to improve the heat transfer capacity of the working fluid). Figure 3 As shown, each heat source exchanger 306 and each radiator exchanger 308 has a corresponding bypass conduit 310. In some examples, other numbers of heat source exchangers 306 and / or radiator exchangers 308 may have corresponding bypass conduits 310. In some examples, the heat source valve 312 and the radiator valve 314 are hydraulically controlled based on the pressure of the working fluid within the heat transfer bus 302. Thus, for example, when the pressure of the heat exchange fluid meets a certain pressure threshold, the heat source valve 312 opens to allow a portion of the working fluid to bypass the heat source exchanger 306.

[0064] We now discuss certain relationships that the inventors unexpectedly discovered after building and operating several TMS systems and jet engines as described above. These relationships are combined Figure 3 The TTB pump 304 shown in FIG. Figure 1 and Figure 2 It will be appreciated that these relationships can be used to design an engine 100 that can be used in Figure 1-3 TMS102, Figure 1-2 Among other advantages, when implementing the relationships set forth below to select or design a pump, the relationships discovered identify TTB pumps that improve SFC for certain engine thrust classes without exceeding the benefits of SFC.

[0065] By designing and constructing several engines and TTB pumps, the inventors of the present disclosure have discovered that a TTB pump can be designed to generate a desired amount of pump power based on the thrust level, thereby improving the SFC of a gas turbine engine at that thrust level. Prior to this discovery, the inventors set out to design various TTB pumps that generate different pump powers and include different rotor diameters, while determining the effectiveness of the TTB pump to improve the SFC of different thrust levels, as described above. In the process of designing various TTB pumps, the inventors discovered that the relationship between i) the pump power of the TTB pump and ii) the fuel flow rate of the engine during cruise conditions leads to a compact and lightweight TTB pump design (relative to the engine size), thereby generating fuel consumption benefits. Relatedly, the inventors unexpectedly discovered that there is a relationship between i) the pump rotor diameter and ii) the fuel flow rate to provide a radially compact pump and generate fuel consumption benefits (i.e., lower SFC). The inventors discovered that when one or both of these relationships are applied to the pump design, a lower SFC of a gas turbine engine used to generate aircraft thrust can be achieved while avoiding a TTB pump that is larger and / or heavier than required.

[0066] The unexpected discovery between the pump mass flow rate (the mass flow rate of the working fluid leaving the TTB pump) and the aircraft fuel mass flow rate (the mass flow rate of the fuel injected into the combustion chamber) identifies the pump power and rotor diameter that are appropriate for a given engine thrust rating. The selected TTB pump design neither oversizes nor undersizes the pump for a given engine thrust rating and provides the desired SFC benefit. The disclosed relationship also eliminates inefficient TTB pump designs, such as TTB pumps that provide too little SFC benefit or TTB pumps that are too large for a given engine thrust rating.

[0067] The disclosed relationship enables the TTB pump to pressurize the working fluid to a mass flow rate that improves the SFC of the engine while maintaining the effective size of the TTB pump. Figure 6 and Figure 7 Some example TTB pump embodiments disclosed herein can increase the pressure of the working fluid to offset the pressure drop on the TMS. In addition, some example TTB pump embodiments disclosed herein can improve SFC while taking into account the effectiveness of the radiator exchanger connected to the TTB and the fuel supply pipeline, the type of working fluid used in the TTB (e.g., supercritical carbon dioxide (sCO2), liquid helium (He), helium xenon, etc.), the type of fuel used in the aircraft (e.g., hydrocarbon fuel (Jet-A), hydrogen fuel (H2), etc.), the efficiency of the gas turbine engine, etc.

[0068] Reference now Figure 4, which illustrates the above findings in graphical form, reveals a relationship between the properties of a pressurized working fluid in a heat transfer bus (TTB) and the fuel properties of the fuel consumed by an aircraft thrust producing gas turbine engine. The inventors discovered this relationship after developing several embodiments taking into account different types, sizes and performance characteristics of pumps for various engine thrust classes. If the demand for working fluid pressure for heat transfer is not carefully matched to the fuel rate, volume and temperature environment of a particular engine during operation (cruise conditions), the fuel combustion benefits (i.e., lower SFC) of selecting an appropriately sized pump for a particular engine size cannot be achieved. A pump for the working fluid that is too large (capable of providing more pumping power than is required to transfer heat to the top of the fuel) or too small (inability to achieve efficient heat transfer from the working fluid to the fuel) will result in an unsatisfactory improvement in the SFC of the gas turbine engine during aircraft cruise conditions.

[0069] Figure 4 A first graph 400 is shown relating a corrected pump power 402 (eg, Y-axis) to a corrected fuel flow rate 404 (eg, X-axis). The corrected fuel flow rate 404 represents the engine thrust level during cruise flight operation. Figure 3 In the example of , the corrected pump power 402 is the pump power output of the TTB pump 304 normalized based on the properties (specific heat capacity and density) of the selected working fluids of water and the TTB 302. Figure 2 In the example of FIG. 4 , the corrected fuel flow rate 404 is normalized based on the properties of the fuel (eg, Jet-A, H2, etc.) and the selected working fluid (specific heat capacity). Figure 1 and Figure 2 The inventors of the present disclosure have discovered that Figure 4 The first design space 406 (shaded area) of represents a range of TTB pump designs that were found to be effective in improving SFC across various engine thrust classes. The first design space 406 defines a range of corrected pump powers that can provide a working fluid (e.g., sCO2, liquid helium, helium xenon, etc.) at a mass flow rate that effectively heats the fuel for a given corrected fuel flow rate. Through their work, the inventors found that when the corrected pump power of the TTB pump 304 does not correspond to the first design space 406, the TTB pump 304 is either too small to provide an SFC benefit for a given engine thrust class, or too large to effectively improve SFC without being oversized.

