Gas turbine engine with third flow

Abstract

A gas turbine engine defining a centerline and a circumferential direction, the gas turbine engine comprising: a turbomachine comprising, in serial flow order, a compressor section, a combustion section, and a turbine section, the turbomachine defining a working gas flow path and a fan duct flow path; a primary fan driven by the turbomachine, the primary fan defining a primary fan tip radius R1 and a primary fan hub radius R2; a secondary fan downstream of the primary fan and driven by the turbomachine, at least a portion of airflow from the primary fan being configured to bypass the secondary fan, the secondary fan defining a secondary fan tip radius R3 and a secondary fan hub radius R4, wherein the secondary fan is configured to provide, during operation, a fan duct airflow through the fan duct flow path to generate a fan duct thrust, wherein the fan duct thrust is equal to a %Fn of a total engine thrust during operation of the gas turbine engine at rated speed during standard day operating conditions 3S ; wherein a ratio of R1 to R3 is equal to wherein EFP is between 1.5 and 11, wherein RqR Prim.‑Fan is a ratio of R2 to R1, and wherein RqR Sec.‑Fan is a ratio of R4 to R3.

Description

[0001] Cross-reference to related applications

[0002] This application is a non-provisional application claiming priority to U.S. Provisional Application No. 63 / 240,796, filed September 3, 2021, pursuant to 35 U.SC §119(e), the entire contents of which are incorporated herein by reference. Technical Field

[0003] This disclosure relates to a gas turbine engine with a third flow. Background Technology

[0004] 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. These combustion gases are directed to the turbine, which extracts energy from the combustion gases to power the compressor and generate useful work to propel the aircraft in flight. The turbine is mechanically coupled to the fan to drive it during operation. Attached Figure Description

[0005] The complete and effective disclosure of this disclosure, including its best mode, is set forth in the description with reference to the accompanying drawings, and is intended for use by those skilled in the art, wherein:

[0006] Figure 1 This is a schematic cross-sectional view of a three-flow engine according to an exemplary embodiment of the present disclosure.

[0007] Figures 2A to 2D This is a table of example embodiments of this disclosure.

[0008] Figure 3 It is a graph depicting the range of the radius ratio (R1 vs. R3) of the third stream versus the percentage of thrust according to various exemplary embodiments of the present disclosure. Detailed Implementation

[0009] Reference will now be made in detail to the present embodiments of this disclosure, one or more examples of which are illustrated in the accompanying drawings. The detailed description uses numerals and letter reference numerals to denote features in the drawings. Similar or analogous reference numerals in the drawings and description have been used to denote similar or analogous portions of this disclosure.

[0010] The term "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 superior or better than other implementations. Furthermore, unless explicitly stated otherwise, all embodiments described herein should be considered exemplary.

[0011] As used herein, the terms “first,” “second,” and “third” are used interchangeably to distinguish one component from another and are not intended to indicate the location or importance of the individual components.

[0012] The terms "front" and "rear" refer to relative positions within a gas turbine engine or vehicle, and specifically to the normal operating posture of the gas turbine engine or vehicle. For example, in the case of a gas turbine engine, "front" refers to the position closer to the engine inlet, while "rear" refers to the position closer to the engine nozzle or exhaust port.

[0013] The terms "upstream" and "downstream" refer to the relative directions of fluid flow within a fluid path. For example, "upstream" refers to the direction from which the fluid flows, and "downstream" refers to the direction from which the fluid flows.

[0014] Unless otherwise stated herein, the terms “connection,” “fixed,” “attached to,” etc., refer to both direct connection, fixation, or attachment, and indirect connection, fixation, or attachment via one or more intermediate components or features.

[0015] Unless the context clearly indicates otherwise, the singular forms “a,” “a,” and “the” include plural references.

[0016] As used throughout the specification and claims, approximate language is applied to modify any quantitative expression that may allow for variation without altering its underlying function. Therefore, values ​​modified by terms such as “about,” “approximately,” “roughly,” and “substantially” are not limited to specified precise values. In at least some cases, approximate language may correspond to the precision of the instrument used to measure the value, or the precision of the method or machine used to construct or manufacture the component and / or system. For example, approximate language may refer to a margin of 1%, 2%, 4%, 5%, 10%, 15%, or 20% of a single value, a range of values, and / or the endpoints of a defined range of values. Scope limitations are combined and interchanged herein and throughout the specification and claims, and such scopes are identified and include all subscopes contained herein unless otherwise indicated by context or language. For example, all scopes disclosed herein include endpoints, and endpoints may be combined independently of each other.

[0017] As used herein, "third flow" refers to a non-primary airflow capable of increasing fluid energy to generate a small amount of total propulsion thrust. The pressure ratio of the third flow can be higher than that of the primary propulsion flow (e.g., bypass or propeller-driven propulsion flow). Thrust can be generated through a dedicated nozzle or by mixing the airflow through the third flow with, for example, the primary propulsion flow or core airflow entering a common nozzle.

[0018] In some exemplary embodiments, the operating temperature of the airflow through the third flow can be below the engine's maximum compressor discharge temperature, and more specifically, below 350 degrees Fahrenheit (e.g., below 300 degrees Fahrenheit, below 250 degrees Fahrenheit, below 200 degrees Fahrenheit, and at least as high as the ambient temperature). In some exemplary embodiments, these operating temperatures can facilitate heat transfer to or from the airflow through the third flow and the separate fluid flow. Furthermore, in some exemplary embodiments, under takeoff conditions, or more specifically, operating at sea level at rated takeoff power, static flight speed, and an ambient temperature of 86 degrees Fahrenheit, the airflow through the third flow can contribute less than 50% (and at least, for example, 2%) of the total engine thrust.

[0019] Furthermore, in some exemplary embodiments, the aspects of the third flow (e.g., airflow, mixing, or exhaust characteristics) and thus the aforementioned exemplary percentage contribution to the total thrust can be passively adjusted during engine operation or purposefully modified by using engine control features (e.g., fuel flow, motor power, variable stator, variable inlet guide vanes, valves, variable exhaust port geometry, or fluid characteristics) to adjust or optimize overall system performance under a wide range of potential operating conditions.

