Gas turbine engine with third flow

Abstract

A gas turbine engine is provided, comprising a turbine having an inlet splitter and a fan duct splitter, the inlet splitter partially defining an inlet to a working gas flow path, and the fan duct splitter partially defining an inlet to a fan duct flow path. The gas turbine engine also includes a main fan driven by the turbine, defining a main fan tip radius R1, a main fan hub radius R2, and a main fan specific thrust rating T. P ; and a secondary fan, located downstream of the main fan and driven by a turbine, the secondary fan defining a secondary fan tip radius R3, a secondary fan hub radius R4, and a secondary fan specific thrust rating T. S The gas turbine engine defines the effective bypass area, and the ratio of R1 to R3 is equal to...

Description

Technical Field

[0001] This disclosure relates to a gas turbine engine with a third flow. Background Technology

[0002] A gas turbine engine typically comprises 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 directed to the combustor, where it is mixed with fuel. The mixture is then ignited to produce hot combustion gases. The combustion gases are directed to the turbine, which extracts energy from the combustion gases to power the compressor and to generate useful work to propel the aircraft in flight. The turbine is mechanically coupled to the fan to drive the fan during operation. Attached Figure Description

[0003] The complete and enabling disclosure of this disclosure, including its best mode, is set forth in the specification with reference to the accompanying drawings, for those skilled in the art, wherein:

[0004] Figure 1 This is a schematic cross-sectional view of a three-flow engine according to an exemplary embodiment of the present disclosure.

[0005] Figure 2 yes Figure 1 A close-up schematic diagram of an exemplary three-flow engine.

[0006] Figure 3 yes Figure 2 A close-up view of the area surrounding the leading edge of the core cowling of an exemplary three-flow engine.

[0007] Figure 4 This describes an exemplary main fan specific thrust rating T according to an exemplary aspect of this disclosure. P Values ​​and exemplary secondary fan specific thrust ratings T S A graph of the values.

[0008] Figures 5A to 5F This is a table representing an example embodiment of this disclosure.

[0009] Figures 6A to 6C It is a graph depicting the range of radius ratios versus effective bypass areas according to various exemplary embodiments of the present disclosure.

[0010] Figure 7 This is a schematic diagram of a turboprop engine according to an exemplary aspect of this disclosure.

[0011] Figure 8 This is a schematic diagram of a direct-drive, inline turbofan engine according to an exemplary aspect of this disclosure.

[0012] Figure 9This is a schematic diagram of a geared, duct-type turbofan engine according to an exemplary aspect of this disclosure. Detailed Implementation

[0013] 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. Detailed description uses numerals and letter designations to refer to features in the drawings. Similar or analogous reference numerals in the drawings and description have been used to refer to similar or analogous portions of this disclosure.

[0014] 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 to or advantageous to other implementations. Furthermore, unless explicitly stated otherwise, all embodiments described herein should be considered exemplary.

[0015] 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.

[0016] 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.

[0017] 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.

[0018] Unless otherwise stated herein, the terms “connection,” “fixed,” “attached to,” etc., refer to direct connection, fixation, or attachment, as well as indirect connection, fixation, or attachment via one or more intermediate components or features.

[0019] Unless the context clearly indicates otherwise, the singular forms “a,” “one,” and “the” include plural references.

[0020] As used herein, "third flow" refers to a small portion of non-primary airflow that can increase fluid energy to generate 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 dedicated nozzles or by mixing the airflow through the third flow with the primary propulsion flow or core airflow, for example, by introducing it into a common nozzle.

[0021] In some exemplary embodiments, the operating temperature of the airflow through the third flow can be below the engine's maximum compressor discharge temperature, 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 ambient temperature). In some exemplary embodiments, these operating temperatures can facilitate heat transfer to or from the airflow via the third flow and separate fluid flows. Furthermore, in some exemplary embodiments, under takeoff conditions, or more particularly, operating at rated takeoff power at sea level, 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.

[0022] Furthermore, in some exemplary embodiments, aspects of the third flow (e.g., airflow, mixing, or exhaust characteristics) and thus the aforementioned exemplary percentage contribution to 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 geometry, or fluid characteristics) to adjust or optimize overall system performance under a wide range of potential operating conditions.

[0023] The term "disk load" refers to the average pressure variation across multiple rotor blades of a rotor assembly, such as the average pressure variation across multiple fan blades of a fan.

[0024] The term "rated speed" refers to the engine's operating conditions, that is, the engine operating under the manufacturer's rated maximum full-load operating conditions.

[0025] The term "standard day operating conditions" refers to environmental conditions of sea level, 59 degrees Fahrenheit, and 60% relative humidity.

[0026] The term "propulsion efficiency" refers to the efficiency of converting the energy contained in the engine fuel into the kinetic energy of the vehicle containing the engine, thereby accelerating it or compensating for losses caused by aerodynamic drag or gravity.

[0027] The term "bypass ratio" refers to the ratio of the airflow rate through a ducted inlet that bypasses the engine to the airflow rate through the ducted inlet of the engine. For example, as discussed below... Figure 1 In the embodiments, the bypass ratio refers to the ratio of the airflow from the fan 152 through the fan shroud 170 to the airflow from the fan 152 through the engine inlet 182.

[0028] For fans with blades, the term "corrected tip speed" refers to the speed of the fan blade at its outer tip in the radial direction, corrected to correspond to standard day conditions (i.e., the rotational speed of the fan blade at its outer tip if the upstream temperature corresponds to standard day conditions). The corrected tip speed of the rotor (e.g., the fan) can be calculated by dividing the physical speed by the square root of the average rotor inlet temperature (in Rankine) at a reference temperature of 518.67 Rankine.

[0029] Generally, turbofan engines include relatively large fans to provide the desired thrust without overloading the fan blades (i.e., without increasing the disk load on the fan blades beyond a certain threshold), thus maintaining the desired overall propulsive efficiency of the turbofan engine. Traditional turbofan engine design practice involves providing a large fan, or more precisely, a large-diameter fan, to provide as much total thrust as possible to the turbofan engine. The goal of designing traditional turbofan engines is to maximize their propulsive efficiency. However, including such large fans can lead to problems, such as issues with mounting the turbofan engine on an aircraft, and relatively heavy turbofan engines (especially for ducted turbofan engines). Furthermore, as the demand for greater thrust from turbofan engines continues, the thermal demands on them also increase accordingly.

[0030] However, the inventors of this disclosure have discovered that for a three-flow turbofan engine with a main fan and a secondary fan, where the secondary fan is a ducted fan that supplies airflow to the third flow of the engine, the thrust required by the main fan can be reduced, with the secondary fan providing the difference through the third flow. This configuration can effectively maintain the desired overall propulsive efficiency of the turbofan engine, or unexpectedly, can actually increase the over-propulsive efficiency of the turbofan engine.

[0031] The following references are included in the design process of several different types of turbofan engines. Figure 1 , 2The gas turbine engines described in 7, 8, and 9 are constructed by the inventors through the following methods: designing an engine with given primary fan characteristics, secondary fan characteristics, and turbine characteristics; examining the propulsion efficiency of the designed turbofan engine; redesigning a turbofan engine with different primary fan, secondary fan, and turbine characteristics; and re-examining the propulsion efficiency of the redesigned turbofan engine. In the process of studying / evaluating various primary fan characteristics, secondary fan characteristics, and turbine characteristics that are considered to best meet mission requirements, a relationship was found 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, or more specifically, the radius ratio of the primary and secondary fans.

