Optimization of open rotor propulsion systems including unducted and ducted fans

Abstract

The present disclosure relates to an open rotor propulsion system (10) comprising: an unducted fan (20); an inlet duct (71); a ducted fan (40) positioned behind the unducted fan (20) within the inlet duct (71); an engine positioned aft of the inlet duct (71), comprising a shaft (25) connecting the fan drive turbine (60) and the ducted fan (40); and four bearings configured to support the shaft (25) with respect to the engine housing (49), in which the bearings include: a first bearing (61) positioned radially inward of the ducted fan (40); and a second bearing (62) located rearward of the first bearing (61) and including a center of gravity (G2) positioned axially between a first plane (P3) including the center of gravity of the low-pressure compressor (45) and a second plane (P4) including the center of gravity of the intermediate housing (51).

Description

Technical Field

[0001] The technology described in this article relates to open rotor and ducted propulsion systems, and specifically to the architecture used in such systems. This technology is particularly advantageous when applied to gas turbine engines used for propulsion in aircraft. Background Technology

[0002] Gas turbine engines employing an open rotor design are known. Turbofan engines operate on the principle that a central gas turbine core drives a bypass fan located radially within the engine nacelle and upstream of the core. This allows the fan to operate within a "duct" formed by the inner surface of the nacelle, but the air driven by the fan "bypasses" the central gas turbine core. In contrast, open rotor propulsion systems operate on the principle that the bypass fan is located outside the engine nacelle—in other words, "without a duct." This allows for the use of larger fan blades compared to turbofan engines, enabling the application of a larger volume of air and thus improving propulsive efficiency compared to conventional ducted engine designs.

[0003] Optimal performance has been achieved through an open rotor design with a fan provided by two counter-rotating rotor assemblies, each carrying an array of airfoil blades located outside the engine nacelle. As used herein, "counter-rotational relationship" means that the blades of the first rotor assembly and the blades of the second rotor assembly are arranged to rotate in opposite directions to each other. Typically, the blades of the first rotor assembly and the blades of the second rotor assembly are arranged to rotate in opposite directions about a common axis and are axially spaced along that axis. For example, the blades of the first rotor assembly and the corresponding blades of the second rotor assembly can be mounted coaxially and spaced apart, wherein the blades of the first rotor assembly are configured to rotate clockwise about the axis and the blades of the second rotor assembly are configured to rotate counterclockwise about the axis (or vice versa). Externally, the fan blades of the open rotor engine resemble the propeller blades of a conventional turboprop engine.

[0004] To reduce design complexity while achieving propulsion efficiency comparable to counter-rotating propulsion designs with significantly reduced weight and length, document US 2021 / 0108597 proposes a propulsion system comprising a fan section, which includes a rotating element, a stationary element, and an inlet located between the rotating and stationary elements, wherein the inlet radially passes through the interior of the stationary element. The inlet leads to an inlet duct, which contains a ducted fan having a rotation axis and multiple blades. Furthermore, the inlet duct is divided into a first duct and a second duct separate from the first duct. The rotating element of the fan section is driven by a fan-driven turbine section via a fan drive shaft.

[0005] However, ducted fans are large and heavy in the case of large diameters, making them vulnerable to external attacks. Therefore, ducted fans may become unbalanced.

[0006] Furthermore, the applicant discovered that the fan-driven turbine shaft of the propulsion system is in a supercritical state. In fact, the fan-driven turbine shaft exhibits at least one bending mode within its operating range and can resonate during stable flight phases. In some cases, the bending mode of the fan-driven turbine shaft can even overlap and couple with the bending mode of the unbalanced ducted fan (radial translation of the ducted fan blades). Summary of the Invention

[0007] The purpose of this application is to provide a propulsion system including a fan section comprising an open rotating element and a ducted fan, thereby providing improved dynamic response during stable flight phases, and more specifically, wherein the energy contribution of the ducted fan to the fan-driven turbine shaft and the amount of potential energy in the fan-driven turbine shaft are reduced.

[0008] To this end, according to a first aspect, this disclosure proposes an open rotor propulsion system comprising: The fan section includes a ductless rotating element and a non-rotating stationary element. The ductless rotating element includes a first array of fan airfoil blades. The front frame houses the inlet duct, which includes an inlet located between a ductless rotating element and a non-rotating fixed element. Ducted fan, the ducted fan is positioned at the rear of the ductless rotating element, located in the inlet duct, wherein the ducted fan includes a second array of fan airfoil blades; The core duct and fan duct extend from the rear of the ducted fan; An engine, positioned within a core duct and including a low-pressure compressor, a high-pressure compressor, an intermediate housing located between the low-pressure and high-pressure compressors, a fan-driven turbine, and a shaft connected to the fan-driven turbine and configured to drive a ducted fan; and Four bearings are configured to support the shaft relative to the engine housing, wherein the bearings include: The first bearing is connected to one of the front frame and the intermediate housing; and The second bearing is located behind the first bearing and includes a center of gravity, which is axially positioned between a first plane including the center of gravity of the low-pressure compressor and a second plane including the center of gravity of the intermediate housing.