[0020] The term "disk load" refers to the average pressure change across multiple rotor blades of a rotor assembly, such as the average pressure change across multiple fan blades of a fan.

[0021] The term "rated speed" refers to the operating conditions of an engine, whereby the engine operates under the manufacturer's rated maximum full-load operating conditions.

[0022] The term "standard day operating conditions" refers to environmental conditions of sea level, 59 degrees Fahrenheit, and 60% relative humidity.

[0023] The term "propulsion efficiency" refers to the efficiency with which the energy contained in the engine fuel is converted into kinetic energy for the vehicle containing the engine, in order to accelerate the vehicle or compensate for losses due to aerodynamic drag or gravity.

[0024] The term "bypass ratio" refers to the ratio of the airflow that bypasses a piped inlet in an engine to the amount that passes through that inlet. For example, as discussed below... Figure 1 In the embodiments, the bypass ratio refers to the ratio of the airflow from fan 152 through fan shroud 170 to the airflow from fan 152 through engine inlet 182.

[0025] For fans with fan blades, the term "corrected tip speed" refers to the speed of the fan blade at the outer tip of the fan blade that has been corrected radially to correspond to the standard day conditions (i.e., the speed at which the fan blade rotates at its outer tip if the upstream temperature corresponds to the standard day conditions).

[0026] Typically, turbofan engines include relatively large fans to provide the desired amount of thrust without overloading the fan blades (i.e., without increasing the disk load on the fan blades beyond a certain threshold), and thus maintain the desired overall propulsive efficiency of the turbofan engine. Conventional turbofan engine design practice involves providing a large fan, or more precisely, a large-diameter fan, to deliver as much total thrust as possible. The goal in designing conventional turbofan engines is to maximize propulsive efficiency. However, including such large fans can lead to problems such as encapsulating the turbofan engine in an aircraft, resulting in relatively heavy turbofan engines (especially for ducted turbofan engines). Furthermore, as the demand for greater thrust from turbofan engines continues, the thermal requirements for them increase accordingly.

[0027] However, the inventors of this disclosure have discovered that for a three-flow turbofan engine with a primary fan and a secondary fan, where the secondary fan is a ducted fan that supplies airflow to the third flow of the engine, the required thrust generation from the primary fan can be reduced, while the secondary fan provides the difference through the third flow. Such a configuration can maintain the desired overall propulsive efficiency of the turbofan engine, or unexpectedly, can actually increase the super-propulsive efficiency of the turbofan engine.

[0028] In several different types of turbofan engines (including references below) Figure 1During the design of the described gas turbine engine, the inventors proceeded in the following manner: designing an engine with given primary fan characteristics, secondary fan characteristics, and turbine characteristics; checking the propulsion efficiency of the designed turbofan engine; redesigning a turbofan engine with different primary fan, secondary fan, and turbine characteristics; re-checking the propulsion efficiency of the redesigned turbofan engine; and so on. During the practice of studying / evaluating various primary fan characteristics, secondary fan characteristics, and turbine characteristics deemed feasible to best meet mission requirements, a relationship was found between the percentage of total turbofan engine thrust provided by the third flow (as defined herein) and the relative dimensions of the primary and secondary fans of the turbofan, or more specifically, the radius ratio of the primary to secondary fans. As described herein, the obtained radius ratio relationship with the third flow thrust can be considered an indicator of the turbofan engine's ability to maintain or even improve the desired propulsion efficiency via the third flow, and also indicates improvements in turbofan engine packaging and weight issues, as well as thermal management capabilities.

[0029] Now for reference Figure 1 A schematic cross-sectional view of a gas turbine engine according to another exemplary embodiment of the present disclosure is provided. Specifically, Figure 1 An engine with a rotor assembly having single-stage non-ducting rotor blades is provided. In this way, the rotor assembly may be referred to herein as a "non-ducting fan," or the entire engine 100 may be referred to herein as a "non-ducting engine." Furthermore, Figure 1 The engine includes a third flow that extends from the compressor section to the rotor assembly flow path on the turbine, which will be explained in more detail below.

[0030] For reference, engine 100 defines an axial direction A, a radial direction R, and a circumferential direction C. Furthermore, engine 100 defines an axial centerline or longitudinal axis 112 extending along the axial direction A. Generally, the axial direction A extends parallel to the longitudinal axis 112, the radial direction R extends outward and inward toward the longitudinal axis 112 in a direction perpendicular to the axial direction A, and the circumferential direction extends 360° around the longitudinal axis 112. Engine 100 extends, for example, between a front end 114 and a rear end 116 along the axial direction A.

[0031] Engine 100 includes a turbine 120 and a rotor assembly, also referred to as a fan section 150, positioned upstream therefrom. Typically, turbine 120 includes a compressor section, a combustion section, a turbine section, and an exhaust section in a sequential flow order. Specifically, as... Figure 1As shown, turbine 120 includes a core shroud 122 defining an annular core inlet 124. The core shroud 122 further at least partially surrounds the low-pressure system and the high-pressure system. For example, the shown core shroud 122 at least partially surrounds and supports a booster or low-pressure (“LP”) compressor 126 for pressurizing air entering turbine 120 through core inlet 124. A high-pressure (“HP”) multi-stage axial compressor 128 receives the pressurized air from the LP compressor 126 and further increases the air pressure. The pressurized airflow flows downstream to a combustor 130 in the combustion section, where fuel is injected into the pressurized airflow and ignited to increase the temperature and energy level of the pressurized air.

[0032] 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 do not imply any absolute speed and / or pressure values.

[0033] High-energy combustion products flow downstream from combustor 130 to high-pressure turbine 132. High-pressure turbine 128 drives high-pressure compressor 128 via high-pressure shaft 136. At this point, high-pressure turbine 128 is drivably connected to high-pressure compressor 128. The high-energy combustion products then flow to low-pressure turbine 134. Low-pressure turbine 134 drives components of low-pressure compressor 126 and fan section 150 via low-pressure shaft 138. At this point, low-pressure turbine 134 is drivably connected to components of low-pressure compressor 126 and fan section 150. In this example embodiment, LP shaft 138 is coaxial with HP shaft 136. After driving each of turbines 132 and 134, combustion products exit turbine 120 through turbine exhaust nozzle 140.