[0032] Furthermore, it has been found that instead of calculating the percentage of total turbofan engine thrust provided by the third flow for each design, which can be difficult to calculate accurately under various operating conditions, environmental conditions, engine designs, etc., the area-weighted specific thrust rating of the main fan and secondary fan (based on the area-weighted sum of the inlet area of ​​the third flow and the inlet area of ​​the bypass channel) can effectively and accurately represent the percentage of total turbofan engine thrust provided by the third flow.

[0033] As described in this paper, the resulting radius ratio and the relationship between third-flow thrust can be considered an indicator of the turbofan engine's ability to maintain or even improve the desired propulsion efficiency through the third flow. In addition, it also indicates improvements in the turbofan engine's packaging and weight issues, as well as its thermal management capabilities.

[0034] Now for reference Figure 1 A schematic cross-sectional view of a gas turbine engine 100 is provided according to an exemplary embodiment of this disclosure. Specifically, Figure 1 A turbofan engine with a rotor assembly having single-stage ductless rotor blades is provided. In this way, the rotor assembly may be referred to herein as a "ductless fan," or the entire engine 100 may be referred to as a "ductless turbofan engine." Furthermore, Figure 1 The engine 100 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.

[0035] 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. Typically, the axial direction A extends parallel to the longitudinal axis 112, the radial direction R extends outward and inward from the longitudinal axis 112 in a direction orthogonal to the axial direction A, and the circumferential direction extends 360° around the longitudinal axis 112. Engine 100 extends between a front end 114 and a rear end 116, for example, along the axial direction A.

[0036] Engine 100 includes a turbine 120 and a rotor assembly (also referred to as a fan section 150) located upstream therefrom. Typically, turbine 120 includes a compressor section, a combustion section, a turbine section, and an exhaust section in a series flow sequence. Specifically, as... Figure 1 As 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 turbocharger or low-pressure (“LP”) compressor 126 for pressurizing air entering turbine 120 through the core inlet 124. A high-pressure (“HP”), multi-stage axial compressor 128 receives pressurized air from the LP compressor 126 and further increases the air pressure. The pressurized air flows downstream to a combustor 130 in 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.

[0037] 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.

[0038] 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. In this respect, high-pressure turbine 128 is drivably coupled to high-pressure compressor 128. 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. In this respect, low-pressure turbine 134 is drivably coupled 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, 134, combustion products exit turbine 120 through turbine exhaust nozzle 140.

[0039] 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 generally positioned 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.

[0040] Fan section 150 includes fan 152, which in this example embodiment is the main fan. For Figure 1 In the embodiment shown, fan 152 is an open rotor or ductless fan 152. In this way, engine 100 can be referred to as an open rotor engine.

[0041] As shown in the figure, fan 152 includes an array of fan blades 154. Figure 1 (Only one is shown in the image). The fan blades 154 are, for example, rotatable 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 embodiment shown, fan 152 is connected to LP shaft 138 via reduction gearbox 155, for example, in an indirect drive or gear drive configuration.

[0042] Furthermore, the array of fan blades 154 can be arranged at equal intervals around the longitudinal axis 112. Each fan 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 along the radial direction R from the longitudinal axis 112 to the tip, and a hub radius (or inner radius) R2 along the radial direction R from the longitudinal axis 112 to the base. Furthermore, each fan blade 154 of fan 152, or more precisely, fan 152, defines a fan radius ratio RqR equal to R2 divided by R1. Since fan 152 is the main fan of engine 100, the fan radius ratio RqR of fan 152 can be referred to as the main fan radius ratio RqR. Prim.-Fan。

[0043] Furthermore, each fan blade 154 defines a central blade axis 156. In this embodiment, each fan blade 154 of the fan 152 is rotatable about its respective central blade axis 156, for example, in sync with each other. One or more actuators 158 are provided to facilitate such rotation, and thus can be used to change the pitch of the fan blades 154 about their respective central blade axes 156.

[0044] Fan section 150 further includes a fan guide vane array 160, which includes fan guide vanes 162 disposed around a longitudinal axis 112. Figure 1(Only one is shown in the image). 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. Figure 1 As shown, the fan guide vane 162 may be uncovered, or alternatively, it may be covered by, for example, 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.

[0045] Each fan guide vane 162 defines a central blade axis 164. In this embodiment, each fan guide vane 162 of the fan guide vane array 160 is rotatable about its respective central blade axis 164, for example, aligned with each other. One or more actuators 166 are provided to facilitate such rotation and can therefore be used to change the pitch of the fan guide vane 162 about its corresponding central blade axis 164. However, in other embodiments, each fan guide vane 162 may be fixed or not tiltable about its central blade axis 164. The fan guide vanes 162 are mounted to a fan shroud 170.

[0046] like Figure 1 As shown, in addition to the ductless fan 152, a ducted fan 184 is also included at the rear of fan 152, such that engine 100 includes both ducted and ductless 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 described embodiment, without passing through HP compressor 128 and combustion chamber). The ducted fan 184 is rotatable about the same axis as fan blades 154 (e.g., longitudinal axis 112). In the described embodiment, the ducted fan 184 is driven by a low-pressure turbine 134 (e.g., coupled to LP shaft 138). In the described embodiment, as described above, fan 152 may be referred to as the primary fan, and the ducted fan 184 may be referred to as the secondary fan. It should be understood that these terms "primary" and "secondary" are for convenience and do not imply any particular importance, power, etc.

[0047] The duct fan 184 includes multiple fan blades arranged in a single stage (in Figure 1 Not marked separately; see also Figure 2The fan blades 185 (marked in the middle) make the duct fan 184 a single-stage fan. The fan blades of the duct fan 184 can 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 along the radial direction R from the longitudinal axis 112 to the tip, and a hub radius (or inner radius) R4 along the radial direction R from the longitudinal axis 112 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 can be referred to as the secondary fan radius ratio RqR. Sec.-Fan .

[0048] The fan shroud 170 annularly surrounds at least a portion of the core shroud 122 and is generally positioned radially outward of at least a portion of the core shroud 122. Specifically, a downstream section of the fan shroud 170 extends above the front portion of the core shroud 122 to define a fan duct flow path, or simply to define a 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.

[0049] Incoming air enters fan duct 172 through fan duct inlet 176 and exits through fan exhaust nozzle 178 to generate propulsive 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 by a plurality of substantially radially extending, circumferentially spaced fixed struts 174. Figure 1 (Only one is shown in the diagram) They are connected together and supported. Each of the fixed struts 174 may be aerodynamically profiled to guide airflow therefrom. Other struts besides the fixed 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 (typically axially) on opposite sides (e.g., opposite radial sides) of the core shroud 122. For example, each of the fan duct 172 and the core duct 142 may extend directly from the leading edge 144 of the core shroud 122 and may extend substantially axially and partially together on opposite radial sides of the core shroud.

