Inlet to the ductless propulsion system

Abstract

A propulsion system is provided, comprising a ductless rotary fan defining a fan axis; and a turbine disposed downstream of the ductless rotary fan, wherein the turbine defines a working gas flow path flowing therethrough; wherein the propulsion system defines a third flow path and an inlet passage, the inlet passage having an inlet offset from the fan axis, wherein the inlet passage is configured to provide an inlet airflow to the working gas flow path, and wherein the third flow path bypasses at least a portion of the turbine.

Description

Technical Field

[0001] The present disclosure generally relates to an unducted fan aircraft engine having an offset inlet. Background Art

[0002] A gas turbine engine generally includes a turbine and a rotor assembly. A gas turbine engine, such as a turbofan engine, can be used for aircraft propulsion. In the case of a turbofan engine, the rotor assembly can be configured as a fan assembly.

[0003] Existing turbofan engines typically include large diameter fans and large gearboxes, which can lead to problems with underhood packaging and thermal management. Other problems with existing turbofan engines typically include poor deformation limits and poor debris extraction due to size and inlet design limitations. Improving turbofan engines to address these issues would be welcome in the art. BRIEF DESCRIPTION OF THE DRAWINGS

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

[0005] Figure 1 is a cross-sectional view of a first propulsion system configured as an unducted turbofan engine with an offset inlet according to an exemplary aspect of the present disclosure.

[0006] Figure 2 According to an exemplary aspect of the present disclosure Figure 1 An enlarged isolated view of the portion marked 2 - 2 and showing the height of the third inlet and the blade span of the middle fan rotor blade.

[0007] Figure 3 is a cross-sectional view of a second propulsion system configured as a turboprop engine with an offset inlet according to an exemplary aspect of the present disclosure.

[0008] Figure 4 is a front view facing rearward of a propulsion system according to an exemplary aspect of the present disclosure and illustrating a first inlet configuration.

[0009] Figure 5 is a front view facing rearward of a propulsion system according to an exemplary aspect of the present disclosure and illustrating a second inlet configuration.

[0010] Figure 6 is a graph illustrating the curvature of an inlet surface offset from the inlet of a propulsion system according to an exemplary aspect of the present disclosure.

[0011] Figure 7 is a front view facing rearward at the inlet of a third flow of a propulsion system according to an exemplary aspect of the present disclosure.

[0012] Figure 8 is a cross-sectional view of a turbine having a first counter-flow configuration according to an exemplary aspect of the present disclosure.

[0013] Fig. 9 is a cross-sectional view of a turbine having a second counter-flow configuration according to an exemplary aspect of the present disclosure.

[0014] Fig.10 is a cross-sectional view of a turbine having a third counter-flow configuration according to an exemplary aspect of the present disclosure.

[0015] Fig.11 is a cross-sectional view of a propulsion system having a turbine including a fourth reverse flow configuration according to an exemplary aspect of the present disclosure.

[0016] Repeat use of reference characters in the present specification and drawings is intended to represent the same or analogous features or elements of the present disclosure. DETAILED DESCRIPTION

[0017] Reference will now be made in detail to the present embodiments of the present disclosure, one or more examples of which are illustrated in the accompanying drawings. The detailed description uses numerical and letter references to refer to features in the drawings. Similar or similar reference numbers in the drawings and description have been used to refer to similar or similar parts of the present disclosure.

[0018] The word "exemplary" is used herein to mean "serving as an example, instance, or illustration." Any implementation described herein as "exemplary" is not necessarily to be construed as preferred or advantageous over other implementations. Furthermore, unless expressly stated otherwise, all embodiments described herein are to be considered exemplary.

[0019] As used herein, the terms “first,” “second,” and “third” may be used interchangeably to distinguish one component from another, and are not intended to indicate the position or importance of each component.

[0020] The terms "front" and "rear" refer to relative positions within a gas turbine engine or vehicle and refer to the normal operating attitude of the gas turbine engine or vehicle. For example, for a gas turbine engine, front refers to a position closer to the engine inlet, while rear refers to a position closer to the engine nozzle or exhaust.

[0021] The terms "upstream" and "downstream" refer to the relative directions of the flow of fluid in a fluid path. For example, "upstream" refers to the direction from which the fluid flows, and "downstream" refers to the direction to which the fluid flows.

[0022] Unless otherwise specified herein, the terms "coupled," "fixed," "attached to," and the like refer to both direct coupling, fixing, or attachment as well as indirect coupling, fixing, or attachment through one or more intermediate components or features.

[0023] The singular forms "a," "an," and "the" include plural references unless the context clearly dictates otherwise.

[0024] The term "distal" refers to being located farthest from the point of origin or attachment, while the term "proximal" refers to being located closest to the point of origin or attachment.

[0025] Approximate language used in the entire specification and claims herein is applied to modify any quantitative representation that can allow variation without causing a change in the basic function associated therewith. Therefore, the value modified by one or more terms (such as "about", "approximately" and "substantially") is not limited to the specified exact value. In at least some cases, approximate language can correspond to the accuracy of the instrument used to measure the value, or the accuracy of the method or machine used to build or manufacture parts and / or systems. For example, approximate language can refer to within a margin of 1, 2, 4, 10, 15 or 20%. These approximate margins can be applied to a single value, any endpoint or two endpoints that define a numerical range, and / or the margin of the range between the endpoints.

[0026] Here and throughout the specification and claims, range limitations are combined and interchanged, and unless context or language dictates otherwise, such ranges are identified and include all sub-ranges contained therein. For example, all ranges disclosed herein are inclusive of the endpoints, and the endpoints are independently combinable with each other.

[0027] As used herein, "tertiary flow" refers to a non-primary air flow that can add fluid energy to produce a small amount of total propulsion system thrust. The pressure ratio of the tertiary flow can be higher than the pressure ratio of the primary propulsion flow (e.g., the propulsion flow driven by a fan or propeller). Thrust can be generated by a dedicated nozzle or by mixing the airflow through the tertiary flow with the primary propulsion flow or core air flow, such as into a common nozzle.

[0028] In certain exemplary embodiments, the operating temperature of the airflow through the third stream can be below the maximum compressor discharge temperature of the engine, and more specifically, can be below 350 degrees Fahrenheit (e.g., below 300 degrees Fahrenheit, such as below 250 degrees Fahrenheit, such as below 200 degrees Fahrenheit, and at least as high as the ambient temperature). In certain exemplary embodiments, these operating temperatures can facilitate heat transfer from or to the airflow through the third stream and the separate fluid streams. Furthermore, in certain exemplary embodiments, the airflow through the third stream can contribute less than 50% of the total engine thrust (and at least, for example, 2% of the total engine thrust) under takeoff conditions, or more specifically, when operating at sea level at rated takeoff power, static flight speed, and 86 degrees Fahrenheit ambient temperature operating conditions.

[0029] Moreover, in certain exemplary embodiments, aspects of the airflow through the third stream (e.g., airflow, mixing, or exhaust characteristics), and thereby the above-described exemplary percentage contribution to total thrust, can be passively adjusted during engine operation or purposefully modified through the use of engine control features (e.g., fuel flow, motor power, variable stators, variable inlet guide vanes, valves, variable exhaust geometry, or flow characteristics) to adjust or optimize overall system performance over a wide range of potential operating conditions.

[0030] The present disclosure generally relates to a gas turbine engine with a ductless fan, and more specifically to the relative positioning between an engine inlet and a fan axis of a ductless fan of the engine. The proposed engine configuration includes a third flow architecture to reduce the diameter of the ductless fan. In addition, the engine inlet is offset from the fan axis. The offset engine inlet configuration can improve the underhood packaging arrangement and reduce the diameter of the fan. In addition, this combination of the third flow and the offset inlet can provide improved deformation limit and inlet debris extraction.

[0031] The disclosed third flow engine configuration operates by modifying the fan pressure ratio of the unducted fan so that the fan pressure ratio supports takeoff and climb using a smaller diameter fan, while maintaining a low fan pressure ratio at cruise to improve propulsive efficiency.