[0034] Therefore, turbine 120 defines a working gas flow path or core duct 142 extending between core inlet 124 and turbine exhaust nozzle 140. Core duct 142 is an annular duct positioned approximately inside core casing 122 in the radial direction R. Core duct 142 (e.g., through the working gas flow path of turbine 120) may be referred to as a second flow.

[0035] Fan section 150 includes fan 152, which, in this example embodiment, is a primary fan. For Figure 1 In the illustrated embodiment, fan 152 is an open rotor or non-ducted fan 152. As depicted, fan 152 includes an array of fan blades 154. Figure 1(Only one is shown in the image). The fan blades 154 are rotatable, for example, about the longitudinal axis 112. As described above, the fan 152 is drivenly connected to the low-pressure turbine 134 via the LP shaft 138. For Figure 1 In the illustrated embodiment, for example in an indirect drive or gear drive configuration, the fan 152 is connected to the LP shaft 138 via a reduction gearbox 155.

[0036] Furthermore, the fan blades 154 may be arranged at equal intervals around the longitudinal axis 112. Each blade 154 has a root and a tip, and a span defined between them. Additionally, each fan blade 154 defines a fan blade tip radius R1 in the radial direction R from the longitudinal axis 12 to the tip, and a hub radius (or inner radius) R2 in the radial direction R from the longitudinal axis 12 to the base. Furthermore, each fan blade 154 of the fan 152, or more precisely, the fan 152, is defined with a fan radius ratio RqR equal to R2 divided by R1. Since the fan 150 is the primary fan of the engine 100, the fan radius ratio RqR of the fan 152 can be referred to as the primary fan radius ratio RqR. Prim.-Fan .

[0037] Furthermore, each blade 154 defines a central blade axis 156. In this embodiment, each blade 154 of the fan 152 can rotate about its respective central blade axis 156, for example, in unison with each other. One or more actuators 158 are provided to facilitate this rotation and are therefore used to change the pitch of the blades 154 about their respective central blade axes 156.

[0038] Fan section 150 also includes a fan guide vane array 160, which includes fan guide vanes 162 arranged around a longitudinal axis 112. Figure 1 (Only one is shown). In this embodiment, the fan guide vane 162 cannot rotate about the longitudinal axis 112. Each fan guide vane 162 has a root and a tip, and a span defined between them. The fan guide vane 162 can be as follows: Figure 1 The area shown is not covered, or alternatively, it can be covered, for example, by an annular shroud spaced outward from the tip of the fan guide vane 162 in the radial direction R or attached to the fan guide vane 162.

[0039] Each fan guide vane 162 defines a central blade axis 164. In 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 this rotation and are therefore used to change the pitch of the fan guide vanes 162 about their respective central blade axes 164. However, in other embodiments, each fan guide vane 162 may be fixed or unable to change its pitch about its central blade axis 164. The fan guide vanes 162 are mounted to the fan shroud 170.

[0040] like Figure 1 As shown, in addition to the non-ductless fan 152, a ducted fan 184 is also included downstream of fan 152, such that engine 100 includes both ducted and non-ducted fans, both used to generate thrust by the movement of air without passing through at least a portion of turbine 120 (e.g., in the depicted embodiment, without passing through HP compressor 128 and combustion chamber). The ducted fan is rotatable about the same axis as fan blade 154. In the depicted embodiment, ducted fan 184 is driven by low-pressure turbine 134 (e.g., coupled to LP shaft 138). In the depicted embodiment, as described above, fan 152 may be referred to as the primary fan, and ducted fan 184 may be referred to as the secondary fan. It should be understood that these terms "primary" and "secondary" are convenient terms and do not imply any particular importance, power, etc.

[0041] The duct fan 184 includes multiple fan blades ( Figure 1 (Not separately marked). The fan blades of the duct fan 184 may be arranged at equal intervals around the longitudinal axis 112. Each blade of the duct fan 184 has a root and a tip, and a span defined between them. Furthermore, each fan blade of the duct fan 184 defines a fan blade tip radius R3 in the radial direction R from the longitudinal axis 12 to the tip, and a hub radius (or inner radius) R4 in the radial direction R from the longitudinal axis 12 to the base. Moreover, the duct fan 184, or more precisely, each fan blade of the duct fan 184, defines a fan radius ratio RqR equal to R4 divided by R3. Since the duct fan 184 is a secondary fan of the engine 100, the fan radius ratio RqR of the duct fan 184 may be referred to as the secondary fan radius ratio RqR. Sec.-Fan .

[0042] The fan shroud 170 annularly surrounds at least a portion of the core shroud 122 and is positioned generally radially R outside at least a portion of the core shroud 122. Specifically, a downstream section of the fan shroud 170 extends over the front portion of the core shroud 122 to define a fan flow path or fan duct 172. According to this embodiment, the fan flow path or fan duct 172 can be understood as forming at least a portion of a third flow of the engine 100.

[0043] Incoming air enters fan duct 172 through fan duct inlet 176 and exits through fan exhaust nozzle 178 to generate thrust. Fan duct 172 is an annular duct positioned radially R generally outside core duct 142. Fan shroud 170 and core shroud 122 are connected together and are supported by a plurality of substantially radially extending, circumferentially spaced stationary struts 174. Figure 1 Only one support is shown in the diagram. The stationary struts 174 may each be aerodynamically shaped to guide airflow therefrom. Other struts besides the stationary struts 174 may be used to connect and support the fan shroud 170 and / or the core shroud 122. In many embodiments, the fan duct 172 and the core duct 142 may extend at least partially together (generally axially) on opposite sides (e.g., opposite radial sides) of the core shroud 122. For example, the fan duct 172 and the core duct 142 may each extend directly from the leading edge 144 of the core shroud 122 and may extend generally axially and partially together on opposite radial sides of the core shroud.

[0044] Engine 100 also defines or includes an inlet duct 180. The inlet duct 180 extends between engine inlet 182 and core inlet 124 / fan duct inlet 176. Engine inlet 182 is generally defined at the front end of fan shroud 170 and positioned along the axial direction A between fan 152 and fan guide vane array 160. Inlet duct 180 is an annular duct positioned radially R inside fan shroud 170. Air flowing downstream along inlet duct 180 is diverted (not necessarily uniformly) by a splitter or leading edge 144 of core shroud 122 into core duct 142 and fan duct 172. Inlet duct 180 is wider radially R than core duct 142. Inlet duct 180 is also wider radially R than fan duct 172.