[0050] Engine 100 also defines or includes an inlet duct 180. 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 along the radial direction R inside fan shroud 170. Air flowing downstream along inlet duct 180 is not necessarily uniformly distributed into core duct 142 and fan duct 172 via a fan duct splitter or leading edge 144 of core shroud 122. Inlet duct 180 is wider than core duct 142 along the radial direction R. Inlet duct 180 is also wider than fan duct 172 along the radial direction R.

[0051] It is worth noting that, for the described embodiment, engine 100 includes one or more features to increase the third-flow thrust Fn. 3S The efficiency (e.g., generated by the airflow exiting through the fan exhaust nozzle 178 and through the fan duct 172, at least partially by the thrust generated by the ducted fan 184). 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 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 is rotatable about its respective central blade axis, for example, aligned with each other. In this way, the inlet guide vanes 186 can be considered variable geometry components. 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 or not tiltable about its central blade axis.

[0052] Furthermore, located downstream of the ducted 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 described embodiment, unlike the array of inlet guide vanes 186, the array of outlet guide vanes 190 is configured as fixed-pitch outlet guide vanes.

[0053] Furthermore, it should be understood that, in the described embodiment, the fan exhaust nozzle 178 of the fan duct 172 is further 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 can 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 amount of 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 can also be used.

[0054] During one or more engine operating conditions, the combination of the array of inlet guide vanes 186 upstream of the duct fan 184, the array of outlet guide vanes 190 downstream of the duct fan 184, and the fan exhaust nozzle 178 can result in a more efficient generation of third-flow thrust Fn. 3S Furthermore, by introducing variability in the geometry of the inlet guide vane 186 and the fan exhaust nozzle 178, the engine 100 can operate under 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 total engine thrust Fn) Totall A more effective third-flow thrust Fn is generated within a smaller range. 3S .

[0055] In addition, still refer to Figure 1 In an exemplary embodiment, the air passing through fan duct 172 may be relatively cold (e.g., at a lower temperature) than one or more fluids used in turbine 120. In this way, 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 utilize the air passing through fan duct 172 to cool one or more fluids from the core engine as a source of heat removal from the fluids (e.g., compressor exhaust air, oil, or fuel).

[0056] Although not shown, heat exchanger 200 may be an annular heat exchanger extending substantially 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 exhaust air, electrical components, etc.). Heat exchanger 200 uses air passing through duct 172 as a radiator, thereby increasing the temperature of the air downstream of heat exchanger 200 and exiting fan exhaust nozzle 178.

[0057] During operation of engine 100 under operating conditions, engine 100 generates total thrust Fn. Total Operating conditions can be engine 100 operating at rated speed during standard day operating conditions. Total thrust is the first-rate thrust Fn. 1S (For example, the main fan thrust generated by the airflow through the bypass channel 194 above the fan shroud 170 and the core shroud 122 produced by fan 152), third-flow thrust Fn 3S Second-rate thrust Fn 2S (For example, the sum of thrust generated by the airflow discharged through the core duct 142 via the turbine exhaust nozzle 140).

[0058] It should be understood that, as briefly noted above, under various operating conditions, environmental conditions, engine design, and other conditions, it may be difficult to accurately calculate the total thrust Fn provided by the third flow (or more precisely, the airflow exiting the fan duct 172 through the fan exhaust nozzle 178). Total The percentage. However, the inventors discovered that the specific thrust rating of the main fan (main fan specific thrust rating T) is... P ) and the specific thrust rating of the secondary fan (specific thrust rating of the secondary fan T) S Area-weighted calculus provides the total thrust Fn from the third flow. Total The precise representation of the expected percentage.

[0059] More specifically, see reference Figure 2 Provided Figure 1 A simplified close-up schematic diagram of a turbofan engine 100. As described above, the turbofan engine 100 includes a main fan, or more precisely, a fan 152 with fan blades 154, and a secondary fan, or more precisely, a duct fan 184 with fan blades 185. Airflow from fan 152 is split between bypass passage 194 and inlet duct 180 via inlet splitter 196. Airflow from duct fan 184 is split between fan duct 172 and core duct 142 via leading edge 144 (sometimes also called fan duct splitter).

[0060] Figure 2 The exemplary turbofan engine 100 shown further defines the main fan outer fan area A. P_Out The area of ​​the main fan (A) P_In Secondary fan external fan area A S_Out The area A of the secondary fan S_In .

[0061] Main fan external fan area A P_OutThis refers to the area defined by the torus representing a portion of the inlet splitter 196 located outside the fan shroud 170, which represents fan 152. Specifically, the turbofan engine 100 further defines the fan shroud splitter radius R5. The fan shroud splitter radius R5 is defined along the radial direction R from the longitudinal axis 112 to the inlet splitter 196. The main fan outer fan area refers to the area defined by the formula... Limited area.

[0062] Main fan internal fan area A P-In This refers to the area defined by the toroidal surface representing a portion of the inlet splitter 196 located inside the fan shroud 170, which represents fan 152. Specifically, the turbofan engine 100 further defines an engine inlet inner radius R6. This engine inlet inner radius R6 is defined along the radial direction R from the longitudinal axis 112 to the inner housing, which defines the engine inlet 182 directly inward from the inlet splitter 196 along the radial direction R. The main fan inner fan area refers to the area defined by the formula... Limited area.

[0063] The area of ​​the secondary fan's outer fan, A S_Out This refers to the area representing a portion of the airflow supplied from the duct fan 184 to the fan duct 172. Specifically, the leading edge 144 defines a leading edge radius R7, and the turbofan engine 100 defines an effective fan duct inlet outer radius R8 (see...). Figure 3 The leading edge radius R7 is defined along the radial direction R from the longitudinal axis 112 to the leading edge 144.

[0064] Brief Reference Figure 3 A close-up view of the area surrounding the leading edge 144 is provided. The fan duct 172 defines a lateral height 198 measured from the leading edge 144 to the fan shroud 170 in a direction perpendicular to the average flow direction 204 of the first 10% of the airflow through the fan duct 172. Notably, the angle 206 is defined by the average flow direction 204 relative to a reference line 208 extending parallel to the longitudinal axis 112. The angle 206 is referred to as θ. In some embodiments, the angle 206 may be between 5 degrees and 80 degrees, for example, between 10 degrees and 60 degrees. The effective fan duct inlet outer radius R8 is defined along the radial direction R from the longitudinal axis 112 to the point where it intersects the lateral height 198 with the fan shroud 170. The secondary fan outer fan area A S_Out It refers to the formula Limited area.

[0065] Return to reference Figure 2 The area of ​​the secondary fan, A S_InThis refers to the area defined by the toroidal surface representing a portion of the leading edge 144 of the core shroud 122, which is the inner side of the ducted fan 184. Specifically, the turbofan engine 100 further defines the inner radius R9 of the core inlet. The inner radius R9 of the core inlet is defined along the radial direction R from the longitudinal axis 112 to the inner shroud, which defines the core inlet 124 directly inward from the leading edge 144 along the radial direction R. The fan area inside the main fan refers to the area defined by the formula... Limited area.

[0066] These areas, especially the area A of the main fan and the external fan. P_Out and the area A of the secondary fan and the external fan S_Out It can be used to measure the specific thrust rating T of the main fan. P Compared to the secondary fan, the rated thrust T S Weighting is applied to accurately represent the total thrust Fn provided by the third stream. Total The expected percentage, which will be explained in more detail below. It is worth noting that, as used in this article, the main fan external fan area A... P_Out The area A of the secondary fan and the external fan S_Out The ratio is referred to in this paper as the effective bypass area or EBA.