[0070] The Y-axis of the first graph 400 represents the corrected pump power 402 of the TTB pump 304 and is measured in horsepower (HP). During operation, the TTB pump 304 generates a pump power P pump, to pressurize the working fluid to the pressure head in the TTB 302. The corrected pump power 402 represents the normalized pump power that the TTB pump 304 can generate to pressurize water to the same pressure head in the TTB 302. The pump power P pump is normalized to the corrected pump power or POW. The inventors have found that based on the different working fluids considered for the TTB pump 304 design and / or when the pump power P pump The corrected pump power POW is calculated based on the view that POW does not change dramatically when different working fluid properties are used. pump is more suitable for defining the first design space 406. is associated with the working fluid and is Figure 4 The fluid properties expressed in the relationship and implemented in the further aspects below include a first specific heat capacity c p_pump and the first density ρ pump The fluid properties associated with water and implemented in the disclosed relationship also include a second specific heat capacity c p_water and the second density ρ water The corrected pump power POW is obtained from (1):

[0071]

[0072] The corrected fuel flow rate 404 is represented by Figure 4 The x-axis of the first graph 400 in FIG. 4 is in pounds mass per second (lbm / s). The corrected fuel flow rate 404 represents the behavior of the engine across various thrust levels. It has been found that this well explains the differences between engines of different thrust levels. During cruise operation, the engine 100 generates engine power P engine The fuel is burned at a fuel flow rate of 1.5 %. When the fuel burns, the chemical reaction releases a net heat of combustion NHC fuel It has been found that the fuel flow rate of the engine 100 can be controlled by the engine power P engine Net heat of combustion with fuel NHC fuel The fuel flow rate is well approximated by the ratio of the working fluid in the engine 100 and the working fluid in the TTB 302. The fuel flow rate is normalized using the properties of the fuel in the engine 100 and the working fluid in the TTB 302 to provide a corrected fuel flow rate FFR. The inventors have found that the corrected fuel flow rate FFR can more accurately account for changes in engine thrust levels than the fuel flow rate because the corrected fuel flow rate FFR does not change dramatically when different working fluids are used to transfer heat to the fuel and / or when different fuels are used to power an engine (such as the engine 100). The fluid property associated with the fuel is called the third specific heat capacity c p_fuel and the third density ρ fuel The corrected fuel flow rate FFR is obtained from (2):

[0073]

[0074] Figure 4 Graph 400 identifies a baseline corrected pump power curve 408 (baseline curve 408). A TTB pump that generates a corrected pump power POW and follows baseline curve 408 may provide an SFC benefit for a given engine thrust level (represented by an associated corrected fuel flow rate FFR). Figure 4 The baseline curve 408 is shown as follows:

[0075] POW=FFR 5 / 3 (3)

[0076] In developing an embodiment of the first design space 406, the inventors discovered that a baseline operating curve that produces a desired SFC benefit while avoiding an oversized (or undersized) pump can vary between an upper limit corrected pump power curve 410 (upper curve 410) and a lower limit corrected pump power curve 412 (lower curve 412). Figure 4 The upper curve 410 of is represented as follows:

[0077] POW=12*FFR 5 / 3 (4)

[0078] Figure 4 The lower curve 412 of is represented as follows:

[0079] POW=0.008*FFR 5 / 3 (5)

[0080] The first design space 406 is defined as follows:

[0081] 0.008≤POW / FFR 5 / 3 ≤12 (6)

[0082] Figure 5 A second graph 500 is shown relating a corrected rotor diameter 502 (eg, Y-axis) to a corrected fuel flow rate 404 (eg, X-axis). The corrected rotor diameter 502 is based on water and TTB 302 ( Figure 3 ) The properties (specific heat capacity) of the working fluid in Figure 3 The inventors found that Figure 5The second design space 506 of aptly represents a range of TTB pump designs that effectively improve SFC for various engine thrust classes. The second design space 506 defines a range of rotor diameters that can provide a working fluid (e.g., sCO2, liquid helium, helium xenon, etc.) at a mass flow rate that effectively heats the fuel for a given corrected fuel flow rate. Through their work in developing embodiments of the second design space 506, the inventors discovered that when the outlet rotor diameter of the TTB pump 304 does not correspond to the second design space 506, the TTB pump 304 is either too small to provide SFC benefits for a given engine thrust class, or too large to effectively improve SFC without being oversized.

[0083] The Y-axis of the second graph 500 representing the corrected rotor diameter 502 is sized in inches. During operation, the TTB pump 304 rotates to have an outlet rotor diameter D rotor 302 to pressurize the working fluid to a pressure head in the TTB 302. The corrected rotor diameter ERD represents the normalized outlet rotor diameter of the TTB pump 304 rotor. The normalized outlet rotor diameter is the diameter of the rotor of the selected pump (e.g., axial, centrifugal, single stage, or multi-stage) that would produce an equivalent desired pressure head if water were used as the pumped fluid instead. In some examples, the TTB pump 304 includes a single-stage centrifugal compressor having an impeller with a rotor diameter D rotor Corresponds to the outlet diameter of the impeller. In some examples, the TTB pump 304 includes a multi-stage centrifugal compressor having multiple impellers, wherein the rotor diameter D rotor Corresponds to the outlet diameter of the most upstream impeller of the plurality of impellers. In some examples, the TTB pump 304 includes an axial compressor having a plurality of rotors, wherein the rotor diameter D rotor Corresponds to the outlet diameter of the most upstream rotor among the multiple rotors.