[0032] Referring now to the drawings, wherein like numerals refer to like elements throughout the several views, Figure 1is a cross-sectional view of a propulsion system 10 and illustrates an unducted fan 12 (having a nose 13, fan blades 14, a fan axis 16, and a fan flow 18), guide vanes 20, a row 22 of guide vanes 20, a casing 24, an inlet 26, an inlet passage 28, an inlet surface 30, an inlet flow 32, a tertiary stream flow path 34, a tertiary stream outlet flow 36, a working gas flow path 38, a turbine 40 (having a compressor section 42 (including a mid-fan 44 having a diameter 46 and a high pressure compressor 47), a combustor 48, a turbine section 50, a turbine section 50, and an exhaust section 52), an exhaust flow 54, a load device 56, a shaft 58 (having an axis 60), an axial direction A, a turbine direction R T and the fan radial direction RF. For reference purposes, Figure 1 A forward direction F is also depicted by arrow F, which in turn indicates the front and rear of the propulsion system 10 .

[0033] The propulsion system 10 is a gas turbine engine. In this example, the propulsion system 10 defines an axial direction A. Also in this example, the unducted fan 12 defines a fan radial direction R F And the turbine 40 defines a turbine radial direction R T .like Figure 1 As shown, propulsion system 10 takes the form of an open rotor propulsion system having an unducted fan 12 including a set or stage of fan blades 14 arranged about a fan axis 16 of unducted fan 12. In this manner, propulsion system 10 may be referred to as a single unducted rotor propulsion system 10.

[0034] In one example, the propulsion system 10 may be connected to an aircraft, such as via a pylon to a wing of the aircraft. The propulsion system 10 may be configured to provide thrust to the aircraft on which the propulsion system 10 is installed.

[0035] The ductless fan 12 is a ductless rotary fan configured to rotate about a fan axis 16. The ductless fan 12 is mounted at the upstream end of the propulsion system 10. The ductless fan 12 is operably coupled to a load device 56 and driven by the torque transmitted by the load device 56. Figure 1 As shown, the ductless fan 12 is located in front of the turbine 40 in a "puller" configuration. As shown, the ductless fan 12 is driven by the turbine 40, more specifically, by the shaft 58. More specifically, Figure 1The propulsion system 10 in the illustrated embodiment includes a load device 56, and the unducted fan 12 is driven across the load device 56 by a shaft 58 of the turbine 40. In this manner, the fan blades 14 of the unducted fan 12 can rotate about the fan axis 16 and generate thrust to propel the propulsion system 10, and thus propel the aircraft to which the propulsion system 10 is installed, in the forward direction F. The load device 56 may include a gear set for increasing or decreasing the rotational speed of the shaft 58 relative to the turbine section 50, so that the unducted fan 12 can rotate at a slower or faster rotational speed than the shaft 58.

[0036] As the unducted fan 12 rotates, the fan blades 14 generate a fan stream 18 and push air in the fan stream to provide propulsive thrust to an aircraft, for example.

[0037] The fan blades 14 and the guide vanes 20 are airfoils. Throughout this disclosure, the fan blades 14 will be referred to in the plural form as a plurality of fan blades 14. The fan blades 14 are arranged along the outer surface of the fan 12, generally arranged in an equally spaced relationship in a circumferential direction around a fan axis 16. The fan blades 14 are arranged to rotate about the fan axis 16 together with the ductless fan 12. The fan blades 14 are arranged upstream of the guide vanes 20. In some examples, the fan blades 14, the guide vanes 20, or both may be combined with a pitch change mechanism so that the airfoils (e.g., the fan blades 14, the guide vanes 20) may be rotated independently or in conjunction with each other relative to the pitch rotation axis. Such pitch variations may be used to vary thrust and / or vortex effects under various operating conditions, including adjusting the magnitude or direction of thrust generated at the fan blades 14, or providing a thrust reversal feature that may be useful under certain operating conditions, such as when an aircraft is landing, or desirably adjusting at least in part the acoustic noise generated by the fan blades 14, the guide vanes 20, or the aerodynamic interaction of the fan blades 14 relative to the guide vanes 20.

[0038] The fan axis 16 is the centerline axis of the ductless fan 12. The fan axis 16 is disposed to pass through the axial center point of the ductless fan 12.

[0039] The fan flow 18 is the air flow or airflow generated from the fan blades 14 of the ductless fan 12. The fan flow 18 is at least partially defined by an exterior surface of the housing 24.

[0040] The fan flow 18, the third stream flow path 34, and the working gas flow path 38 may each include an exhaust nozzle structure. Such structures may include variable area structures, fixed area structures, convergent-divergent nozzles, thrust vectoring structures, lobe exhaust mixers, or other suitable exhaust structures.

[0041] The guide vanes 20 extend from the housing 24 and are positioned behind the ductless fan 12. The guide vanes 20 are fixed guide vanes. For example, the guide vanes 20 may be mounted to the housing 24 (e.g., a fixed frame or other mounting structure) and do not rotate relative to the fan axis 16. The guide vanes 20 are generally arranged around the fan axis 16 in an equidistant relationship along a circumferential direction (e.g., a circumferential direction extending around the fan axis 16). As will be appreciated, the guide vanes 20 may be configured to straighten the fan flow 18 from the ductless fan 12 (e.g., by reducing vortices in the fan flow 18) to increase the efficiency of the propulsion system 10. For example, the size, shape, and configuration of the guide vanes 20 may be designed to impart a reverse swirl to the fan flow 18 from the fan blades 14, so that in the downstream direction behind the two rows of airfoils (e.g., fan blades 14, guide vanes 20), the degree of vortex in the fan flow 18 is greatly reduced, which may translate into an improved level of induced efficiency.

[0042] The row 22 is a row or stage of guide vanes 20 . The row 22 of guide vanes 20 is arranged downstream of the fan blades 14 .

[0043] The housing 24 is an outer shell or cowling that defines the exterior of the propulsion system 10. The turbine 40 is typically mounted in the housing 24. Furthermore, it should be appreciated that the housing 24 at least partially defines the inlet 26 and the exhaust section 52, and includes the working gas flow path 38 that extends between the inlet 26 and the exhaust section 52. The housing 24 provides structural and aerodynamic support for the components of the propulsion system 10.

[0044] Inlet 26 is a fluid port or opening. Figure 1 As shown, the inlet 26 is disposed behind the guide vanes 20 (in the axial direction A) and provides a path for the incoming atmosphere (and a portion of the fan flow 18) to enter the inlet passage 28. Such a location is advantageous for a variety of reasons, including the management of icing performance and protecting the inlet 26 from various objects and materials that may be encountered during operation. The inlet 26 receives the air flow and provides it to the propulsion system 10, particularly the third stream flow path 34 and the turbine 40. As will be described in more detail below (e.g., with reference to Figure 4 and Figure 5 ), the inlet 26 is an offset inlet, defining a non-annular shape.

[0045] The inlet channel 28 is a channel or passage configured to convey a fluid (e.g., the inlet flow 32). The inlet channel 28 is fluidly connected to the inlet 26 and extends between the third stream flow path 34 and the working gas flow path 38. The inlet channel 28 receives the inlet flow 32 from the inlet 26 and conveys the inlet flow 32 to the third stream flow path 34 and the working gas flow path 38.

[0046] The inlet surface 30 is an inner surface of the inlet passage 28. The inlet surface 30 is disposed along a front side of the inlet passage 28. The inlet surface 30 helps to direct the inlet flow 32 as it flows through the inlet passage 28.

[0047] Inlet flow 32 is the air flow or inlet airflow from inlet 26 through inlet passage 28. Inlet flow 32 passes through and is contained within inlet passage 28.

[0048] The third flow path 34 is a flow path that at least partially bypasses the turbine 40. The third flow path 34 is disposed downstream of the inlet passage 28 and extends from the working gas flow path 38 in the turbine radial direction R t In some embodiments, the third stream flow path 34 is in fluid communication with the fan flow 18 via the inlet 26 and the inlet passage 28. In the depicted embodiment, the third stream flow path 34 is configured to receive a portion of the inlet flow 32 from the inlet passage 28 at a position downstream of the middle fan 44. More specifically, in the depicted embodiment, the third stream flow path 34 is configured to receive a portion of the inlet flow 32 from the inlet passage 28 at a position downstream of the middle fan 44 and upstream of the high pressure compressor 47.

