Gas turbine engine with a heat exchanger located in an annular duct

Abstract

A gas turbine engine is provided, comprising: a turbine having a compressor section, a combustion section, and a turbine section arranged in a serial flow sequence; a rotor assembly driven by the turbine, the rotor assembly, the turbine, or both including a substantially annular duct relative to a centerline of the gas turbine engine, the substantially annular duct defining a flow path; a heat exchanger positioned within the annular duct and extending substantially continuously in a circumferential direction, the heat exchanger including a first material defining a heat exchange surface exposed to the flow path, wherein the first material defines a heat exchange coefficient, and wherein the heat exchange surface defines a surface area (A), and wherein, for operating conditions, the heat exchanger has an effective transmission loss (ETL) between 5 decibels and 1 decibel.

Description

Technical Field

[0001] The present subject matter generally relates to heat exchangers for gas turbine engines. Background Art

[0002] A gas turbine engine typically consists of a fan and a turbine. The turbine typically includes an inlet, one or more compressors, a combustor, and at least one turbine. The compressor compresses air, which is then directed to the combustor, where it is mixed with fuel. The mixture is then ignited to produce hot combustion gases. The combustion gases are then directed to the turbine, which extracts energy from the combustion gases to power the compressor and generate useful work to propel an aircraft in flight or power a load, such as an electrical generator.

[0003] During operation of a gas turbine engine, various systems may generate relatively large amounts of heat. For example, significant amounts of heat may be generated during operation of the thrust generation system, lubrication system, electric motors and / or generators, hydraulic systems, or other systems. Therefore, a device for dissipating the heat generated by various systems would be advantageous in the art. BRIEF DESCRIPTION OF THE DRAWINGS

[0004] A full and enabling disclosure of the present disclosure, including the best mode thereof, to one of ordinary skill in the art is set forth in the specification with reference to the accompanying drawings, in which:

[0005] Figure 1 is a schematic cross-sectional view of a gas turbine engine according to an exemplary embodiment of the present disclosure.

[0006] Figure 2 is a schematic cross-sectional view of a three-flow engine according to an exemplary embodiment of the present disclosure.

[0007] Figure 3 is a schematic cross-sectional view of a heat exchanger and flow paths according to an exemplary embodiment of the present disclosure.

[0008] Figure 4 is a schematic cross-sectional view of a heat exchanger and flow paths according to another exemplary embodiment of the present disclosure.

[0009] Figure 5 is an exploded perspective view of a heat exchanger according to another exemplary embodiment of the present disclosure.

[0010] Figure 6 is a flow path according to an exemplary embodiment of the present disclosure Figure 5 Schematic cross-sectional view of an exemplary heat exchanger.

[0011] Figure 7 is a schematic perspective view of a heat exchanger according to yet another exemplary embodiment of the present disclosure.

[0012] Figure 8 is a schematic perspective view of a heat exchanger according to yet another exemplary embodiment of the present disclosure.

[0013] Figure 9 is a schematic perspective view of a heat exchanger according to yet another exemplary embodiment of the present disclosure.

[0014] Figure 10 is a schematic perspective view of a heat exchanger according to yet another exemplary embodiment of the present disclosure.

[0015] Figure 11 is a graph illustrating the relationship between ETL and UA at low mass flow rates for a heat exchanger according to one or more exemplary embodiments of the present disclosure.

[0016] Figure 12 Provided include corresponding Figure 11 Table of numerical values ​​for several ETL values ​​plotted in .

[0017] Figure 13 is a graph illustrating the relationship between ETL and UA at a medium mass flow rate for a heat exchanger according to one or more exemplary embodiments of the present disclosure.

[0018] Figure 14 Provided include corresponding Figure 13 Table of numerical values ​​for several ETL values ​​plotted in .

[0019] Figure 15 is a graph illustrating the relationship between ETL and UA of a high mass flow rate for a heat exchanger according to one or more exemplary embodiments of the present disclosure.

[0020] Figure 16 Provided include corresponding Figure 15 Table of numerical values ​​for several ETL values ​​plotted in . DETAILED DESCRIPTION

[0021] Reference will now be made in detail to the present embodiments of the present disclosure, one or more examples of which are illustrated in the accompanying drawings. The detailed description uses numerical and letter designations to refer to features in the drawings. Like or similar designations in the drawings and the description have been used to refer to like or similar parts of the disclosure.

[0022] The word "exemplary" is used herein to mean "serving as an example, instance, or illustration." Any implementation described herein as "exemplary" is not necessarily to be construed as preferred or advantageous over other implementations. Furthermore, unless expressly stated otherwise, all embodiments described herein should be considered exemplary.

[0023] As used herein, the terms “first,” “second,” and “third” may be used interchangeably to distinguish one component from another and are not intended to indicate the position or importance of each component.

[0024] The terms "fore" and "aft" refer to relative positions within a gas turbine engine or vehicle and refer to the normal operating attitude of the gas turbine engine or vehicle. For example, for a gas turbine engine, the front position refers to the position closer to the engine inlet, while the aft position refers to the position closer to the engine nozzle or exhaust.

[0025] The terms "upstream" and "downstream" refer to the relative directions of flow in a path. For example, with respect to fluid flow, "upstream" refers to the direction from which the fluid is flowing, and "downstream" refers to the direction toward which the fluid is flowing. However, as used herein, the terms "upstream" and "downstream" may also refer to electrical current.

[0026] The term "fluid" may be a gas or a liquid. The term "fluid communication" means that a fluid is able to establish a connection between designated areas.

[0027] The singular forms "a," "an," and "the" include plural references unless the context clearly dictates otherwise.

[0028] As used throughout the specification and claims, approximating language is applied to modify any quantitative expression that can be allowed to vary without causing a change in the basic function to which it is related. Therefore, values ​​modified by terms such as "about," "approximately," "roughly," and "substantially" are not limited to the precise values ​​specified. In at least some cases, approximate language can correspond to the precision of an instrument used to measure a value, or the precision of a method or machine used to construct or manufacture a component and / or system. In at least some cases, approximate language can correspond to the precision of an instrument used to measure a value, or the precision of a method or machine used to construct or manufacture a component and / or system. For example, approximate language can refer to a margin of 1%, 2%, 4%, 5%, 10%, 15%, or 20% of a single value, a range of values, and / or an endpoint of a defined range of values. Here and throughout the specification and claims, range limitations are combined and interchanged, and unless the context or language indicates otherwise, such ranges are identified and include all subranges contained therein. For example, all ranges disclosed herein include the endpoints, and the endpoints can be combined independently of each other.

[0029] With respect to a conduit or flow path (e.g., a conduit or flow path in which a heat exchanger is positioned), "substantially annular" refers to a conduit or flow path that is completely annular (i.e., extending continuously and uninterrupted in the circumferential direction except for the heat exchanger), or a partially annular conduit or flow path that has a void volume percentage of at least 50% except for the heat exchanger (e.g., at least 60%, such as at least 70%, such as at least 80%, such as at least 90% void volume percentage except for the heat exchanger). For example, in certain embodiments, a conduit or flow path may include struts or other similar structures extending therethrough, thereby forming a partially annular conduit.

[0030] As used herein, "transmission loss" or "TL" refers to a measurement of the reduction in sound level from a source as it passes through an acoustic barrier. TL is expressed in decibels (dB) and indicates the reduction in sound intensity (at a given frequency) when the pressure wave generating the sound encounters a structure or acoustic barrier, such as a heat exchanger located within an annular flow path.

[0031] The "effective transmission loss," or "ETL," of a component of a gas turbine engine refers to the amount of TL expected from the component of the gas turbine engine during specified operating conditions. ETL is defined in greater detail below. The ETL and TL of the disclosed embodiments are more specifically expressed as the average ETL or TL, respectively, over a frequency bandwidth (e.g., between 1,000 Hertz ("Hz") and 5,000 Hz), or, if the context indicates, at a specific frequency. According to the present disclosure, the range of ETL and TL is at least 1 dB and less than 5 dB.

[0032] As used herein, "UA" refers to the product of the overall heat transfer coefficient (U) of a portion of a heat exchanger exposed to a fluid (e.g., air) passing through a flow path in which the heat exchanger is positioned, and the total surface area (A) of the heat exchanger positioned within the flow path. Units may be expressed as British thermal units per hour per degree Fahrenheit (Btu / (hr-℉)). The ability of that portion of the heat exchanger to reject heat to or accept heat from the fluid is related to the heat transfer properties of the material (e.g., aluminum, steel, metal alloy, etc.) forming the portion exposed to the fluid, or more specifically, to the overall heat transfer coefficient (CTE) of the portion of the heat exchanger exposed to the fluid and the surface area of ​​the portion. The parameter "UA" represents the effect of the CTE and the surface area exposed to the fluid.

[0033] As used herein, "porosity" refers to the void volume of a heat exchanger positioned within a flow path. For example, a heat exchanger may define a flow area at one location and a flow path may define a flow area at the same location (i.e., without the flow area of ​​the heat exchanger). The porosity of the heat exchanger is the ratio of the flow area of ​​the heat exchanger to the flow area of ​​the flow path at that location.

[0034] As used herein, "fan frequency" or "fan pass frequency" refers to the product of the rotational rate (in revolutions per minute or RPM) and the number of fan blades. The unit of fan pass frequency is kilohertz (kHz). A fan may refer to a fan external to a turbine (e.g., a fan located within a duct of a turbofan, such as a fan located in a turbine). Figure 1 14 ), or a fan inside the turbine, such as a fan located downstream of the turbine inlet and upstream of at least one compressor of the turbine (e.g., Figure 2 fan 184).