[0084] Based on the properties of water and working fluid, the rotor diameter D rotor Normalized to the corrected rotor diameter ERD. The inventors found that the corrected rotor diameter ERD is larger than the rotor diameter D rotor is more suitable for defining the second design space 506 because the corrected rotor diameter ERD does not change drastically when different working fluids are considered for the TTB pump 304 design. The fluid properties used for ERD include the first specific heat capacity c p_pump and the second specific heat c p_water The corrected rotor diameter ERD is obtained from (7):

[0085]

[0086] Figure 5The second graph 500 of FIG. 5 identifies a baseline corrected rotor diameter curve 508 (baseline curve 508). A TTB pump along the baseline curve 508 (represented by points along the curve 508) with a corrected rotor diameter ERD may provide an SFC benefit for a given engine thrust level (represented by an associated corrected fuel flow rate FFR). Figure 5 The baseline curve 508 is represented as follows:

[0087] ERD=0.88*FFR 0.1 (8)

[0088] In designing embodiments within the second design space 506, the inventors discovered that a rotor diameter baseline operating curve that produces a desired SFC benefit (and while avoiding an oversized pump) can vary between an upper limit corrected rotor diameter curve 510 (upper curve 510) and a lower limit corrected rotor diameter curve 512 (lower curve 512). Figure 5 The upper curve 510 of is represented as follows:

[0089] ERD=3.0*FFR 0.1 (9)

[0090] Figure 5 The lower curve 512 of is represented as follows:

[0091] ERD=0.4*FFR 0.1 (10)

[0092] The second design space 506 is defined as follows:

[0093] 0.4≤ERD / FFR 0.1 ≤3.0 (11)

[0094] Referring to the baseline curves 408, 508, a larger corrected fuel flow rate 404 is associated with a larger corrected pump power 402 and / or a larger corrected rotor diameter 502 to increase the SFC for the corresponding engine thrust level. However, in some examples, for a given engine thrust level (corrected fuel flow rate 404), based on various performance factors of the TTB pump 304, TMS 102, aircraft 10 and / or gas turbine engine 100, the TTB pump 304 can be designed to have a higher or lower pump power or rotor diameter, as described below. In some examples, the first and second design spaces 406, 506 in which the desired TTB pump 304, power source 304b design (corrected pump power POW) and / or compressor 304a design (corrected rotor diameter ERD) are located are constrained by 0.05≤FFR≤16lbm / s on the X-axis. In some other examples, the first and second design spaces 406, 506 are constrained by 0.25≤FFR≤15lbm / s on the X-axis. In other examples, for lower thrust rating engines, the first and second design spaces 406, 506 are constrained on the X-axis by 0.05≤FFR≤3.5lbm / s. In other examples, for higher thrust rating engines, the first and second design spaces 406, 506 are constrained on the X-axis by 3.5≤FFR≤16lbm / s. An efficient TTB pump 304 (e.g., compressor 304a, power source 304b, etc.) can be designed and / or constructed to provide a desired mass flow rate of the working fluid while remaining below an inefficiently large pump size to achieve improved heat transfer to the fuel upstream of the combustion chamber, which provides an SFC benefit for a given engine thrust rating.

[0095] The TTB pump embodiments defined by the first and / or second design spaces 406, 506 (disclosed below) may implement Figure 3 The TTB pump 304. Figure 4 The embodiment indicated by the baseline curve 408 in represents a specific example of certain TTB pump configurations with a corrected pump power POW for certain engine architectures. The embodiment indicated along the baseline curve 508 represents a TTB pump with a corrected rotor diameter ERD for the same or similar engine architecture. Other embodiments that are not along the baseline curves 408, 508 but fall within the first and second design spaces 406, 506 (e.g., embodiments that satisfy the relationship in (5) or (10)) may also be applicable to a given engine design and thrust level. For example, a TTB pump embodiment with a pump power POW corresponding to the upper curve 410 of the first design space 406 (e.g., an embodiment that satisfies the relationship in (4)) outputs greater pressure and flow rate to provide a certain SFC benefit to the engine without becoming too large. The following provides an example engine architecture that can affect the design of a TTB pump embodiment, but does not depart from the scope of the present invention and benefits from the above relationship.

[0096] like Figure 4 and Figure 5 As shown, for a given thrust level, the corrected pump power POW and the corrected rotor diameter ERD can be determined based on the corrected fuel flow rate FFR. engine It can be used to determine the pump power P of the TTB pump 304 pump and / or rotor diameter D impeller A guide to improving the SFC of the engine 100 is provided.

[0097] The effective corrected pump power POW of the TTB pump 304 with an effectively sized power source 304b (e.g., electric motor, turbine, etc.) to improve the SFC of the engine 100 is defined by the above expressions (1)-(6). The ranges of physical properties or characteristics of the engine and TMS for which these relationships are defined and effective are set forth below in Table 1:

[0098]

[0099] Table 1: Parameter variable range

[0100] The effective corrected rotor diameter ERD of a TTB pump 304 having an effective size of compressor 304a (e.g., rotor, impeller, etc.) to improve the SFC of the engine 100 is defined by the above expressions (2) and (7)-(11). The ranges of physical properties or characteristics of the engine and TMS for which these relationships are defined and effective are set forth below in Table 2:

[0101]

[0102] Table 2: Parameter variable range

[0103] In addition to identifying the above relationships between pump properties and fluid characteristics and the engine properties to which these relationships apply, further considerations may be made when designing the TTB pump 304. Such considerations include performance factors of the TMS 102 and / or the gas turbine engine 100. With reference to these embodiments, various TTB pump embodiments (disclosed below) conform to the size and performance constraints set within the first and second design spaces 406, 506, while also satisfying the performance factors mentioned below. For example, the pressure drop across the TMS 102 may affect the selection of the disclosed TTB pump. Certain designs or configurations of the TMS 102 may affect the pressure drop across the system and define the required pressure increase (or pressure differential) of the TTB pump 304. In some examples, the length of the TTB 302, the number and size of the heat source exchangers 306, and / or the number and size of the radiator exchangers 308 may affect the pressure differential that the TTB pump 304 needs to provide. For example, a large number (e.g., five, ten, etc.) of radiator exchangers 308 in a TMS 102 may result in a significant pressure drop across the TTB 302, which requires the TTB pump 304 to provide a larger pressure differential. Figure 6 and Figure 7 ) can overcome certain pressure drops in the TMS while improving SFC and complying with the first and second design spaces 406, 506. Through the design of these TTB pumps, the inventors unexpectedly discovered the parameter relationships disclosed above.