[0049] The third stream flow path 34 is configured to receive fluid energy and provide such fluid energy to the fan flow 18 to generate a portion of the total thrust of the propulsion system 10. During operation, the propulsion system 10 is configured so that the pressure of the airflow through the third stream flow path 34 during operation is higher than the pressure of the fan flow 18. In one embodiment, the third stream flow path 34 may include a dedicated exhaust nozzle at the outlet end. In the embodiment shown, the airflow through the third stream flow path 34 may be mixed with the fan flow 18. Various embodiments of the third stream flow path 34 are configured to generate less than 50% of the total thrust of the propulsion system 10. In certain embodiments, during operation, the third stream flow path 34 is configured to generate 2% or more of the total thrust of the propulsion system 10. In one embodiment, the third stream flow path 34 is configured to generate 2% or more of the total thrust and up to 50% of the total thrust under takeoff conditions, full load conditions, or rated takeoff power conditions. Exemplary rated takeoff power conditions may be associated with sea level static flight under 86 degrees Fahrenheit ambient temperature operating conditions.

[0050] In certain embodiments, the operating temperature of the air passing through the third stream flow path 34 is less than the maximum compressor discharge temperature of the propulsion system 10. In a particular embodiment, the operating temperature of the air passing through the third stream flow path 34 is less than approximately 350 degrees Fahrenheit. In another embodiment, the operating temperature of the air passing through the third stream flow path 34 is less than approximately 250 degrees Fahrenheit. In yet another embodiment, the operating temperature of the air passing through the third stream flow path 34 is less than approximately 200 degrees Fahrenheit. In various embodiments, the operating temperature of the air passing through the third stream flow path 34 is at least the ambient temperature, or at least the temperature of the air entering the compressor section 42. It should be understood that the operating temperature range through the third stream flow path 34 can allow heat to be transferred to or from the third stream flow path 34 and another flow path (e.g., the working gas flow path 38).

[0051] Those skilled in the art will appreciate that the tertiary flow path 34 extends from the working gas flow path 38 upstream of the combustor 48. In certain embodiments, the tertiary flow path 34 extends downstream of the ductless fan 12. In various embodiments, the tertiary flow path 34 is configured to allow air flow to exit the propulsion system 10 to generate a portion of the total thrust of the propulsion system 10, such as described herein.

[0052] Furthermore, it will be appreciated by those skilled in the art that the third stream flow path 34 is at least partially distinct from the exhaust circuit or other flow paths via the outlet of air as a propulsive thrust.

[0053] The propulsion system 10 may be configured to passively adjust thrust output from the third stream flow path 34. In certain embodiments, the propulsion system 10 may be configured to actively adjust thrust output from the third stream flow path 34, such as via adjustment of fuel flow, motor load, variable stator, variable inlet guide vanes, variable exhaust area or geometry, or general flow characteristics, and based on desired performance relative to operating conditions.

[0054] The tertiary outflow flow 36 is the air flow that flows out of the tertiary outflow flow path 34. The tertiary outflow flow 36 is discharged from the tertiary outflow flow path 34. The tertiary outflow flow 36 provides a portion of the thrust of the propulsion system 10.

[0055] The working gas flow path 38 is the flow of air from the inlet passage 28 through the turbine 40. The working gas flow path 38 is disposed within the interior of the turbine 40 and is fluidly connected to the inlet passage 28. The working gas flow path 38 extends through at least the mid-fan 44, the compressor section 42, the combustor 48, the turbine section 50, and the exhaust section 52. The working gas flow path 38 provides an air and / or gas flow through the turbine 40.

[0056] exist Figure 1 In the depicted embodiment, the propulsion system 10 includes a turbine 40 with a compressor section 42 having a mid-fan 44, a combustor 48, a turbine section 50, an exhaust section 52, and a shaft 58 extending therebetween and connecting the compressor section 42 and the turbine section 50. The turbine 40 is disposed in the housing 24 and downstream of the ductless fan 12. The turbine 40 provides rotational power to drive the ductless fan 12.

[0057] The compressor section 42 is a portion of the turbine 40 that compresses the airflow through the working gas flow path 38 to deliver a high-pressure air flow to the combustor 48. In one example, the compressor section 42 includes a mid-fan 44 (which can be considered part of the low-pressure compressor of the compressor section 42) and a high-pressure compressor 47. The compressor section 42, the combustor 48, the turbine section 50, and the exhaust section 52 are generally positioned in a series aerodynamic flow arrangement. The compressor section 42 compresses the incoming air and increases the pressure of the air before delivering the pressurized air to the combustor 48.

[0058] The mid-fan 44 is arranged to rotate about an axis 60 to compress the airflow through the working gas flow path 38. The mid-fan 44 is arranged on the upstream end of the turbine 40 at least partially within the working gas flow path 38, and is in fluid communication with the inlet passage 28 and the third stream flow path 34, as will be further explained below. In one example, the mid-fan 44 can be operably coupled to the shaft 58. The mid-fan 44 compresses and pushes air through the working gas flow path 38 and into the third stream flow path 34 in a downstream direction.

[0059] Diameter 46 is the diameter of the middle fan 44 and is defined by the distance between the distal ends of the tips of the fan blades of the middle fan 44 opposite the axis 60 .

[0060] The combustor 48 is a portion of the turbine 40 that is configured to combust the gas flow received from the compressor section 42. The combustor 48 is disposed between and fluidly connected to the compressor section 42 and the turbine section 50. The combustor 48 may include one or more configurations for receiving a mixture of fuel and air and for providing a flow of combustion gases through the turbine section 50 to drive a shaft 58.

[0061] In various embodiments, the combustor 48 may be configured as a deflagration combustion system or a detonation combustion system. The combustor 48 may include any suitable type of system for receiving a liquid and / or gas fuel flow and generating hot gases, including but not limited to annular, can, trapped vortex, volute or vortex, rotating detonation, pulse detonation, subsonic, or supersonic combustion systems.

[0062] The turbine section 50 is a portion of a turbine that is configured to convert and transfer energy from an airflow flowing through the turbine section 50 into rotational energy or torque in a shaft 58. The turbine section 50 is disposed between and fluidly connected to the combustor 48 and the exhaust section 52. The turbine section 50 is operably coupled to the ductless fan 12 via the shaft 58 and the load device 56 to drive one or more stages of fan blades 14 of the ductless fan 12. The turbine section 50 receives a high-speed airflow from the combustor 48 and converts energy from the high-speed air into rotational energy of the shaft 58.

[0063] The exhaust section 52 is a portion of the turbine 40 that is configured to exhaust an exhaust flow 54 from the propulsion system 10. The exhaust section 52 is disposed downstream of and fluidly connected to the turbine section 50. The exhaust section 52 exhausts air from the propulsion system 10 to generate thrust for the aircraft in which the propulsion system 10 is installed.

[0064] Exhaust flow 54 is the flow of air that flows out of propulsion system 10 via exhaust section 52. Exhaust flow 54 flows out of exhaust section 52. Exhaust flow 54 provides a portion of the total thrust provided by propulsion system 10.

[0065] In this example, the load device 56 is a gearbox. In other examples, the load device 56 may include one or more of a motor, a mechanical drive device, or a fluid flow device. For example, the load device 56 may form an accessory gearbox, a reduction gear assembly, a fan pitch assembly, or a main gearbox assembly. The load device 56 may include one or more fuel pumps, motors (e.g., motors and / or generators, constant frequency or variable frequency machines, hybrid power systems, etc.), lubricant pumps, hydraulic pumps, air compressors, engine starters, sensor drivers, and auxiliary gearbox drives, or combinations thereof. The load device 56 is operably coupled to the propulsion system 10 via a shaft 58 or, in particular, a high-speed spool connecting the compressor section 42 and the turbine section 50. The load device 56 transfers rotational energy from the shaft 58 to the ductless fan 12.

