Large aspect ratio rotor configuration with split turbine

Abstract

The application discloses a large-length-diameter-ratio rotor structure with a separated turbine, comprising a four-stage fan, a two-stage front turbine, a single-stage rear turbine, four bearing supporting points, a connecting structure and the like. A 1-2-1 four supporting point supporting scheme is adopted, and an intermediate supporting point is additionally arranged at the front turbine to control the bending deformation of the large-length-diameter-ratio rotor. Meanwhile, the two-stage front turbine in the middle section of the rotor is designed as an integrated ring cavity structure, the structure angular rigidity is enhanced, and the overall gyroscopic moment effect is exerted. The front turbine and the rear turbine are connected with the rotor shaft through a large-space double-centering-face stop mouth-toothed sleeve connection, the bending moment bearing capacity is improved, the gyroscopic moment of the wheel disc is fully applied to the shaft section, the rotor resonance speed safety margin is improved, and the bending deformation is controlled. The application is the first rotor structure proposed for a new engine with an inter-stage combustion chamber, and meets the dynamic design requirements of the rotor with the characteristics of the multi-stage separated turbine and the large-length-diameter ratio.

Description

Technical Field

[0001] This invention belongs to the field of aero-engine rotor structure design, specifically involving a rotor structure with a large length-to-diameter ratio and a split turbine. Background Technology

[0002] An aero engine is a complex and precise machine. During operation, air flows in through the intake, is compressed and pressurized by the compressor, and then enters the combustion chamber. There, it mixes with the injected fuel and burns. The resulting high-temperature, high-pressure gas flow expands through the turbine, driving the compressor, and finally exits through the exhaust nozzle to generate thrust, propelling the aircraft forward. However, after the gas has done work in the high-pressure turbine, its temperature and pressure decrease, and its work-capacity decreases as it enters the low-pressure turbine. To increase the work-capacity of the low-pressure turbine, multi-stage turbine designs are required. For example, high-bypass engines typically use multi-stage turbines to drive large-diameter fans. However, this significantly increases the engine's structural weight and reduces its structural efficiency.

[0003] In the exploration of future aero-engines, an interstage combustor scheme has been proposed to fully enhance the work capacity of the low-pressure turbine. This involves designing an interstage combustor after the high-pressure turbine outlet, where secondary combustion increases the temperature and pressure of the gas entering the low-pressure turbine, thereby improving the work capacity without increasing the number of stages. For example, Chinese patent application CN115451427A, "An Interstage Combustor and a Turbine Fan Engine Having the Same," designs a highly integrated, high-efficiency interstage combustor positioned between the high-pressure and low-pressure turbines to enhance engine thrust. This highly integrated interstage combustor does not alter the overall engine structural layout. While the axial length of the low-pressure rotor is increased, its length-to-diameter ratio can be appropriately increased to maintain a similar dynamic characteristic to that of a traditional engine's low-pressure rotor.

[0004] Highly integrated interstage combustion chambers achieve temperature and pressure increases within a limited range, enhancing the low-pressure turbine's work capacity to a certain extent without altering characteristics such as flow rate and bypass ratio. Chinese patent application CN114776473A, "A Variable Cycle Engine Configuration Based on a Multi-Bypass Inlet Interstage Combustion Chamber," proposes a variable cycle engine scheme with dual variable-stage combustion chambers. By controlling the intake of the bypass airflow, it achieves different modes such as turbojet and turbofan, significantly increasing the engine's flow rate and thrust adjustment capabilities. However, its unique configuration requires multiple turbine components, including a high-pressure turbine, an intermediate-pressure turbine, and a low-pressure turbine. Furthermore, the large-size dual variable-stage combustion chambers significantly increase the length of its low-pressure rotor, posing challenges to structural layout and dynamic design. Chinese patent application CN115405421B, "An Overall Structure of a Three-Spindle Variable Cycle Engine with an Interstage Combustion Chamber," proposes a three-spindle layout scheme. It employs variable stiffness supports for the low-pressure rotor, controlling the critical speed margin and deformation of the ultra-long-span low-pressure rotor by adjusting its support stiffness and designing the low-pressure turbine journal configuration. However, its three-rotor structure layout and variable stiffness support structure inevitably bring about structural complexity, making practical application difficult.