[0104] Figure 6 and Figure 7 Various examples of TTB pumps that fall within the first and second design spaces 406, 506 disclosed herein are shown. Such TTB pump embodiments can improve SFC and account for a range of differential pressures across the TMS. Figure 6 A first graph 400 is shown relating a corrected pump power 402 (eg, Y-axis) and a corrected fuel flow rate 404 (eg, X-axis). Figure 6 The first graph 400 identifies a particular TTB pump embodiment within a first design space 406 discovered by the inventors. Figure 7 A second graph 500 is shown relating the corrected rotor diameter 502 (eg, Y-axis) and the corrected fuel flow rate 404 (eg, X-axis). Figure 7 The second graph 500 identifies a particular TTB pump embodiment within a second design space 506 .

[0105] Figure 6 Example corrected pump power POW values ​​corresponding to various TTB pump embodiments disclosed with respect to the first design space 406 are shown. Figure 7Example corrected rotor diameter ERD values ​​corresponding to various TTB pump embodiments disclosed with respect to the second design space 506 are shown. The TTB pump embodiments shown reflect engines associated with certain engine thrust levels. Different example engines (e.g., engine 100, etc.) can be associated with different thrust levels based on the corrected fuel flow rate FFR of the engine consumed fuel. In some examples, the first engine 100a, the second engine 100b, the third engine 100c, the fourth engine 100d, and the fifth engine 100e burn Jet-A fuel at certain corrected fuel flow rates. In some examples, the sixth engine 100f, the seventh engine 100g, the eighth engine 100h, the ninth engine 100i, and the tenth engine 100j burn hydrogen (H2) fuel at certain corrected fuel flow rates. Table 3 below describes the characteristics of the fuel associated with the example TTB pump embodiments disclosed herein, engine 100a-100j (when operating in cruise flight) and the example TTB pump embodiments disclosed herein.

[0106]

[0107] Table 3: Engine and fuel values

[0108] Figure 6 Example corrected pump power POW values ​​for a TTB pump embodiment for the first design space 406 are shown. Figure 6 The data points marked in refer to the first example TTB pump 602, the second example TTB pump 604, the third example TTB pump 606, the fourth example TTB pump 608, the fifth example TTB pump 610, the sixth example TTB pump 612, the seventh example TTB pump 614, the eighth example TTB pump 616, the ninth example TTB pump 618, and the tenth example TTB pump 620. Table 4 below provides example values ​​corresponding to the TTB pump embodiments disclosed herein. The first, second, third, and fourth TTB pumps 602-608 have corrected pump power POW values ​​that define the lower curve 412 of the first design space 406. The fifth, sixth, seventh, eighth, and ninth TTB pumps 610-618 have corrected pump power POW values ​​that define the upper curve 410 of the first design space 406.

[0109] Figure 7 Example corrected rotor diameter ERD values ​​for a TTB pump embodiment for the second design space 506 are shown. Figure 7 The data points marked in are those related to Figure 6 The same example TTB pumps 602-620 shown in FIG. The TTB pump designs disclosed herein can be configured to pump a variety of working fluids (e.g., sCO2, He, etc.), or can include larger or smaller pump powers, rotor diameters, etc., while maintaining the effectiveness levels represented by relationships (1)-(11).

[0110]

[0111] Table 4: TTB Pump Examples and Working Fluid Values

[0112] The TTB pumps 602-620 shown in Table 4 are configured to operate within a range of calibrated pump powers POW to pressurize sCO2 to a range of total pressures (PTA), such as between 1000 and 4000 pounds per square inch (psi), etc. Additional embodiments (in Figure 6-7 602-620). Such TTB pump embodiments may be configured to pump sCO2, He, and / or other working fluids with a PTA range that may be similar to the PTA of TTB pumps 602-620. Table 5 below provides additional example values ​​corresponding to TTB pump embodiments disclosed herein and within design spaces 406 and 506.

[0113]

[0114] Table 5: TTB Pump Examples and Working Fluid Values

[0115] The TTB pumps of the aforementioned embodiments satisfying the relationships (1)-(11) may include a single-stage centrifugal compressor (sometimes referred to as a single-stage radial compressor), a multi-stage centrifugal compressor (sometimes referred to as a multi-stage radial compressor), a single-stage axial compressor, a multi-stage axial compressor, a rotary screw compressor, etc. Therefore, various TTB pump embodiments of various configurations may be designed based on the disclosed relationships (1)-(11), and the TTB pump embodiments have a corrected pump power POW falling within the first design space 406 and / or a corrected rotor diameter ERD falling within the second design space 506. The engines 100a-100j of Table 3 are those that can be used in Figure 1 The exemplary engine implemented in the aircraft 10 or another exemplary aircraft, and the exemplary TTB pump embodiments of Tables 4 and 5 may be implemented in the exemplary engines 100a-100j. However, the relationships disclosed herein are not limited to the engines 100a-100j represented in Table 3, the TTB pumps 602-620 represented in Table 4, or the TTB pump embodiments of Table 5. Instead, other TTB embodiments may be designed for other engine thrust classes based on the disclosed relationships that still fall within the first and second design spaces 406, 506 and provide the physical characteristics of the system in accordance with Table 2.