[0066] Shaft 58 is a rotating shaft. In one example, shaft 58 can be a low-speed or low-pressure shaft. For example, shaft 58 can be operably coupled to a low-pressure turbine of the turbine section 50 of the turbine 40 for receiving rotational energy from the low-pressure turbine and providing such rotational energy to a low-pressure compressor including a mid-section fan 44, and to the fan 12 via a load device 56. Axis 60 is a centerline axis of shaft 58. Axis 60 is disposed through a center point of shaft 58 and is parallel to the fan axis 16. Shaft 58 can rotate with the compressor section 42, the turbine section 50, or both. Axial direction A is a direction disposed parallel to the fan axis 16 and axis 60. Axial direction A is in Figure 1 The fan radial direction R F The radial direction R of the turbine is arranged perpendicular to the axial direction A and extends 360° around the fan axis 16. T It is arranged in a direction perpendicular to the axial direction A and extends 360° around the fan axis 60 .

[0067] The proposed configuration of the propulsion system 10 including a turbofan engine having an unducted fan 12 and a third-stream architecture enables a reduction in the diameter of the unducted fan 12. For example, the third-stream flow path 34 generates an additional amount of thrust via the third-stream outlet flow 36, and the diameter of the unducted fan 12 can be smaller because the amount of thrust required by the unducted fan 12 is reduced approximately by the amount of thrust provided by the third-stream flow path 34.

[0068] In addition, the propulsion system 10 including the inlet 26 offset from the fan axis 16 can improve the underhood packaging arrangement. Other benefits of the combination of the third stream flow path 34 and the inlet 26 offset from the fan axis 16 include improved deformation limit and debris extraction entering the inlet 26. For example, when debris enters the inlet passage 28, the debris will continue to enter the third stream flow path 34, because the momentum of the debris causes the debris to continue along its line of motion and / or the vortex of the airflow from the mid-fan 44 causes the heavier debris to be centrifuged outward relative to the axis 60. In contrast, the air flowing into the inlet passage 28 with the inlet flow 32 is bent and drawn into the working gas flow path 38. In this way, the amount of debris entering the turbine 40 can be reduced compared to a configuration without the third stream flow path 34. In addition, the third stream flow path 34 provides thermal management functions as well as additional thrust for the propulsion system 10.

[0069] Figure 2 yes Figure 1 An enlarged isolated view of portion 2 - 2 is indicated and illustrates inlet passage 28 , tertiary stream flow path 34 , working gas flow path 38 , mid-section fan 44 , blades 62 (having span 64 ), and tertiary stream inlet 66 (defining height 68 ).

[0070] The blades 62 are airfoils of the mid-section fan 44. Figure 2 , a single blade 62 is shown. However, the mid-fan 44 may include a plurality of blades 62 positioned annularly about the axis 60. In one example, the blade 62 is mounted to a rotating hub disposed as part of the compressor section 42. During operation, the blade 62 (and other blades) rotate about the axis 60 (see, e.g. Figure 1 ) rotates to push air through the working gas flow path 38 and, in the illustrated embodiment, into the tertiary flow path 34.

[0071] The span 64 is the length or height of the blade 62. In one example, the span 64 may be defined by the length between the interface of the root and airfoil portion of the blade 62 and the distal end or tip of the blade 62. The span 64 is defined by the length of the blade 62 in the radial direction R.

[0072] The third inlet 66 is the starting point of the third flow path 34. The third inlet 66 is disposed at the most upstream end of the third flow path 34 and is disposed downstream of the blades 62 of the middle fan 44. The third inlet 66 receives the accelerated airflow from the blades 62. In one example, the third inlet 66 is an annular inlet. In another example, the third inlet 66 is a non-annular inlet.

[0073] The height 68 is the length or distance of the third inlet 66 along the radial direction R of the turbine 40. The height 68 is the radial height of the third inlet 66. The height 68 is the height defined by the third inlet 66 along the radial direction R. In one example, the height 68 of the third inlet 66 is at least 5% of the span 64 of the blade 62 and up to 50% of the span 64 of the blade 62 of the mid-fan 44. In another example, the height 68 of the third inlet 66 is at least 10% of the span 64 of the blade 62 and up to 40% of the span 64 of the blade 62 of the mid-fan 44.

[0074] Because the height 68 of the third flow inlet 66 is at least 5% of the height 68 of the blade 62 and up to 50% of the span 64 of the blade 62, the third flow path can receive and absorb a large amount of debris contained in the airflow passing through the inlet passage, which in other existing systems without the third flow path 34 would be sucked into the internal turbine.

[0075] In certain exemplary embodiments, the inlet passage may include an inlet guide vane 69 and the third stream flow path 34 may include an outlet guide vane 70. Here, the inlet guide vane 69 and the outlet guide vane 70 are shown in phantom. In this example, the inlet guide vane 69 and the outlet guide vane 70 may be fixed geometry airfoils. In other examples, at least one of the inlet guide vane 69 and the outlet guide vane 70 may be a variable geometry guide vane.

[0076] It should be understood that Figure 1 and Figure 2The exemplary single ductless rotor propulsion system 10 depicted in FIG. 1 is provided as an example only, and in other exemplary embodiments, the propulsion system 10 may have any other suitable configuration, including, for example, any other suitable number of shafts or spools, turbines, compressors, etc.; direct drive configuration (i.e., without a gearbox), etc. For example, in other exemplary embodiments, the propulsion system 10 may be a three-shaft engine with a medium-speed compressor and / or turbine. In such a configuration, it should be understood that the terms "high" and "low" used herein with respect to the speed and / or pressure of the turbine, compressor, or spool are for convenience in distinguishing the components, but do not require any particular relative speed and / or pressure, and do not exclude additional compressors, turbines, and / or spools or shafts. Additionally or alternatively, in other exemplary embodiments, the propulsion system 10 may include a multi-stage open rotor configuration. Furthermore, in yet another exemplary embodiment, the turbine of the propulsion system may include one or more "reverse flow" sections, in which the gas passing through the working gas flow path flows in a forward direction. For example, in one or more of such embodiments, the turbine may include a reverse flow burner, or the entire working gas flow path may be configured in a reverse flow orientation.

[0077] For example, now refer to Figure 3 , a cross-sectional view of a propulsion system 10 ′ is provided according to another exemplary embodiment of the present disclosure. Figure 3 The exemplary propulsion system 10' may be configured with Figure 1 The exemplary propulsion systems are configured in substantially the same manner. For example, Figure 3 The propulsion system of FIG. 1 shows a propeller 12′ (having propeller blades 14′, propeller axis 16′ and propeller flow 18′), a housing 24, an inlet 26, an inlet passage 28, an inlet surface 30, an inlet flow 32, a third stream flow path 34, a third stream outlet flow 36, a working gas flow path 38, a turbine 40 (having a compressor section 42 (including a mid-fan 44 with a diameter 46 and a high-pressure compressor 47), a combustor 48, a turbine section 50, a turbine section 50 and an exhaust section 52), an exhaust flow 54, a load device 56, a shaft 58 (having an axis 60), an axial direction A, a propeller radial direction R P and the turbine radial direction R T For reference purposes, Figure 3 A forward direction F is also depicted by arrow F, which in turn defines the front and rear of the propulsion system 10 ′.

[0078] However, for Figure 3 In an exemplary embodiment of the invention, the propulsion system is configured as a turboprop engine, wherein the propeller 12 ′ is instead configured as a propeller 12 without the first stage exit guide vanes.

[0079] The propulsion system 10' is a gas turbine engine. In this example, the propulsion system 10' is configured as a single, ductless rotor propulsion system 10' defining an axial direction A. Figure 3 As shown, the propulsion system 10' takes the form of an open rotor propulsion system having a propeller 12' that includes an array or stage of propeller blades 14' arranged about a propeller axis 16' of the propeller 12'.