[0035] As used herein, "mass flow" or "mass flow rate" refers to the rate of mass flow of a fluid through a heat exchanger, the mass flow through piping upstream or downstream of a heat exchanger, or the mass flow through an enclosed volume. The units are pounds mass per second (lbm / sec.).

[0036] The "pressure drop" across an obstacle is the change in fluid pressure that occurs as the fluid passes through the obstacle. The pressure drop is the static pressure of the fluid immediately upstream of the obstacle minus the static pressure of the fluid immediately downstream of the obstacle, divided by the static pressure of the fluid immediately upstream of the obstacle, expressed as a percentage.

[0037] The present disclosure provides various examples of heat exchangers, including "plate-fin" heat exchangers, "tube" heat exchangers, "counter-flow" heat exchangers, "onion" heat exchangers, and "any dedicated channel" for heat exchange.

[0038] As used herein, the term "fin-based" heat exchanger refers to a heat exchanger that utilizes one or more fins that extend into a cooling or heating fluid stream to increase the surface area exposed to the cooling or heating fluid stream, thereby increasing the efficiency of the heat exchanger. Examples of fin-based heat exchangers include plate-fin heat exchangers and pin-fin heat exchangers.

[0039] As used herein, a "plate fin" heat exchanger refers to a heat exchanger having a surface with fins extending therefrom, the fins being configured to increase heat transfer between the surface and a fluid passing over the fins. Figure 5 Describe an example of this type of heat exchanger.

[0040] As used herein, a "pin-fin" heat exchanger refers to a heat exchanger having a first surface and a second surface. Fins and pins extend from the first surface, the second surface, or both surfaces to increase heat transfer between the first and / or second surface and a fluid passing through the fins and pins.

[0041] As used herein, a "tubular" heat exchanger refers to a heat exchanger that includes one or more tubes or other conduits extending through a fluid flow path. Such a heat exchanger can facilitate heat transfer between a fluid passing through the tubes or other conduits and a fluid passing through the fluid flow path. Figure 4 An example of this type of heat exchanger is described.

[0042] As used herein, a "tube-and-sheet" heat exchanger refers to a heat exchanger having a plurality of tubes and a plate with a plurality of holes through which the tubes extend.

[0043] A "shell and tube" heat exchanger is one that consists of a shell containing a large number of tubes. Figures 8 to 10 An example of this type of heat exchanger is described.

[0044] As used herein, a "counterflow" heat exchanger refers to a heat exchanger in which the direction of flow of one working fluid is opposite to the direction of flow of the other working fluid.

[0045] As used herein, an "onion-style" heat exchanger refers to a heat exchanger having converging and diverging sections, wherein the heat exchange features extend through the sections. Examples of this type of heat exchanger can be seen in U.S. Patent Application No. 15 / 858,453, filed on December 29, 2017, and published as U.S. Patent Application Publication No. 2019 / 0204010 (the "'453 Application"), which is incorporated herein by reference in its entirety for all purposes. For example, an embodiment of this type of heat exchanger can be seen in the '453 Application. Figures 2 to 8 and more specifically, in e.g. Figures 2 to 4 (described, for example, in paragraphs

[0024] -

[0040] ), Figure 5 (described, for example, in paragraphs

[0047] -

[0050] ), Figure 6 (described, for example, in paragraphs

[0041] -

[0044] ), Figure 7 (described, for example, in paragraph

[0041] ) and Figure 8 (as described in, for example, paragraphs

[0051] -

[0054] ).

[0046] As used herein, the term "any dedicated channel" heat exchanger refers to any channel created specifically for conveying fluid for the purpose of exchanging thermal energy.

[0047] The term "length" as used herein with respect to a heat exchanger refers to a measurement in the direction of average fluid flow through the heat exchanger from the most upstream edge of the heat exchanger to the most downstream edge of the heat exchanger, the heat exchanger positioned within the fluid flow path.

[0048] The term "medium power operating condition" refers to the operating condition of the engine during the flight phase that occurs after the aircraft climbs to a set altitude and remains level before the aircraft begins to descend (i.e., a cruise operating condition). In addition, the medium power operating condition may refer to a descent operating condition.

[0049] The phrase "low power operating condition" refers to an operating condition of the engine at a power level that is less than the cruise power level during a cruise operating condition. For example, the low power operating condition may refer to a flight idle operating condition, a ground idle operating condition, an approach idle operating condition, etc., wherein the engine is operated at a power level that is less than approximately 85% of the rated power of the engine (e.g., less than approximately 80% of the rated power of the engine).

[0050] The phrase "high power operating condition" refers to an operating condition of the engine at a power level greater than the cruise power level during the cruise operating condition. For example, the high power operating condition may refer to a takeoff operating condition, a climb operating condition, etc.

[0051] As used herein, the terms “first flow” and “second flow” refer to the working gas flow path of the turbine through the core (high pressure compressor, combustor, and high pressure turbine) of the turbine and the fan flow or bypass flow, respectively.

[0052] As used herein, "tertiary flow" refers to a non-primary flow that can increase fluid energy to produce a small amount of total propulsion system thrust. The pressure ratio of the tertiary flow is higher than the pressure ratio of the primary propulsion flow (e.g., bypass or propeller-driven propulsion flow). Thrust can be generated by a dedicated nozzle or by mixing the airflow through the tertiary flow with the primary propulsion flow or core flow, for example, entering a common nozzle.

[0053] In certain exemplary embodiments, the operating temperature of the airflow through the third stream can be below the maximum compressor discharge temperature of the engine, and more specifically, can be below 350 degrees Fahrenheit (e.g., below 300 degrees Fahrenheit, such as below 250 degrees Fahrenheit, such as below 200 degrees Fahrenheit, and at least as high as the ambient temperature). In certain exemplary embodiments, these operating temperatures can facilitate heat transfer to and from the airflow through the third stream and the separate fluid stream. Furthermore, in certain exemplary embodiments, the airflow through the third stream can contribute less than 50% of the total engine thrust (and at least, for example, 2% of the total engine thrust) under takeoff conditions, or more specifically, under operating conditions of rated takeoff power at sea level, static flight speed, and an ambient temperature of 86 degrees Fahrenheit.

[0054] Furthermore, in certain exemplary embodiments, aspects of the airflow through the tertiary stream (e.g., airflow, mixing, or exhaust characteristics), and thereby the aforementioned exemplary percentage contributions to total thrust, may be passively adjusted or purposefully modified during engine operation through the use of engine control features (e.g., fuel flow, motor power, variable stators, variable inlet guide vanes, valves, variable exhaust port geometry, or flow characteristics) to adjust or optimize overall system performance over a wide range of potential operating conditions.

[0055] References to "noise," "noise level," or "perceived noise," or variations thereof, should be understood to include the sound pressure level (SPL) outside the fuselage, the fuselage external noise level, the perceived noise level, the effective perceived noise level (EPNL), the instantaneous perceived noise level (PNL(k)), or the pitch-corrected perceived noise level (PNLT(k)), or one or more duration correction factors, pitch correction factors, or other applicable factors, as defined by the U.S. Federal Aviation Administration (FAA), the European Union Aviation Safety Agency (EASA), the International Civil Aviation Organization (ICAO), the Swiss Federal Office of Civil Aviation (FOCA), or its committees, or other equivalent regulatory or administrative bodies. Where certain ranges of noise levels are provided herein (e.g., in decibels or dB), it should be understood that those skilled in the art will understand methods for measuring and determining such levels without ambiguity or undue experimentation. Methods for measuring and determining one or more noise levels provided herein with reasonable certainty and without undue experimentation by one skilled in the art include, but are not limited to, understanding the reference frames (including, but not limited to, distances, positions, angles, etc.) between the measurement system, the engine, and / or the aircraft relative to the measurement system or other sensor, or atmospheric conditions (including, but not limited to, temperature, humidity, dew point, wind speed and vector, and reference points for their measurement), as may be defined by the FAA, EASA, ICAO, FOCA, or other regulatory or administrative agencies.

[0056] As used herein, the term "community noise" refers to the amount of noise produced by engines and / or aircraft observed on the ground (typically in the community surrounding an airport) during takeoff or landing.

[0057] As provided herein, embodiments of the engines included herein define noise levels between 5 decibels (dB) and 10 dB below the ICAO Annex 16, Volume 1, Chapter 14 noise standards applicable on or after December 31, 2017, for aircraft with a maximum takeoff weight of at least 55 tons. Additionally or alternatively, embodiments of the engines provided herein can attenuate low-frequency noise, such as those that may be transmitted to the ground when the engine is at cruising altitude, or what may be referred to as en-route noise or community noise.

[0058] In certain exemplary embodiments of the present disclosure, a gas turbine engine defining a centerline and a circumferential direction is provided. The gas turbine engine may generally include a turbine and a rotor assembly. The rotor assembly may be driven by the turbine. The turbine, the rotor assembly, or both may define a substantially annular flow path relative to a centerline of the gas turbine engine. The gas turbine engine includes a heat exchanger positioned within the flow path and extending in a circumferential direction, such as extending substantially continuously in the circumferential direction. The heat exchanger may be fully annular, meaning a complete annular structure, or partially annular, such that a portion of the fluid passing through the conduit will not pass through the flow area of ​​the heat exchanger flow, while another portion will pass through the flow area of ​​the heat exchanger flow.