[0045] During operation of engine 100 under operating conditions, engine 100 generates total thrust Fn. Total Operating conditions can be that the engine operates at rated speed at 100 during standard day operating conditions. Total thrust is the first-rate thrust Fn. 1S (For example, primary fan thrust generated by fan 152, generated by airflow above fan shroud 170 and core shroud 122), third-stream thrust Fn 3S(For example, the thrust generated by the airflow through the fan duct 172 exiting through the fan exhaust nozzle 178, at least partially generated by the duct fan 184), and the second-flow thrust Fn. 2S (For example, the sum of thrust generated by the airflow through the core duct 142 exiting through the exhaust nozzle 140).

[0046] It is worth noting that, for the depicted embodiment, engine 100 includes one or more features to increase third-flow thrust Fn. 3S Efficiency. Specifically, the engine 100 also includes an array of inlet guide vanes 186 positioned in an inlet duct 180 upstream of the ducted fan 184 and downstream of the engine inlet 182. The array of inlet guide vanes 186 is arranged about a longitudinal axis 112. In this embodiment, the fan inlet guide vanes 186 cannot rotate about the longitudinal axis 112. Each inlet guide vane 186 defines a central blade axis (not labeled for clarity) and can rotate about its respective central blade axis, for example, in unison with each other. One or more actuators 188 are provided to facilitate this rotation and can therefore be used to change the pitch of the inlet guide vanes 186 about their respective central blade axes. However, in other embodiments, each inlet guide vane 186 may be fixed about its central blade axis or may not pitch about its central blade axis.

[0047] Furthermore, downstream of the duct fan 184 and upstream of the fan duct inlet 176, the engine 100 includes an array of outlet guide vanes 190. Like the array of inlet guide vanes 186, the array of outlet guide vanes 190 cannot rotate about the longitudinal axis 112. However, in the depicted embodiment, unlike the array of inlet guide vanes 186, the array of outlet guide vanes 190 is configured as fixed-pitch outlet guide vanes.

[0048] Furthermore, it should be understood that, for the depicted embodiment, the fan exhaust nozzle 178 of the fan duct 172 is also configured as a variable geometry exhaust nozzle. In this way, the engine 100 includes one or more actuators 192 for adjusting the variable geometry exhaust nozzle. For example, the variable geometry exhaust nozzle may be configured to change the total cross-sectional area (e.g., the area of ​​the nozzle in a plane perpendicular to the longitudinal axis 112) to adjust the thrust generated based on one or more engine operating conditions (e.g., the temperature, pressure, mass flow rate, etc. of the airflow through the fan duct 172). A fixed geometry exhaust nozzle may also be used.

[0049] The combination of the array of inlet guide vanes 186 located upstream of the duct fan 184, the array of outlet guide vanes 190 located downstream of the duct fan 184, and the exhaust nozzle 178 can result in more efficient generation of third-flow thrust Fn during one or more engine operating conditions. 3S Furthermore, by introducing variability into the geometry of the inlet guide vane 186 and the exhaust nozzle 178, the engine 100 can operate under a relatively wide range of engine operating conditions (including takeoff and climb, which typically require maximum total engine thrust Fn). Total And cruise (which typically requires a smaller amount of total engine thrust Fn) Total This generates a more effective third-flow thrust Fn. 3S .

[0050] Still referencing Figure 1 In an exemplary embodiment, the air passing through fan duct 172 may be relatively colder (e.g., at a lower temperature) than one or more fluids used in turbine 120. Thus, one or more heat exchangers 200 may be positioned in thermal communication with fan duct 172. For example, 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 fluid (e.g., compressor bleed air, oil, or fuel).

[0051] Although not depicted, heat exchanger 200 may be an annular heat exchanger extending approximately 360 degrees (e.g., at least 300 degrees, such as at least 330 degrees) within fan duct 172. In this way, heat exchanger 200 can effectively utilize air passing through fan duct 172 to cool one or more systems of engine 100 (e.g., lubrication system, compressor bleed air, electrical components, etc.). Heat exchanger 200 uses air passing through duct 172 as a radiator and accordingly increases the temperature of the air downstream of heat exchanger 200 and exiting nozzle 178.

[0052] As previously stated, the inventors unexpectedly discovered, during the design of turbofan engines (i.e., designing turbofan engines (in-line and out-of-line turbofan engines) with various primary and secondary fan characteristics (physical and operational characteristics) and evaluating overall propulsion efficiency), a significant relationship exists between the percentage of total turbofan engine thrust provided by the third stream (as defined herein) and the relative dimensions of the primary and secondary fans of the turbofan. As described herein, the resulting radius ratio relationship with the third stream thrust can be considered an indicator of the turbofan engine's ability to maintain or even improve the desired propulsion efficiency via the third stream, and further indicates improvements in turbofan engine packaging and weight issues, as well as thermal management capabilities.

[0053] As will be understood, higher and lower third-flow thrust alters the packing capacity and radiator capacity of a turbofan engine. For example, increased thrust from the airflow through the third flow typically means more airflow (on a mass flow rate basis) through the third flow, which in turn means a greater heat capacity of that airflow. Furthermore, the inventors found that if the thrust provided from the third flow is too small, the turbofan engine may not have to be large (and therefore more difficult to pack) and heavy, and may further fail to provide the desired amount of radiator capacity. If too much thrust is provided through the third flow, the engine may not be able to fully utilize the relatively effective thrust that can be generated by the primary fan.

[0054] The aforementioned relationship can be a function of the bypass ratio of a turbofan engine, which is typically limited by reasonable engine temperatures (including operating temperatures, such as exhaust gas temperature (EGT)). For example, as will be understood in light of the foregoing teachings, the radius of the primary fan relative to the radius of the secondary fan, and the percentage of total turbofan engine thrust generated during operation by the airflow through the third flow, are each in part a function of the bypass ratio and together characterize the balance in the aforementioned relationship.