[0005] The turbines located before and after the interstage combustion chamber can be placed on the same rotor to reduce the number of rotors and lower structural complexity. This multi-stage turbine layout with separate front and rear turbines on the same rotor is an innovative layout (called a split turbine), and no similar rotor structure design has been found. The low-pressure rotor is longer and passes through the interior of the small-sized core engine rotor, limiting its diameter. How to design the rotor structure and its support scheme to effectively control rotor deformation and dynamic characteristics without increasing structural complexity is the difficulty and key to this type of rotor design with split turbines. Summary of the Invention

[0006] To address the aforementioned technical problems, this invention provides a rotor structure with a large length-to-diameter ratio and a split turbine. The aim is to achieve precise control of the engine rotor's critical speed and deformation when the length-to-diameter ratio of the low-pressure rotor cannot be reduced. This invention employs a 1-2-1 four-point support scheme, with an intermediate support point located near the front turbine in the middle of the turbine shaft to control the bending deformation of the large length-to-diameter ratio low-pressure rotor, thus achieving resonance speed control based on rotor bending deformation. A high-bending-resistance, large-pitch double-centering surface stop-gear connection structure is used to ensure a reliable connection between the turbine and the turbine shaft. The front turbine section of the split turbine is connected via an interstage sealing drum to form an integrated annular cavity structure, utilizing its overall gyroscopic torque effect to control the deformation of the middle section of the turbine shaft.

[0007] To achieve the above-mentioned objectives, the technical solution adopted by this invention is as follows:

[0008] The novel high aspect ratio rotor structure with a split turbine includes a four-stage fan, two-stage front turbines, a single-stage rear turbine, four bearing supports, and connecting structures. The four-stage fan is located at the front end of the rotor and has an overall arched structure. It is assembled onto the turbine shaft through a fan rear journal arc end tooth connection structure and a fan rear sleeve tooth connection structure. The two-stage front turbines are located in the middle of the rotor, forming a closed annular cavity structure between the turbine disks, and are connected to the turbine shaft through a front turbine conical shell-sleeve tooth connection structure. The single-stage rear turbine is located at the rear end of the rotor and is connected to the turbine shaft through a rear turbine stop-sleeve tooth connection structure. The four bearing supports are located at the front and rear of the four-stage fan, at the middle of the turbine shaft near the two-stage front turbines, and at the rear side of the single-stage rear turbine.

[0009] Furthermore, the four-stage fan includes a front fan journal, a fan blade, and a rear fan journal. The fan consists of four stages of blades and a rotor, with the rotors connected by a drum with a diameter of 512mm. The rear fan journal extends rearward from the edge of the third-stage rotor.

[0010] Furthermore, the two-stage front turbine includes a turbine shaft, a front turbine, an interstage sealing drum, and a front turbine journal, while the single-stage rear turbine structure includes a rear turbine, a rear turbine front journal, and a rear turbine rear journal. The two-stage front turbines are connected by the interstage sealing drum to form an integrated annular cavity structure, which can fully utilize its overall gyroscopic torque effect. The rear turbine is coaxially arranged with the front turbine but axially separated. The design of the rear turbine front journal and the rear turbine rear journal being parallel to the turbine shaft improves the overall structural integrity and increases the local rigidity of the rotor's rear end. The turbine shaft is slender, resulting in a large rotor length-to-diameter ratio of 24, which is 50% higher than the length-to-diameter ratio (13-17) of a conventional high thrust-to-weight ratio engine's low-pressure rotor.

[0011] Furthermore, the four bearing supports include the front support of the fan, the rear support of the fan, the middle support of the turbine shaft, and the rear support of the turbine.

