A direct-drive medium-bypass-ratio aero-engine

Abstract

A direct drive type middle-high bypass ratio aero-engine comprises a fan case, an inner channel inlet inner case, a middle case arranged in the middle of the engine, a fan structure arranged in the front of the engine, and a turbine support arranged in the tail of the engine, wherein the fan structure comprises a ring motor, fan blades, a fan wheel and a fan main mounting ring; the stator of the ring motor is mounted on the fan case; the fan main mounting ring is fixed in the inner hole of the rotor of the ring motor; the blade crown of the fan blade is connected with the fan main mounting ring; the blade root of the fan blade is connected with the fan wheel; and the fan wheel is rotatably mounted on the connecting component inside the inner channel inlet inner case. The stress condition of the fan blade root and the fan wheel is greatly improved, the fan is directly driven by the ring motor, the turbine system is no longer needed to directly work on the fan, and the fan and the rotors of each stage of the engine can work at the optimal rotating speed.

Description

Technical Field

[0001] This invention relates to the field of aero-engine technology, and in particular to a direct-drive medium-high bypass ratio aero-engine. Background Technology

[0002] In traditional aero engines, the fan is directly mounted on the low-pressure shaft of the low-pressure turbine system inside the engine. This causes the fan's rotation to follow the low-pressure shaft, making it impossible to adjust the fan speed to achieve optimal operation. Furthermore, the fan's rotation requires the low-pressure turbine system to directly perform work, affecting the operation of the engine's low-pressure turbine system. Mounting the fan on the low-pressure shaft also complicates the engine's shaft system design, hindering its simplification. Additionally, in traditional aero engine fan structures, the blade roots are only connected to the fan rotor, resulting in a cantilevered blade mounting. Centrifugal force, aerodynamic forces, and other loads on the fan blades are transmitted to the fan rotor via the blade roots, resulting in a single force transmission path. This single force transmission path makes the connection between the blade roots and the fan rotor prone to stress concentration, increasing the risk of cracks and fan damage. Moreover, in traditional center-cantilever mounting structures, when the airflow in front of the fan is unstable, the cantilevered fan blade roots are subjected to unstable torque and bending moments, causing complex aerodynamic vibrations. Summary of the Invention

[0003] The main objective of this invention is to propose a direct-drive medium-high bypass ratio aero-engine to solve the aforementioned technical problems.

[0004] To achieve the above objectives, this invention proposes a direct-drive medium-high bypass ratio aero-engine, including a fan casing, an inner inlet casing, an intermediate casing located in the middle of the engine, a fan structure located directly in front of the engine, and a turbine support located at the tail of the engine. The fan structure includes a ring motor, fan blades, a fan wheel, and a main fan mounting ring; the stator of the ring motor is mounted on the fan casing.

[0005] The main fan mounting ring is fixedly connected to the inner hole of the annular motor rotor; the blade crown of the fan blade is connected to the main fan mounting ring; the blade root of the fan blade is connected to the fan wheel; the fan wheel is rotatably mounted on the connecting component inside the inner casing of the inner inlet.

[0006] Preferably, the aero-engine also includes a flow divider ring disposed behind the fan structure, the outer ring of which is connected to the fan casing via a fan support; the flow divider ring divides the flow path behind the fan into an inner duct and an outer bypass duct; the inner duct is formed by the inner ring of the flow divider ring and the outer wall of the inner casing of the inner duct inlet; the outer bypass duct is formed by the outer ring of the flow divider ring and the inner wall of the fan casing.

[0007] Preferably, an inner intake deflector is provided in the inner ring of the air distribution ring. When the aircraft lands, the inner intake deflector opens to connect the outer bypass duct with the inner bypass duct.

[0008] Preferably, the aero-engine further includes a low-pressure rotor system, which includes a low-pressure shaft, a low-pressure turbine, and a booster stage compressor. The low-pressure shaft is located inside the engine, with its rear end rotatably mounted on a connecting component inside the turbine support and its front end rotatably mounted on a connecting component inside the intermediate casing. The booster stage compressor is located behind the inner duct and at the front end of the intermediate casing, and is connected to the front end of the low-pressure shaft. The low-pressure turbine is located near the tail of the engine and is connected to the rear end of the low-pressure shaft.

[0009] Preferably, the aero-engine further includes a high-pressure rotor system, which includes a high-pressure turbine, a high-pressure compressor, and a high-pressure shaft; the high-pressure compressor is located at the rear of the intermediate casing, and the high-pressure turbine is located at the front of the low-pressure turbine; a combustion chamber and a high-pressure turbine guide are disposed between the high-pressure compressor and the high-pressure turbine, and the high-pressure turbine guide is located at the rear end of the combustion chamber; the front end of the high-pressure shaft is rotatably mounted on a connecting component inside the intermediate casing, and the rear end is rotatably mounted on a connecting component of the high-pressure turbine guide; the high-pressure compressor is connected to the front end of the high-pressure shaft; and the high-pressure turbine is connected to the rear end of the high-pressure shaft.

[0010] Preferably, the rear end of the low-pressure shaft is mounted on a connecting component inside the turbine support via bearing A, and the front end is mounted on a connecting component inside the intermediate casing via bearing B.