[0067] As previously stated, the inventors, in the process of designing gas turbine engines—that is, designing gas turbine engines (e.g., pipe-mounted and non-pipe-mounted turbofan engines and turboprop engines) with various primary and secondary fan characteristics (including physical and operational characteristics)—and evaluating overall propulsion efficiency, unexpectedly discovered a significant relationship between the percentage of total gas turbine thrust provided by the third flow (as defined herein) and the relative dimensions of the primary and secondary fans of the gas turbine engine. The resulting radius ratio and relationship to the third flow thrust, as described herein, can be considered an indicator of the gas turbine engine's ability to maintain or even improve the desired propulsion efficiency through the third flow, and further, indicates improvements in gas turbine engine packaging and weight issues, as well as thermal management capabilities.

[0068] As will be understood, higher and lower third-flow thrust alters the gas turbine engine's packing capacity and radiator capacity. For example, increased thrust from the airflow through the third flow typically means more airflow (based on mass flow rate) through the third flow, which in turn means more heat capacity of that airflow. Furthermore, the inventors have found that if too little thrust is provided from the third flow, the gas turbine engine may be unnecessarily 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 the main fan can generate.

[0069] The aforementioned relationship can be a function of the bypass ratio of a gas turbine engine, which is typically limited by reasonable engine temperatures, including operating temperatures such as exhaust gas temperature (EGT). For example, given the above teachings, it will be understood that the radius of the main fan relative to the radius of the secondary fan, and the percentage of total gas turbine 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.

[0070] Many aspects of the structure determine the bypass ratio of a gas turbine engine. For example, the bypass ratio is partly a function of the corrected tip speed of the main fan relative to the corrected tip speed of the secondary fan, and the corresponding specific thrust of the main and secondary fans. The specific thrust of the main and secondary fans is, in turn, a function of the pressure ratio of the main and secondary fans, and the disk load (also known as the power load) on the main 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.

[0071] As described above, the inventors of this disclosure have discovered a relationship between the percentage of engine thrust provided by the airflow through a third flow and the radius ratio of the main fan and the secondary fan. This relationship allows the gas turbine engine to maintain or even improve the desired propulsion efficiency, while also taking into account gas turbine engine packaging and weight issues and providing desired thermal management capabilities. The relationship discovered below can determine improved engine configurations suitable for specific mission requirements, taking into account mounting, packaging and load, radiator requirements, and other factors influencing the optimal selection of engine configuration.

[0072] In addition to resulting in an improved gas turbine engine, as explained in detail above, the inventors discovered, utilizing this relationship, that many suitable or feasible gas turbine designs, combining a main fan and a secondary fan and defining a third airflow, can meet propulsion efficiency requirements and reduce packaging, weight, and radiator requirements significantly. This facilitates a faster downward selection of the design to consider when developing a gas turbine engine. This benefit provides greater insight into the requirements of a given gas turbine engine before the specific technology, integration, and system requirements are fully developed. This benefit avoids later redesign.

[0073] The inventors discovered a desirable relationship for improving gas turbine engines, expressed as follows:

[0074]

[0075] Where R1 is the tip radius of the main fan, R2 is the hub radius of the secondary 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, RqRSec-Fan It is the ratio of R4 to R3, EFP is the effective fan parameter, and T is the ratio of R4 to R3. P It is the rated specific thrust of the main fan, T S This refers to the fan's specific thrust rating, and EBA is the effective bypass area.

[0076] EBA will use the area A of the main fan and the external fan. P_Out The area A of the secondary fan and the external fan S_Out Related and expressed as a percentage. Specifically, EBA is expressed as the following ratio: A P_Out / A S_Out A higher EBA corresponds to a larger primary fan bypass ratio and a lower secondary fan bypass ratio. Conversely, a lower EBA corresponds to a smaller primary fan bypass ratio and a higher secondary fan bypass ratio.

[0077] EFP is a function of the corrected tip speed of the main fan, the corrected tip speed of the secondary fan, the disk load of the main fan, and the disk load of the secondary fan. By taking into account the corrected tip speeds of the main and secondary fans, EFP accounts for factors such as specific engine configurations (e.g., geared, direct drive, etc.), which for a turbofan engine with desired propulsive efficiency might affect the tip radius ratio (R1 vs. R3) and the percentage thrust (%Fn) through the third flow. 3S The relationship between these factors has some influence. For turbofan engines, the factors affecting EFP are described in more detail above in relation to the tip radius ratio (R1 to R3) and the percentage thrust through the third flow (%Fn). 3S ) relationship.

[0078] Main fan specific thrust rating T P Compared to the secondary fan, the rated thrust T S This reflects the main fan speed at rated speed during standard day operating conditions. Figure 1 and Figure 2 Fan 152 and secondary fan (in the embodiment) Figure 1 and Figure 2 The corrected tip speed of the duct fan 184 in the embodiment. (See below...) Figure 4 The main fan specific thrust rating T is provided and can be incorporated into expression (1). P Compared to the secondary fan, the rated thrust T S Example values.

[0079] In particular, Figure 4 An exemplary main fan specific thrust rating T along the Y-axis 302 is provided. P And the exemplary secondary fan thrust rating T along the X-axis 304 S The curve 300. It is worth noting that curve 300 is plotted from the following reference... Figures 5A to 5FThe example main fan specific thrust rating T of at least some of the example engines described P Compared to the secondary fan, the rated thrust T S .from Figure 4 The example shown in the diagram illustrates that, for a given engine, the secondary fan has a specific thrust rating T. S It can be considered to be typically higher than the main fan's specific thrust rating T. P This is because the secondary fan is downstream of the main fan and is built upon the flow from the main fan.

[0080] for Figure 4 In one embodiment, the main fan specific thrust rating T P It ranges from 0.08 to 0.59, and is 0.1 to 0.5 in at least some exemplary aspects. Similarly, the secondary fan's specific thrust rating T S It ranges from 0.21 to 0.6, and in at least some exemplary respects from 0.26 to 0.5. These wider ranges correspond to Figure 4 The area in the middle is 306, while the narrower range corresponds to Figure 4 The area is 308.

[0081] In particular, Figure 4 The curve 300 in the figure determines the rated specific thrust T of the main fan. P Compared to the secondary fan, the rated thrust T S The value has four sub-regions. Specifically, Figure 4 The curve 300 identifies the first sub-region 310, the second sub-region 312, the third sub-region 314, and the fourth sub-region 316.

[0082] The first sub-region 310 includes two example engines drawn—Example 318 and Example 320. Example 318 corresponds to a turboprop engine (see reference below for example). Figure 7 The exemplary turboprop engine 526 described, example 320 corresponds to an open rotor engine, such as Figure 1 and Figure 2 An exemplary engine 100. Each of these engines is a geared engine (i.e., including a gearbox between the drive shaft and the fan shaft; see, for example, see...). Figure 1 The gearbox 155), and further, the engine is a ductless engine (i.e., excluding the outer nacelle surrounding the main fan). The main fan specific thrust rating T of the first sub-region 310. P Within the range of 0.1 to 0.23, the secondary fan specific thrust rating T in the first sub-region S In the range of 0.35 to 0.45.