[0116] Although the TTB pump embodiments of Tables 4 and 5 are designed to pump sCO2 or He as a working fluid, other TTB pump embodiments can be designed based on the disclosed relationships to pump other working fluids (e.g., helium xenon, water, oil, etc.). Although the example engines 100a-100j of Table 3 are designed to combust Jet-A or H2 to generate thrust, other fuels (e.g., biodiesel, etc.) can be combusted by the engines 100a-100j or other engines at a corrected fuel flow rate FFR corresponding to the first and second design spaces 406, 506 developed by the inventors.

[0117] Figure 8 8 is a flow chart illustrating a method 800 for operating a gas turbine engine (e.g., engines 100a-100e, etc.) used in an aircraft according to the teachings disclosed herein. The method 800 ensures that a TTB pump (e.g., TTB pumps 602-620, etc.) operates at a pump power that provides an SFC benefit to the gas turbine engine during a cruise phase of flight of an aircraft (e.g., aircraft 10, etc.). The method 800 begins at box 802, where the gas turbine engine provides a fuel flow rate to generate an engine power associated with a cruise condition of the aircraft. At box 804, the gas turbine engine uses a pump to circulate a working fluid (e.g., sCO2, etc.) through a heat transfer bus (e.g., TTB 302, etc.) to heat the fuel in the gas turbine engine. The pump generates and combines Figure 4 The described parameter relationships correspond to the pump power and / or have a combination with Figure 5 The parameter relationships described correspond to the outlet rotor diameter. After block 804, the gas turbine engine continues to provide the fuel flow rate and use the pump until operation of the gas turbine engine of the aircraft during cruise conditions ends.

[0118] Based on the above, it should be understood that the disclosed relationship enables the manufacture of a heat transfer bus (TTB) pump that can pressurize a working fluid in a thermal management system (TMS) to transfer heat to the fuel in an aircraft engine before combustion. Transferring heat to the fuel via a working fluid pressurized with a TTB pump disclosed herein can improve the specific fuel consumption (SFC) corresponding to the engine thrust level of the aircraft engine during cruise operation. The disclosed parameter relationship can be used to accelerate the design or selection process of a TTB pump for a certain engine thrust level. In addition, the relationship between the corrected pump power and the engine thrust level (corrected fuel flow rate) enables the manufacture of a TTB pump that operates with an effective pump power to pressurize the working fluid without being too large. In other words, the disclosed parameter relationship between the corrected rotor diameter and the engine thrust level enables the manufacture of a TTB pump that provides an effective mass flow rate of the working fluid without being too large to improve the SFC (associated with the thrust level). Considering the importance of limiting the additional weight of the aircraft, the parameter relationship ensures that the TTB pump is not too large to exceed the performance advantages provided. Thus, the example TTB pump disclosed herein can improve the operability and efficiency of a turbine engine while limiting the size of the TTB pump in the TMS. In other words, the disclosed parameter relationships can determine the pump power for a thrust class, which results in the TTB pump being appropriately sized so that the TTB pump provides SFC benefits to an engine of that thrust class without negatively impacting the weight of the engine.

[0119] Further aspects of the disclosure are provided by the subject matter of the following clauses:

[0120] An aircraft, comprising: a gas turbine engine, the gas turbine engine being configured to generate an engine power (P engine ) of a fuel flow rate to burn fuel, the fuel having a first specific heat capacity (c p_fuel ) and net heat of combustion (NHC fuel ) characterizes; and a thermal management system comprising a heat transfer bus configured to transfer heat from a working fluid to the fuel, the working fluid being characterized by a second specific heat capacity (c p_pump ) and the first density (ρ pump ) characterizes that the heat transfer bus includes a pump configured to generate a pump power (P pump ) to pressurize the working fluid in the heat transfer bus, and wherein

[0121]

[0122] 0.008≤POW / FFR 5 / 3 ≤12, FFR is between 0.05 lb-mass / s and 16 lb-mass / s, and ρwater and c p_water are the density and specific heat of water respectively.

[0123] An aircraft according to any preceding clause, wherein the pump comprises a rotor having a diameter (D rotor ) of the rotor, and wherein And 0.4≤ERD / FFR 0.1 ≤3.0.

[0124] An aircraft as described in any preceding clause, wherein the FFR is between 0.05 lbs-mass / second and 3.5 lbs-mass / second.

[0125] An aircraft as described in any preceding clause, wherein the FFR is between 3.5 lbs. mass / second and 16 lbs. mass / second.

[0126] An aircraft as claimed in any preceding clause, wherein the pump is a single stage radial compressor.

[0127] An aircraft as claimed in any preceding clause, wherein the pump is a multi-stage radial compressor.

[0128] An aircraft as claimed in any preceding clause, wherein the net heat of combustion (NHC fuel ) is between 1.0e6 ft-lbf / lb-mass and 1.0e8 ft-lbf / lb-mass, and wherein the first specific heat capacity (c p_fuel ) is between 100 foot-pounds-force / pound-mass times Rankine and 5000 foot-pounds-force / pound-mass times Rankine.

[0129] An aircraft as described in any preceding clause, wherein the fuel is a hydrocarbon fuel.

[0130] An aircraft as described in any preceding clause, wherein the fuel is hydrogen fuel.

[0131] An aircraft according to any preceding clause, wherein the second specific heat capacity (c p_pump ) is between 100 foot-pounds per pound-mass times Rankine and 5000 foot-pounds per pound-mass times Rankine, and wherein the first density (ρ pump ) is between 0.1 lbs / ft3 and 100 lbs / ft3.

[0132] An aircraft as described in any preceding clause, wherein the working fluid is supercritical carbon dioxide.