[0059] For example, during descent of an aircraft including a gas turbine engine, a heat exchanger design for the gas turbine engine may be designed for flight idle conditions. When designing a heat exchanger, the objective can generally be stated as meeting a minimum heat transfer capability from the hot fluid to the cold fluid to achieve an acceptable pressure drop across the heat exchanger. Key factors to consider include the mass flow rate through the piping at flight idle conditions and the type or characteristics of the heat exchanger selected.

[0060] However, a heat exchanger optimized for flight idle conditions may prove to be unacceptable during other flight conditions, such as during high power operating conditions where maximum thrust may be required (e.g., takeoff, climb, turns during descent, etc.). During such time periods, a heat exchanger optimized for flight idle may allow an unacceptable amount of noise, whether cabin noise or community noise, to be attenuated therethrough. Given the complex nature of sound propagation through fluids, standard engineering practice to date has been to evaluate the acoustic environment under different flight conditions for a selected heat exchanger, or a heat exchanger optimized for maximum heat transfer with an acceptable pressure drop. And if a selected heat exchanger (i.e., one optimized for pressure drop and heat transfer between fluids) is not expected to provide the desired amount of noise reduction as air passes through the ducts and interior surfaces of the heat exchanger, then the heat exchanger may need to be redesigned to produce less noise during flight conditions (e.g., takeoff). Therefore, standard practice is to optimize the heat exchanger for flight idle, assess whether the heat exchanger produces an acceptable noise level throughout the flight envelope (or, more precisely, allows for an acceptable amount of noise attenuation across the heat exchanger), and if not, redesign, i.e., essentially start over and reoptimize the heat exchanger to reduce the amount of noise produced during the affected flight conditions while still meeting the heat transfer and / or maximum pressure drop requirements. It is best to establish an initial design or design requirements for the heat exchanger at the beginning to avoid this iterative process; that is, establish conditions or constraints on the heat exchanger that meet the engine structural requirements, taking into account the acceptable pressure drop, the expected transmission losses of the air through the annular duct, and the heat transfer requirements at flight idle.

[0061] In several different types of turbines (e.g. Figure 1 and Figure 2 During the design of a heat exchanger (e.g., those shown in FIG), the inventor's practice has been to design a heat exchanger, modify the heat exchanger, and redesign the heat exchanger to meet acoustic requirements, then re-examine the acoustic response, etc. The types of heat exchangers considered in these design iterations (i.e., heat exchanger optimization versus the resulting acoustic environment) include heat exchanger designs using one or more of "fin-based" heat exchangers, "plate-fin" heat exchangers, "shell-and-tube" heat exchangers, "counterflow" heat exchangers, "onion" heat exchangers, "any dedicated channel" heat exchangers, etc. The following are examples of turbine engines and heat exchanger types developed by the inventor.

[0062] Referring now to the accompanying drawings, Figure 1 is a schematic partial cross-sectional side view of an exemplary gas turbine engine 10 that may incorporate various embodiments of the present disclosure. The engine 10 may be configured as a gas turbine engine for an aircraft. Although further described herein as a turbofan engine or a non-ducted engine ( Figure 2 ), but with reference to several examples including engines 10 and 100, the principles set forth in this specification may alternatively be applied to a turboshaft engine, a turboprop engine, or a turbojet gas turbine engine according to the present disclosure.

[0063] like Figure 1 As shown, the engine 10 has a longitudinal or axial centerline 12 extending therethrough, for reference. An axial direction A extends in the same direction as the axial centerline 12, for reference. The engine 10 further defines an upstream end 99 (or front end) and a downstream end 98 (or rear end) for reference. Generally, the engine 10 includes a fan assembly 14 and a turbine 16 disposed downstream of the fan assembly 14. For reference, the engine 10 defines an axial direction A, a radial direction R, and a circumferential direction C. Generally, the axial direction A extends parallel to the axial centerline 12, the radial direction R extends outward from the axial centerline 12 and inward toward the axial centerline 12 in a direction perpendicular to the axial direction A, and the circumferential direction extends three hundred and sixty degrees (360°) around the axial centerline 12.

[0064] The turbine 16 includes a substantially tubular casing 18 that defines an annular inlet 20 of the turbine 16. The casing 18 surrounds or at least partially forms, in serial flow relationship: a compressor section having a boost or low pressure (LP) compressor 22, a high pressure (HP) compressor 24; a heat addition system 26; an expansion section or turbine section including a high pressure (HP) turbine 28 and a low pressure (LP) turbine 30; and an ejection exhaust nozzle section 32. A high pressure (HP) rotor shaft 34 drivingly connects the HP turbine 28 to the HP compressor 24. A low pressure (LP) rotor shaft 36 drivingly connects the LP turbine 30 to the LP compressor 22. The LP rotor shaft 36 may also be connected to a fan shaft 38 of the fan assembly 14. In certain embodiments, as shown Figure 1 As shown, the LP rotor shaft 36 is connected to the fan shaft 38 via a reduction gear 40 , for example in an indirect drive or gear drive configuration.

[0065] like Figure 1 As shown, fan assembly 14 includes a plurality of fan blades 42 coupled to and extending radially outward from fan shaft 38. An annular fan case or nacelle 44 circumferentially surrounds fan assembly 14 and / or at least a portion of turbine 16. It should be understood that nacelle 44 is configured to be supported relative to turbine 16 by a plurality of circumferentially spaced outlet guide vanes or struts 46. Furthermore, at least a portion of the depicted nacelle 44 extends over an outer portion of turbine 16 to define a first flow or fan flow path 48 therebetween.

[0066] During operation of the engine 10, an air flow, schematically indicated by arrow 74, enters an inlet 76 of the engine 10 defined by the fan case or nacelle 44. A portion of the air, schematically indicated by arrow 80, enters the turbine 16 through the inlet 20, which is at least partially defined by the outer casing 18. The air flow is provided in a serial flow through the compressor, the heat addition system 26, and the expansion section. In particular, for the illustrated embodiment, the turbine 16, and more specifically, the compressor section, the heat addition section 26, and the turbine section, together at least partially define the working gas flow path 70 or second stream. As the air flow 80 flows through successive stages of the compressors 22, 24 (e.g., schematically indicated by arrow 82), the air flow 80 becomes increasingly compressed. The compressed air 82 enters the heat addition system 26 and is mixed with liquid and / or gaseous fuel and ignited to produce combustion gases 86. It should be understood that the heat addition system 26 may include any suitable system for generating combustion gases, including but not limited to a deflagration or detonation combustion system, or a combination thereof. The heat addition system 26 may include annular, can, can-annular, trapped vortex, involute or vortex, rich burn, lean burn, rotating detonation, or pulse detonation configurations, or combinations thereof.

[0067] The combustion gases 86 release energy before being discharged from the exhaust nozzle section 32 to drive the rotation of the HP turbine 28 and shaft 34, and the LP turbine 30 and shaft 36. The energy released from the combustion gases 86 further drives the rotation of the fan assembly 14, including the fan blades 42. A portion of the air 74 bypasses the turbine 16 and flows through the fan flow path 48, as schematically indicated by arrows 78.

[0068] It should be understood that Figure 1 A dual flow engine having a fan flow path 48 (first flow) and a turbine flow path 70 (second flow) is depicted and described. Figure 1 The embodiment depicted in has a nacelle 44 surrounding the fan blades 42 to provide noise attenuation, blade-off protection, and other benefits known for nacelles, and may be referred to herein as a "ducted fan," or the entire engine 10 may be referred to as a "ducted engine."

[0069] Notably, in the illustrated embodiment, the engine 10 also includes a heat exchanger 200 in the second flow / bypass passage 48. As will be appreciated, the bypass flow 48 is an annular flow path relative to the centerline 12. The heat exchanger 200 is positioned in the bypass flow 48 and extends in a circumferential direction C within the flow path 48 (although depicted schematically only at the top for clarity).

[0070] However, in additional or alternative embodiments, the heat exchanger 200 can be positioned in any other annular or substantially annular passage, such as within the exhaust section 32, as shown in phantom, such as a waste heat recovery heat exchanger. The heat exchanger 200 in the exhaust section 32 can also be an annular heat exchanger and can be configured to receive heat from the airflow 86.

[0071] In this manner, it should be understood that in one or more of these example embodiments, the exchanger 200 can extend at least about 30 degrees, such as at least 90 degrees, such as at least 150 degrees, such as at least 180 degrees, such as at least 240 degrees, such as at least 300 degrees, such as at least 330 degrees, of an annular or substantially annular passage within the flow path in the circumferential direction C. Additionally or alternatively, in certain example embodiments, the exchanger 200 can extend substantially continuously within the flow path in the circumferential direction C (e.g., at least about 345 degrees of the annular or substantially annular passage), or extend continuously within the flow path in the circumferential direction C (e.g., 360 degrees of the annular passage).

[0072] Now refer to Figure 2 , provides a schematic cross-sectional view of a gas turbine engine according to another example embodiment of the present disclosure. In particular, Figure 2An engine is provided having a rotor assembly with a single stage of non-ducted rotor blades. In this manner, the rotor assembly may be referred to herein as a "non-ducted fan," or the entire engine 100 may be referred to as a "non-ducted engine." Additionally, Figure 2 The engine includes a third flow path extending from the compressor section to the rotor assembly on the turbine, as will be explained in more detail below.

[0073] For reference, the engine 100 defines an axial direction A, a radial direction R, and a circumferential direction C. Furthermore, the engine 100 defines an axial centerline or longitudinal axis 112 extending along the axial direction A. Generally speaking, the axial direction A extends parallel to the longitudinal axis 112, the radial direction R extends outward from and inward toward the longitudinal axis 112 in a direction perpendicular to the axial direction A, and the circumferential direction extends three hundred and sixty degrees (360°) about the longitudinal axis 112. The engine 100 extends, for example, along the axial direction A between a forward end 114 and a rearward end 116.