[0083] It should be understood that, with Figure 4Compared to other examples drawn in the diagram, the main fan of the first sub-region 310 has a specific thrust rating T. P The relatively low thrust is at least because the engine is a non-ducting engine, allowing for a larger fan diameter. For a larger diameter fan, the lower main fan thrust is relative to the thrust rating T. P This is acceptable, while still providing the desired thrust to the engine (since the engine may rotate more slowly, resulting in less thrust per unit area). In this way, it will be appreciated that the main fans of Examples 318 and 320 can be relatively efficient because they typically limit a relatively small pressure ratio during operation.

[0084] Furthermore, as previously mentioned, the engines illustrated as examples 318 and 320 are geared engines and can be characterized by relatively high gear ratios (e.g., compared to at least some of the examples in the second sub-region 312, as described below). Nevertheless, the subfan specific thrust rating T... S It can still be smaller (e.g., compared to at least some examples in the second sub-region 312). In particular, due to the secondary fan's thrust rating T... S Based on the main fan's specific thrust rating T P Based on (and compared with the example in the second sub-region 312), the main fan specific thrust rating T P Smaller), therefore the secondary fan has a higher thrust rating than the T S It's also smaller.

[0085] The second sub-region 312 includes two example engines drawn—Example 322 and Example 324. Examples 322 and 324 correspond to a geared and a ducted turbofan engine, respectively (see reference below for example). Figure 9 The exemplary geared, ducted turbofan engine 544 described has a gearbox 546. More specifically, each of these engines is a geared engine and further is a ducted engine (i.e., including an outer nacelle surrounding the main fan). The main fan specific thrust rating T of the second sub-region 312. P Within the range of 0.23 to 0.35, and the secondary fan specific thrust rating T of the second sub-region 312. S In the range of 0.4 to 0.5.

[0086] It should be understood that the main fan thrust rating T of the engine in the second sub-region 312 PThe thrust is higher than that of the engine in the first sub-region 310. This is often a result of the engine being a tubular engine, as the fans may be smaller and require them to rotate faster to generate the desired thrust. Furthermore, the engines drawn as examples 322, 324 may have a lower gear ratio than the engines drawn as examples 318, 320. While this typically results in a lower sub-fan thrust rating T... S The thrust is reduced, but because the engines drawn as examples 322 and 324 have a higher main fan specific thrust rating T P This could drive the secondary fan to a thrust rating of T. S higher.

[0087] Still referencing Figure 4 The third sub-region 314 and the fourth sub-region 316 include example engines drawn entirely as direct-drive engines (i.e., excluding the reduction gearbox between the drive shaft and the fan shaft). Typically, for each of these engines, since the secondary fan rotates at the same speed as the primary fan, for a given primary fan specific thrust rating T... P The specific thrust rating of the secondary fan T S It will be lower.

[0088] Specifically, first refer to the third sub-region 314, which includes examples 326 and 328. The engines drawn as examples 326 and 328 are direct-drive tubular turbofan engines (see, for example, reference below). Figure 8 The exemplary direct-drive turbofan engine 538 described. In particular, these engines can be configured with a relatively high bypass ratio for subsonic flight operation. The main fan specific thrust rating T of the third sub-region 314. P Within the range of 0.23 to 0.38, and the secondary fan specific thrust rating T in the third sub-region 314 S It is in the range of 0.26 to 0.4.

[0089] Conversely, referring now to sub-region 316, which includes examples 330 and 332, the engines drawn as examples 330 and 332 are also direct-drive tubular turbofan engines, but can be limited to relatively low bypass ratios, for example, for high-speed operation favoring smaller fan diameters. In this way, the engines drawn as examples 330 and 332 can have smaller, faster-rotating fans compared to those drawn as examples 326 and 328, resulting in higher efficiency and lower drag. The main fan thrust rating T of sub-region 316 is... P Within the range of 0.35 to 0.5, and the secondary fan specific thrust rating T of the fourth sub-region 316. S It is in the range of 0.37 to 0.5.

[0090] Furthermore, the values ​​of various other parameters of the engine influence characteristics defined by expression (1) are shown in Table 3 below:

[0091]

[0092] Now for reference Figures 5A to 5F and Figures 6A to 6C This illustrates an exemplary gas turbine engine according to one or more exemplary embodiments of the present disclosure, showing the relationship between various parameters of expression (1). Specifically, Figures 5A to 5F Provided including with Figures 5A to 5C The table shows the corresponding values ​​for several drawn gas turbine engines. Figures 6A to 6C This is a graph of a gas turbine engine according to one or more exemplary embodiments of the present disclosure, showing the relationship between the tip radius ratio (R1 to R3; Y-axis) and EBA, effective bypass area (X-axis). Figure 6A The subranges based on EFP values ​​are highlighted. Figure 6B The sub-range of pipeless engines is highlighted. Figure 6C The subrange of the pipeline engine is highlighted. It can be understood, at least from expression (1) and the discussion herein, that EBA multiplied by the main fan specific thrust rating T P Compared to the secondary fan, the rated thrust T S The ratio (see) Figure 4 This can typically be related to the percentage of thrust provided by the engine's third flow. In this way, it will be understood that... Figure 6B The subranges (e.g., 406, 408) can be with Figure 4 The sub-range combinations in sub-regions 310 and 312, and similarly, Figure 6C Subregions (e.g., 410, 412) can be compared with those in... Figure 4 The sub-range combinations in sub-regions 314 and 316. Furthermore, Figure 4 The area of ​​the sub-range is 308. Figure 6A The subscopes in this document, or both, can be combined with any other subscope in this document (e.g., Figure 6B (Or a combination of 6C).

[0093] Special Reference Figure 6A A first range 402 and a second range 404 are provided. The first range 402 may correspond to an EFP greater than or equal to 0.15 and less than or equal to 33. This range corresponds to the main specific thrust rating T listed in Table 1. P and the rated thrust T S Combined with EBA's wider range, it can be used with percentage thrust %Fn between 2% and 50% through the third stream. 3SThis is related. This can result in the engine having the desired propulsive efficiency.

[0094] The second range 404 can correspond to an EFP greater than or equal to 2 and less than or equal to 20. This range corresponds to the main specific thrust rating T listed in Table 1. P and the rated thrust T S Combined with the wider range of EBA, it can be used with the narrower range of percentage thrust %Fn through the third stream. 3S (For example, between 5% and about 20%). This can result in the engine having a better propulsion efficiency.

[0095] Now for reference Figure 6B The third scope 406 is provided together with the fourth scope 408, whereby the fourth scope 408 is a subscope of the third scope 406. The third scope 406 and the fourth scope 408 relate to pipeless engines, such as turboprop engines (see...). Figure 7 ) or open rotary engine (see Figure 1 and 2 The third range 406 can correspond to an EBA between 0.8% and 6.5% and an R1 to R3 value between 3.4 and 6.5. The fourth range 408 can correspond to an EBA between 0.9% and 2.9% and an R1 to R3 value between 3.5 and 6.4.