[0009] Preferred, but not limiting, features of the open rotor propulsion system according to the first aspect are as follows: - The first bearing is positioned radially inside the ducted fan; -The first bearing is connected to the intermediate housing; - The center of gravity of the first bearing is axially positioned between the third plane and the fourth plane. The third plane intersects with the foremost point of the leading edge of the ductless fan, and the fourth plane intersects with the lowest point of the trailing edge of the ductless fan. - The first bearing is located in front of the ducted fan; - The first bearing includes a roller bearing, and the second bearing includes a ball bearing; - The second bearing includes a squirrel cage and a squeeze diaphragm damper; - The center of gravity of the second bearing extends between the fifth plane and the second plane, the fifth plane including the center of gravity of the rotating components of the last stage of the low-pressure compressor; - The open rotor propulsion system further includes: a first support member that connects a first bearing to one of an intermediate housing and an inlet housing; and a second support member that connects a second bearing to an engine housing, optionally connecting the second bearing to the intermediate housing, wherein the radial stiffness of the first support member is at least twice that of the radial stiffness of the second support member. - The bearing also includes: a third bearing, which is located radially inside the fan drive turbine; and a fourth bearing, which is located behind the third bearing; - The center of gravity of the third bearing extends axially between the sixth and seventh planes. The sixth plane intersects with the center of gravity of the fan-driven turbine, and the seventh plane intersects with the rotating component of the fan-driven turbine located in front of the sixth plane. - The open rotor propulsion system includes: a turbine rear frame located at the rear of the fan-driven turbine, wherein a fourth bearing extends radially inward on the turbine rear frame. - The open rotor propulsion system also includes: a support structure that connects the third and fourth bearings to the turbine rear frame; - The open rotor propulsion system also includes: a high-pressure turbine that drives a high-pressure compressor; and a turbine housing that extends between the high-pressure turbine and the fan-driven turbine, wherein the center of gravity of the third bearing extends axially between an eighth plane and a ninth plane, the eighth plane intersecting the front portion of the turbine housing and the ninth plane intersecting the rotating components of the first stage of the fan-driven turbine. - The center of gravity of the fourth bearing extends between the sixth and tenth planes. The sixth plane intersects with the center of gravity of the fan-driven turbine, and the tenth plane intersects with the foremost part of the turbine's rear frame. - The open rotor propulsion system also includes: a third support member that connects the third bearing to the turbine housing; and a fourth support member that connects the fourth bearing to the turbine rear frame; - The third bearing includes a squeeze diaphragm damper; and / or - Between the first and second bearings, the shaft has a nominal thickness, while below the fan drive turbine and up to the fourth bearing, the shaft thickness is 40% to 60% greater than the nominal thickness.

[0010] According to the second aspect, an aircraft is proposed, the aircraft including at least one open rotor propulsion system according to the first aspect, wherein the open rotor propulsion system is connected to the aircraft via a pylon.

[0011] According to a third aspect, this disclosure proposes a method for disassembling an open rotor propulsion system, the open rotor propulsion system being based on the open rotor propulsion system of the first aspect, the method comprising the following steps: - Remove the first module, which includes the fan-driven turbine and shaft, from the open rotor propulsion system; -Remove the second module, which includes the high-pressure compressor, combustion chamber, and high-pressure turbine; - Remove the third module, which includes the fan section, inlet housing, ducted fan, and low-pressure compressor.

[0012] In one embodiment, the method further includes removing the second module by detaching the shaft from the shaft connected to the low-pressure compressor.

[0013] The first bearing may be connected to the inlet housing, and the step of removing the third module may include the following sub-steps: - Remove the fourth module, which includes the ducted fan and low-pressure compressor; and - Remove the fifth module, which includes the fan section and the inlet housing.

[0014] This disclosure applies to propulsion systems, for example, whose redline (maximum speed achievable by the propulsion system) is between 8,000 rpm and 15,000 rpm, for example, about 10,000 rpm. Attached Figure Description

[0015] Other features, objects, and advantages of this disclosure will become apparent when read in the following detailed description with reference to the accompanying drawings, which are given by way of non-limiting example and in which: Figure 1 This is a cross-sectional schematic diagram of an exemplary embodiment of an open rotor propulsion system; Figure 2 This refers to an exemplary aircraft that includes an open rotor propulsion system; Figure 3a This is a schematic diagram of a ducted fan, which shows an exemplary construction of a first embodiment based on a first bearing and a second bearing for driving the turbine shaft of the fan; Figure 3b This is a schematic diagram of a ducted fan, which shows an exemplary construction of a second embodiment based on a first bearing and a second bearing driving the turbine shaft of the fan; Figure 4 This is a schematic diagram of an open rotor propulsion system, in which a first exemplary configuration of the third and fourth bearings of the fan-driven turbine shaft is shown. Figure 5 This is a schematic diagram of an open rotor propulsion system, in which a second exemplary configuration of the third and fourth bearings of the fan-driven turbine shaft is shown; Figure 6 This is a schematic diagram of an example of an open rotor propulsion system fixed to a pylon of an aircraft according to an embodiment; and Figure 7 The steps of an example of a method for disassembling an open rotor propulsion system according to an embodiment are shown.

[0016] The same reference numerals in the specification and drawings are used to refer to the same parts of this disclosure. Detailed Implementation

[0017] Figure 1 A front cross-sectional view of an exemplary embodiment of an open rotor propulsion system 10 for propulsing an aircraft 100 is shown. (See from...) Figure 1 As seen, the open rotor propulsion system 10 includes a fan section comprising a rotating element 20, which includes an array of fan airfoil blades 21 surrounding a central longitudinal axis 11 of the open rotor propulsion system 10. The blades 21 are generally arranged at equal intervals around the centerline 11, and each blade 21 has a root 23, a tip 24, a span defined between the root 23 and the tip 24, and a central blade axis 22. The open rotor propulsion system 10 includes a gas turbine engine having a low-pressure (LP) compressor 45 or a supercharger, a gas turbine core, and a low-pressure (LP) turbine 60. The gas turbine core includes a high-pressure (HP) compressor 27, a combustor 28, and a high-pressure (HP) turbine 29 in a series flow relationship. A high-pressure (HP) shaft 26 enables the HP turbine 29 to drive the HP compressor 27. A low-pressure (LP) shaft 25 enables the LP turbine 60 to drive the rotating element 20 and the supercharger 45.