[0012] Furthermore, the front support point of the fan adopts a roller bearing and is installed on the front journal of the fan; the rear support point of the fan adopts a ball bearing and is located on the radially outer side of the rear gear connection structure of the fan; the middle support point of the turbine shaft adopts a roller bearing, located in the axial middle of the turbine shaft and close to the center of mass of the front turbine; the rear support point of the turbine adopts a roller bearing and is installed on the rear journal of the rear turbine.

[0013] Furthermore, the connection structure includes a fan stage bolt connection structure, a fan rear journal arc end tooth connection structure, a fan rear sleeve tooth connection structure, a front turbine disc shaft bolt connection structure, a front turbine cone shell-sleeve tooth connection structure, and a rear turbine stop-sleeve tooth connection structure.

[0014] Furthermore, the fan stage bolt connection structure is located between the second-stage and third-stage fan blades; the fan rear journal arc end tooth connection structure connects the fan and the transition shaft and is located radially outside the fan rear sleeve tooth connection structure, making it easy to assemble and install; the fan rear sleeve tooth connection structure connects the transition shaft and the turbine shaft.

[0015] Furthermore, the front turbine and turbine shaft are connected via a front turbine conical shell-gear connection structure with a conical shell angle of 50°. This appropriate conical shell angle ensures the bending stiffness of the front turbine disk. The rear turbine extends out of the front journal and is connected to the turbine shaft via a rear turbine stop-gear connection structure. Centering is achieved through the cylindrical surface of the stop, reducing the impact of disk deformation on centering. The large distance between the centering surfaces of the two gear connection structures effectively improves the bending moment bearing capacity, fully applies the disk gyroscopic torque to the shaft section, increases the rotor resonant speed safety margin, controls bending deformation, and achieves a highly robust and reliable connection between the turbine and turbine shaft.

[0016] The present invention has the following beneficial effects:

[0017] (1) The present invention adopts a four-point support scheme, with an intermediate support point set at the front turbine of the two-stage low pressure. Through support constraint optimization, the problem of deformation runaway caused by the non-coordination of mass and stiffness distribution of the front turbine of the low pressure is solved, local bending deformation is controlled, the safety margin of rotor bending resonance is improved, and the problem of rotor dynamics design with large length-to-diameter ratio with split turbine is solved well.

[0018] (2) In view of the connection requirements of the rotor structure system brought about by the multi-stage split turbine, the present invention adopts a 50° conical shell-tooth connection for the two-stage front turbine to ensure the journal and connection bending stiffness; and adopts a large-pitch double centering surface stop-tooth connection for the single-stage rear turbine to improve the bending moment bearing capacity and realize a high robustness and high reliability connection between the turbine and the turbine shaft.

[0019] (3) This invention is applicable to rotor structure systems with turbine front and rear separation and large length-to-diameter ratio characteristics, such as interstage combustion chambers. The two front turbines are connected by interstage sealing drums and bolts to form a high bending stiffness integral annular cavity structure, which fully utilizes its overall gyroscopic torque effect. Through the high bending stiffness journal and connecting structure, the gyroscopic torque is applied to the rotor shaft section, effectively controlling the bending deformation of the low-pressure shaft with large length-to-diameter ratio and improving the bending resonance safety margin of the shaft section. Attached Figure Description

[0020] Figure 1 This is a schematic diagram of the overall structure of the rotor structure with a large length-to-diameter ratio and a split turbine of the present invention.

[0021] Figure 2 This is the Campbell diagram of the present invention;

[0022] Figure 3 This is a diagram of the bolted connection structure between fan stages;

[0023] Figure 4 This is a diagram of the fan disc shaft connection structure;

[0024] Figure 5 This is a diagram of the front turbine disk shaft connection structure;

[0025] Figure 6 This is a diagram of the rear turbine stop-gear connection structure.