[0096] As mentioned above, ductless engines typically have larger fans, and therefore may have larger R1 to R3 values. Similarly, due to the larger fan, EBA (Electronic Bypass Area) is generally likely to be lower because of the larger bypass area of ​​the main fan. It is worth noting that turboprop engines can typically have higher R1 to R3 values ​​than open rotor turbofan engines. For example, the R1 to R3 value for turboprop engines can be between 4.5 and 6.5, while the R1 to R3 value for open rotor engines can be between 3.5 and 5.

[0097] Now for reference Figure 6C Fifth scope 410 is provided together with sixth scope 412, which is a subscope of fifth scope 410. Fifth scope 410 and sixth scope 412 relate to tubular engines, such as geared tubular turbofan engines (see...). Figure 9 ) or directly driven tubular turbofan engines (see Figure 8 The fifth range 410 can correspond to an EBA between 1.25% and 8.9% and an R1 to R3 value between 1.7 and 4. The sixth range 412 can correspond to an EBA between 1.35% and 3.8% and an R1 to R3 value between 1.7 and 3.9.

[0098] As mentioned above, inline turbofan engines can typically have smaller fans, and therefore may have smaller R1 to R3 values. Due to the smaller fan, EBA (Efficient Bypass Area) is often higher because the main fan's bypass area is not as large as the potential bypass area of ​​the secondary fan. It is worth noting that geared turbofan engines can typically have higher R1 to R3 values ​​than direct-drive turbofan engines because the gearbox allows the main fan to rotate slower than the secondary fan, thus allowing for a larger fan without significantly increasing the main fan's pressure ratio (and therefore without significantly reducing efficiency). For example, the R1 to R3 value of a geared turbofan engine can be between 2.3 and 4, while the R1 to R3 value of a direct-drive turbofan engine can be between 1.7 and 2.8.

[0099] It should be understood that, although the above discussion usually refers to the references above. Figure 1 and 2 The open rotary engine 10 is described, but in various embodiments of this disclosure, the relationships outlined above with respect to, for example, expression (1), can be applied to any other suitable engine architecture. For example, reference will now be made to... Figures 7 to 9 Each figure schematically depicts the engine architecture associated with this disclosure.

[0100] Figures 7 to 9 Each gas turbine engine typically includes a rotor 502 rotatable about a rotor axis 504 and a turbine 506 rotatable about a longitudinal axis 508. The rotor 502 corresponds to the "main fan" described herein. The turbine 506 is at least partially surrounded by a core shroud 510 and includes a compressor section 512, a combustion section 514, and a turbine section 516 in a series flow sequence. In addition to the rotor 502, Figures 7 to 9 Each gas turbine engine also includes a ducted intermediate fan or secondary fan 518. Each gas turbine engine includes a fan shroud 520 surrounding the secondary fan 518.

[0101] Still referencing Figures 7 to 9 Each gas turbine engine further defines a bypass passage 522 downstream of the respective rotor 502 and above the respective fan shroud 520 and core shroud 510, and further defines a third stream 524 extending from a location downstream of the respective secondary fan 518 to the respective bypass passage 522 (at least in the illustrated embodiment; in other embodiments, the third stream 522 may alternatively extend to a location downstream of the bypass passage 522).

[0102] Special Reference Figure 7The exemplary gas turbine engine shown is configured as a turboprop engine 526. In this way, the rotor 502 (or main fan) is configured as a propeller, defining a relatively large diameter. Furthermore, the turboprop engine 526 includes an engine shaft 528 driven by a turbine 506, a fan shaft 530 rotatable with the rotor 502, and a gearbox 532 mechanically connecting the engine shaft 528 and the fan shaft 530. The gearbox 532 is an offset gearbox such that the rotor axis 504 is radially offset from the longitudinal axis 508 of the turboprop engine 526.

[0103] It is worth noting that in other embodiments of this disclosure, the turboprop engine may be equipped with a counter-current combustor.

[0104] refer to Figure 8 and Figure 9 All gas turbine engines are configured as turbofan engines, and more specifically, as ducted turbofan engines. In this way, each gas turbine engine includes an outer nacelle 534 surrounding a rotor 502, and each rotor 502 (or main fan) is therefore configured as a ducted fan. Furthermore, each gas turbine engine includes an outlet guide vane 536 extending from a fan shroud 520, a core shroud 510, or both, through a bypass passage 522 to the outer nacelle 534.

[0105] More specifically, Figure 8 The gas turbine engine is configured as a direct-drive inline turbofan engine 538. Specifically, the direct-drive inline turbofan engine 538 includes an engine shaft 540 driven by a turbine section 516 and a fan shaft 542 that rotates together with a rotor 502. The fan shaft 542 is configured to rotate directly with the engine shaft 540 (i.e., at the same speed as the engine shaft 540).

[0106] on the contrary, Figure 9 The gas turbine engine is configured as a geared, in-line, turbofan engine 544. Specifically, the geared, in-line turbofan engine 544 includes an engine shaft 540 driven by a turbine section 516 and a fan shaft 542 rotatable with a rotor 502. However, the exemplary geared, in-line turbofan engine 544 also includes a gearbox 546 mechanically connecting the engine shaft 540 to the fan shaft 542. The gearbox 546 allows the rotor 502 to rotate at a slower speed than the engine shaft 540, and therefore at a slower speed than the secondary fan 518.

[0107] It is worth noting that, Figure 9The exemplary geared, inline turbofan engine 544 also includes a pitch mechanism 548 that operates in conjunction with the rotor 502 to change the pitch of the rotor blades of the rotor 502. This allows for increased efficiency of the gas turbine engine.

[0108] As will be understood from the description herein, various embodiments of gas turbine engines are provided. Some of these embodiments may be ductless single-rotor gas turbine engines (see...). Figure 1 and 2 ), turboprop engine (see Figure 7 ) or ducted turbofan engine (see Figure 8 and 9 Another example of a ducted turbofan engine can be found in U.S. Patent Application No. 16 / 811,368, filed March 6, 2020 (published as U.S. Patent Application Publication No. 2021 / 0108597) (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, which is 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 figures.

[0109] For example, in some embodiments of this disclosure, the engine may include a heat exchanger located in an annular duct, such as in a third flow. The heat exchanger may extend substantially continuously in the circumferential direction of the gas turbine engine (e.g., at least 300 degrees, such as at least 330 degrees).

[0110] In one or more of these embodiments, during cruise operation mode, at cruise altitude, the threshold power or disk load of the fan (e.g., a ductless single-rotor or main forward-facing fan) can be 25 horsepower per square foot (hp / ft). 2 (or a wider range). In a specific embodiment of the engine, the structure and method provided herein produce 80 hp / ft at cruise altitude during cruise operating mode. 2 and 160hp / ft 2 The power load can be between or higher, depending on whether the engine is an open rotary engine or a tubular engine.

[0111] In various embodiments, the engine of this disclosure is applied to a vehicle with a cruising altitude up to approximately 65,000 feet. In some embodiments, the cruising altitude is between approximately 28,000 feet and approximately 45,000 feet. In still other embodiments, the cruising altitude is expressed as a flight altitude based on sea-level standard atmospheric pressure, 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 range of pressure-defined cruising altitudes can be adjusted based on different reference sea-level pressures and / or sea-level temperatures.