[0074] The engine 100 includes a turbine 120 and a rotor assembly positioned upstream thereof, also referred to as a fan section 150. Generally, the turbine 120 includes, in serial flow order, a compressor section, a combustion section, a turbine section, and an exhaust section. In particular, as Figure 2 As shown, the turbine 120 includes a core casing 122 that defines an annular core inlet 124. The core casing 122 further at least partially surrounds a low pressure system and a high pressure system. For example, the core casing 122 is shown at least partially surrounding and supporting a boost or low pressure ("LP") compressor 126, which is used to pressurize air entering the turbine 120 through the core inlet 124. A high pressure ("HP") multi-stage axial flow compressor 128 receives the pressurized air from the LP compressor 126 and further increases the pressure of the air. The pressurized air flow flows downstream to the combustor 130 of the combustion section, where fuel is injected into the pressurized air flow and ignited to increase the temperature and energy level of the pressurized air.

[0075] It should be understood that, as used herein, the terms "high / low speed" and "high / low pressure" are used interchangeably with respect to high pressure / high speed systems and low pressure / low speed systems. Furthermore, it should be understood that the terms "high" and "low" are used in the same context to distinguish between the two systems and are not meant to imply any absolute speed and / or pressure values.

[0076] The high-energy combustion products flow downstream from the combustor 130 to the high-pressure turbine 132. The high-pressure turbine 128 drives the high-pressure compressor 128 via a high-pressure shaft 136. At this point, the high-pressure turbine 128 is drivingly coupled to the high-pressure compressor 128. The high-energy combustion products then flow to the low-pressure turbine 134. The low-pressure turbine 134 drives the low-pressure compressor 126 and components of the fan section 150 via a low-pressure shaft 138. At this point, the low-pressure turbine 134 is drivingly coupled to components of the low-pressure compressor 126 and the fan section 150. In this example embodiment, the LP shaft 138 is coaxial with the HP shaft 136. After driving each of the turbines 132, 134, the combustion products exit the turbine 120 through a turbine exhaust nozzle 140.

[0077] Thus, the turbine 120 defines a working gas flow path or core duct 142 extending between the core inlet 124 and the turbine exhaust nozzle 140. The core duct 142 is an annular duct positioned generally inwardly of the core shroud 122 in the radial direction R. The core duct 142 (e.g., the working gas flow path through the turbine 120) may be referred to as a secondary flow.

[0078] Fan section 150 includes fan 152, which in this example embodiment is the main fan. Figure 2 In the illustrated embodiment, the fan 152 is an open rotor or non-ducted fan 152. As depicted, the fan 152 includes an array of fan blades 154 ( Figure 2 Only one is shown in FIG. 1 ). Fan blades 154 are rotatable, for example, about longitudinal axis 112. As described above, fan 152 is drivingly coupled to low pressure turbine 134 via LP shaft 138. For example, in a direct drive configuration, fan 152 may be directly coupled to LP shaft 138. However, for Figure 2 In the illustrated embodiment, the fan 152 is coupled to the LP shaft 138 via a reduction gearbox 155 , for example in an indirect drive or gear drive configuration.

[0079] Furthermore, fan blades 154 can be arranged at equal intervals about longitudinal axis 112. Each blade 154 has a root and a tip, and a span defined therebetween. Each blade 154 defines a central blade axis 156. For this embodiment, each blade 154 of fan 152 can rotate about its respective central blade axis 156, for example, in unison with one another. One or more actuators 158 are provided to facilitate such rotation and can therefore be used to change the pitch of blades 154 about their respective central blade axis 156.

[0080] The fan section 150 also includes a fan guide vane array 160 including fan guide vanes 162 ( Figure 2For this embodiment, the fan guide vanes 162 are not rotatable about the longitudinal axis 112. Each fan guide vane 162 has a root and a tip, and a span defined therebetween. The fan guide vanes 162 may be configured as follows: Figure 2 It is shown uncovered, or alternatively, may be covered by, for example, an annular shroud spaced outwardly in the radial direction R from the tips of the fan guide vanes 162 or attached to the fan guide vanes 162 .

[0081] Each fan guide vane 162 defines a central blade axis 164. For this embodiment, each fan guide vane 162 in the fan guide vane array 160 can rotate about its respective central blade axis 164, for example, in unison with each other. One or more actuators 166 are provided to facilitate such rotation and, therefore, can be used to change the pitch of the fan guide vanes 162 about their respective central blade axis 164. However, in other embodiments, each fan guide vane 162 can be fixed or unable to change pitch about its central blade axis 164. The fan guide vanes 162 are mounted to a fan housing 170.

[0082] like Figure 2 As shown, in addition to the non-ducted fan 152, a ducted fan 184 is included behind the fan 152, so that the engine 100 includes both ducted and non-ducted fans, both of which are used to generate thrust by moving air without passing through at least a portion of the turbine 120 (e.g., the HP compressor 128 and the combustion section of the illustrated embodiment). The ducted fan is shown at approximately the same axial position as the fan blades 154 and radially inward of the fan blades 154. For the illustrated embodiment, the ducted fan 184 is driven by the low-pressure turbine 134 (e.g., coupled to the LP shaft 138).

[0083] Fan cowl 170 annularly surrounds at least a portion of core cowl 122 and is positioned generally outside of at least a portion of core cowl 122 in radial direction R. In particular, a downstream section of fan cowl 170 extends over a forward portion of core cowl 122 to define a fan flow path or fan duct 172. Fan flow path or fan duct 172 may be referred to as a tertiary flow of engine 100.

[0084] Incoming air may enter the fan duct 172 through the fan duct inlet 176 and may exit through the fan exhaust nozzle 178 to generate propulsive thrust. The fan duct 172 is an annular duct positioned generally outboard of the core duct 142 in the radial direction R. The fan shroud 170 and the core shroud 122 are connected together and are supported by a plurality of substantially radially extending, circumferentially spaced stationary struts 174 ( Figure 2The fan duct 172 and the core duct 142 are supported by the fan duct 170 and the core duct 142. The fan duct 172 and the core duct 142 are supported by the fan duct 170 and the core duct 142. The fan duct 172 and the core duct 142 are supported by the fan duct 170 and the core duct 142. The fan duct 172 and the core duct 142 are supported by the fan duct 170 and the core duct 142. The fan duct 172 and the core duct 142 are supported by the fan duct 170 and the core duct 142. The fan duct 172 and the core duct 142 can each extend directly from the leading edge 144 of the core duct 122 and can be partially extended approximately axially on the opposite radial sides of the core duct.

[0085] The engine 100 further defines or includes an inlet duct 180. The inlet duct 180 extends between an engine inlet 182 and the core inlet 124 / fan duct inlet 176. The engine inlet 182 is generally defined at the forward end of the fan cowl 170 and is positioned between the fan 152 and the fan guide vane array 160 in the axial direction A. The inlet duct 180 is an annular duct positioned inboard of the fan cowl 170 in the radial direction R. Air flowing downstream along the inlet duct 180 is split (not necessarily evenly) by the splitter or leading edge 144 of the core cowl 122 into the core duct 142 and the fan duct 172. The inlet duct 180 is wider in the radial direction R than the core duct 142. The inlet duct 180 is also wider in the radial direction R than the fan duct 172.

[0086] In an exemplary embodiment, the air passing through fan duct 172 may be relatively cooler (e.g., lower temperature) than one or more fluids used in turbine 120. As such, one or more heat exchangers 200 may be disposed within fan duct 172 and used to cool one or more fluids from the core engine (where air passes through fan duct 172) as a resource for removing heat from the fluids (e.g., compressor bleed air, oil, or fuel).

[0087] Although not depicted, in certain exemplary embodiments, the engine 100 may also include one or more heat exchangers 200 in other annular ducts or flow paths of the engine 100 (e.g., in the inlet duct 180, in the turbomachinery flow path / core duct 142, within the turbine section and / or turbine exhaust nozzle 140, etc.).

[0088] In at least some exemplary embodiments, Figure 2 (and Figure 1 ) can extend in the circumferential direction C. For example, now briefly referring to Figure 3 , provides Figure 2FIG. 1 is a partial cross-sectional view of the heat exchanger 200 , and it will be appreciated that the heat exchanger 200 may extend substantially continuously in the circumferential direction C about the centerline 112 , for example, extending substantially 360 degrees in the circumferential direction C. FIG.

[0089] In addition, still refer to Figure 3 It should be understood that, as described above, the fan duct 172 in which the heat exchanger 200 is positioned is an annular duct, or more precisely, a fully annular duct, because it extends continuously and uninterrupted in the circumferential direction C. However, in other embodiments, the fan duct 172 in which the heat exchanger 200 is positioned, or a portion of the fan duct 172 , or another duct or flow path in which the heat exchanger 200 is positioned, may be a partially annular duct.

[0090] More specifically, reference is now also made to Figure 4 , provides a close-up cross-sectional view of a heat exchanger 300 positioned within a flow path 302. In at least some exemplary embodiments, the heat exchanger 300 and the flow path 302 may be arranged in a manner similar to that described above with reference to FIG. Figure 1 or Figure 2 The exemplary heat exchanger 200 and flow paths (eg, flow path 172 ) are depicted as being configured in a similar manner.