[0018] In this application, upstream and downstream are defined relative to the normal flow direction of gas through the propulsion system 10. Furthermore, the axial direction corresponds to the direction of the longitudinal axis 11, and the radial direction is the direction perpendicular to and through the axis 11. The reference radial directions are used for internal and external, respectively, such that the internal portions or surfaces of the element are closer to the axis 11 than the external portions or surfaces of the same element. Finally, all planes defined herein are perpendicular to the longitudinal axis 11.

[0019] exist Figure 1In an exemplary embodiment, the fan section of the open rotor propulsion system 10 further includes a non-rotating stationary element 30, which comprises an array of blades 31 also arranged around a central axis 11, each blade 31 having a root 33, a tip 34, and a span defined between the root 33 and the tip 34. These blades 31 may be arranged such that they are not all equidistant from the rotating components and are without shields (e.g., Figure 1 (as shown), or optionally may include an annular shroud or conduit remote from axis 11 (axis 11 is in Figure 1 (As shown in the diagram). These blades are mounted to a fixed frame and do not rotate relative to the central axis 11, but may include mechanisms for adjusting the orientation of these blades relative to their axis 35 and / or relative to blade 21. For reference purposes, Figure 1 It also depicts the forward direction, indicated by arrow F, thus defining the front and rear portions of the system. For example... Figure 1 As shown, the rotating element 20 is located in front of the gas turbine core in a "puller" configuration, and the exhaust portion 80 is located behind the stationary element 30.

[0020] The open rotor propulsion system 10 may also include a power gearbox 12, which may include a gear set for reducing the rotational speed of the rotating element 20 relative to the low-pressure turbine 60. The blades 21 of the open ductless rotating element 20 may have a fixed pitch or blade angle, or alternatively may have a variable pitch or blade angle to change thrust and blade load during operation, and in some configurations, a reverse thrust configuration is provided to decelerate the aircraft upon landing.

[0021] An annular 360-degree inlet 70 is located between the rotating element 20 and the stationary or stationary element 30 and includes blades 36 that receive a path that allows incoming air to enter the gas turbine core located radially inside the stationary element 30. This location may be advantageous for various reasons, including the management of icing performance and the protection of the inlet 70 from various objects and materials that may be encountered during operation.

[0022] Figure 1 A configuration that may be referred to as a "puller" configuration is shown, wherein the rotating element 20 that generates thrust is located in front of the gas turbine core. Other configurations, such as embodiments that may be referred to as a "pusher" configuration, are possible and contemplated within the scope of this disclosure, wherein the gas turbine core is located in front of the rotating element 20. For example, various architectures are shown and described in publications WO 2022 / 069834, WO 2022 / 018380, and US 2013 / 0098066.

[0023] In addition to the open rotor or ductless rotating element 20 having multiple fan airfoil blades 21, a ducted fan 40 is included at the rear of the open rotor rotating element 20, such that the open rotor propulsion system 10 includes both a ducted fan and a ductless fan, both used to generate thrust by the movement of air at atmospheric temperature (but not through the gas turbine core). The ducted fan 40 is shown located at approximately the same axial position as the blade 31 and radially inward of the blade root 33. Alternatively, the ducted fan 40 may be located between the blade 31 and the core duct 72, or further forward of the blade 31. The ducted fan 40 may be driven by the low-pressure turbine 60 or any other suitable rotating source and may be used as the first stage of the supercharger 45, or may operate independently.

[0024] The ducted fan 40 includes an array of fan airfoil blades 41 surrounding the central longitudinal axis 11 of the open rotor propulsion system 10. The blades 41 are generally arranged in a relationship of equal spacing around the centerline 11, and each blade 41 has a root 43, a tip 44, and a span defined between the root 43 and the tip 44.

[0025] In the following description, the invention will be illustrated with reference to a ducted fan 40 driven by a low-pressure turbine 60. Furthermore, this same low-pressure turbine 60 drives an unducted rotating element 20 via a gearbox 12. However, in one embodiment, the unducted rotating element 20 may be driven by any other suitable rotational source.

[0026] Air entering the gas turbine core flows through the inlet duct 71 and then splits behind the ducted fan 40, such that a portion of the air flows through the core duct 72 and a portion flows through the fan duct 73. The fan duct 73 may contain a heat exchanger 74 and is discharged to the atmosphere through independent fixed or variable nozzles 75 located behind the stationary element 30 and outside the gas generator core shroud 721. Thus, the air flowing through the fan duct 73 “bypasses” the engine core without passing through it. Therefore, the open rotor propulsion system 10 includes a ductless fan formed by the rotating element 20, followed by a ducted fan 40 that directs airflow into two concentric or non-concentric ducts 72 and 73, thereby forming a three-flow engine architecture with three paths for the air to pass through the rotating element 20.