[0055] Many aspects of the architecture determine the bypass ratio of a turbofan engine. For example, the bypass ratio is partly a function of the primary fan's corrected tip speed relative to the secondary fan's corrected tip speed, and the specific thrust of the corresponding primary and secondary fans. The specific thrust of the primary and secondary fans is, in turn, a function of the pressure ratio of the primary and secondary fans, and the disk load (also known as the power load) on the primary and secondary fans, respectively. These factors also affect the balance in the above relationships, as will be described in more detail below with reference to the effective fan parameter EFP.

[0056] As described above, the inventors of this disclosure have discovered a relationship between the percentage of engine thrust provided by the airflow through the third flow and the radius ratio of the primary and secondary fans, which can result in the turbofan engine maintaining or even improving the desired propulsion efficiency, while also improving the packaging and weight issues of the turbofan engine, and providing the desired thermal management capabilities.

[0057] Using this relationship, the inventors discovered that the number of suitable or feasible turbofan engine designs combining primary and secondary fans and defining a third flow can meet propulsion efficiency requirements, and that packaging, weight, and radiator requirements can be significantly reduced, thus facilitating faster selection of designs to consider when developing turbofan engines. This benefit allows for a deeper understanding of the requirements of a given turbofan engine before fully developing specific technologies, integration, and system requirements. It avoids later redesign. The desired relationship is expressed by the following equation (1):

[0058] Equation (1):

[0059] Where R1 is the tip radius of the primary fan, R2 is the hub radius of the primary fan, R3 is the tip radius of the secondary fan, R4 is the hub radius of the secondary fan, and RqR Prim.-Fan It is the ratio of R2 to R1, RqR Sec.-Fan It is the ratio of R4 to R3, %Fn 3S It is the percentage of the third-stream thrust relative to the engine's total thrust (e.g., for...). Figure 1 In an embodiment, Fn 3S Divide by Fn Total EFP is called the effective fan parameter. In terms of equation (1), %Fn 3S This is limited when the engine is operated at rated speed during standard day operating conditions.

[0060] EFP is a function of the corrected tip speed of the primary fan, the corrected tip speed of the secondary fan, the disk load of the primary fan, and the disk load of the secondary fan. By taking into account the corrected tip speeds of the primary and secondary fans, EFP accounts for factors such as the specific engine configuration (e.g., geared, direct drive, etc.), which may affect the tip radius ratio (R1 vs. R3) and the percentage of thrust through the third flow (%Fn) of a turbofan engine with the desired propulsive efficiency. 3S The relationship between these factors and the tip radius ratio (R1 vs. R3) of the turbofan engine and the percentage of thrust through the third flow (%Fn) is described in more detail above. 3S The relationship between ).

[0061] R l The values ​​of / R3 and the corresponding values ​​of the engine's influence characteristics defined by equation (1) are listed in Table 1:

[0062]

[0063] Figures 2A to 2D and Figure 3 A gas turbine engine according to one or more exemplary embodiments of the present disclosure is shown, illustrating the relationship between the tip radius ratio and the percentage of thrust through the third flow. Specifically, Figures 2A to 2D Provided including corresponding Figure 3 A table showing the numerical values ​​of several gas turbine engines. Figure 3 This is a graph of a gas turbine engine according to one or more exemplary embodiments of the present disclosure, showing the tip radius ratio (R1 vs. R3; Y-axis) and the percentage of thrust through the third flow (%Fn). 3S The relationship between the X-axis and the X-axis.

[0064] It is worth noting that, in Figure 3 The document provides a first range and a second range. The first range can correspond to EFP between 1.5 and 11, where %Fn 3S It is between approximately 2% and approximately 50%. This can result in the engine having the desired propulsive efficiency.

[0065] The second range can correspond to approximately 2.5 and approximately 4 EFP, where %Fn 3S It is between approximately 5% and approximately 20%. This can result in the engine having a better propulsion efficiency.

[0066] As will be understood from the description herein, various embodiments of gas turbine engines are provided. Some of these embodiments may be non-ducting single-rotor gas turbine engines or ducting turbofan engines. Examples of ducting turbofan engines can be found in U.S. Patent Application Publication No. 16 / 811,368 (published as U.S. Patent Application Publication No. 2021 / 0108597), filed March 6, 2020 (Figure 10, paragraph

[0062] , etc.; including an annular fan housing 13 comprising airfoil blades 21 surrounding a rotating element 20 and wheel blades 31 surrounding a stationary element 30; and including a third flow / fan duct 73 (as shown in Figure 10, extensively described throughout the application)). Various additional aspects of one or more of these embodiments are discussed below. These exemplary aspects may be combined with one or more of the exemplary gas turbine engines discussed above with respect to the drawings.

[0067] For example, in some embodiments of this disclosure, the engine may include a heat exchanger located in an annular duct (e.g., in a third flow). The heat exchanger may extend substantially continuously in the circumferential direction of the gas turbine engine (e.g., at least about 300 degrees, such as at least about 330 degrees).

[0068] In one or more of these embodiments, at the cruise altitude during cruise operation mode, the threshold power or disk load of the fan (e.g., a non-ducted single-rotor or front primary fan) can be 25 hp / ft. 2 (or a wider range). In a specific embodiment of the engine, at the cruise altitude during cruise operation mode, the structures and methods provided herein generate at 80 hp / ft 2 and 160hp / ft 2 The power load can be between or higher, depending on whether the engine is an open rotor or a tubular engine.

[0069] In various embodiments, the engine of this disclosure is applied to a vehicle with a cruising altitude up to approximately 65,000 ft. In some 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 conditions are between FL280 and FL650. In another embodiment, the cruising flight conditions are between FL280 and FL450. In yet another embodiment, the cruising altitude is defined at least based on atmospheric pressure, wherein the cruising altitude is between approximately 4.85 psia and approximately 0.82 psia based on a sea level pressure of approximately 14.70 psia and a sea level temperature of approximately 59 degrees Fahrenheit. In another embodiment, the cruising altitude is between approximately 4.85 psia and approximately 2.14 psia. It should be understood that in some embodiments, the pressure-defined range of cruising altitude may be adjusted based on different reference sea level pressures and / or sea level temperatures.

[0070] Therefore, it should be understood that an engine of this configuration can be configured to generate at least about 25,000 pounds and less than about 80,000 pounds of thrust during operation at rated speed, for example, between about 25,000 and 50,000 pounds of thrust during operation at rated speed, for example, between about 25,000 and 40,000 pounds of thrust during operation at rated speed.