[0091] for Figure 4 In an embodiment, the heat exchanger 300 is configured as a tube-based heat exchanger 300, including a plurality of channels or tubes 304 extending through a flow path 302. The heat exchanger 300 also includes a plurality of manifolds 306, wherein each manifold 306 is fluidly coupled to a hot fluid line 308, which can be a supply line or a return line. In this manner, the heat exchanger 300 can be configured to exchange heat from the hot fluid passing through the plurality of tubes 304 to the air flow passing through the flow path 302.

[0092] It should be understood that the number, size, and configuration of tubes 304, manifolds 306, etc. are provided as examples only, and in other exemplary embodiments, heat exchanger 300 may have any other suitable configuration. Figure 4The exemplary heat exchanger 300 depicted in FIG extends continuously in the circumferential direction C, but it should be understood that in other exemplary embodiments, the heat exchanger 300 can be a plurality of discrete heat exchangers 300 arranged in the circumferential direction C. The plurality of discrete heat exchangers 300 can collectively extend substantially continuously in the circumferential direction C, with only relatively small gaps or spacings between adjacent heat exchangers 300. With this configuration, the plurality of discrete heat exchangers 300 can collectively extend at least about 180 degrees (e.g., at least 240 degrees, such as at least 300 degrees, such as at least 330 degrees, such as at least about 345 degrees) of an annular or substantially annular channel along the circumferential direction C within the flow path, or extend continuously along the circumferential direction C within the flow path (e.g., 360 degrees of an annular channel). It is noteworthy that the porosity ranges described herein and provided below take into account any small gaps or spacings between adjacent heat exchangers 300, as well as arrangements in which the heat exchanger 300 otherwise does not fully extend through the flow path in the circumferential direction C.

[0093] In addition, despite the Figure 4 In the embodiment of FIG. 1 , a single row of channels or tubes 304 is depicted as extending in the circumferential direction C, but it should be understood that the heat exchanger 300 may have multiple channels or tubes 304 in each layer (e.g., for FIG. 1 ). Figure 4 Each of the three layers depicted in FIG includes a plurality of channels or tubes 304 arranged along the axial direction A. In addition, although Figure 4 The channels or tubes 304 are depicted as extending generally in the circumferential direction C, but in other embodiments, the tubes 304 may additionally or alternatively extend in the axial direction A such that the heat exchanger 300 includes multiple tubes at each layer arranged along the circumferential direction C. The number of axially extending channels at a particular layer of the heat exchanger 300 may be referred to as the channel density of the heat exchanger 300 .

[0094] It will also be understood that the flow path 302 defines a flow path flow area Af. The flow path flow area Af generally refers to the cross-sectional area of ​​the flow path 302, and more specifically refers to the cross-sectional area of ​​the flow path 302 at the location where the heat exchanger 300 is located, excluding the heat exchanger 300. For a perfect annular flow path 302, the flow path flow area Af can be given by (R2 2 –R1 2 )xπ, where R2 is the outer radius of the flow path 302 and R1 is the inner radius of the flow path 302. In addition, the heat exchanger 300 defines a heat exchanger flow area Ah. The heat exchanger flow area Ah can refer to the minimum cross-sectional area of ​​the open path through the heat exchanger 300. For the embodiment shown, the heat exchanger flow area Ah can be calculated as the flow path flow area Af minus Figure 4The cross-sectional area of ​​each of the tubes 304 and manifold 306 of the heat exchanger 300 depicted in FIG. The ratio of the heat exchanger flow area Ah to the flow path flow area Af may generally be referred to as the porosity of the heat exchanger 300 .

[0095] However, it should be understood that in other exemplary embodiments, the heat exchanger 300 may have any other suitable configuration. Figure 5 , provides a schematic perspective view of a heat exchanger 300 according to another exemplary embodiment of the present disclosure. The heat exchanger 300 defines an axial direction A, a radial direction R, and a circumferential direction C. When installed in a gas turbine engine, the axial direction A, radial direction R, and circumferential direction C of the heat exchanger can be aligned with the axial direction A, radial direction R, and circumferential direction C of the gas turbine engine. As will be seen from Figure 5 As will be appreciated from the embodiments of FIG. 3 , in other exemplary embodiments, the heat exchanger 300 may be a fin-based heat exchanger 300. Specifically, for Figure 5 In an embodiment, heat exchanger 300 includes a plurality of plates 310, a first plurality of fins 312 extending between adjacent plates 310, and a second plurality of fins 314 also extending between adjacent plates 310 and opposite one of plates 310 from the first plurality of fins 312. A first fluid stream can pass through first plurality of fins 312, and a second fluid stream can pass through second plurality of fins 314. Heat can flow from the first fluid stream through first plurality of fins 312, through plates 310 positioned between first plurality of fins 312 and second plurality of fins 314, and into the second fluid stream (or alternatively, heat can flow in the opposite direction). As shown, there can be several layers of first plurality of fins 312 and second plurality of fins 314, as well as plates 210.

[0096] Now also briefly refer to Figure 6 , provides a position within the flow path 302 when viewed along the centerline of the engine Figure 5 The schematic diagram of one layer of the heat exchanger 300 is shown, it should be understood that Figure 5 The heat exchanger 300 defines a relatively large heat exchanger flow area Ah (at least Figure 4 ). compared to the exemplary heat exchanger 300 of FIG. Figure 6 The layer shown in FIG3 is a first plurality of fins 312. The heat exchanger 300 may also include a second plurality of fins 314 opposite the plate 310 and, for example, outside the flow path.

[0097] However, returning a reference Figure 5It should also be understood that, for the illustrated embodiment, the fins in the first plurality of fins 312 may define a relatively long length in the direction of the flow path along the axial direction A. As the length of the fins 314 increases, the efficiency E of the heat exchanger 300 may generally also increase because the increase in length provides a greater surface area to facilitate heat exchange with the airflow passing through the flow path 302.

[0098] However, it should be understood that in other exemplary embodiments, the heat exchanger 300 may have other suitable configurations. For example, in other exemplary embodiments, the heat exchanger 300 may be one or more of a pin-fin heat exchanger, a shell-and-tube heat exchanger, a tube-and-sheet heat exchanger, or a counter-flow heat exchanger.

[0099] More specifically, refer to Figure 7 , provides a perspective partial view of a heat exchanger 300 according to another exemplary embodiment of the present disclosure. In other exemplary embodiments, the heat exchanger 300 may be a pin-fin heat exchanger 300. With such a configuration, the heat exchanger 300 includes a plate 316 and a plurality of fins 318 extending from the plate 316, the plurality of fins 318 being spaced apart along a circumferential direction C. However, for Figure 7 In the exemplary heat exchanger 300 , the fins 318 are further separated into discrete “pins 320 ” spaced apart along the axial direction A. In this manner, the fins 318 may create more turbulence in the airflow through the heat exchanger 300 , thereby increasing the amount of heat exchange with the airflow through the heat exchanger 300 .

[0100] refer to Figures 8 to 10 , provides schematic diagrams of three separate heat exchangers according to various other exemplary embodiments of the present disclosure. More specifically, Figures 8 to 10 The heat exchangers 300 are each configured as a shell and tube heat exchanger. Each of these heat exchangers 300 includes a shell 322 and one or more tubes 324 positioned within the shell 322. In addition, each of the heat exchangers 300 defines a first fluid inlet 326 and a first fluid outlet 328 in fluid communication with the interior of the shell 322, and a second fluid inlet 330 and a second fluid outlet 332 in fluid communication with the one or more tubes 324. Figure 8 In FIG, the heat exchanger 300 includes one or more tubes 324 in a "U-tube" configuration. Figure 9 In FIG, the heat exchanger 300 includes one or more tubes 324 in a single-pass configuration. Figure 10 , the heat exchanger 300 includes one or more tubes 324 in a dual-pass configuration.

[0101] In this way, it should be understood that Figures 8 to 10 The heat exchanger 300 can be arranged in a parallel flow configuration (where the second fluid flows in the same direction as the first fluid (see e.g. Figure 9 ), arranged in a counter-flow configuration (wherein the second fluid flows in a direction opposite to the first fluid), or arranged in a combination of parallel and counter-flow configurations (see e.g. Figure 8 and Figure 10 ).

[0102] It will also be understood that each heat exchanger 300 is configured to transfer heat from a heating fluid (e.g., a fluid rejecting heat) to a cooling fluid (e.g., a fluid receiving heat). Figure 2 When heat exchanger 200 is used in engine 100 (e.g., in fan duct 172 as heat exchanger 200), the cooling fluid may be the airflow through fan duct 172, and the heating fluid may be, for example, compressor bleed air (air-to-air heat exchanger), fuel (fuel-to-air heat exchanger), or lubricating oil (oil-to-air heat exchanger).

[0103] As previously mentioned, standard practice is to optimize the heat exchanger for flight idle (or other conditions) and then, after selecting the optimal heat exchanger, verify whether it will operate acceptably across the flight envelope from a heat transfer perspective. Furthermore, the inventors have discovered that it is also beneficial to verify whether it will operate acceptably across the flight envelope from the perspective of the noise generated when air flows through the annular duct. This can be a labor- and time-intensive process, as it is iterative and involves selecting a heat exchanger designed for flight idle and demonstrating thermal efficiency with an acceptable pressure drop, then evaluating whether the annular duct location at other times in flight (non-flight idle) would produce unacceptable noise levels, necessitating a redesign of the heat exchanger to increase the transfer losses of air through the annular duct. In other words, a heat exchanger is selected based on size, type, etc. until a heat exchanger is found that meets all three key requirements (heat transfer, acceptable pressure drop, and acceptable noise generation across all flight conditions). It is desirable to have a limited or reduced scope embodiment defined for an engine architecture that meets mission requirements, including heat transfer, pressure ratio, and noise transmission level requirements when a heat exchanger is selected and located within the engine.