[0027] Because the open rotor propulsion system 10 includes an open rotor rotating assembly 20 and a ducted fan assembly 40, the thrust output of these two assemblies and the workload sharing between them can be customized to achieve specific thrust, fuel combustion, thermal management, and acoustic characteristic targets that are superior to those achieved by typical ducted fan gas turbine propulsion assemblies with comparable thrust levels. By reducing the proportion of thrust required from the ductless fan assembly 20, the ducted fan assembly 40 allows for a smaller overall fan diameter for the ductless fan assembly, thus providing installation flexibility and weight reduction.

[0028] Operationally, the open rotor propulsion system 10 may include a control system that manages the loads on the respective open fan and ducted fan, and potentially manages the exit region of the variable fan nozzles to provide different thrust, noise, cooling capacity, and other performance characteristics for different parts of the flight envelope and various operating conditions associated with the aircraft's operation. For example, in climb mode, the ducted fan can operate at its maximum pressure ratio to maximize the thrust of the flow, while in cruise mode, the ducted fan can operate at a lower pressure ratio to improve overall efficiency by relying on thrust from the ductless fan. Nozzle actuation adjusts the operating line of the ducted fan and the overall engine fan pressure ratio, independent of the overall engine airflow.

[0029] The open rotor propulsion system 10 includes an engine housing 49 that houses a gas turbine engine located behind a ducted fan 40. The engine housing 49 includes a front frame 78 that extends at the level of the inlet 70 behind the ductless fan 20. More specifically, the front frame 78 extends between the ductless fan 20 and the ducted fan 20 and includes blades 36 facing a path that allows incoming air to enter the gas turbine core located radially inward of a stationary element 30. Thus, the inlet duct 71 is housed within the front frame 78. The front frame 78 is a structural component of the propulsion system 10 and is supported by bearings of the ductless fan 20. In one embodiment, a gearbox 12 may be attached to the front frame 78, optionally suspended to the front housing.

[0030] The engine housing 49 also includes an OGV housing 79 (i.e., an open guide vane housing) located at the rear of the front frame 78, which houses the ducted fan 40. The OGV housing 79 is a structural housing that supports the mounting element 30. Note that the fan duct 73 is contained within the OGV housing 79.

[0031] The engine housing 79 also includes, in a series flow relationship, a supercharger housing 50 housing the supercharger 45, an intermediate housing 51, a high-pressure housing 52 housing the gas turbine core, a turbine housing 53, a turbine housing 54, and a turbine rear frame 55. The intermediate housing 51 (or intermediate frame) extends between the supercharger housing 50 and the high-pressure housing 52 and is supported by at least one front bearing 61, 62 of the LP shaft 25. The intermediate housing 51 is also a structural housing of the propulsion system 10. In one embodiment, the OGV housing 79 is configured to be connected to the pylon 37 of the aircraft 100 and structurally supported by the intermediate housing 51.

[0032] It should be noted here that in this application, "structure" is used to define the housing or frame of the engine, which is configured to transmit the loads of the propulsion system 10. In other words, the structural frame or housing is the housing through which the axial and radial forces of the propulsion system 10 are transmitted (e.g., the loads of the bearings supporting the shaft are transmitted via the structural housing toward the engine's suspension (e.g., pylon 37)). As previously described, the front frame 78, OGV housing 79, and intermediate housing 51 are structural housings. The front frame 78 is configured to receive thrust generated by the ductless fan 20 and transmit that thrust to the aircraft via pylon 37. The intermediate housing 51 is configured to receive thrust generated by the ducted fan 40 and transmit that thrust to the aircraft via the OGV housing and pylon 37. Conversely, the supercharger housing 50 and high-pressure housing 52 support the corresponding compressor sections and define the flow paths within those compressor sections; however, these housings 50 and 52 are not structural housings in the sense of this application.

[0033] As previously described, the LP shaft 25 is connected to and driven by the LP turbine rotor 60, and drives the supercharger 45, the ducted fan 40, and the unducted fan 40. The LP shaft 25 drives the supercharger 45 via the connector 76.

[0034] The LP shaft 25 is connected to the engine housing 49 of the gas turbine engine via exactly four bearings 61, 62, 63, and 64. The first bearing 61 and the second bearing 62 support the front portion of the shaft 25. The third bearing 63 and the fourth bearing 64 support the rear portion of the shaft 25.

[0035] The first bearing 61 is connected to one of the front frame 78 and the intermediate housing 51 and is positioned near the ducted fan 40.

[0036] The second bearing 62 is axially located behind the first bearing 61. Preferably, taking into account the available space below the supercharger housing 50 and the intermediate housing 51, and more specifically the connection 76 between the LP shaft 25 and the supercharger 45, the second bearing 62 is positioned as far away from the first bearing 61 as possible. The center of gravity G2 of the second bearing 62 is axially positioned between a first plane P3, which includes the center of gravity of the supercharger 45, and a second plane P4, which includes the center of gravity of the intermediate housing 51. In one embodiment, the center of gravity G2 of the second bearing 62 extends between plane P3' and the second plane P4, where plane P3' includes the center of gravity of the rotating component of the last stage of the supercharger 45. The second bearing 62 helps control the LP shaft mode position and the stability of the LP shaft 25.

[0037] It can be noted here that only the bearings 61, 62, 63, and 64 themselves (including the inner and outer rings) are considered to determine the center of gravity of a given bearing. Therefore, the support members that connect the outer rings of bearings 61, 62, 63, and 64 to the engine housing 49 are not part of the bearings 61, 62, 63, and 64 themselves. Furthermore, the front portion of the intermediate housing 51 corresponds to the front connecting flange of the intermediate housing 51, which connects to the rear portion of the turbocharger housing 50.