[0071] In various exemplary embodiments, the fan may include twelve (12) fan blades. From a load perspective, such a number of blades allows for a reduction in the span of each blade, thereby reducing the overall diameter of the primary fan (e.g., to about 12 feet in one exemplary embodiment). That is, in other embodiments, the fan may have any suitable number of blades and any suitable diameter. In some suitable embodiments, the fan includes at least eight (8) blades. In another suitable embodiment, the fan may have at least twelve (12) blades. In yet another suitable embodiment, the fan may have at least fifteen (15) blades. In yet another suitable embodiment, the fan may have at least eighteen (18) blades. In one or more of these embodiments, the fan includes twenty-six (26) or fewer blades, such as twenty (20) or fewer blades.

[0072] Furthermore, in some exemplary embodiments, the rotor assembly may be defined with a rotor diameter (or fan diameter) of at least 10 feet (e.g., at least 11 feet, at least 12 feet, at least 13 feet, at least 15 feet, at least 17 feet, up to 28 feet, up to 26 feet, up to 24 feet, up to 18 feet).

[0073] In various embodiments, it should be understood that the engine includes a ratio of the number of impeller blades to the number of blades that may be less than, equal to, or greater than 1:1. For example, in a particular embodiment, the engine includes twelve (12) fan blades and ten (10) impeller blades. In other embodiments, the impeller assembly includes a greater number of impeller blades than fan blades. For example, in a particular embodiment, the engine includes ten (10) fan blades and twenty-three (23) impeller blades. For example, in some embodiments, the engine may include a ratio of the number of impeller blades to the number of blades that is between 1:2 and 5:2. This ratio may be adjusted based on a variety of factors, including impeller blade size, to ensure that the desired amount of swirl is removed from the airflow from the primary fan.

[0074] Furthermore, in some exemplary embodiments, when the engine includes a third flow and a middle fan (a ducted fan following the primary fan), the ratio R1 / R2 can be between about 1 and 10, or between 2 and 7, or at least about 3.3, at least about 3.5, at least about 4 and less than or equal to about 7, where R1 is the radius of the primary fan and R2 is the radius of the middle fan.

[0075] It should be understood that various embodiments of the engine (such as the single non-ducting rotary engine depicted and described herein) can allow operation at normal subsonic aircraft cruise altitudes equal to or higher than Mach 0.5. In some 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 some embodiments, the engine allows rotor blade tip velocities equal to or less than 750 feet per second (fps). In other embodiments, rotor blade tip velocities under cruise flight conditions can be 650 to 900 fps, or 700 to 800 fps.

[0076] As measured across the fan blades under cruise flight conditions, the fan pressure ratio (FPR) of the fan assembly can be from 1.04 to 1.20, or in some embodiments from 1.05 to 1.1, or in some embodiments less than 1.08.

[0077] To enable the gas turbine engine to operate in conjunction with a fan having the aforementioned characteristics to define the aforementioned FPR, a gear assembly can be provided to reduce the rotational speed of the fan assembly relative to the drive shaft (e.g., the low-pressure shaft coupled to the low-pressure turbine). In some embodiments, the gear ratio of the input speed to the output speed is greater than 4.1. For example, in certain embodiments, the gear ratio is in the range of 4.1 to 14.0, in the range of 4.5 to 14.0, or in the range of 6.0 to 14.0. In some embodiments, the gear ratio is in the range of 4.5 to 12, or in the range of 6.0 to 11.0. Thus, in some embodiments, the fan can be configured to rotate at a speed of 700 to 1500 rpm under cruise flight conditions, while the power turbine (e.g., the low-pressure turbine) is configured to rotate at a speed of 2,500 to 15,000 rpm under cruise flight conditions. In a particular embodiment, the fan may be configured to rotate at a speed of 850 to 1,350 rpm under cruise flight conditions, while the power turbine may be configured to rotate at a speed of 5,000 to 10,000 rpm under cruise flight conditions.

[0078] Regarding the turbine, compressor, and / or turbine of a gas turbine engine, the number of stages can include various numbers of stages. As disclosed herein, the number of stages includes the number of rotor or blade stages in a particular component (e.g., compressor or turbine). For example, in some embodiments, a low-pressure compressor can include 1 to 8 stages, a high-pressure compressor can include 8 to 15 stages, a high-pressure turbine can include 1 to 2 stages, and / or a low-pressure turbine (LPT) can include 3 to 7 stages. In particular, an LPT can have 4 stages, or 4 to 7 stages. For example, in some embodiments, the engine can include a single-stage low-pressure compressor, an 11-stage high-pressure compressor, a two-stage high-pressure turbine, and an LPT with 4 stages or between 4 and 7 stages. As another example, the engine can include a three-stage low-pressure compressor, a 10-stage high-pressure compressor, a two-stage high-pressure turbine, and a 7-stage low-pressure turbine.

[0079] The core engine is typically encapsulated in a housing that defines half the core diameter (Dcore), which can be considered as the maximum range from the centerline axis (a reference to R). In some embodiments, the engine comprises a length (L) from the longitudinal (or axial) front end to the longitudinal rear end. In various embodiments, the engine defines an L / Dcore ratio that provides reduced mounting resistance. In one embodiment, the L / Dcore is at least 2. In another embodiment, the L / Dcore is at least 2.5. In some embodiments, the L / Dcore is less than 5, less than 4, and less than 3. In various embodiments, it should be understood that the L / Dcore is for a single non-inductor rotary engine.

[0080] Reduced installation drag can further provide improved efficiency, such as improved fuel consumption rate. Additionally or alternatively, reduced drag can provide improved cruise altitude for engines and aircraft operating at the aforementioned Mach numbers. Certain embodiments may still offer the benefit of reduced noise from the interaction between blade assemblies and wheel blade assemblies, and / or reduced total noise generated by the engine, through structures located in the engine's annular duct.

[0081] Furthermore, it should be understood that the range of power loads and / or rotor blade tip velocities may correspond to certain structures, core dimensions, thrust outputs, etc., or other structures at the core engine. However, as previously stated, where one or more structures provided herein may be known in the art, it should be understood that this disclosure may include combinations of structures previously unknown, at least in part due to conflict of interest, desired operating modes, or other forms of teaching in the art.