[0112] Therefore, it should be understood that an engine of this configuration can be configured to produce at least 25,000 pounds and less than 80,000 pounds of thrust during operation at rated speed, for example, between 25,000 pounds and 50,000 pounds of thrust during operation at rated speed; for example, between 25,000 pounds and 40,000 pounds of thrust during operation at rated speed. Alternatively, in other exemplary aspects, the engine of this disclosure can be configured to produce much less power during operation at rated speed, for example, at least 2,000 pounds of thrust.

[0113] In various exemplary embodiments, the fan (or rotor) 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 main fan (e.g., to twelve 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. Alternatively, in some suitable embodiments, the fan may include only four (4) blades, such as the fan of a turboprop engine.

[0114] Furthermore, in some exemplary embodiments, the rotor assembly may be defined with a rotor diameter (or fan diameter) of at least 10 feet, for example at least 11 feet, for example at least 12 feet, for example at least 13 feet, for example at least 15 feet, for example at least 17 feet, for example up to 28 feet, for example up to 26 feet, for example up to 24 feet, for example up to 18 feet.

[0115] In various embodiments, it should be understood that the engine includes a ratio of the number of impeller blades to the number of blades, which 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 between 1:2 and 5:2. This ratio can be adjusted based on a variety of factors, including impeller blade size, to ensure that the desired amount of turbulence is removed from the airflow from the main fan.

[0116] Furthermore, in some exemplary embodiments, when the engine includes a third flow and an intermediate fan (a ducted fan behind the main forward fan), the ratio R1 / R2 may be between 1 and 10, or between 2 and 7, or at least 3.3, at least 3.5, at least 4 and less than or equal to 7, where R1 is the radius of the main fan and R2 is the radius of the intermediate fan.

[0117] It should be understood that various embodiments of the engine, such as the single tubeless rotor engine depicted and described herein, can allow operation at normal subsonic aircraft cruise altitudes of Mach 0.5 or higher. In some embodiments, the engine allows normal aircraft operation at cruise altitudes between Mach 0.55 and Mach 0.85. In still other 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 of 750 feet per second (fps) or less. In other embodiments, rotor blade tip velocities under cruise flight conditions can be 650 to 900 fps, or 700 to 800 fps. Alternatively, in some suitable embodiments, the engine allows normal aircraft operation at at least Mach 0.3, for example, using a turboprop engine.

[0118] As measured on the fan blades of the main fan under cruise flight conditions, the fan pressure ratio (FPR) of the main fan of the fan assembly can be from 1.04 to 2.20, or from 1.05 to 1.2 in some embodiments, or less than 1.08 in some embodiments.

[0119] To enable the gas turbine engine to operate in conjunction with a fan having the aforementioned characteristics, thereby defining 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 between 3.0 and 4.0, between 3.2 and 3.5, or between 3.5 and 4.5. In some embodiments, the gear ratio of the input speed to the output speed is greater than 4.1. For example, in a particular embodiment, 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.

[0120] The turbine, compressor, and / or turbine of a gas turbine engine can include a variety of numbers of stages. As disclosed herein, the number of stages refers to 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 4 to 15 stages, a high-pressure turbine can include 1 to 2 stages, and / or a low-pressure turbine (LPT) can include 1 to 7 stages. In particular, an LPT can have 4 stages, or between 4 and 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 4 or 4 to 7 stages of LPT. 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.

[0121] The core engine is typically enclosed in a housing that defines half the core diameter (Dcore), which can be considered as the maximum extent from the centerline axis (a reference to R). In some embodiments, the engine includes 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 used for a single, pipeless rotary engine.

[0122] The reduction in mounting drag can further improve efficiency, such as increasing specific fuel consumption. Alternatively, the reduced mounting drag can provide the engine and aircraft with the aforementioned Mach numbers at cruise altitude. Certain other embodiments may offer the benefit of reduced noise from the interaction between blade assemblies and wheel blade assemblies, and / or reduced total engine noise due to structures located in the engine's annular duct.

[0123] 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, or other structures of the core engine. However, as previously stated, within the scope of one or more structures provided herein that may be known in the art, it should be understood that this disclosure may include combinations of structures not previously known to be combined, at least in part based on conflict of interest, desired operating modes, or other forms of teaching in the art.

[0124] Although described in the above embodiments as an uncovered or open rotor engine, it should be understood that aspects of the disclosure provided herein are applicable to covered or tubular engines, partially tubular 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 uncovered or open rotor engines, such as, but not limited to, factors related to gear ratios, fan diameter, fan speed, engine length (L), maximum diameter of the engine's core (Dcore), engine L / Dcore, desired cruise altitude, and / or desired operating cruise speed, or combinations thereof.

[0125] 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 performing any combination. The patentable 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 be 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.

[0126] The following items provide further details:

[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 series flow sequence; the turbine including an inlet splitter and a fan duct splitter, the inlet splitter partially defining an inlet to a working gas flow path, and the fan duct splitter partially defining an inlet to a fan duct flow path; and a main fan driven by the turbine, defining a main fan tip radius R1, a main fan hub radius R2, and a main fan specific thrust rating T. PA secondary fan, located downstream of the main fan and driven by the turbine, wherein at least a portion of the airflow from the main fan is configured to bypass the secondary fan, the secondary fan defining a secondary fan tip radius R3, a secondary fan hub radius R4, and a secondary fan specific thrust rating T. S The gas turbine engine defines an effective bypass area, and the ratio of R1 to R3 is equal to...

[0128] Furthermore, the EFP values ​​are between 0.15 and 33, and RqR... Prim.-Fan It is the ratio of R2 to R1, where RqR Sec.-Fan It is the ratio of R4 to R3, wherein the main fan has a specific thrust rating T. P Between 0.08 and 0.59, the specific thrust rating T of the secondary fan is... S Between 0.21 and 0.6, and wherein the effective bypass area is between 0.2% and 15%.

[0129] In any of the preceding clauses, the ratio of R1 to R3 in the gas turbine engine is between 1.35 and 10.

[0130] Gas turbine engines according to any of the foregoing clauses, wherein the EFP is between 2 and 20.

[0131] The gas turbine engine according to any of the foregoing clauses, wherein the effective bypass area is between 2% and 10%.

[0132] According to any of the foregoing clauses, the gas turbine engine, wherein RqR Prim.-Fan Between 0.2 and 0.4.

[0133] According to any of the foregoing clauses, the gas turbine engine, wherein RqR Prim.-Fan Between 0.25 and 0.35.

[0134] According to any of the foregoing clauses, the gas turbine engine, wherein RqR Sec.-Fan Between 0.2 and 0.9.

[0135] According to any of the foregoing clauses, the gas turbine engine, wherein RqR Sec.-Fan Between 0.35 and 0.7.

[0136] The gas turbine engine according to any of the foregoing clauses, wherein the gas turbine engine is a pipeless gas turbine engine, wherein the effective bypass area is between 0.8% and 6.5%, wherein R1 to R3 is between 3.4 and 6.5, wherein the main fan specific thrust rating T PBetween 0.1 and 0.35, wherein the secondary fan specific thrust rating T S Between 0.35 and 0.5.

[0137] According to any of the foregoing clauses, the effective bypass area is between 0.9% and 2.9%, and R1 is between 3.5 and 6.4 compared to R3.