[0026] In the diagram: 1-Fan front journal; 2-Fan; 2a-Fan second-stage impeller; 2b-Fan third-stage impeller; 3-Fan interstage bolt connection structure; 3a-Bolt connection stop; 3b-Short bolt; 4-Fan rear journal; 5-Fan rear journal arc-shaped end tooth connection structure; 5a-Arch-shaped end tooth bolt; 5b-Arch-shaped end tooth; 5c-Arch-shaped surface; 6-Transition shaft; 7-Turbine shaft; 8-Turbine shaft intermediate support point; 9-Front turbine; 9a-Front turbine flange; 10-Interstage sealing drum; 11-Rear turbine; 12-Turbine rear support point; 13-Fan front support point; 14-Fan rear sleeve Toothed connection structure; 14a-Fan rear sleeve tooth; 14b-Fan rear nut; 14c-Fan rear sleeve tooth cylindrical surface; 15-Fan rear pivot point; 16-Front turbine disc shaft bolt connection structure; 17-Front turbine rear journal; 18-Front turbine cone shell-sleeve tooth connection structure; 18a-Front turbine sleeve tooth; 18b-Front turbine sleeve tooth cylindrical surface; 18c-Front turbine sleeve tooth nut; 19-Rear turbine stop-sleeve tooth connection structure; 19a-Rear turbine stop cylindrical surface; 19b-Rear turbine sleeve tooth; 19c-Rear turbine sleeve tooth nut; 20-Rear turbine front journal; 21-Rear turbine rear journal. Detailed Implementation

[0027] The following will be based on embodiments of the present invention. Figures 1-5 The technical solutions in the embodiments of the present invention will be clearly and completely described. Obviously, the described embodiments are only some embodiments of the present invention, and not all embodiments. Unless otherwise specified, the technical means used in the embodiments are conventional means well known to those skilled in the art.

[0028] It needs to be explained that, in Figure 1 The image shows a schematic diagram of a cross-section, with the rotation center line below it. It can be understood that this invention is a three-dimensional structure symmetrical about the rotation center line. Figure 1 For ease of demonstration, a cross-section is used for illustration.

[0029] This invention is applicable to the structure of aero-engines with interstage combustion chambers. Due to the presence of interstage combustion chambers and the structural design of high-pressure rotors, low-pressure rotors are characterized by being slender and having low deformation control capabilities. This invention solves the dynamic design problem caused by these rotor characteristics. Compared with the prior art, the overall structural scheme of this invention has a simple force transmission path, high structural efficiency, sufficient critical speed margin, and strong deformation control capability.

[0030] like Figure 1 As shown, the high aspect ratio rotor structure with a split turbine of the present invention, from front to back, is as follows: Figure 1 From left to right, the overall structure includes, in sequence, the fan front pivot 13, fan front journal 1, fan 2, fan rear pivot 15, turbine shaft 7, turbine shaft middle pivot 8, front turbine 9, rear turbine front journal 20, rear turbine 11, rear turbine rear journal 21, and turbine rear pivot 12.

[0031] The low-pressure rotor adopts an "arch + disc shaft" configuration and a 1-2-1 support scheme, that is, a support point is set in front of the fan 2, two support points are set between the fan 2 and the rear turbine 11, located at the rear journal 4 of the fan and the front turbine 9 respectively, and a support point is set behind the rear turbine 11, so as to compensate for the deformation runaway problem caused by the non-coordination of mass and stiffness distribution through support constraints.

[0032] Fan 2 is supported by front support point 13 and rear support point 15. Front support point 13 is a roller bearing, mounted on front journal 1, which transmits part of the radial load of low-pressure rotor fan 2 outward. Rear support point 15 is a ball bearing, mounted on transition shaft 6 connected to rear journal 4, which transmits all axial load of low-pressure rotor and part of radial load of fan 2 outward. Turbine shaft intermediate support point 8 is a roller bearing, which transmits radial load of front turbine 9 outward. Its axial position is close to the center of mass of front turbine 9 to control the lateral displacement of front turbine 9 and the bending deformation of turbine shaft 7. Turbine rear support point 12 is a roller bearing, which transmits radial load of rear turbine 11 outward. Its axial position is close to the center of mass of rear turbine 11 to control the lateral deformation of rear turbine 11.