[0080] The propeller 12' is a fan configured to rotate about a propeller axis 16'. The propeller 12' is mounted at the upstream end of the propulsion system 10'. The propeller 12' is operably coupled to the load device 56 and is driven by the torque transmitted by the load device 56. Figure 3 As shown, propeller 12' is located in front of turbine 40 in a "puller" configuration. As depicted, propeller 12' is driven by turbine 40, and more specifically, by shaft 58. More specifically, Figure 3 The propulsion system 10' in the illustrated embodiment includes a load device 56, and the propeller 12' is driven across the load device 56 by a shaft 58 of the turbine 40. In this manner, the propeller blades 14' of the propeller 12' can rotate about the propeller axis 16' and generate thrust to propel the propulsion system 10', and thus the aircraft to which the propulsion system 10' is installed, in a forward direction F. The load device 56 may include a gear set for increasing or decreasing the rotational speed of the shaft 58 relative to the turbine section 50, so that the propeller 12' can rotate at a slower or faster rotational speed than the shaft 58.

[0081] As the propeller 12' rotates, the propeller blades 14' generate a propeller flow 18' and push air in a fan flow to provide propulsive thrust to the aircraft.

[0082] The propeller blades 14' are airfoils. Throughout this disclosure, propeller blades 14' will be referred to in the plural form as a plurality of propeller blades 14'. The propeller blades 14' are arranged along the outer surface of the propeller 12', generally in an equally spaced relationship in a circumferential direction around the propeller axis 16'. The propeller blades 14' are arranged to rotate around the propeller axis 16' together with the propeller 12'. The propeller blades 14' are arranged upstream of the inlet 26. In some examples, the propeller blades 14' can be combined with a pitch change mechanism so that the airfoils (e.g., propeller blades 14') can be rotated independently or in conjunction with each other relative to the pitch rotation axis. Such pitch changes can be used to change thrust and / or vortex effects under various operating conditions, including adjusting the magnitude or direction of the thrust generated at the propeller blades 14', or providing a thrust reversal feature that is useful under certain operating conditions, such as when the aircraft is landing, or it is desirable to adjust at least part of the acoustic noise generated by the propeller blades 14'.

[0083] The propeller axis 16' is the centerline axis of the propeller 12'. The propeller axis 16' is arranged to pass through the axial center point of the propeller 12'. The propeller flow 18' is the air flow or airflow generated by the propeller blades 14' of the propeller 12'. The propeller flow 18', the third stream flow path 34 and the working gas flow path 38 may each include an exhaust nozzle structure. Such a structure may include a variable area structure, a fixed area structure, a convergent-divergent nozzle, a thrust vectoring structure, a lobe exhaust mixer or other suitable exhaust structure.

[0084] In certain embodiments, the tertiary stream flow path 34 is in fluid communication with the propeller stream 18 ′ via the inlet 26 and the inlet passage 28 .

[0085] The third stream flow path 34 is configured to recover fluid energy to generate a portion of the total thrust of the propulsion system 10'. During operation, the pressure ratio of the third stream flow path 34 is higher than the pressure ratio at the propeller flow 18'. In one embodiment, a portion of the total thrust generated by the third stream flow path 34 may include a dedicated exhaust nozzle located at the outlet end. In another embodiment, a portion of the total thrust generated by the third stream flow path 34 may be mixed with the propeller flow 18'. In yet another embodiment, a portion of the total thrust generated by the third stream flow path 34 may be mixed with the working gas flow path 38 downstream of the combustor 48 and discharged through the exhaust section 52. Various embodiments of the third stream flow path 34 are configured to generate less than 50% of the total thrust of the propulsion system 10'. In certain embodiments, during operation, the third stream flow path 34 is configured to generate 2% or more of the total thrust of the propulsion system 10'. In an embodiment, the third stream flow path 34 is configured to generate 2% or more of the total thrust and up to 50% of the total thrust under takeoff conditions, full load conditions, or rated takeoff power conditions. Exemplary rated takeoff power conditions may relate to sea level static flight at an 86 degree Fahrenheit ambient temperature operating condition.

[0086] The turbine section 50 is operably coupled to the propeller 12' via a shaft 58 and a load device 56 to drive one or more stages of propeller blades 14' of the propeller 12'. The load device 56 transfers rotational energy from the shaft 58 to the propeller 12'.

[0087] The shaft 58 is a rotating shaft. In one example, the shaft 58 may be a low speed or low pressure shaft. The shaft 58 is operably coupled to the turbine 40 to drive the load device 56. The shaft 58 transfers rotational energy from the turbine section 50 to the load device 56. The axis 60 is the centerline axis of the shaft 58. The axis 60 is disposed through the center point of the shaft 58 and is parallel to the propeller axis 16'. The axial direction A is a direction disposed parallel to the propeller axis 16' and the axis 60. The axial direction A is Figure 3The propeller radial direction RP is a direction arranged perpendicular to the axial direction A and the two propeller axes 16 ′. The turbine radial direction RT is a direction arranged perpendicular to the axial direction A and the axis 60 .

[0088] The proposed configuration of the propulsion system 10′ including the turbofan engine having the propeller 12′ and the third flow architecture enables a reduction in the diameter of the propeller 12′. For example, since the third flow flow path 34 generates an additional amount of thrust via the third flow outlet flow 36, the diameter of the propeller 12′ can be smaller because the amount of thrust required for the propeller 12′ is reduced approximately by the amount of thrust provided by the third flow flow path 34.

[0089] As described above, the propulsion system 10 of the present disclosure includes an inlet 26 that is configured to be offset from the inlet, defining a non-annular shape. Figure 4 , provides a front view facing the rear of the propulsion system 10 and illustrates the ductless fan 12 (having a nose 13, a fan axis 16, a housing 24, an inlet 26 (having a centerline axis 72), and an angular range θ.

[0090] In this example, the fan blades 14 are omitted from the ductless fan 12 for clarity. As shown here, the propulsion system 10 includes a single inlet 26 disposed at the bottom of the propulsion system 10 (the bottom is at Figure 4 In other examples, the inlet 26 may be offset into a single discrete inlet, such as Figure 4 As shown, or multiple discrete inlets (e.g., bifurcated). As described above, the inlet 26 is fluidly connected to the third stream flow path 34 and the working gas flow path 38 (see, e.g., Figure 1 ).like Figure 4 As shown, the inlet 26 defines a non-annular shape. In some embodiments, a single inlet (e.g., Figure 4 ) or multiple entries (e.g. Figure 5 ) does not circumscribe or surround the fan axis 16. In some embodiments, the single inlet or multiple inlets are close to the axis and away from the fan axis (e.g., at Figure 4 and Figure 5 In the embodiment, the inlets 26, 26A, and 26B are respectively close to the axis 60 and away from the axis 15.)

[0091] The head 13 is the rotator of the ductless fan 12. The head 13 is configured to rotate with the ductless fan 12 around the fan axis 16. The centerline axis 72 is the inlet centerline axis or axial centerline of the inlet 26. In this example, the centerline axis 72 enters and exits the page. Similarly, the fan axis 16 and the axis 60 are shown as pointing in and out of the page. In addition, the centerline axis 72 is located at the midpoint of the inlet 26 along the circumferential direction of the ductless fan 12 and is located at the midpoint of the inlet 26 along the radial direction of the ductless fan 12.

[0092] like Figure 4 As shown, the inlet 26 extends along a circumferential direction C of the propulsion system 10. The inlet 26 extends along a first portion of the circumference of the propulsion system 10, the first portion being represented by an angular range θ. The angular range θ, Figure 4 In , it refers to the angle between an imaginary reference line extending between opposite ends of the inlet 26 in a circumferential direction and the axis 60. In addition, it should be understood that the inlet 26 provides substantially all of the air flow to the working gas flow path 38 and the third flow flow path 34. In one example, the angular range is less than or equal to 90° relative to the axis 60. In another example, the angular range θ is less than or equal to 180° relative to the axis 60 (e.g., less than or equal to half of the total circumference of the propulsion system 10). In another example, the angular range θ is less than or equal to 270° relative to the axis 60. In such an example, the centerline axis 72 on the inlet 26 is disposed offset from the fan axis 16 in a radial direction relative to the fan axis 16 of the ductless fan 12. In other words, the centerline axis 72 is offset from the fan axis 16 so that the fan axis 16 and the centerline axis 72 are not coaxial.