[0104] During the engine design process (i.e., designing a heat exchanger and evaluating its impact on the acoustic environment at off-design points, a time-consuming, iterative process just described), the inventors unexpectedly discovered a relationship between the heat exchanger's expected acoustic transmission loss (ETL) and its heat transfer capacity for a given level of pressure drop across the heat exchanger. Pressure drop is incorporated into the parameter UA because it is a function of porosity, which is a function of area A. Utilizing this relationship, the inventors discovered that the number of suitable or feasible heat exchangers positioned within the engine's generally annular duct that can meet both heat transfer and acoustic requirements can be significantly reduced, thereby facilitating a more rapid selection of designs to consider during engine development. This benefit provides a deeper understanding of a given engine's requirements before specific technical, integration, and system requirements are fully developed. It avoids late-stage redesigns. It also provides a heat exchanger design that incorporates both the acoustic and heat exchanger considerations of an aircraft's gas turbine engine, taking into account the unique environment of the aircraft. The desired relationship is represented by the effective transmission loss ("ETL"):

[0105] Equation (1):

[0106] Wherein C1, C2 and C3 are constants depending on the mass flow rate through the annular duct. EOC illustrates factors affected by engine size and operating conditions, which will be explained in more detail below. Constants C1, C2 and C3 and EOC each depend on flight conditions, and more specifically, on the mass flow rate ("W") of the airflow through the annular duct occupied by the heat exchanger. ETL represents the transmission loss level (in decibels, dB) that can be expected from the heat exchanger for a given mass flow rate W and UA. Once the engine architecture is more fully defined, a more detailed fluid model may be required later to more accurately determine the transmission loss under specific flight conditions. For the purpose of ETL, the mass flow rates of interest are characterized as low, medium and high mass flow rate conditions. The lowest mass flow rate can correspond to the low power operating condition of the engine (e.g., ground idle, flight idle), the medium mass flow rate can correspond to the medium power operating condition (e.g., cruising or descending), and the high mass flow rate can correspond to the high power operating condition (e.g., takeoff operating condition or climb operating condition).

[0107] Table 1 provides the values ​​of C1, C2, and C3 and EOC for three flight regimes, defined by the mass flow rate through the annular duct in which the heat exchanger is located:

[0108]

[0109] C1, C2, and C3, along with the EOC, reflect the variation in mass flow through the engine's annular duct during various operating conditions (typically, low-power, medium-power, and high-power operating conditions, as described above). The EOC also accounts for the variability of specific engine operating conditions within each of these flow regimes (low / medium / high). The EOC accounts for factors such as the specific engine type operating within the flow regime, expected variations in transient thrust, environmental conditions, tolerances, and / or engine cycling or degradation, all of which may have some impact on the transmission losses of the flow through the heat exchanger located in the annular duct. It will be appreciated that, based on the teachings herein, for the expressed EOC ranges, the ETL provides a good approximation of available heat exchanger design options suitable for meeting mission requirements from a thermal management and acoustics perspective. If desired, more accurate knowledge of transmission losses can be gathered later by performing a full 3D CFD analysis of the acoustic field. However, this level of analysis may not be necessary when the objective is to assess the acoustic environment at off-design points before proceeding with heat exchanger optimization. As described above, the ETL eliminates unfeasible designs at an early stage, prior to optimizing the heat exchanger located in the annular duct. Therefore, in one aspect, ETL can be considered as an alternative to performing a comprehensive 3D CFD analysis of the flow field before performing heat exchanger optimization within annular ducts.

[0110] Furthermore, it should be understood that transmission losses through a heat exchanger are also affected by the length of the heat exchanger, the porosity of the heat exchanger, the pressure drop across the heat exchanger, the mass flow rate through the annular duct in which the heat exchanger is positioned, and the power spectral density (PSD) distribution of the air immediately upstream of the heat exchanger.

[0111] For example, as the length of a heat exchanger increases, the amount of acoustic transmission loss generally increases. This factor influences the value of C2. The length of a heat exchanger (sometimes also called the channel length) directly affects the volume through which the fluid passes (along with the area of ​​the heat exchanger). As this volume increases, the amount of transmission loss generally increases.

[0112] The pressure drop across the heat exchanger is incorporated into Equation 1 (ETL) via the UA parameter, as described above. The ETL envisions a maximum pressure drop of 15%, for example, up to 10% and at least 1%. Generally, as the area of ​​the heat exchanger increases (and as the porosity of the heat exchanger increases), the pressure drop also increases. Generally, higher pressure drops are also associated with greater heat transfer. However, pressure drops above these levels may have too great an impact on the thrust generated by the airflow through the duct to justify the thermal benefit.

[0113] More specifically, it was found that for low power operating conditions (e.g., for flow rates less than or equal to about 50 lbm / s), an ETL of between 1 and 5 dBs can be achieved at relatively low pressure drops (e.g., a pressure drop of less than or equal to about 5%, such as a pressure drop of less than or equal to about 2.5%). It was also found that for medium power operating conditions (e.g., for flow rates greater than or equal to about 50 lbm / s and less than or equal to about 150 lbm / s), an ETL of between 1 and 5 dBs can be achieved at a pressure drop within design limits (e.g., less than or equal to about 15% (and, for example, greater than or equal to about 2%)). It was further found that for high power operating conditions (e.g., for flow rates greater than or equal to about 150 lbm / s and less than or equal to about 300 lbm / s), an ETL of between 1 and 3 dBs can be achieved while maintaining a pressure drop of less than about 15%. As described above, pressure drop is a function of UA, as it is a function of the area of ​​the heat exchanger. It was found that the effect of heat exchanger area on pressure drop increased with increasing mass flow rate, resulting in a larger pressure drop for a given ETL amount compared to lower mass flow rates.

[0114] The PSD is determined by the upstream fan or turbine characteristics (e.g. Figure 2 The middle fan 184 upstream of the heat exchanger 200, or Figure 2 The noise characteristics associated with the upstream fan are expressed in terms of a fan pass frequency, which is defined as the number of rotations per second of the immediately upstream fan or turbine multiplied by the number of fan blades or rotor blades in the turbine stage. For example, referring to Figure 2 In the illustrated embodiment, the fan pass frequency of the noise source associated with the heat exchanger 200 located in the third flow annular duct (or more specifically, the fan flow duct 172) will be obtained by multiplying the number of rotations per second of the fan 184 by the number of blades of the fan 184. In another example, still referring to Figure 2 In the embodiment shown in , the fan pass frequency of the noise source associated with the heat exchanger 140 located in the aft frame will be obtained by multiplying the number of rotations per second of the low pressure turbine 134 by the number of turbine rotor blades associated with the last stage of the low pressure turbine 134.

[0115] Sound transmission through a heat exchanger is often a byproduct of many complex interactions between sound waves and the heat exchanger's internal surfaces. This often requires detailed flow modeling of the air traveling through the heat exchanger to fully assess the sound transmission environment for specific flight conditions (e.g., takeoff or full-power flight conditions), as previously discussed. Furthermore, the fan or rotor speeds that generate the most noise don't necessarily occur when the engine is operating at full power. Therefore, the noise environment is typically modeled for a variety of flight conditions, not just full power. Nevertheless, the inventors have discovered that it is possible to make assumptions about the expected transmission loss levels of the heat exchanger during non-flight idle periods of flight (optimized for flight idle conditions), where the noise is most generated. As a result, feasible heat exchanger embodiments can be found for a given engine operating environment, using ETLs while meeting both thermal and acoustic requirements. These heat exchanger embodiments consider the competing interests associated with transmission loss requirements, maximum acceptable pressure drop, and heat transfer efficiency. By defining the embodiments in this manner, extensive heat exchanger redesign can be avoided, as previously discussed. For example, a heat exchanger located in an annular duct is optimized for engine performance during flight idle conditions. When the acoustic performance of the engine was later evaluated, for example using 3D CFD analysis, it was found that this configuration did not produce a sufficient amount of transmission losses when the air passed through the annular duct. Such a heat exchanger then needed to be redesigned because too much noise was generated.

[0116] The ETL was discovered by evaluating the impact of different pressure drop levels on transmission loss and overall heat exchanger efficiency, as well as the heat exchanger's geometry and its relationship to transmission loss. Based on these relationships, the heat exchanger's ETL was found to provide a good approximation of the expected transmission loss through the heat exchanger for a given mass flow rate as a function of the UA and general heat exchanger characteristics, as outlined in Table 2, which define the operating environment and heat exchanger characteristics used to derive the ETL. Therefore, for a heat exchanger located in an annular duct and confined within these ranges, the ETL can predict the transmission loss from the heat exchanger for a specified mass flow rate and UA.

[0117]

[0118] Figures 11 to 16 FIG2 shows a heat exchanger according to one or more exemplary embodiments of the present disclosure, illustrating the relationship between ETL and UA. In particular, Figure 11 is a graph illustrating the relationship between ETL and UA at low mass flow rates for a heat exchanger according to one or more exemplary embodiments of the present disclosure, and Figure 12 Provided include corresponding Figure 11 A table of values ​​for several of the ETL values ​​plotted in . Figure 13is a graph illustrating the relationship between ETL and UA at a medium mass flow rate for a heat exchanger according to one or more exemplary embodiments of the present disclosure, and Figure 14 Provided include corresponding Figure 13 A table of values ​​for several of the ETL values ​​plotted in . Figure 15 is a graph illustrating the relationship between ETL and UA of a high mass flow rate for a heat exchanger according to one or more exemplary embodiments of the present disclosure, and Figure 16 Provided include corresponding Figure 15 A table of values ​​for several of the ETL values ​​plotted in .