[0133] An aircraft as described in any preceding clause, wherein the working fluid is liquid helium.

[0134] A gas turbine engine comprises: a combustion section, the combustion section being engine ) is operated at a fuel flow rate to burn fuel having a first specific heat capacity (c p_fuel ) and net heat of combustion (NHC fuel ) characterization; and a heat transfer bus pump, the heat transfer bus pump with a pump power (P based on the fuel flow rate pump ) is operated to pump a working fluid through a heat transfer bus to heat the fuel in the gas turbine engine prior to combustion, the working fluid being a second specific heat capacity (c p_pump ) and the first density (ρ pump ) characterization, where

[0135]

[0136] 0.008≤POW / FFR 5 / 3 ≤12, FFR is between 0.05 lb-mass / s and 16 lb-mass / s, and ρ water and c p_water are the density and specific heat of water respectively.

[0137] A gas turbine engine according to any preceding clause, wherein the heat transfer bus pump comprises a rotor having a diameter (D rotor ) of the rotor, and wherein And 0.4≤ERD / FFR 0.1 ≤3.0.

[0138] A gas turbine engine as in any preceding clause, wherein the FFR is between 0.05 lbs-mass / second and 3.5 lbs-mass / second.

[0139] A gas turbine engine according to any preceding clause, wherein the FFR is between 3.5 lbs-mass / second and 16 lbs-mass / second.

[0140] A gas turbine engine according to any preceding clause, wherein the net heat of combustion (NHC fuel ) is between 1.0e6 ft-lbf / lb-mass and 1.0e8 ft-lbf / lb-mass, and wherein the first specific heat capacity (c p_fuel ) is between 100 foot-pounds-force / pound-mass times Rankine and 5000 foot-pounds-force / pound-mass times Rankine.

[0141] A gas turbine engine according to any preceding clause, wherein the second specific heat capacity (c p_pump ) is between 100 foot-pounds per pound-mass times Rankine and 5000 foot-pounds per pound-mass times Rankine, and wherein the first density (ρ pump) is between 0.1 lbs / ft3 and 100 lbs / ft3.

[0142] A method for operating a gas turbine engine for use in an aircraft, the method comprising: providing a method for generating an engine power (P engine ) of the fuel flow rate, the fuel of the gas turbine engine is composed of a first specific heat capacity (c p_fuel ) and net heat of combustion (NHC fuel ) characterization; and circulating a working fluid through a heat transfer bus to heat the fuel in the gas turbine engine using a pump, the pump having a pump power (P) based on the fuel flow rate of the cruise condition pump ), the working fluid has a second specific heat capacity (c p_pump ) and the first density (ρ pump ) characterization, where

[0143]

[0144] 0.008≤POW / FFR 5 / 3 ≤12, FFR is between 0.05 lb-mass / s and 16 lb-mass / s, and ρ water and c p_water are the density and specific heat of water respectively.

[0145] A method according to any preceding clause, wherein operation of the pump comprises rotating a rotor having a rotor diameter (D rotor ) of the rotor, and wherein And 0.4≤ERD / FFR 0.1 ≤3.0.

[0146] A thermal management system, the thermal management system comprising: a pump, the pump being configured to pump ) operates to pressurize the working fluid in the thermal management system, the pump including a rotor with a defined diameter (D rotor ) of the rotor, the working fluid is composed of a first specific heat capacity (c p_pump ) and the first density (ρ pump ) characterizes; a heat transfer bus coupled to the pump and a heat exchanger, the heat exchanger being configured to transfer heat to the fuel during operation of the gas turbine engine, the gas turbine engine being configured to generate engine power (P) under cruise conditions engine ) of the fuel flow rate to burn the fuel, the fuel having a third specific heat capacity (c p_fuel ) and net heat of combustion (NHC fuel ) characterizes, and where

[0147]

[0148] 0.008≤POW / FFR 5 / 3 ≤12,0.4≤ERD / FFR 0.1 ≤3.0, and the FFR is between 0.05 lb-mass / sec and 16 lb-mass / sec, and ρ water and c p_water are the density and specific heat of water respectively.

[0149] A thermal management system as in any preceding clause, wherein the FFR is between 0.05 lb-mass / second and 3.5 lb-mass / second.

[0150] A thermal management system as in any preceding clause, wherein the FFR is between 0.25 lb-mass / second and 15 lb-mass / second.

[0151] A thermal management system as in any preceding clause, wherein the FFR is between 0.05 lb-mass / second and 3.5 lb-mass / second.

[0152] A thermal management system as in any preceding clause, wherein the FFR is between 3.5 lb-mass / second and 16 lb-mass / second.

[0153] A thermal management system as claimed in any preceding clause wherein P engine Between 1.0e6 ft-lbf / sec and 1.0e8 ft-lbf / sec.

[0154] A thermal management system as claimed in any preceding clause wherein P pump Between 0.0020 hp and 1200 hp.

[0155] A thermal management system as described in any preceding clause, wherein the pump is a single stage radial compressor.

[0156] A thermal management system as described in any preceding clause, wherein the pump is a multi-stage radial compressor.

[0157] A thermal management system as described in any preceding clause wherein the pump is a single stage axial compressor.

[0158] A thermal management system as described in any preceding clause, wherein the pump is a multi-stage axial compressor.

[0159] A thermal management system as described in any preceding clause, wherein the pump is a rotary screw compressor.

[0160] A thermal management system according to any preceding clause, wherein the rotor diameter (D rotor ) between 0.3 inches and 4.4 inches.

[0161] A thermal management system according to any preceding clause, wherein the net heat of combustion (NHC fuel ) is between 1.0e6 ft-lbf / lb-mass and 1.0e8 ft-lbf / lb-mass, and wherein the first specific heat capacity (c p_fuel ) is between 100 foot-pounds-force / pound-mass times Rankine and 5000 foot-pounds-force / pound-mass times Rankine.