[0033] The turbine shaft 7 is a slender hollow shaft that connects the fan 2, the front turbine 9, and the rear turbine 11, transmitting loads such as torque, bending moment, and axial force. The slender turbine shaft 7 results in a large rotor length-to-diameter ratio, preferably 24, which is 50% higher than the length-to-diameter ratio (13-17) of a conventional high thrust-to-weight ratio engine's low-pressure rotor, making it prone to bending deformation. Figure 2As shown, compared to low-pressure rotors with conventional configurations and aspect ratios, this invention utilizes the weak bending stiffness of the turbine shaft 7 with a large aspect ratio to control the overall first-bend critical speed below the operating speed range, while providing sufficient safety margin. By positioning the front turbine 9 in the middle of the turbine shaft 7, the gyroscopic torque effect of the front turbine 9 is used to suppress bending deformation of the turbine shaft 7, raising the overall second-bend critical speed above the operating speed range, while also providing sufficient safety margin.

[0034] like Figure 3 As shown, the fan stage 2a and the fan stage 3b are connected by a fan stage bolted structure 3. The bolted connection stop 3a is used for centering, and the short bolts 3b are used for tightening. This facilitates processing, assembly, and disassembly. The flange edge is locally thickened, which helps to improve the structural stability.

[0035] like Figure 4 As shown, fan 2 and transition shaft 6 are connected by a fan rear journal arc-shaped end tooth connection structure 5. Axial force is transmitted through the arc-shaped end tooth bolt 5a, torque is transmitted through the arc-shaped end tooth 5b, and the arc surface 5c is used for centering, facilitating installation and disassembly. Transition shaft 6 and turbine shaft 7 are connected by a fan rear sleeve tooth connection structure 14. Torque is transmitted through the fan rear sleeve tooth 14a, axial positioning is achieved through the fan rear nut 14b, and centering is achieved through the fan rear sleeve tooth cylindrical surface 14c. This connection provides strong rigidity and improves interface stability.

[0036] like Figure 5 As shown, the two-stage low-pressure front turbine 9 extends from the front turbine flange edge 9a at the rim and connects to the front turbine rear journal 17 via the interstage sealing drum 10 and the front turbine disk shaft bolt connection structure 16, forming an integrated annular cavity structure between the two-stage turbine disks. This improves the angular stiffness of the turbine disks, leverages their overall gyroscopic torque effect, and enhances the structural robustness. The front turbine rear journal 17 assembles the front turbine 9 and turbine shaft 7 together via the front turbine conical shell-sleeve tooth connection structure 18. The front turbine conical shell-sleeve tooth connection structure 18 transmits torque through the front turbine sleeve teeth 18a. The outer surface of the connecting journal is provided with a precision front turbine sleeve tooth cylindrical surface 18b for centering. The increased distance between the two centering surfaces allows the gyroscopic torque effect of the front turbine 9 to fully act on the shaft section. The front turbine sleeve tooth nut 18c provides axial locking and positioning.

[0037] like Figure 6 As shown, the rear turbine 9 extends from the rear turbine front journal 20 at the end edge of the wheel disk and is connected to the turbine shaft 7 through the rear turbine stop-sleeve connection structure 19. The rear turbine stop cylindrical surface 19a is used for centering and positioning to reduce the influence of wheel disk deformation on centering. The rear turbine sleeve 19b transmits torque, and the rear turbine sleeve nut 19c is used for axial compression. The use of a large-pitch double centering surface improves the bending moment bearing capacity, allows the gyroscopic torque effect to be fully utilized in the shaft section, controls bending deformation, and achieves a highly robust and reliable connection between the turbine and the turbine shaft.

[0038] The embodiments described above are merely preferred embodiments of the present invention and are not intended to limit the scope of the present invention. Any modifications, alterations, or substitutions made by those skilled in the art to the technical solutions of the present invention without departing from the spirit of the present invention should fall within the protection scope defined by the claims of the present invention.