[0119] exist Figure 11 、 13 In each of Figures 1 and 15, the solid line surrounding the embodiment represents the range of TL and UA, as provided by the range of variable EOC. The TL range is 5 dB to 1 dB. The UA range varies between low, medium, and high mass flow rates, but is generally between 7,500 and 45,000 Btu / (hr-°F). Embodiments within this range include heat exchangers having a length between 3 inches and 9 inches (measured in the flow direction, which corresponds to the cold flow length characteristic of the heat exchanger, depending on the embodiment) and a heat exchanger porosity between 23% and 51%.

[0120] This disclosure is not limited to Figures 11 to 16 For example, in other embodiments, the heat exchangers of the present disclosure may be, for example, up to 15 inches in length and may define a porosity of up to 80%.

[0121] As will be understood from the description herein, embodiments of a gas turbine engine (e.g., a non-ducted, single-spool gas turbine engine) are provided. Some embodiments of the engine that include a heat exchanger located in an annular duct and that are considered to be within the scope of the present disclosure may further include one or more of the following features. At cruising altitude during a cruise operating mode, the threshold power or disk load of the fan (e.g., fan 154) may be 25 horsepower per square foot (hp / ft 2 ) or greater. In a specific embodiment of the engine, at a cruising altitude during a cruise mode of operation, the structures and methods provided herein generate power at 80 hp / ft 2 and 160hp / ft 2or higher power loads, depending on whether the engine is an open rotor or ducted engine. In various embodiments, the engine is applied to a vehicle with a cruising altitude of up to approximately 65,000 ft. In certain embodiments, the cruising altitude is between approximately 28,000 ft and approximately 45,000 ft. In still other embodiments, the cruising altitude is expressed as a flight altitude based on standard atmospheric pressure at sea level, wherein the cruising flight condition is between FL280 and FL650. In another embodiment, the cruising flight condition is between FL280 and FL450. In yet other embodiments, the cruising altitude is defined at least based on atmospheric pressure, wherein based on a sea level pressure of approximately 14.70 psia and a sea level temperature of approximately 59 degrees Fahrenheit, the cruising altitude is between approximately 4.85 psia and approximately 0.82 psia. In another embodiment, the cruising altitude is between approximately 4.85 psia and approximately 2.14 psia. It should be understood that in certain embodiments, the cruising altitude range defined by pressure can be adjusted based on different reference sea level pressures and / or sea level temperatures.

[0122] Furthermore, in certain exemplary embodiments, the rotor assembly can define a rotor diameter (or fan diameter) of at least 10 feet (e.g., at least 11 feet, such as at least 12 feet, such as at least 13 feet, such as at least 15 feet, such as at least 17 feet, such as up to 28 feet, such as up to 26 feet, such as up to 24 feet, such as up to 18 feet). Figure 2 In an embodiment, the ratio R1 / R2 may be between about 1 and 6, or 2 and 4, or about 1.5 to 3, wherein Figure 2 R1 is the span from the root to the tip of the fan blade 154 , and R2 is the span from the root to the tip of the fan 184 .

[0123] It should be understood that various embodiments of the engine (e.g., a single non-ducted rotor engine depicted and described herein) can allow normal subsonic aircraft cruise altitude operation at or above Mach 0.5. In certain embodiments, the engine allows normal aircraft operation at cruise altitudes between Mach 0.55 and Mach 0.85. In still specific embodiments, the engine allows normal aircraft operation between Mach 0.75 and Mach 0.85. In certain embodiments, the engine allows rotor blade tip speeds equal to or less than 750 feet per second (fps).

[0124] Furthermore, based on the configurations provided herein, certain embodiments of the engines provided herein may allow for normal subsonic aircraft cruise altitude operation at or above Mach 0.5, or above Mach 0.75. In certain embodiments, the engines allow for normal aircraft operation at cruise altitudes between Mach 0.55 and Mach 0.85, or between Mach 0.75 and Mach 0.85. In certain embodiments, the engines allow for rotor blade tip speeds of 750 feet per second (fps) or less. Still certain embodiments may provide benefits whereby, by virtue of configurations located within the annular duct of the engine, interaction noise between the blade and vane assemblies and / or overall noise generated by the engine is reduced. Furthermore, it should be understood that a range of power loads and / or rotor blade tip speeds may correspond to certain configurations, core sizes, thrust outputs, etc., or other configurations of the core engine and rotor assembly. However, as previously mentioned, where one or more configurations provided herein may be known in the art, it should be understood that the present disclosure may include combinations of configurations not previously known due, at least in part, to competing advantages and disadvantages, desired operating modes, or other forms of teaching in the art.

[0125] This written description uses examples to disclose the present disclosure, including the best mode, and also to enable any person skilled in the art to practice the present disclosure, including making and using any device or system and performing any combined methods. The patentable scope of the present disclosure is defined by the claims and may include other examples that occur to those skilled in the art. If these other examples include structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements that do not differ substantially from the literal language of the claims, then these other examples are intended to fall within the scope of the claims.

[0126] Further aspects are provided by the subject matter of the following clauses:

[0127] A gas turbine engine defining a centerline and a circumferential direction, the gas turbine engine comprising: a turbine including a compressor section, a combustion section, and a turbine section arranged in a serial flow order; a rotor assembly driven by the turbine, the rotor assembly, the turbine, or both including a substantially annular duct relative to the centerline of the gas turbine engine, the substantially annular duct defining a flow path; a heat exchanger positioned within the annular duct and extending substantially continuously in the circumferential direction, the heat exchanger including a heat exchanger defining a heat exchanger exposed to the flow path a first material for an exchange surface, wherein the first material defines a heat exchange coefficient, and wherein the heat exchange surface defines a surface area (A), wherein a product of the heat exchange coefficient and the surface area, UA, is between 7,500 British thermal units per hour per degree Fahrenheit (Btu / (hr-℉)) and 45,000 Btu / (hr-℉); wherein an effective transmission loss (ETL) of the heat exchanger positioned within the annular conduit is between 5 decibels and 1 decibel for an operating condition, wherein the operating condition is one of a low power operating condition, a medium power operating condition, or a high power operating condition, wherein the ETL is equal to

[0128]

[0129] wherein, when the operating condition is the low-power operating condition, C1 is equal to 19.22, C2 is equal to 0.222, C3 is equal to 956.3, and the EOC is between 41,467 and 19,965; wherein, when the operating condition is the medium-power operating condition, C1 is equal to 19.64, C2 is equal to 0.67, C3 is equal to 298, and the EOC is between 52,809 and 16,677; and wherein, when the operating condition is the high-power operating condition, C1 is equal to 21.02, C2 is equal to 0.027, C3 is equal to 107, and the EOC is between 50,347 and 12,587.

[0130] A gas turbine engine according to one or more of these clauses, wherein the heat exchanger defines a length between 3 inches and 15 inches and a porosity between 20% and 80%, wherein the gas turbine engine defines a fan pass frequency within the turbine, the rotor assembly, or both, the fan pass frequency being between 1 kHz and 5 kHz during the operating conditions.

[0131] The gas turbine engine according to one or more of these clauses, wherein the length of the heat exchanger is between 4 inches and 9 inches.

[0132] The gas turbine engine according to one or more of these clauses, wherein the heat exchanger defines a pressure drop of 15% or less during operation of the gas turbine engine.

[0133] The gas turbine engine of one or more of these clauses, wherein the gas turbine engine defines a mass flow rate through the heat exchanger during the low power operating condition that is less than or equal to 50 lbm / s, and wherein ETL is equal to: The EOC ranges from 41,467 to 19,965.

[0134] The gas turbine engine of one or more of these clauses, wherein the gas turbine engine defines a mass flow rate through the heat exchanger during the medium power operating condition of greater than or equal to 50 pounds-mass / second (lbm / s) and less than or equal to 150 lbm / s, and wherein ETL equals: The EOC ranges from 52,809 to 16,677.

[0135] The gas turbine engine of one or more of these clauses, wherein the gas turbine engine defines a mass flow rate through the heat exchanger during the high power operating condition that is greater than or equal to 150 pounds-mass / second (lbm / s) and less than or equal to 300 lbm / s, and wherein ETL is equal to: The EOC ranges from 50,347 to 12,587.

[0136] A gas turbine engine according to one or more of these clauses, wherein the annular duct is defined by the turbine and includes an inlet for a third flow, wherein the compressor section includes a fan located upstream of the inlet for the third flow, wherein the gas turbine engine defines a fan pass frequency within the turbine, wherein the fan pass frequency is that of a mid-fan, and wherein the heat exchanger is positioned within the third flow.

[0137] The gas turbine engine according to one or more of these clauses, wherein the rotor assembly of the gas turbine engine is configured as a non-ducted rotor assembly including a single stage of rotor blades.

[0138] The gas turbine engine of one or more of these clauses, wherein the single stage rotor blades define a blade diameter greater than or equal to 10 feet and less than or equal to 28 feet, optionally less than 18 feet, optionally less than 15 feet.

[0139] The gas turbine engine according to one or more of these clauses, wherein the heat exchanger has one of the following architectures: fin-based, pin-fin, tube, shell and tube, tube-sheet, counter-flow, or a combination thereof.

[0140] The gas turbine engine according to one or more of these clauses, wherein the rotor assembly of the gas turbine engine is configured as a ducted rotor assembly.

[0141] The gas turbine engine according to one or more of these clauses, wherein the heat exchanger extends substantially continuously within the flow path.