[0162] A thermal management system as described in any preceding clause wherein the fuel is a hydrocarbon fuel.

[0163] A thermal management system as described in any preceding clause, wherein the fuel is hydrogen fuel.

[0164] A thermal management system according to any preceding clause, wherein the second specific heat capacity (c p_pump ) is between 100 foot-pounds per pound-mass times Rankine and 5000 foot-pounds per pound-mass times Rankine, and wherein the first density (ρ pump ) is between 0.1 lbs / ft3 and 100 lbs / ft3.

[0165] A thermal management system as described in any preceding clause, wherein the working fluid is supercritical carbon dioxide.

[0166] A thermal management system as described in any preceding clause, wherein the working fluid is liquid helium.

[0167] A system comprising: a gas turbine engine configured to generate an engine power (P engine ) of a fuel flow rate to burn fuel, the fuel having a first specific heat capacity (c p_fuel ) and net heat of combustion (NHC fuel ) characterizes; a thermal management system comprising a heat transfer bus configured to transfer heat from a working fluid to the fuel, the working fluid being characterized by a second specific heat capacity (c p_pump ) and the first density (ρ pump ) characterizes that the heat transfer bus includes a defined rotor diameter (D rotor ) of a pump, and wherein

[0168] 0.4≤ERD / FFR 0.1 ≤3.0, FFR between 0.05 lb-mass / s and 16 lb-mass / s, and ρ water and c p_water are the density and specific heat of water respectively.

[0169] A system according to any preceding clause, wherein the pump is operated at a pump power (Ppump ) operation, and where

[0170]

[0171] 0.008≤POW / FFR 5 / 3 ≤12, FFR is between 0.05 lb-mass / s and 16 lb-mass / s, and ρ water and c p_water are the density and specific heat of water respectively.

[0172] A system as in any preceding clause, wherein the FFR is between 0.05 lb-mass / sec and 3.5 lb-mass / sec.

[0173] A system as in any preceding clause, wherein the FFR is between 3.5 lbs-mass / sec and 16 lbs-mass / sec.

[0174] A system according to any preceding clause, wherein the pump power (P pump ) between 0.0020 hp and 1200 hp.

[0175] A system according to any preceding clause, wherein the rotor diameter (D rotor ) between 0.3 inches and 4.4 inches.

[0176] A system as in any preceding clause, wherein the FFR is between 0.25 lb-mass / sec and 15 lb-mass / sec.

[0177] A system as in any preceding clause, wherein the FFR is between 0.05 lb-mass / sec and 3.5 lb-mass / sec.

[0178] A system as in any preceding clause, wherein the FFR is between 3.5 lbs-mass / sec and 16 lbs-mass / sec.

[0179] A system as described in any preceding clause, wherein the pump is a single stage radial compressor.

[0180] A system as described in any preceding clause, wherein the pump is a multi-stage radial compressor.

[0181] A system as described in any preceding clause, wherein the pump is a single stage axial compressor.

[0182] A system as described in any preceding clause, wherein the pump is a multi-stage axial compressor.

[0183] A system as described in any preceding clause, wherein the pump is a rotary screw compressor.

[0184] A system according to any preceding clause, wherein the power (P engine ) is between 1.0e6 ft-lbf / s and 1.0e8 ft-lbf / s.

[0185] A system according to any preceding clause, wherein the net heat of combustion (NHC fuel ) is between 1.0e6 ft-lbf / lb-mass and 1.0e8 ft-lbf / lb-mass, and wherein the first specific heat capacity (c p_fuel ) is between 100 foot-pounds-force / pound-mass times Rankine and 5000 foot-pounds-force / pound-mass times Rankine.

[0186] A system as described in any preceding clause, wherein the fuel is a hydrocarbon fuel.

[0187] A system as described in any preceding clause, wherein the fuel is hydrogen fuel.

[0188] A system according to any preceding clause, wherein the second specific heat capacity (c p_pump ) is between 100 foot-pounds per pound-mass times Rankine and 5000 foot-pounds per pound-mass times Rankine, and wherein the first density (ρ pump ) is between 0.1 lbs / ft3 and 100 lbs / ft3.

[0189] A system as described in any preceding clause, wherein the working fluid is supercritical carbon dioxide.

[0190] The system of any preceding clause, wherein the working fluid is liquid helium.

[0191] The following claims are incorporated into the Detailed Description by reference, with each claim standing on its own as a separate embodiment of the disclosure.

Claims

1. An aircraft, characterized in that: include: A gas turbine engine configured to generate an engine power P engine The fuel flow rate burns fuel having a first specific heat capacity c p_fuel and net heat of combustion NHC fuel Characterize the engine power P engine In units of foot-pound force / second, the first specific heat capacity c p_fuel The net heat of combustion NHC is expressed in foot-pound force / pound mass times degrees Rankine. fuel Measured in foot-pounds force per pound mass; as well as A thermal management system comprising a heat transfer bus configured to transfer heat from a working fluid having a second specific heat capacity c to the fuel. p_pump and the first density ρ pump Characterized in that the heat transfer bus includes a pump configured to generate a pump power P pump To pressurize the working fluid in the heat transfer bus, the second specific heat capacity c p_pump The first density ρ is expressed in ft-lbf / lbm times Rankine. pump In pounds per cubic foot, the pump power P pump In horsepower, and where 0.008≤POW / FFR 5 / 3 ≤12, POW is the corrected pump power, FFR is the corrected fuel flow rate and is between 0.05 lb-mass / sec and 16 lb-mass / sec, and ρ water and c p_water are the density and specific heat of water respectively.