Claims

1. A rotor structure with a large length-to-diameter ratio and a split turbine, characterized in that: It includes a four-stage fan, a two-stage front turbine, a single-stage rear turbine, four bearing supports, and a connecting structure. The four-stage fan is located at the front end of the rotor and has an overall arched structure. It is assembled on the turbine shaft (7) through the fan rear journal arc end tooth connection structure (5) and the fan rear sleeve tooth connection structure (14). The two-stage front turbine is located in the middle of the rotor. A closed annular cavity structure is formed between the turbine disks and it is connected to the turbine shaft (7) through the front turbine cone shell-sleeve tooth connection structure (18). The single-stage rear turbine is located at the rear end of the rotor and is connected to the turbine shaft (7) through the rear turbine stop-sleeve tooth connection structure (19). The four bearing supports are located in front of the four-stage fan, behind the four-stage fan, in the middle of the turbine shaft near the two-stage front turbine, and on the rear side of the single-stage rear turbine. The two-stage front turbine includes a turbine shaft (7), a front turbine (9), an interstage sealing drum (10), and a front turbine rear journal (17). The single-stage rear turbine includes a rear turbine (11), a rear turbine front journal (20), and a rear turbine rear journal (21). The two-stage front turbines are connected by the interstage sealing drum (10) to form an integrated annular cavity structure. The rear turbine (11) is coaxially arranged with the front turbine (9) and axially separated. The rear turbine front journal (20) and the rear turbine rear journal (21) are parallel to the turbine shaft (7). The turbine shaft (7) is slender, resulting in a rotor length-to-diameter ratio of 24.

2. The rotor structure with a large length-to-diameter ratio and a split turbine according to claim 1, characterized in that, The four-stage fan includes a front journal (1), a fan (2), a rear journal (4), and a transition shaft (6); the fan (2) consists of four blades and a wheel, with the wheels connected by a drum with a diameter of 512mm, and the rear journal (4) extends backward from the edge of the third-stage wheel.

3. The rotor structure with a large length-to-diameter ratio and a split turbine according to claim 1, characterized in that, The four bearing supports include the front fan support (13), the rear fan support (15), the turbine shaft center support (8), and the turbine rear support (12).

4. The rotor structure with a large length-to-diameter ratio and a split turbine according to claim 3, characterized in that, The front support point (13) of the fan is a roller bearing and is installed on the front journal (1) of the fan; the rear support point (15) of the fan is a ball bearing located on the radial outside of the rear gear connection structure (14) of the fan; the middle support point (8) of the turbine shaft is a roller bearing located in the axial middle of the turbine shaft (7) and close to the center of mass of the front turbine (9); the rear support point (12) of the turbine is a roller bearing and is installed on the rear journal (21) of the rear turbine.

5. The rotor structure with a large length-to-diameter ratio and a split turbine according to claim 1, characterized in that, The connection structure includes a fan stage bolt connection structure (3), a fan rear journal arc end tooth connection structure (5), a fan rear sleeve tooth connection structure (14), a front turbine disc shaft bolt connection structure (16), a front turbine cone shell-sleeve tooth connection structure (18), and a rear turbine stop-sleeve tooth connection structure (19).

6. The rotor structure with a large length-to-diameter ratio and a split turbine according to claim 5, characterized in that, The fan stage bolt connection structure (3) is located between the second-stage fan blade disk (2a) and the third-stage fan blade disk (2b); the fan rear journal arc end tooth connection structure (5) connects the fan (2) and the transition shaft (6) and is located radially outside the fan rear sleeve tooth connection structure (14); the fan rear sleeve tooth connection structure (14) connects the transition shaft (6) and the turbine shaft (7).

7. The rotor structure with a large length-to-diameter ratio and a split turbine according to claim 5, characterized in that, The front turbine (9) is connected to the turbine shaft (7) through the front turbine cone shell-sleeve tooth connection structure (18), the radius of which is higher than the rear turbine stop-sleeve tooth connection structure (19), thereby ensuring that the front turbine (9) and the front turbine rear journal (17) can be installed from the turbine shaft (7) axially to the middle of the turbine shaft (7); the rear turbine (11) is connected to the turbine shaft (7) through the rear turbine stop-sleeve tooth connection structure (19).

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