[0142] The gas turbine engine of one or more of these clauses, wherein the flow path is a turbine flow path, and wherein the duct is at least partially positioned in the compressor section, the combustion section, the turbine section, or a combination thereof.

[0143] The gas turbine engine according to one or more of these clauses, wherein the heat exchanger is a waste heat recovery heat exchanger.

[0144] The gas turbine engine according to one or more of these clauses, wherein the rotor assembly limits a fan pass frequency to between 1 kHz and 5 Khz during the operating conditions, and wherein the heat exchanger is located downstream of the rotor assembly.

[0145] The gas turbine engine according to one or more of these clauses, wherein the gas turbine engine limits a fan pass frequency within the turbine to between 1 kHz and 5 Khz during the operating conditions, and wherein the heat exchanger is located within the turbine.

[0146] The gas turbine engine according to one or more of these clauses, wherein the heat exchanger has an ETL of between 5 decibels and 1 decibel during the operating conditions.

[0147] A gas turbine engine defining a centerline and a circumferential direction, the gas turbine engine comprising: a turbine including a compressor section, a combustion section, and a turbine section arranged in a serial flow order; a rotor assembly driven by the turbine, the rotor assembly, the turbine, or both including a substantially annular duct relative to the centerline of the gas turbine engine, the annular duct defining a flow path; a heat exchanger positioned within the annular duct and extending substantially continuously along the circumferential direction, the heat exchanger defining a length between 3 inches and 15 inches and a porosity between 20% and 80%, the heat exchanger comprising a first material defining a heat exchange surface exposed to the flow path, wherein the first material defines a heat exchange coefficient, and wherein the heat exchange surface defines a surface area (A), wherein a product of the heat exchange coefficient and the surface area, UA, is between 7,500 British thermal units per hour per degree Fahrenheit (Btu / (hr-°F)) and 45,000 Btu / (hr-°F), wherein the gas turbine engine defines a fan pass frequency within the turbine, the rotor assembly, or both, between 1 kHz and 5 kHz during operating conditions, and wherein, for the operating conditions, the heat exchanger has an effective transmission loss (ETL) between 5 decibels and 1 decibel.

[0148] The gas turbine engine according to one or more of these clauses, wherein ETL is equal to:

[0149]

[0150] wherein, when the operating condition is a low-power operating condition, C1 is equal to 19.22, C2 is equal to 0.222, C3 is equal to 956.3, and the EOC is between 41,467 and 19,965; wherein, when the operating condition is a medium-power operating condition, C1 is equal to 19.64, C2 is equal to 0.67, C3 is equal to 298, and the EOC is between 52,809 and 16,677; and wherein, when the operating condition is a high-power operating condition, C1 is equal to 21.02, C2 is equal to 0.027, C3 is equal to 107, and the EOC is between 50,347 and 12,587.

[0151] A gas turbine engine according to one or more of these clauses, wherein when the operating condition is a low power operating condition, UA is greater than 7500 Btu / (hr-°F) and less than 45000 Btu / (hr-°F), for example, greater than 10000 Btu / (hr-°F) and less than 35000 Btu / (hr-°F), when the operating condition is a medium power operating condition, UA is, for example, greater than 14000 Btu / (hr-°F) and less than 5000 Btu / (hr-°F), or when the operating condition is a high power operating condition, UA is greater than 15000 Btu / (hr-°F) and less than 44000 Btu / (hr-°F).

[0152] The gas turbine engine according to one or more of these clauses, wherein the pressure drop is less than 15%, such as less than 10%, such as less than 8%, such as greater than 1%.

[0153] The gas turbine engine according to one or more of these clauses, wherein when the operating condition is a low power operating condition, the pressure drop is less than or equal to about 5%, such as less than or equal to about 2.5%.

[0154] The gas turbine engine according to one or more of these clauses, wherein when the operating condition is a medium power operating condition, the pressure drop is less than or equal to about 15%.

[0155] The gas turbine engine according to one or more of these clauses, wherein the pressure drop is less than or equal to about 15%, wherein the ETL is between 1 and 3 dB, and wherein the operating condition is a high power operating condition.

[0156] The gas turbine engine according to one or more of these clauses, wherein the length of the heat exchanger is between 3 inches and 15 inches, such as between 4 inches and 9 inches.

[0157] Gas turbine engine according to one or more of these clauses, wherein the porosity of the heat exchanger is 20% to 80%, such as 30% to 55%.

Claims

1. A gas turbine engine defining a centerline and a circumferential direction, characterized in that The gas turbine engine comprises: a turbomachine comprising a compressor section, a combustion section, and a turbine section arranged in serial flow order; a rotor assembly driven by the turbine, the rotor assembly, the turbine, or both including a substantially annular duct relative to the centerline of the gas turbine engine, the substantially annular duct defining a flow path; a heat exchanger positioned within the annular conduit and extending substantially continuously in the circumferential direction, the heat exchanger comprising a first material defining a heat exchange surface exposed to the flow path, wherein the first material defines a heat exchange coefficient, and wherein the heat exchange surface defines a surface area (A), wherein a product UA of the heat exchange coefficient and the surface area is between 7,500 British thermal units per hour per degree Fahrenheit (Btu / (hr-℉)) and 45,000 Btu / (hr-℉); wherein the effective transmission loss (ETL) of the heat exchanger positioned within the annular duct is between 5 decibels and 1 decibel for an operating condition, the operating condition being one of a low power operating condition, a medium power operating condition, or a high power operating condition, Among them, ETL is equal to ； wherein, when the operating condition is the low power operating condition, C1 is equal to 19.22, C2 is equal to 0.222, C3 is equal to 956.3, and EOC is between 41,467 and 19,965; wherein, when the operating condition is the medium power operating condition, C1 is equal to 19.64, C2 is equal to 0.67, C3 is equal to 298, and EOC is between 52,809 and 16,677; and When the operating condition is the high power operating condition, C1 is equal to 21.02, C2 is equal to 0.027, C3 is equal to 107, and EOC is between 50,347 and 12,587.

2. The gas turbine engine according to claim 1, wherein: in, The heat exchanger defines a length between 3 inches and 15 inches and a porosity between 20% and 80%, wherein the gas turbine engine defines a fan pass frequency within the turbine, the rotor assembly, or both, the fan pass frequency being between 1 kHz and 5 Khz during the operating conditions.

3. The gas turbine engine according to claim 2, characterized in that in, The length of the heat exchanger is between 4 inches and 9 inches.

4. The gas turbine engine according to claim 1, wherein: in, The heat exchanger defines a pressure drop of 15% or less during operation of the gas turbine engine.

5. The gas turbine engine according to claim 1, wherein: in, The gas turbine engine defines a mass flow rate through the heat exchanger during the low power operating condition that is less than or equal to 50 lbm / s, and wherein ETL is equal to: ； Among them, the EOC is between 41,467 and 19,965.

6. The gas turbine engine according to claim 1, wherein: in, The gas turbine engine defines a mass flow rate through the heat exchanger during the medium power operating condition that is greater than or equal to 50 pounds mass per second (lbm / s) and less than or equal to 150 lbm / s, and wherein ETL is equal to: ； Among them, the EOC ranged between 52,809 and 16,677.

7. The gas turbine engine according to claim 1, wherein: in, The gas turbine engine defines a mass flow rate through the heat exchanger during the high power operating condition that is greater than or equal to 150 pounds mass per second (lbm / s) and less than or equal to 300 lbm / s, and wherein ETL is equal to: ； Among them, the EOC ranged between 50,347 and 12,587.

8. The gas turbine engine according to claim 1, wherein: in, The annular duct is defined by the turbine and includes a third flow at an inlet, wherein the compressor section includes a fan located upstream of the inlet to the third flow, wherein the gas turbine engine defines a fan pass frequency within the turbine, wherein the fan pass frequency is that of a mid-fan, and wherein the heat exchanger is positioned within the third flow.

9. The gas turbine engine according to claim 1, wherein: in, The rotor assembly of the gas turbine engine is configured as a non-ducted rotor assembly including a single stage of rotor blades.

10. The gas turbine engine according to claim 9, characterized in that in, The single-stage rotor blades define a blade diameter greater than or equal to 10 feet and less than or equal to 28 feet, optionally less than 18 feet, and optionally less than 15 feet.

11. The gas turbine engine according to claim 1, wherein: in, The heat exchanger has one of the following architectures: fin-based, pin-fin, tube, shell and tube, tube-sheet, counter-flow, or a combination thereof.

12. The gas turbine engine according to claim 1, wherein: in, The rotor assembly of the gas turbine engine is configured as a ducted rotor assembly.

13. The gas turbine engine according to claim 1, wherein: in, The heat exchanger extends substantially continuously within the flow path.

14. The gas turbine engine according to claim 1, wherein in, The flow path is a turbine flow path, and wherein the conduit is at least partially positioned in the compressor section, the combustion section, the turbine section, or a combination thereof.

15. The gas turbine engine according to claim 14, characterized in that in, The heat exchanger is a waste heat recovery heat exchanger.

16. The gas turbine engine according to claim 1, wherein in, The rotor assembly limits the fan pass frequency to between 1 kHz and 5 Khz during the operating conditions, and wherein the heat exchanger is located downstream of the rotor assembly.

17. The gas turbine engine according to claim 1, wherein: in, The gas turbine engine limits a fan pass frequency within the turbine to between 1 kHz and 5 Khz during the operating conditions, and wherein the heat exchanger is located within the turbine.

18. The gas turbine engine according to claim 1, wherein: in, The heat exchanger has an ETL of between 5 dB and 1 dB during the operating conditions.

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