[0038] This structural configuration of the first bearing 61 and the second bearing 62 separates the bending mode of the LP shaft 25 from the suspension mode of the ducted fan 40, and shifts the first bending mode of the LP shaft 25 outside of transitional, unstable flight phases (e.g., climb phases) or operating speeds. Therefore, the LP shaft 25 is less sensitive to imbalances in the ducted fan 40, and the deformation of the LP shaft 25 (more specifically, the maximum bending of the LP shaft 25 at the level of the HP compressor) is reduced. Furthermore, since the first bearing 61 is arranged adjacent to the ducted fan 40, the static clearance (caused by the rotation and gravity of the LP shaft 25) is improved. In contrast, a structure including only one bearing would be much more complex to achieve the same advantages, as using two different bearings provides more degrees of freedom to separate the effects.

[0039] Furthermore, the farther the second bearing 62 is from the first bearing 61 (generally towards the second plane), the higher the rotational speed corresponding to the bending mode of the LP axis 25. However, it should be noted that the LP axis mode remains within an unstable speed range (beyond which resonance occurs). This is why the second bearing 62 should be positioned such that its center of gravity G2 is axially positioned between the first plane P3 and the second plane P4. The position of the second bearing 62 within this axial range can then be optimized to account for stability and resonance (depending on component weight, inertia, stiffness, etc. for each similar architecture).

[0040] In the first embodiment ( Figure 3aIn this configuration, a first bearing 61 is connected to an intermediate housing 51 and is positioned substantially below the ducted fan 40 (that is, radially inward of the ducted fan 40). More specifically, the center of gravity G1 of the first bearing 61 is axially positioned between planes P1 and P2, where plane P1 intersects the foremost point of the leading edge of the blade 41 of the ducted fan 40, and plane P2 intersects the lowermost point of the trailing edge of the blade 41. The leading edge extends in the opposite direction to the airflow entering the ducted fan 40. The leading edge corresponds to the front portion of the fan airfoil 41, facing the airflow and splitting it into a pressure flow and a suction flow. The trailing edge corresponds to the rear portion of the fan airfoil 41, where the pressure flow and suction flow meet. The first bearing 61 helps control the ducted fan suspension mode, the amount of clearance closure under maneuver, and the ducted fan imbalance.

[0041] Second embodiment ( Figure 3b In this configuration, the first bearing 61 is connected to the front frame 78 and positioned in front of the ducted fan 40. More specifically, the center of gravity G1 of the first bearing 61 is axially positioned between plane P1 and the plane intersecting with the trailing edge of the blade 36 of the front frame 78.

[0042] The first bearing 61 and the second bearing 62 are connected to the engine housing 49 via a first support member 65 and a second support member 66, respectively. The first support member 65 and the second support member 66 are each fixed to the outer ring of the corresponding bearing and may have a truncated conical shape. In the first embodiment described above, the first support member 65 is connected to the intermediate housing 51, while in the second embodiment, the first support member 65 is connected to the front frame 78.

[0043] To further separate the bending pattern from the ducted fan 40 from the LP shaft 25, the radial stiffness of the first support 65 is at least twice that of the second support 66. For example, the radial stiffness of the first support 65 can be between 2 and 5 × 10⁻⁶. -9 The ratio of m / N is between 5 and 10, and the radial stiffness of the second support member 66 can be between 5 and 10*10. -9 Between m / N.

[0044] It can be noted here that the radial stiffness of the first support 65 and the second support 66 is inherently defined, that is, inherently defined by considering the support itself outside the propulsion system 10. Therefore, the value of the radial stiffness is an absolute rather than a relative value, and does not depend on the environment in which the radial stiffness is measured, making it possible to integrate these supports into any propulsion system 10. The radial stiffness of a given support 65, 66 can be determined by wrapping one end of the support 65, 66, allowing the other end to move freely radially. This configuration actually reflects the configuration of the supports 65, 66 in the propulsion system 10 (the wrapped end corresponds to the end connected to the engine housing 49, and the free end corresponds to the end connected to the corresponding bearing). The stiffness of the support 65, 66 is then defined as the ratio between the radial force applied to the free end and the radial displacement of the free end relative to the wrapped end generated by that radial force.

[0045] The second support member 66 can be connected to the intermediate housing 51. For example, in the first embodiment, the first support member 65 and the second support member 66 can be connected to the same attachment of the intermediate housing 51. It can be noted that the position of the second bearing 26 allows the front portion of the engine to deform. In fact, during operation, the ductless fan 20 pulls on the intermediate housing 51, creating relative movement between the ductless fan 20 and the stationary element 30, which creates a gap. Therefore, by positioning the second bearing 62 between planes P3 and P4 and by connecting the support member 62 to the intermediate housing 51, the support member 66 follows the stationary element 30 and reduces the gap in the front portion.

[0046] The first bearing 61 may include a roller bearing, and the second bearing 62 may include a ball bearing, to improve the dynamic behavior of the front portion of the shaft 25.

[0047] Optionally, the second bearing 62 may include a soft squirrel cage and a squeeze film damper to significantly reduce the impact of the first bending mode of the LP shaft on the architecture. However, for the first bearing 61, a squirrel cage and a squeeze film damper are not required.

[0048] The third bearing 63 is located substantially below the LP turbine 60. The fourth bearing 64 is located behind the third bearing 63.