2. The aircraft according to claim 1, characterized in that in, The pump comprises a rotor having a diameter D rotor The rotor diameter D rotor In inches, and where ERD is the corrected rotor diameter, and 0.4 ≤ ERD / FFR 0.1 ≤3.

0.

3. The aircraft according to claim 2, characterized in that: in, The rotor diameter D rotor Between 0.3 inches and 4.4 inches.

4. The aircraft according to claim 1, characterized in that: in, The FFR is between 0.05 lbs-mass / sec and 3.5 lbs-mass / sec.

5. The aircraft according to claim 1, characterized in that: in, The FFR is between 3.5 lbs-mass / sec and 16 lbs-mass / sec.

6. The aircraft according to claim 1, characterized in that: in, The pump is a single-stage radial compressor.

7. The aircraft according to claim 1, characterized in that: in, The pump is a multi-stage radial compressor.

8. The aircraft according to claim 1, characterized in that: in, The net heat of combustion NHC fuel between 1.0e6 ft-lbf / lb-mass and 1.0e8 ft-lbf / lb-mass, and wherein the first specific heat c p_fuel Between 100 foot-pounds force / pound-mass times Rankine and 5000 foot-pounds force / pound-mass times Rankine.

9. The aircraft according to claim 1, characterized in that: in, The fuel is a hydrocarbon fuel.

10. The aircraft according to claim 1, characterized in that: in, The fuel is hydrogen fuel.

11. The aircraft according to claim 1, characterized in that: in, The second specific heat capacity c p_pump Between 100 foot-pounds per pound-mass times Rankine and 5000 foot-pounds per pound-mass times Rankine, and wherein the first density p pump Between 0.1 lbs / ft3 and 100 lbs / ft3.

12. The aircraft according to claim 1, characterized in that in, The working fluid is supercritical carbon dioxide.

13. The aircraft according to claim 1, characterized in that in, The working fluid is liquid helium.

14. A gas turbine engine, characterized in that: include: The combustion section is operated at an engine power P engine The fuel flow rate of the operation is to burn the fuel, the fuel having a first specific heat capacity c p_fuel and net heat of combustion NHC fuel Characterize the engine power P engine In units of foot-pound force / second, the first specific heat capacity c p_fuel The net heat of combustion NHC is expressed in foot-pound force / pound mass times degrees Rankine. fuel Measured in foot-pounds force per pound mass; as well as A heat transfer bus pump, the heat transfer bus pump with a pump power P based on the fuel flow rate pump The invention is operable to pump a working fluid through a heat transfer bus to heat the fuel in the gas turbine engine prior to combustion, the working fluid having a second specific heat capacity c p_pump and the first density ρ pump Characterize the pump power P pump In horsepower, the second specific heat capacity c p_pump The first density ρ is expressed in ft-lbf / lbm times Rankine. pump The units are pounds per cubic foot, where 0.008≤POW / FFR 5 / 3 ≤12, POW is the corrected pump power, FFR is the corrected fuel flow rate and is between 0.05 lb-mass / sec and 16 lb-mass / sec, and ρ water and c p_water are the density and specific heat of water respectively.

15. The gas turbine engine according to claim 14, characterized in that in, The heat transfer bus pump includes a rotor having a diameter D rotor The rotor diameter D rotor In inches, and where ERD is the corrected rotor diameter, and 0.4 ≤ ERD / FFR 0.1 ≤3.

0.

16. The gas turbine engine according to claim 14, characterized in that in, The FFR is between 0.05 lbs-mass / sec and 3.5 lbs-mass / sec.

17. The gas turbine engine according to claim 14, characterized in that in, The FFR is between 3.5 lbs-mass / sec and 16 lbs-mass / sec.

18. The gas turbine engine according to claim 14, characterized in that in, The net heat of combustion NHC fuel between 1.0e6 ft-lbf / lb-mass and 1.0e8 ft-lbf / lb-mass, and wherein the first specific heat c p_fuel Between 100 foot-pounds force / pound-mass times Rankine and 5000 foot-pounds force / pound-mass times Rankine.

19. The gas turbine engine according to claim 14, characterized in that in, The second specific heat capacity c p_pump Between 100 foot-pounds per pound-mass times Rankine and 5000 foot-pounds per pound-mass times Rankine, and wherein the first density p pump Between 0.1 lbs / ft3 and 100 lbs / ft3.

20. A method for operating a gas turbine engine for use in an aircraft, characterized in that The method comprises: Provided to generate an engine power P associated with the cruising condition of the aircraft engine The fuel flow rate of the gas turbine engine is determined by the first specific heat capacity c p_fuel and net heat of combustion NHC fuel Characterize the engine power P engine In units of foot-pound force / second, the first specific heat capacity c p_fuel The net heat of combustion NHC is expressed in foot-pound force / pound mass times degrees Rankine. fuel In units of foot-pounds-force per pound-mass; and A working fluid is circulated through a heat transfer bus to heat the fuel in the gas turbine engine using a pump having a pump power P based on the fuel flow rate at the cruise condition. pump , the working fluid has a second specific heat capacity c p_pump and the first density ρ pump Characterize the pump power P pump In horsepower, the second specific heat capacity c p_pump The first density ρ is expressed in ft-lbf / lbm times Rankine. pump The units are pounds per cubic foot, where 0.008≤POW / FFR 5 / 3 ≤12, POW is the corrected pump power, FFR is the corrected fuel flow rate and is between 0.05 lb-mass / sec and 16 lb-mass / sec, and ρ water and c p_water are the density and specific heat of water respectively.

21. The method according to claim 20, characterized in that in, The operation of the pump comprises rotating a rotor having a diameter D rotor The rotor diameter D rotor In inches, and where ERD is the corrected rotor diameter, and 0.4 ≤ ERD / FFR 0.1 ≤3.0.

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