[0093] Also like Figure 4 As shown, Figure 4 is the relative positioning between the fan axis 16 and the axis 60. Figure 4 As can be seen, the fan axis 16 and the axis 60 of the shaft 58 (see for example Figure 1 and Figure 3 ) are offset from each other along the radial direction of the propulsion system 10, and so that the fan axis 16 and the axis 60 are not coaxial.

[0094] It should be understood that Figure 4 The exemplary propulsion system 10 depicted in FIG. 1 is provided as an example only, and in other exemplary embodiments, Figure 4 A propulsion system 10' may alternatively be depicted having a propeller 12' but without guide vanes 20 (see e.g. Figure 3 ).

[0095] Furthermore, it should be understood that the inlet 26 may have any other suitable configuration. Figure 5, provides a front view facing the rear of the propulsion system 10, which illustrates the propulsion system 10 according to another exemplary embodiment of the present disclosure. Figure 5 In the embodiment, propulsion system 10 includes a ductless fan 12 (having a nose 13), a fan axis 16, a housing 24, a first inlet 26A (having a first centerline axis 72A), a second inlet 26B (having a second centerline axis 72B), and an angular range θ.

[0096] exist Figure 5 In the example shown in , the propulsion system 10 includes two inlets 26 , including a first inlet 26A and a second inlet 26B. In other examples, the propulsion system 10 may include more or less than two discrete inlets 26 spaced circumferentially around the nose 13 .

[0097] The first inlet 26A and the second inlet 26B are two separate and discrete inlets 26 that are spaced a distance from each other along the circumferential direction C of the propulsion system 10. The first centerline axis 72A and the second centerline axis 72B are both disposed offset from the fan axis 16 along the radial direction of the ductless fan 12. In other words, the first centerline axis 72A and the second centerline axis 72B are not coaxial with the fan axis 16. Figure 5 In the embodiment of FIG. 1 , each of the inlets 26A, 26B is located in the lower half of the propulsion system 10 (defined relative to the height of the housing 24 at the axial position shown) and together define a similar configuration to that described with reference to FIG. Figure 4 The angle θ is described for an angular range θ (not labeled). Figure 5 The angular range θ extends from the distal end of the first inlet 26A to the distal end of the second inlet 26B. In other examples, each inlet 26A, 26B may be positioned at any angular position relative to the fan axis 16 or axis 60.

[0098] Figure 6 is a graph illustrating a portion of the inlet passage 28 , and more specifically illustrating the curvature 74 of the inlet surface 30 of the inlet 26 of the propulsion system 10 . Figure 6 A curvature 74 (having a first portion 76 and a second portion 78 ) and an inflection point 80 are shown.

[0099] The curvature 74 is the curved shape of the inlet surface 30. The first portion 76 is the initial or upstream section of the curvature 74. In this example, the first portion 76 is concave (eg, Figure 6 The second portion 78 is a secondary or downstream portion of the curvature 74. In this example, the second portion 78 of the curvature 74 is concave (as shown in FIG. Figure 6The inflection point 80 is a point along the curvature 74 at the interface between the first portion 76 and the second portion 78. In this example, the inflection point 80 is located at the transition point of the first portion 76, which is concave upward, and the second portion 78 is concave downward.

[0100] In the examples disclosed herein (see Figure 1-3 ), the inlet surface 30 defines a curvature 74 and the slope X is defined by equation 1.1.

[0101] X = ΔR / ΔA Equation 1.1

[0102] Here, ΔR is a change in a first length along a turbine radial direction RT of the propulsion system 10 and ΔA is a change in a second length along an axial direction A of the propulsion system 10. It is worth noting that the slope X is a local slope defined at the inflection point 80, more specifically, in the illustrated embodiment, the slope X is the maximum slope of the inlet surface 30.

[0103] As will be understood from the description herein, certain embodiments of the present disclosure having an offset inlet include a third flow inlet at a location immediately downstream of the mid-fan (e.g., downstream of the mid-fan 44, closer to the mid-fan 44 than the first stage of the next stage of compressor rotor blades). The height of the inlet of the third flow flow path may allow for a higher slope compared to existing configurations because an outer portion of the airflow through the inlet passage 28, which may be separated from the inlet surface 30 due to the high slope X (creating a relatively large amount of turbulence in the airflow), is drawn into the third flow flow path, and an inner portion of the airflow through the inlet passage 28, which may have a relatively low amount of turbulence, is drawn into the downstream portion of the turbine. The increased slope X achieved by this configuration may allow for desired packaging of the propulsion system 10.

[0104] In one example, the inflection point 80 is defined by a location along the inlet surface 30 where the second derivative of the best fit equation for the curvature of the inlet surface 30 is zero. In another example, the inflection point 80 is disposed at a point along the curvature where the second derivative of the best fit equation is an isolated zero and changes sign. In such an example, the best fit equation may be a logistic function (e.g., to match Figure 1-3 The shape or curvature of the inlet surface 30 shown).

[0105] Here, because the propulsion system 10 having the offset inlet 26 provides benefits in debris extraction before the airflow enters the working gas flow path 38 , the slope of the inlet passage 28 may be greater than that achieved with prior propulsion systems not having the offset inlet 26 .

[0106] Now briefly refer to Figure 7 ,along Figure 1A view of the axis 60 of the exemplary propulsion system 10 is provided at the third inlet 66. It will be appreciated that Figure 7 In the exemplary embodiment, the third flow inlet 66 is an annular inlet that is substantially symmetrical about the axis 60. However, as depicted by the dashed lines, in certain exemplary aspects, the third flow inlet 66′ can be non-annular, or more specifically asymmetrical about the axis 60. Using this configuration, the third flow inlet 66′ can define a higher height (e.g., between 5% greater and 15% greater, such as at least 10%, 15% greater, or 20% greater) at the top end of the propulsion system 10 compared to the bottom end of the propulsion system 10 in order to accommodate a larger amount of relatively high turbulence airflow at the top end due to potential separation of the airflow from the inlet surface 30.

[0107] Reference now Figure 8 , Figure 8 is a cross-sectional view of a turbine 140A having a first counter-flow configuration according to an exemplary aspect of the present disclosure.

[0108] In certain exemplary embodiments, turbine 140A includes one or more "reverse flow" sections, wherein the working gas flow path flows in a forward direction (eg, to the left, such as Figure 8 ). For example, in one or more such embodiments, the turbine 140A can be configured such that the entire working gas flow path is configured in a counter-flow orientation. The turbine 140A defines an axial direction A and a turbine radial direction RT. In this exemplary embodiment, the turbine 140A can be disposed in the housing 24 of the propulsion system 10.

[0109] The turbine 140A includes a reverser 141. The reverser 141 is configured to change the direction of the working gas flow 146 from a rearward direction (e.g., Figure 8 ) is reversed to the forward direction (e.g., as shown in FIG. Figure 8 In certain exemplary embodiments, the reverser 141 may be mounted to a non-rotating structural component of the propulsion system 10. Additionally or alternatively, in other exemplary embodiments, the reverser 141 may include one or more reversers disposed about the axis 60.

[0110] Turbine 140A further includes a compressor section 142. In the exemplary embodiment, compressor section 142 is disposed at the rear end of turbine 140A behind combustor 148 (e.g., Figure 8During operation, compressor section 142 receives a portion of working gas flow 146 in the form of reverse inlet flow 156. In certain exemplary embodiments, reverse inlet flow 156 is a portion of working gas flow 146 that is fed into and reversed by one or more flow reversers 141.

[0111] Turbine 140A further includes a fan 144A. In certain exemplary embodiments, fan 144A may be referred to as a mid-fan or a third-flow fan. In the exemplary embodiment, fan 144A is disposed behind compressor section 142 (e.g., to the right of compressor section 142 along axial direction A). However, it should be understood that in other exemplary embodiments, fan 144A may be disposed at other locations along axis 60 relative to compressor section 142 (see, e.g., FIG. 1 ). Figure 9-10 ).

[0112] Turbine 140A further includes a combustor 148. In addition to the forward direction (e.g., Figure 8 In addition to receiving and sending a working gas stream 146 (shown on the left side), the burner 148 can be connected to the working gas stream 146 from Figure 1 and Figure 3 The burner 148 is configured in substantially the same manner.