[0082] Although described above in the embodiments as unshielded or open rotor engines, it should be understood that the aspects of this disclosure provided herein can be applied to shielded or ducted engines, partially ducted engines, rear-fan engines, or other gas turbine engine configurations, including those for marine, industrial, or aerospace propulsion systems. Certain aspects of this disclosure are applicable to turbofan engines, turboprop engines, or turboshaft engines. However, it should be understood that certain aspects of this disclosure can address problems that may be specific to unshielded or open rotor engines, such as, but not limited to, problems related to gear ratios, fan diameter, fan speed, engine length (L), the maximum diameter of the engine's core (Dcore), the engine's L / Dcore ratio, desired cruise altitude, and / or desired operating cruise speed, or combinations thereof.

[0083] This written description uses examples to disclose this disclosure, including best practices, and also enables any person skilled in the art to practice this disclosure, including making and using any device or system and methods of making any combination. The patent scope of this disclosure is defined by the claims, but may include other examples that would occur to a person skilled in the art. Such other examples are intended to fall within the scope of the claims if they include structural elements that are not indistinguishable from the literal language of the claims, or if they include equivalent structural elements that are not substantially different from the literal language of the claims.

[0084] Further details are provided by the following topics:

[0085] 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 sequential flow order, the turbine defining a working gas flow path and a fan duct flow path; a primary fan driven by the turbine, the primary fan defining a primary fan tip radius R1 and a primary fan hub radius R2; and a secondary fan located downstream of the primary fan and driven by the turbine, at least a portion of the airflow from the primary fan being configured to bypass the secondary fan, the secondary fan defining a secondary fan tip radius R3 and a secondary fan hub radius R4, wherein the secondary fan is configured to provide fan duct airflow through the fan duct flow path during operation to generate fan duct thrust, wherein the fan duct thrust is equal to %Fn of the total engine thrust during operation of the gas turbine engine at rated speed under standard day operating conditions. 3S The ratio of R1 to R3 is equal to...

[0086]

[0087] EFP is between 1.5 and 11, where RqR Prim.-Fan It is the ratio of R2 to R1, where RqR Sec.-Fan It is the ratio of R4 to R3.

[0088] According to one or more of these clauses, the ratio of R1 to R3 is between about 2 and about 10, for example, between about 2 and about 7.

[0089] According to one or more of these clauses, the ratio of R1 to R3 is between about 3 and about 5.

[0090] According to one or more of these clauses, the gas turbine engine has an EFP between about 2.5 and about 4, and %Fn 3S Greater than or equal to approximately 5% and less than or equal to approximately 20%.

[0091] According to one or more of these clauses, the gas turbine engine, wherein RqR Prim.-Fan Between 0.2 and 0.4.

[0092] According to one or more of these clauses, the gas turbine engine, wherein RqR Prim.-Fan Between 0.25 and 0.35.

[0093] According to one or more of these clauses, the gas turbine engine, wherein RqR Sec.-Fan Between 0.2 and 0.7.

[0094] According to one or more of these clauses, the gas turbine engine, wherein RqR Sec.-Fan Between 0.35 and 0.5.

[0095] According to one or more of these clauses, the gas turbine engine, wherein the EFP is between A2 and B2, wherein the primary fan defines the primary fan correction tip speed during operation of the gas turbine engine at the rated speed under standard day operating conditions, wherein the secondary fan defines the secondary fan correction tip speed during operation of the gas turbine engine at the rated speed under standard day operating conditions, wherein the primary fan correction tip speed is between 500 ft / s and 2,000 ft / s, and wherein the secondary fan correction tip speed is between 500 ft / s and 2,000 ft / s.

[0096] According to one or more of these clauses, the gas turbine engine, wherein %Fn 3S Between 1% and 50%.

[0097] According to one or more of these clauses, the gas turbine engine, wherein %Fn 3S Between 3% and 30%.

[0098] According to one or more of these clauses, the gas turbine engine, wherein %Fn 3S Between 5% and 20%.

[0099] According to one or more of these clauses, the gas turbine engine, wherein the fan duct flow path defines an outlet, and wherein the gas turbine engine further includes: a variable geometry component associated with the secondary fan, wherein the variable geometry component is a variable inlet guide vane stage positioned immediately upstream of the secondary fan, a variable exhaust nozzle located at the outlet of the fan duct flow path, or both.

[0100] According to one or more of these clauses, the gas turbine engine wherein the primary fan is a non-ducted fan.

[0101] According to one or more of these clauses, the gas turbine engine defines a bypass airflow passage, wherein the primary fan is configured to provide a first portion of the primary fan airflow to the bypass airflow passage and a second portion of the primary fan airflow to the secondary fan, and wherein the secondary fan is configured to provide a first portion of the secondary fan airflow as the fan duct airflow to the fan duct flow path and a second portion of the secondary fan airflow to the working gas flow path.

[0102] The gas turbine engine according to one or more of these clauses further includes: a heat exchanger positioned in thermal communication with the fan duct flow path.

[0103] The gas turbine engine according to one or more of these clauses further includes: an inlet guide vane array positioned immediately upstream of the secondary fan.

[0104] The gas turbine engine according to one or more of these clauses further includes: an outlet guide vane array positioned immediately downstream of the secondary fan and upstream of the fan duct.

[0105] The gas turbine engine according to one or more of these clauses further includes: a variable geometry exhaust nozzle located at the outlet of the fan duct.

[0106] The gas turbine engine according to one or more of these clauses further includes: a fan shroud surrounding the secondary fan located downstream of the primary fan, the fan shroud partially defining an engine inlet located downstream of the primary fan; wherein the turbine further includes a core shroud at least partially surrounding the compressor section, the combustion section, and the turbine section, and wherein the fan duct is defined between the core shroud and the fan shroud.