[0049] The following are references respectively. Figure 4 and Figure 5 Two embodiments for positioning the third bearing 63 and the fourth bearing 64 are disclosed. In both embodiments, the third bearing 63 and the fourth bearing 64 are spaced apart to reduce static clearance under operation. For example, the third bearing 63 and the fourth bearing 64 may be spaced apart by a distance of at least 100 mm and at most 500 mm.

[0050] In the first embodiment (see) Figure 4In this configuration, the center of gravity G3 of the third bearing 63 is axially located near the downstream plane P6, which includes the center of gravity of the LP turbine rotor (i.e., the rotating blades of the LP turbine, but excluding the stator blades and the portion of the LP shaft connected to the rotating blades of the LP turbine rotor). For example, to account for component assembly constraints and optimize clearance closure and dynamic modes / stability, the center of gravity G3 of the third bearing 63 can extend axially between the upstream plane P5 and the downstream plane P5', where the upstream plane P5 intersects the foremost point of the stage of the LP turbine 60 located in front of the downstream plane P6, and the downstream plane P5' intersects the last point of the stage of the LP turbine 60 located behind the downstream plane P6. For example, when the center of gravity of the 4-stage LP turbine 60 intersects with the third stage of the LP turbine 60 (see... Figure 4 When the third bearing 63 is axially positioned below the rotating components of the second stage of the LP turbine 60, it can be noted that the position of the third bearing 63 can be limited by the connection 77 between the LP shaft 25 and the rotor of the LP turbine 60, which can be located downstream of the upstream plane P5. In this case, the third bearing 63 can be axially positioned upstream of the center of gravity of the LP turbine 60, up to the connection 77 between the LP shaft 25 and the rotor of the LP turbine 60.

[0051] This construction of the third bearing 63 reduces the influence of the LP shaft 25 on the length portion of the LP shaft pattern, allowing the length of the LP shaft located between the second bearing 62 and the third bearing 63 to be reduced. Therefore, the rotational speed of the LP shaft 25 can be increased, thus reducing the theoretically unstable speed range of the LP shaft 25. Positioning the third bearing 63 radially inside the center of gravity of the LP turbine 60 also reduces the static clearance caused by gravity.

[0052] Preferably, taking into account the available space below the LP turbine 60 housing and the turbine rear frame 55, the fourth bearing 64 should be positioned as far away from the third bearing 63 as possible. For example, the fourth bearing 64 may extend below the turbine rear frame 55, for instance, in the plane P7 containing the center of gravity of the turbine rear frame 55 or even downstream of that plane P7.

[0053] The third bearing 63 affects the LP shaft mode and stability, and helps control the clearance closure under inertial loads. The fourth bearing 61 helps control the clearance closure under dynamic loads.

[0054] The third bearing 63 and the fourth bearing 64 are connected to the engine housing 49 via corresponding supports 67 and 68, each of which is connected to the turbine rear frame 55. Optionally, supports 67 and 68 can both be connected to the same attachments on the turbine rear frame 55.

[0055] Second embodiment ( Figure 5In this configuration, the center of gravity G3 of the third bearing extends axially between planes P8 and P9, with plane P8 intersecting the front portion of the turbine housing 54 and plane P9 intersecting the rotating component of the first stage of the LP turbine 60. Therefore, the third bearing 63 can extend upstream of the connection 77 between the LP shaft 25 and the rotor of the LP turbine 60. In this configuration, the third bearing can be connected to the turbine housing 53. Thus, the constraints associated with the connection between the third bearing 63 and the turbine rear frame 55 are removed. Furthermore, this configuration of the third bearing 63 reduces the length of the portion of the LP shaft 25 that influences the LP shaft pattern, allowing for a reduction in the LP shaft length between the second bearing 62 and the third bearing 63. Reducing the unstable speed range (between the LP shaft pattern and the maximum speed range of the LP rotor) is also made easier due to the reduced distance between bearings 62 and 63.

[0056] In this configuration, the third bearing 63 helps control the LP shaft mode position and instability.

[0057] Here, the fourth bearing 64 is again positioned as far away from the third bearing 63 as possible. However, the fourth bearing 64 should not extend too far below the turbine rear frame 56 to limit the loads applied to the third bearing 63 and the fourth bearing 64. For example, the center of gravity G4 of the fourth bearing 64 can extend axially between plane P3 and plane P10, where plane P3 includes the center of gravity of the LP turbine 60 and plane P10 intersects the foremost portion of the turbine rear frame 56.

[0058] The position of the fourth bearing 64 helps control unbalanced loads and clearance closure (caused by static clearance and imbalance under inertial loads).

[0059] In this embodiment, the support 67 of the third bearing 63 can be connected to the turbine housing 54, and the support 68 of the fourth bearing 64 can be connected to the turbine rear frame 56.

[0060] In both embodiments, the third bearing 63 may include a squirrel cage and a squeeze diaphragm damper (not shown) to further reduce the amplitude of the bending mode of the LP shaft 25 by suppressing the load applied to the bearing and limiting clearance consumption. The squirrel cage and squeeze diaphragm will also suppress the resonance amplitude and avoid potential instability of the LP shaft.