[0113] The turbine 140A further includes a high pressure turbine 152 and a low pressure turbine 154. The high pressure turbine 152 and the low pressure turbine 154 are fluidly connected to the combustor 148 and are configured to receive the working gas flow 146 therefrom. The high pressure turbine 152 and the low pressure turbine 154 together form the turbine section of the turbine 140A.

[0114] The turbine 140A further includes an exhaust duct 158 ​​in fluid communication with the low-pressure turbine 154. The exhaust duct 158 ​​is configured to receive an exhaust flow 160 from the low-pressure turbine 154. In certain exemplary embodiments, the exhaust duct 158 ​​is configured to collect the exhaust flow 160 along 360° of the front end of the low-pressure turbine 154. In yet another exemplary embodiment, the exhaust duct 158 ​​may be configured to discharge the exhaust flow 160 along 360° of the turbine 140A or through one or more discrete openings disposed along the housing 24 (see, e.g., Figure 5 inlets 26A and 26B).

[0115] It should be understood that the turbine 140A may be used with Figure 1 The propulsion system 10 and Figure 3 The illustrated propulsion system 10' is used in conjunction with, for example, replacing, turbine 40. More specifically, in at least some exemplary aspects, the axis 60 (eg, the axis of rotation) of turbine 140A can be offset from the fan axis (eg, Figure 3 10 ′) such that the axis 60 of the turbine 140A is disposed at a first distance from the fan axis. It should also be understood that the counterflow configuration discussed herein may be incorporated into a single shaft or dual shaft machine.

[0116] Here, Figure 8 The embodiment of the turbine 140A provided and discussed in the present invention provides a useful configuration when the core size of the turbine 140A is relatively small relative to larger sized turbines. For example, for a smaller turbine 140A, a correspondingly smaller third flow inlet provides fewer challenges in reversing the flow of the working gas stream 146. In addition, a reverse flow configuration such as that employed by the turbine 140A provides advantages in terms of inlet separation and distortion mitigation of the overall propulsion system 10' (see, e.g., Figure 3 ).

[0117] Reference now Fig. 9 , Fig. 9 is a cross-sectional view of a turbine 140B having a second counter-flow configuration according to an exemplary aspect of the present disclosure. Fig. 9 The provided embodiments can be used with Figure 8 The configuration is described in substantially the same manner, with a different placement of fan 144B.

[0118] In the exemplary embodiment, fan 144B is disposed forwardly from compressor section 142 in axial direction A (eg, Fig. 9 More specifically, in at least some exemplary aspects, fan 144B is aligned with a portion of high pressure turbine 152 and a portion of low pressure turbine 154 along a radial direction R of turbine 140B.

[0119] Reference now Fig.10 , Fig.10 is a cross-sectional view of a turbine 140C having a third counter-flow configuration according to an exemplary aspect of the present disclosure. Fig.10 The embodiments provided in Fig. 9 Configured in substantially the same manner as described, with a different placement of fan 144C.

[0120] exist Fig.10 In this exemplary embodiment provided in , fan 144C is arranged to be aligned with compressor section 142 along axial direction A. In certain exemplary embodiments, fan 144C can be configured as a bladed fan. More specifically, in at least certain exemplary aspects, the blades of fan 144C are radially outward and directly connected to the rotating fan blades of compressor section 142, wherein the fan blades of compressor section 142 are mounted to a disk of compressor section 142, which is mounted to the shaft of turbine 140C.

[0121] Here, the fan blade configuration of fan 144C provides for adjustment of the pressure and temperature of the working fluid flowing through compressor section 142 and fan 144C as may be desired in certain engine types and desired usage conditions.

[0122] Reference now Fig.11 , Fig.11 is a cross-sectional view of a propulsion system 10 ′ having a turbine 140D including a fourth reverse flow configuration according to an exemplary aspect of the present disclosure.

[0123] Fig.11 The embodiment of the propulsion system 10′ provided in Figure 3 The described method is configured in substantially the same manner, with like numerals representing like elements between the figures, with turbine 140D replacing turbine 40 .

[0124] In certain exemplary embodiments, the turbine 140D includes a counter-flow configuration. More specifically, in at least certain exemplary aspects, the combustor 148D of the turbine 140D is configured to flow in a forward direction (e.g., as Fig.11 The compressor section 142 and the turbine section (eg, the high pressure turbine 152 and the low pressure turbine 154) are configured to direct the working gas flow 146 in an aft direction (eg, as shown in FIG. Fig.11 The working gas flow 146 is directed from left to right as shown.

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

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

[0127] A propulsion system comprises: a ductless rotary fan, the ductless rotary fan defining a fan axis; and a turbine, the turbine being disposed downstream of the ductless rotary fan, wherein the turbine defines a working gas flow path flowing therethrough; wherein the propulsion system defines a third flow path and an inlet passage, the inlet passage having an inlet offset from the fan axis, wherein the inlet passage is configured to provide an inlet airflow to the working gas flow path, and wherein the third flow path at least bypasses a portion of the turbine.

[0128] A propulsion system according to one or more of these clauses, wherein the propulsion system further comprises: a gearbox operably coupled to the ductless rotary fan; and a rotating shaft extending from the turbine and connected to the gearbox, wherein the gearbox is operably coupled to the turbine via the rotating shaft, wherein the rotating shaft defines an axis.

[0129] The propulsion system of one or more of these clauses, wherein the rotational axis is offset from the fan axis such that the rotational axis is disposed at a first distance from the fan axis.

[0130] A propulsion system according to one or more of these clauses, wherein the turbine includes a compressor section having a center fan with center fan blades, wherein the propulsion system includes an inlet surface that partially defines the inlet passage, wherein the inlet surface defines a curvature including an inflection point, wherein the inflection point is disposed at a point along the curvature.

[0131] A propulsion system according to one or more of these clauses, wherein the propulsion system includes an inlet surface that partially defines the inlet passage, wherein the inlet surface defines a curvature, wherein the curvature includes an inflection point, wherein the curvature of the inlet surface includes a first portion and a second portion, wherein the first portion is concave, wherein the second portion is concave, and wherein the inflection point is located at a transition point between the concave first portion and the concave second portion.

[0132] The propulsion system of one or more of these clauses, wherein the inlet of the inlet passage defines a non-annular shape.

[0133] A propulsion system according to one or more of these clauses, wherein the propulsion system defines a forward direction and a rearward direction, wherein the ductless rotary fan is configured to propel fan flow in the rearward direction, and wherein at least a portion of the turbine is disposed in a counterflow arrangement such that the turbine is configured to direct a portion of the working gas flow path in the forward direction.

[0134] The propulsion system of one or more of these clauses, further comprising a row of fixed guide vanes disposed downstream of the unducted rotary fan and upstream of the turbine.

[0135] The propulsion system of one or more of these clauses, wherein the turbine includes a compressor section having a mid-fan, wherein the third stream flow path is configured to receive a portion of the inlet airflow from the inlet passage at a location downstream of the mid-fan.

[0136] A propulsion system according to one or more of these clauses, wherein the third flow flow path defines a third flow inlet having a radial height, wherein the mid-fan defines a blade span, and wherein the radial height of the third flow inlet is at least 5% of the blade span of the mid-fan and up to 50% of the blade span of the mid-fan.

[0137] A propulsion system according to one or more of these clauses, wherein the turbine includes a compressor section having a mid-fan and a high-pressure compressor, wherein the third stream flow path is configured to receive a portion of the inlet airflow from the inlet passage at a position downstream of the mid-fan and upstream of the high-pressure compressor.

[0138] The propulsion system of one or more of these clauses, wherein the third stream flow path defines a third stream inlet, wherein the third stream inlet is a non-annular inlet.

[0139] A propulsion system according to one or more of these clauses, wherein the turbine comprises: a compressor section; a combustor, the combustor being fluidly connected to the compressor section and disposed downstream of the compressor section; a turbine section, the turbine section being fluidly connected to the combustor and disposed downstream of the combustor; a rotating shaft, the rotating shaft being capable of rotating with the compressor section, the turbine section, or both; and an exhaust section, the exhaust section being fluidly connected to the turbine section and disposed downstream of the turbine section.