[0107] 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 sequential flow order, the turbine defining a working gas flow path and a fan duct flow path; a primary fan driven by the turbine, the primary fan defining a primary fan tip radius R1 and a primary fan hub radius R2; and a secondary fan located downstream of the primary fan and driven by the turbine, at least a portion of the airflow from the primary fan being configured to bypass the secondary fan, the secondary fan defining a secondary fan tip radius R3 and a secondary fan hub radius R4, wherein the secondary fan is configured to provide fan duct airflow through the fan duct flow path during operation to generate fan duct thrust, wherein the fan duct thrust is equal to %Fn of the total engine thrust during operation of the gas turbine engine at rated speed under standard day operating conditions. 3S The ratio of R1 to R3 is between 2 and 7; optionally between 3 and 7; optionally between 4 and 7; optionally between 5 and 7.

Claims

1. A gas turbine engine defining a centerline and a circumferential direction, characterized in that, The gas turbine engine includes: A turbine comprising a compressor section, a combustion section and a turbine section arranged in a sequential flow order, the turbine defining a working gas flow path and a fan duct flow path; A primary fan, driven by the turbine, the primary fan defining a primary fan tip radius R1 and a primary fan hub radius R2; A secondary fan, located downstream of the primary fan and driven by the turbine, has at least a portion of the airflow from the primary fan configured to bypass it. The secondary fan defines a tip radius R3 and a hub radius R4. The secondary fan is configured to provide fan duct airflow through the fan duct flow path during operation to generate fan duct thrust, wherein the fan duct thrust is equal to %Fn of the total engine thrust during operation at rated speed under standard day operating conditions. 3S ; The ratio of R1 to R3 is equal to EFP is between 1.5 and 11, where RqR Prim.-Fan It is the ratio of R2 to R1, where RqR Sec.-Fan It is the ratio of R4 to R3.

2. The gas turbine engine according to claim 1, characterized in that, in, The ratio of R1 to R3 is between approximately 2 and approximately 10.

3. The gas turbine engine according to claim 1, characterized in that, in, The ratio of R1 to R3 is between approximately 3 and approximately 7.

4. The gas turbine engine according to claim 1, characterized in that, in, EFP is between approximately 2.5 and approximately 4, and %Fn 3S Greater than or equal to approximately 5% and less than or equal to approximately 20%.

5. The gas turbine engine according to claim 1, characterized in that, in, RqR Prim.-Fan Between 0.2 and 0.

4.

6. The gas turbine engine according to claim 1, characterized in that, in, RqR Prim.-Fan Between 0.25 and 0.

35.

7. The gas turbine engine according to claim 1, characterized in that, in, RqR Sec.-Fan Between 0.2 and 0.

7.

8. The gas turbine engine according to claim 1, characterized in that, in, RqR Sec.-Fan Between 0.35 and 0.

5.

9. The gas turbine engine according to claim 1, characterized in that, in, EFP is between A2 and B2, wherein the primary fan defines the primary fan correction tip speed during operation of the gas turbine engine at the rated speed under standard day operating conditions, wherein the secondary fan defines the secondary fan correction tip speed during operation of the gas turbine engine at the rated speed under standard day operating conditions, wherein the primary fan correction tip speed is between 500 ft / s and 2,000 ft / s, and wherein the secondary fan correction tip speed is between 500 ft / s and 2,000 ft / s.

10. The gas turbine engine according to claim 1, characterized in that, in, %Fn 3S Between 1% and 50%.

11. The gas turbine engine according to claim 1, characterized in that, in, %Fn 3S Between 3% and 30%.

12. The gas turbine engine according to claim 1, characterized in that, in, %Fn 3S Between 5% and 20%.

13. The gas turbine engine according to claim 1, characterized in that, in, The fan duct flow path defines an outlet, and the gas turbine engine further includes: A variable geometry component associated with the secondary fan, wherein the variable geometry component is a variable inlet guide vane stage positioned immediately upstream of the secondary fan, a variable exhaust nozzle located at the outlet in the fan duct flow path, or both.

14. The gas turbine engine according to claim 1, characterized in that, in, The primary fan is a non-ducted fan.

15. The gas turbine engine according to claim 14, characterized in that, in, The gas turbine engine defines a bypass airflow passage, wherein the primary fan is configured to provide a first portion of the primary fan airflow to the bypass airflow passage and a second portion of the primary fan airflow to the secondary fan, and wherein the secondary fan is configured to provide a first portion of the secondary fan airflow as the fan duct airflow to the fan duct flow path and a second portion of the secondary fan airflow to the working gas flow path.

16. The gas turbine engine according to claim 1, characterized in that, Further includes: A heat exchanger, which is positioned in thermal communication with the flow path of the fan duct.

17. The gas turbine engine according to claim 1, characterized in that, Further includes: An inlet guide vane array is positioned immediately upstream of the secondary fan.

18. The gas turbine engine according to claim 17, characterized in that, Further includes: An outlet guide vane array is positioned immediately downstream of the secondary fan and upstream of the fan duct.

19. The gas turbine engine according to claim 1, characterized in that, Further includes: A variable geometry exhaust nozzle is located at the outlet of the fan duct.

20. The gas turbine engine according to claim 1, characterized in that, Further includes: A fan shroud surrounds the secondary fan located downstream of the primary fan, and the fan shroud partially defines an engine inlet located downstream of the primary fan; The turbine further includes a core shroud that at least partially surrounds the compressor section, the combustion section, and the turbine section, and the fan duct is defined between the core shroud and the fan shroud.

21. An aircraft comprising a gas turbine engine defining a centerline and a circumferential direction, characterized in that, The gas turbine engine includes: A turbine comprising a compressor section, a combustion section and a turbine section arranged in a sequential flow order, the turbine defining a working gas flow path and a fan duct flow path; A primary fan, driven by the turbine, the primary fan defining a primary fan tip radius R1 and a primary fan hub radius R2; A secondary fan, located downstream of the primary fan and driven by the turbine, has at least a portion of the airflow from the primary fan configured to bypass it. The secondary fan defines a tip radius R3 and a hub radius R4. The secondary fan is configured to provide fan duct airflow through the fan duct flow path during operation to generate fan duct thrust, wherein the fan duct thrust is equal to %Fn of the total engine thrust during operation of the gas turbine engine at rated speed under standard day operating conditions. 3S ; The ratio of R1 to R3 is equal to EFP is between 1.5 and 11, where RqR Prim.-Fan It is the ratio of R2 to R1, where RqR Sec.-Fan It is the ratio of R4 to R3.

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