[0061] Furthermore, the thickness of the LP shaft 25 may increase along the LP turbine 60 to reduce the deformation energy by transferring the deformation energy in the LP shaft 25 to the engine housing 49, and thus prevent the deformation energy from being contained within the LP shaft 25. For example, the LP shaft 25 may have a nominal thickness T0, which corresponds to the thickness of the shaft 25 between the first bearing 61 and the second bearing 62. Then, the thickness T1 of the LP along the LP turbine 60 is 50% greater than the nominal thickness T0. For example, the thickness of the LP shaft may increase along the entire axial length of the LP turbine 60, that is, from the plane intersecting with the first rotating component of the LP turbine 60 to the plane intersecting with the last stationary element of the LP turbine 60. Optionally, in the first embodiment of the third bearing 63 and the fourth bearing 64, the thickness may increase up to the fourth bearing 64.

[0062] For example, the thickness of the LP shaft 25, whose nominal thickness T0 is approximately 35 mm, can be increased by about 50%.

[0063] It can be noted that even though the architecture may be more sensitive to aerodynamic loads due to the support of the third bearing 63 and the fourth bearing 64, the rotor clearance still increases. However, the positions of bearings 61, 62, 63, and 64 are a compromise that reduces rotor clearance by reinforcing the shaft and rotor, regardless of the forces received by the ducted fan 40, which are caused by the non-uniformity of the air entering the inlet duct 71, the impact received by the unducted fan 20, and the magnitude of the maneuvering and bending patterns of the LP shaft 25.

[0064] Regarding the exemplary embodiments described herein, one or more gearboxes may be employed. The gearbox located between the ductless fan and the ducted fan may have a ratio between approximately 2:1 and 12:1, and the second gearbox located between the supercharger and the ducted fan is the same as in the case of a conventional HP turbocharger. Another configuration may utilize two counter-rotating turbines with their two LP shafts 25 extending forward, or counter-rotating turbines driving a common LP shaft 25 extending forward. In each case, the rear gearbox may have a ratio between approximately 2:1 and 5:1.

[0065] The total desired thrust generated by the gas turbine engine can be varied as needed to suit the operating conditions and flight envelope of the relevant aircraft. The maximum design workload distribution between the ducted and unducted fans can also be varied as needed; for example, in some embodiments, the ducted fan can generate up to about 60% of the total thrust to minimize engine diameter or noise, while in another embodiment, if the primary function of the ducted fan flow is as a radiator, the ducted fan can generate only a few percent of the total thrust. The pressure ratio of the subsequent ducted fan can be less than about 2.5.

[0066] The propulsion system 10 can be formed from several modules, which can be disassembled in blocks for inspection and maintenance. The propulsion system may include: - First module, which includes a fan-driven turbine 60 and an LP shaft 25; - The second module includes a high-pressure section, which includes a high-pressure compressor 27, a combustion chamber, and a high-pressure turbine; - The third module includes a fan section, an inlet housing 70, a ducted fan 40, and a low-pressure compressor.

[0067] Disassembly of the propulsion system 10 may include the following steps: - Remove the first module, including the fan-driven turbine 60 and the shaft 25, from the open rotor propulsion system 10; -Remove the second module, which includes the high-pressure compressor 27, combustion chamber, and high-pressure turbine; - Remove the third module, which includes the fan section, front frame 78, ducted fan 40, and low-pressure compressor 45.

[0068] For this purpose, the LP shaft includes a first section connected to the rotor of the LP turbine and a second section connected to the LP compressor. The first and second sections of the LP shaft 25 are connected by fasteners such as splines and nuts. During disassembly, the first module is first removed by pulling the first module away from the second module. Then, the second module is removed from the third module by disconnecting the first and second sections of the LP shaft 25 and pulling the second module away from the third module. Each module can then be inspected, repaired, and / or replaced individually.

[0069] When the first bearing 61 is connected to the intermediate housing 51, the third module can be divided into two sub-modules, namely: - First submodule, the first submodule includes a ducted fan 40 and a low-pressure compressor 45; and - Second submodule, which includes the fan section and front frame 78.

[0070] In fact, the connection between the first support member 65 and the front frame allows the first submodule and the second submodule to be separated without opening the housing containing the first bearing 61.

[0071] The first and second submodules can be connected to each other by means of splines and nuts.

[0072] During disassembly, the first and second sub-modules can be disconnected after the third module is separated from the second module.

Claims

1. An open rotor propulsion system (10), comprising: The fan section includes a ductless rotating element (20) and a non-rotating fixed element (30), the ductless rotating element including a first array of fan airfoil blades (21); A front frame (78) that houses an inlet duct (71) that includes an inlet located between the ductless rotating element (20) and the non-rotating fixed element (30); Ducted fan (40), the ducted fan being positioned at the rear of the ductless rotating element (20) and within the inlet duct (71), wherein the ducted fan (40) includes a second array of fan airfoil blades (41); The core duct (72) and fan duct (73) extend from the rear of the ducted fan (40). An engine, located within the core duct (72) and including a low-pressure compressor (45), a high-pressure compressor (27), an intermediate housing (51) located between the low-pressure compressor (45) and the high-pressure compressor (27), a fan-driven turbine (60), and a shaft (25) connected to the fan-driven turbine (60) and configured to drive the ducted fan (40); and Four bearings configured to support the shaft (25) relative to the engine housing (49), wherein the bearings include: A first bearing (61) is connected to one of the front frame (78) and the intermediate housing (51); and The second bearing (62) is located behind the first bearing (61) and includes a center of gravity (G2), which is axially positioned between a first plane (P3) including the center of gravity of the low-pressure compressor (45) and a second plane (P4) including the center of gravity of the intermediate housing (51).