[0140] A propulsion system defining a radial direction and an axial direction, the propulsion system comprising: a housing defining an inlet and a third stream flow path; a ductless fan defining a fan axis, wherein the ductless fan is arranged to rotate about the fan axis relative to the housing, wherein the fan axis is offset from the inlet along the radial direction; and a turbine, the turbine being arranged downstream of the ductless fan along the radial direction, wherein the turbine defines a working gas flow path, wherein the propulsion system defines an inlet channel connected to the inlet fluid, wherein the inlet channel is configured to provide an inlet airflow to the working gas flow path and the third stream flow path, and wherein the third stream flow path at least bypasses a portion of the turbine.

[0141] The propulsion system of one or more of these clauses, wherein the ductless fan comprises a plurality of fan blades, the propulsion system further comprising a row of fixed guide vanes disposed downstream of the ductless fan and upstream of the inlet.

[0142] The propulsion system of one or more of these clauses, wherein the turbine includes a compressor section having a mid-fan, wherein the third stream flow path is configured to receive a portion of the inlet airflow from the inlet passage at a location downstream of the mid-fan.

[0143] A propulsion system according to one or more of these clauses, wherein the turbine includes a compressor section having a mid-fan, wherein a third flow flow path defines a third flow inlet having a radial height, wherein the mid-fan defines a blade span, and wherein the radial height of the third flow inlet is at least 5% of the blade span of the mid-fan and up to 50% of the blade span of the mid-fan.

[0144] A propulsion system according to one or more of these clauses, wherein the turbine includes a compressor section having a mid-fan, wherein the propulsion system includes an inlet surface that partially defines the inlet passage, wherein the inlet surface defines a curvature including an inflection point, the inflection point being disposed at a point along the curvature.

[0145] A propulsion system according to one or more of these clauses, wherein the curvature of the inlet surface includes a first portion and a second portion, wherein the first portion is concave, wherein the second portion is concave, and wherein the inflection point is located at a transition point between the concave first portion and the concave second portion.

[0146] The propulsion system of one or more of these clauses, wherein the inlet defines an inlet centerline axis, wherein the inlet centerline axis of the inlet is not coaxial with the fan axis.

[0147] The propulsion system of one or more of these clauses, wherein the inlet is a single inlet or a plurality of inlets that do not circumscribe or surround the fan axis.

[0148] A propulsion system according to one or more of these clauses, wherein the inlet is a single inlet or a plurality of inlets, the inlet being proximal to the axis and distal to the fan axis.

Claims

1. A propulsion system, characterized in that: include: a ductless rotary fan, the ductless rotary fan defining a fan axis; and a turbine disposed downstream of the unducted rotary fan, wherein the turbine defines a working gas flow path therethrough; wherein the propulsion system defines a third flow path and an inlet passage, the inlet passage having an inlet offset from the fan axis, wherein the inlet passage is configured to provide an inlet airflow to the working gas flow path, wherein the third flow path bypasses at least a portion of the turbine, wherein the inlet of the inlet passage defines a non-annular shape, wherein the turbine includes a mid-fan, wherein the third flow path defines a third flow inlet having a radial height, wherein the mid-fan defines a blade span, and wherein the radial height of the third flow inlet is at least 5% of the blade span of the mid-fan and up to 50% of the blade span of the mid-fan.

2. The propulsion system according to claim 1, characterized in that: The propulsion system further comprises: a gear box operably coupled to the ductless rotary fan; and A rotating shaft extends from the turbine and is connected to the gearbox, wherein the gearbox is operably coupled to the turbine via the rotating shaft, wherein the rotating shaft defines an axis.

3. The propulsion system according to claim 2, characterized in that: The rotation axis is offset from the fan axis such that the rotation axis is disposed at a first distance from the fan axis.

4. The propulsion system according to claim 1, characterized in that: The turbomachine includes a compressor section having the mid-fan with mid-fan blades, wherein the propulsion system includes an inlet surface that partially defines the inlet passage, wherein the inlet surface defines a curvature including an inflection point disposed at a point along the curvature.

5. The propulsion system according to claim 1, characterized in that: wherein the propulsion system comprises an inlet surface that partially defines the inlet passage, wherein the inlet surface defines a curvature, wherein the curvature comprises an inflection point, wherein the curvature of the inlet surface comprises a first portion and a second portion, wherein the first portion is concave upward, wherein the second portion is concave downward, wherein the inflection point is located at a transition point between the concave first portion and the concave second portion.

6. The propulsion system according to claim 1, characterized in that wherein the propulsion system defines a forward direction and a rearward direction, wherein the ductless rotary fan is configured to propel fan flow in the rearward direction, and wherein at least a portion of the turbine is disposed in a counterflow arrangement such that the turbine is configured to direct a portion of the working gas flow path in the forward direction.

7. The propulsion system according to claim 1, characterized in that It further includes a row of fixed guide vanes, which are arranged downstream of the ductless rotary fan and upstream of the turbine.

8. The propulsion system according to claim 1, characterized in that The turbomachine includes a compressor section having the mid-fan, wherein the third stream flow path is configured to receive a portion of the inlet airflow from the inlet passage at a location downstream of the mid-fan.

9. The propulsion system according to claim 1, characterized in that The turbomachine includes a compressor section having the mid-fan and a high-pressure compressor, wherein the third stream flow path is configured to receive a portion of the inlet airflow from the inlet passage at a location downstream of the mid-fan and upstream of the high-pressure compressor.

10. The propulsion system according to claim 9, characterized in that Wherein the third stream flow path defines a third stream inlet, wherein the third stream inlet is a non-annular inlet.

11. The propulsion system according to claim 1, characterized in that The turbine comprises: compressor section; a combustor in fluid communication with the compressor section and disposed downstream of the compressor section; a turbine section in fluid communication with the combustor and disposed downstream of the combustor; a rotating shaft capable of rotating with the compressor section, the turbine section, or both; and An exhaust section is in fluid communication with the turbine section and is disposed downstream of the turbine section.

12. A propulsion system defining a radial direction and an axial direction, characterized in that: The propulsion system comprises: a housing defining an inlet and a third stream flow path; a ductless fan defining a fan axis, wherein the ductless fan is configured to rotate relative to the housing about the fan axis, wherein the fan axis is offset from the inlet in the radial direction; and A turbine, the turbine being disposed downstream of the ductless fan in the radial direction, wherein the turbine defines a working gas flow path, wherein the propulsion system defines an inlet passage connected to the inlet fluid, wherein the inlet passage is configured to provide an inlet airflow to the working gas flow path and the third flow path, wherein the third flow path at least bypasses a portion of the turbine, wherein the inlet of the inlet passage defines a non-annular shape, wherein the turbine includes a mid-fan, wherein the third flow path defines a third flow inlet having a radial height, wherein the mid-fan defines a blade span, and wherein the radial height of the third flow inlet is at least 5% of the blade span of the mid-fan and up to 50% of the blade span of the mid-fan.

13. The propulsion system according to claim 12, characterized in that The ductless fan comprises a plurality of fan blades, and the propulsion system further comprises a row of fixed guide blades, wherein the row of fixed guide blades is arranged downstream of the ductless fan and upstream of the inlet.

14. The propulsion system according to claim 12, characterized in that The turbomachine includes a compressor section having the mid-fan, wherein the third stream flow path is configured to receive a portion of the inlet airflow from the inlet passage at a location downstream of the mid-fan.

15. The propulsion system according to claim 12, characterized in that The turbomachine includes a compressor section having the mid-fan, wherein the propulsion system includes an inlet surface that partially defines the inlet passage, wherein the inlet surface defines a curvature including an inflection point disposed at a point along the curvature.

16. The propulsion system according to claim 15, characterized in that The curvature of the inlet surface includes a first portion and a second portion, wherein the first portion is concave upward, wherein the second portion is concave downward, and wherein the inflection point is located at a transition point between the concave first portion and the concave second portion.

17. The propulsion system of claim 12, wherein: Wherein the inlet defines an inlet centerline axis, wherein the inlet centerline axis of the inlet is not coaxial with the fan axis.

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