[0138] According to any of the foregoing clauses, the gas turbine engine is a tubular gas turbine engine, wherein the effective bypass area is between 1.25% and 8.9%, wherein R1 to R3 is between 1.7 and 4, and wherein the main fan specific thrust rating T P Between 0.23 and 0.5, wherein the secondary fan specific thrust rating T S Between 0.28 and 0.5.

[0139] According to any of the foregoing clauses, the effective bypass area is between 1.35% and 8.9%, and the ratio of R1 to R3 is between 1.7 and 3.9.

[0140] According to any of the foregoing clauses, the gas turbine engine, wherein the EFP is between 2 and 4.5, wherein the main fan limits the main fan correction tip speed during operation of the gas turbine engine at rated speed under standard day operating conditions, wherein the secondary fan limits the secondary fan correction tip speed during operation of the gas turbine engine at rated speed under standard day operating conditions, wherein the main fan correction tip speed is between 600 ft / s and 1800 ft / s, and wherein the secondary fan correction tip speed is between 1200 ft / s and 1800 ft / s.

[0141] According to any of the preceding 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.

[0142] According to any of the foregoing clauses, the gas turbine engine defines a bypass airflow passage, wherein the main fan is configured to provide a first portion of the main fan airflow to the bypass airflow passage and a second portion of the main fan airflow to the secondary fan, and wherein the secondary fan is configured to provide a first portion of the secondary fan airflow to the fan duct flow path as fan duct airflow and a second portion of the secondary fan airflow to the working gas flow path.

[0143] The gas turbine engine according to any of the foregoing clauses further includes: a heat exchanger positioned in thermal communication with the fan duct flow path.

[0144] The gas turbine engine according to any of the foregoing clauses further includes: an array of inlet guide vanes positioned immediately upstream of the secondary fan.

[0145] The gas turbine engine according to any of the foregoing clauses further includes: an array of outlet guide vanes positioned immediately downstream of the secondary fan and upstream of the fan duct.

[0146] The gas turbine engine according to any of the foregoing clauses further includes: a variable geometry exhaust nozzle located at the outlet of the fan duct.

[0147] The gas turbine engine according to any of the foregoing clauses further includes: a fan shroud surrounding the secondary fan located downstream of the main fan, the fan shroud partially defining an engine inlet located downstream of the main fan; wherein the turbine further includes a core shroud surrounding at least partially 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.

Claims

1. A gas turbine engine, wherein the gas turbine engine defines a centerline, a radial direction, 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 series flow sequence, the turbine including an inlet splitter and a fan duct splitter, the inlet splitter partially defining an inlet to a working gas flow path and the fan duct splitter partially defining an inlet to a fan duct flow path. The main fan, driven by the turbine, defines the main fan tip radius R1, the main fan hub radius R2, and the main fan specific thrust rating T. P ; A secondary fan, located downstream of the main fan and driven by the turbine, has at least a portion of the airflow from the main fan configured to bypass it. The secondary fan defines a secondary fan tip radius R3, a secondary fan hub radius R4, and a secondary fan specific thrust rating T. S ;as well as A fan shroud surrounds the secondary fan located downstream of the main fan, and the fan shroud partially defines an engine inlet located downstream of the main fan. The gas turbine engine defines an effective bypass area, and the ratio of R1 to R3 is equal to... ,and Where EFP is the effective fan parameter, and is a function of the primary fan corrected tip speed, the secondary fan corrected tip speed, the disk load of the primary fan, and the disk load of the secondary fan. EFP is between 0.15 and 33. EBA is the effective bypass area, and is the ratio of the primary fan's external fan area to the secondary fan's external fan area, wherein the primary fan's external fan area refers to the area calculated by the formula... The defined area, R5 is the radius of the fan shroud splitter, and the area of ​​the secondary fan's outer fan refers to the area defined by the formula... The defined area is defined as follows: R7 is the radius of the fan duct splitter, and R8 is the outer radius of the effective fan duct inlet. The outer radius of the effective fan duct inlet is defined along the radial direction from the centerline to the position where the lateral height intersects with the fan shroud. The lateral height is measured from the fan duct splitter to the fan shroud in a direction perpendicular to the average flow direction of the first 10% of the airflow through the fan duct flow path. It is the angle defined by the average flow direction relative to a reference line extending parallel to the centerline, where RqR Prim.-Fan It is the ratio of R2 to R1, where RqR Sec.-Fan It is the ratio of R4 to R3, wherein the main fan has a specific thrust rating T. P Between 0.08 and 0.59, the specific thrust rating T of the secondary fan is... S Between 0.21 and 0.6, and wherein the effective bypass area is between 0.2% and 15%.

2. The gas turbine engine according to claim 1, characterized in that, The ratio of R1 to R3 is between 1.35 and 10.

3. The gas turbine engine according to claim 1, characterized in that, EFP is between 2 and 20.

4. The gas turbine engine according to claim 1, characterized in that, The effective bypass area is between 2% and 10%.

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

4.

6. The gas turbine engine according to claim 5, characterized in that, Where RqR Prim.-Fan Between 0.25 and 0.

35.

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

9.

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

7.

9. The gas turbine engine according to claim 1, characterized in that, The gas turbine engine is a ductless gas turbine engine, the effective bypass area is between 0.8% and 6.5%, the R1 to R3 ratio is between 3.4 and 6.5, and the main fan specific thrust rating T P Between 0.1 and 0.35, wherein the secondary fan specific thrust rating T S Between 0.35 and 0.

5.

10. The gas turbine engine according to claim 9, characterized in that, The effective bypass area is between 0.9% and 2.9%, and R1 is between 3.5 and 6.4 relative to R3.

11. The gas turbine engine according to claim 1, characterized in that, The gas turbine engine is a ducted gas turbine engine, the effective bypass area is between 1.25% and 8.9%, the R1 to R3 ratio is between 1.7 and 4, and the main fan specific thrust rating T P Between 0.23 and 0.5, wherein the secondary fan specific thrust rating T S Between 0.28 and 0.

5.

12. The gas turbine engine according to claim 11, characterized in that, The effective bypass area is between 1.35% and 8.9%, and R1 is between 1.7 and 3.9 compared to R3.

13. The gas turbine engine according to claim 1, characterized in that, The EFP is between 2 and 4.5, wherein the primary fan limits the primary fan correction tip speed during operation of the gas turbine engine at rated speed under standard day operating conditions, wherein the secondary fan limits the secondary fan correction tip speed during operation of the gas turbine engine at rated speed under standard day operating conditions, wherein the primary fan correction tip speed is between 600 ft / s and 1800 ft / s, and wherein the secondary fan correction tip speed is between 1200 ft / s and 1800 ft / s.

14. The gas turbine engine according to claim 1, characterized in that, 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.

15. The gas turbine engine according to claim 1, characterized in that, The gas turbine engine defines a bypass airflow passage, wherein the main fan is configured to provide a first portion of the main fan airflow to the bypass airflow passage and a second portion of the main fan airflow to the secondary fan, and wherein the secondary fan is configured to provide a first portion of the secondary fan airflow to the fan duct flow path as the fan duct airflow 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 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 adjacent to the downstream of the secondary fan and the upstream of the fan duct flow path.

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 flow path.

20. The gas turbine engine according to claim 1, characterized in that, Further includes: 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.

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