2. The open rotor propulsion system (10) according to claim 1, wherein, The first bearing (61) is located radially inside the ducted fan (40).

3. The open rotor propulsion system (10) according to claim 2, wherein, The first bearing (61) is connected to the intermediate housing (51).

4. The open rotor propulsion system (10) according to any one of claims 2 and 3, wherein, The center of gravity (G1) of the first bearing (61) is axially positioned between the third plane (P1) and the fourth plane (P2). The third plane intersects the foremost point of the leading edge of the ductless fan (40), and the fourth plane intersects the lowest point of the trailing edge of the ductless fan (40).

5. The open rotor propulsion system (10) according to claim 1, wherein, The first bearing (61) is located in front of the ducted fan (40).

6. The open rotor propulsion system (10) according to any one of claims 1 to 5, wherein, The first bearing (61) includes a roller bearing, and the second bearing (62) includes a ball bearing.

7. The open rotor propulsion system (10) according to claim 6, wherein, The second bearing (62) includes a squirrel cage and a squeeze diaphragm damper.

8. The open rotor propulsion system (10) according to any one of claims 1 to 7, wherein, The center of gravity (G2) of the second bearing (62) extends between the fifth plane (P3') and the second plane (P4), the fifth plane including the center of gravity of the rotating components of the last stage of the low-pressure compressor (45).

9. The open rotor propulsion system (10) according to any one of claims 1 to 8, further comprising: A first support member connects the first bearing (61) to one of the intermediate housing (51) and the inlet housing (70); And a second support (66) that connects the second bearing (62) to the engine housing (49), optionally connecting the second bearing to the intermediate housing (51), wherein the radial stiffness of the first support is at least twice that of the radial stiffness of the second support (66).

10. The open rotor propulsion system (10) according to any one of claims 1 to 9, wherein, The bearing also includes: A third bearing (63) is positioned radially inside the fan-driven turbine (60); and A fourth bearing (64) is positioned behind the third bearing (63).

11. The open rotor propulsion system (10) according to claim 10, wherein, The center of gravity (G3) of the third bearing (63) extends axially between the sixth plane (P6) and the seventh plane (P5), the sixth plane intersecting the center of gravity of the fan-driven turbine (60), and the seventh plane intersecting the rotating component of the fan-driven turbine (60) located in front of the sixth plane (P6).

12. The open rotor propulsion system (10) according to any one of claims 10 and 11, further comprising: A turbine rear frame (55) is located at the rear of the fan-driven turbine (60), wherein the fourth bearing (64) extends radially inward on the turbine rear frame (55).

13. The open rotor propulsion system (10) according to any one of claims 10 to 12, further comprising: Support members (67, 68) connect the third bearing (63) and the fourth bearing (64) to the turbine rear frame (55).

14. The open rotor propulsion system (10) according to any one of claims 10 to 13, further comprising: A high-pressure turbine (29) that drives the high-pressure compressor (27); and a turbine housing (53) that extends between the high-pressure turbine (29) and the fan-driven turbine (60), wherein the center of gravity (G3) of the third bearing (63) extends axially between an eighth plane (P8) and a ninth plane (P9), the eighth plane intersecting the front portion of the turbine housing (53) and the ninth plane intersecting the rotating component of the first stage of the fan-driven turbine (60).

15. The open rotor propulsion system (10) according to any one of claims 10 to 14, wherein, The center of gravity (G4) of the fourth bearing (64) extends between the sixth plane (P6) and the tenth plane (P10), the sixth plane intersecting the center of gravity of the fan-driven turbine (60), and the tenth plane intersecting the foremost part of the turbine rear frame (55).

16. The open rotor propulsion system (10) according to any one of claims 14 or 15, further comprising: A third support member (67) connects the third bearing (63) to the turbine housing (53). and a fourth support member (68), which connects the fourth bearing (64) to the turbine rear frame (55).

17. The open rotor propulsion system according to any one of claims 10 to 16, wherein, The third bearing (63) includes a squeeze film damper.

18. The open rotor propulsion system (10) according to any one of claims 10 to 17, wherein, Between the first bearing and the second bearing (61, 622), the shaft (25) has a nominal thickness (T0), while below the fan drive turbine (60) and up to the fourth bearing (64), the thickness (T1) of the shaft (25) is 40% to 60% greater than the nominal thickness.

19. An aircraft comprising at least one open rotor propulsion system according to any one of claims 1 to 18, wherein, The open rotor propulsion system is connected to the aircraft via a hangar (37).

20. A method for disassembling an open rotor propulsion system (10), said open rotor propulsion system being an open rotor propulsion system according to any one of claims 1 to 18, the method comprising the following steps: -Remove (S1) the first module comprising the fan-driven turbine (60) and the shaft (25) from the open rotor propulsion system (10); -Remove (S2) a second module including the high-pressure compressor (27), combustion chamber and high-pressure turbine; - Removal (S3) includes the third module comprising the fan section, the inlet housing (70), the ducted fan (40), and the low-pressure compressor.

21. The method of claim 20, further comprising: in, Removal (S2) of the second module includes detaching the shaft from the shaft connected to the low-pressure compressor.

22. The method according to any one of claims 20 and 21, wherein, The first bearing (61) is connected to the inlet housing (70), and the step of removing (S3) the third module includes the following sub-steps: -Remove (S4) the fourth module including the ducted fan (40) and the low-pressure compressor; and - Remove (S5) the fifth module including the fan section and the inlet housing (70).

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