[0011] Preferably, the front end of the high-pressure shaft is mounted on a connecting component inside the intermediate casing via bearing D, and the rear end is mounted on a connecting component of the high-pressure turbine guide via bearing C.

[0012] Preferably, the fan blades, fan disc, and fan main mounting ring are integrally formed.

[0013] Preferably, the fan disc is mounted on a connecting component inside the inner casing via a bearing E.

[0014] Preferably, when the aircraft lands, the ring motor of the fan structure drives the blades to rotate in the opposite direction to generate reverse thrust.

[0015] Due to the adoption of the above technical solution, the beneficial effects of the present invention are as follows:

[0016] (1) Compared with traditional high-bypass turbofan engines, the fan structure of this invention changes the installation form of the fan blades. The traditional cantilevered installation, where the blade root is connected to the fan disk, is changed to a bridge-type installation, where the blade root is connected to the fan disk and the blade crown is connected to the main fan mounting ring. The centrifugal force, aerodynamic force, and other loads on the blades are changed from a single force transmission path that is transmitted from the blade root to the fan disk to a method where the force is transmitted simultaneously from the blade root to the fan disk and from the blade crown to the main fan mounting ring, or from the blade root to the fan disk and from the blade crown to the ring motor rotor, which greatly improves the stress condition of the fan blade root and the fan disk.

[0017] (2) The direct-drive medium-high bypass ratio aero-engine provided by this invention has an aerodynamically simplified fan structure due to the bridge-mounted fan blades, which greatly simplifies the fan impeller structure and reduces the outer diameter of the impeller. This changes the radial distribution of the fan's power delivery to the airflow, shifting the center of gravity of the power delivery distribution towards the fan blade root, improving the flow conditions at the blade crown, eliminating flow losses at the blade crown, increasing the stability margin of the airflow, and enhancing the aerodynamic stability of the fan blades. Simultaneously, when the incoming flow in front of the fan is unstable, the root of a traditional cantilevered fan blade is subjected to unstable torque and bending moment, causing complex aerodynamic vibrations. The bridge-mounted fan blades of this invention are rigidly connected to the fan impeller and the main fan mounting ring, respectively, reducing the degrees of freedom of the fan blades and greatly reducing the vibrations of the fan blades caused by aerodynamic forces.

[0018] (3) The direct-drive medium-to-high bypass ratio aero-engine provided by this invention, in terms of fan transmission, changes the fan rotation from being driven from rear to front by the low-pressure turbine system via the low-pressure shaft to being directly driven by a ring motor arranged around the fan. This eliminates the need for the turbine system to directly perform work on the fan, allowing the fan and each stage of the engine rotor to operate at their optimal speeds. This avoids the complex mechanisms of multi-rotor or gear-driven fans, simplifying the design of the engine shaft system. Furthermore, the ring motor has high power and can adapt to various engine operating conditions.

[0019] (4) The direct-drive medium-high bypass ratio aero-engine provided by the present invention uses a ring motor to drive the fan structure, and the fan blades do not use a low-pressure shaft drive. Therefore, the fan blades can be rotated in reverse by the ring motor. When the aircraft lands, the fan blades driven by the ring motor rotate in reverse to generate a reverse thrust in the opposite direction.

[0020] (5) In this invention, when the ring motor is running, the magnetic field generated on the stator excitation coil winding can not only drive the rotor to rotate, but also provide magnetic levitation force for the rotor, thus avoiding problems such as frictional resistance and vibration transmission caused by mechanical contact between the motor rotor system and the motor stator, and increasing the reliability of the engine. Attached Figure Description

[0021] To more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the drawings used in the description of the embodiments or the prior art will be briefly introduced below. Obviously, the drawings described below are only some embodiments of the present invention. For those skilled in the art, other drawings can be obtained based on the structures shown in these drawings without creative effort.

[0022] Figure 1 This is a schematic diagram of the structure of the direct-drive medium-high bypass ratio aero-engine provided by the present invention.

[0023] Figure 2 This is a schematic diagram of the fan structure in this invention.

[0024] Explanation of reference numerals: 1-Fan structure; 2-Split ring; 3-Fan casing; 4-Support bar; 5-Booster stage compressor; 6-Intermediate casing; 7-High pressure compressor; 8-Combustion chamber; 9-High pressure turbine guide vane; 10-High pressure turbine; 11-Low pressure turbine; 12-Turbine support; 13-Bearing A; 14-Low pressure shaft; 15-Bearing C; 16-High pressure shaft; 17-Bearing D; 18-Bearing B; 19-Inner casing inlet; 20-Bearing E; 21-Stator; 23-Rotor; 22-Coil winding; 24-Fan main mounting ring; 25-Fan blade; 26-Fan impeller; 27-Inner intake deflector. Detailed Implementation

[0025] The technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only a part of the embodiments of the present invention, and not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those of ordinary skill in the art without creative effort are within the scope of protection of the present invention.

[0026] It should be noted that all directional indications (such as up, down, left, right, front, back, etc.) in the embodiments of the present invention are only used to explain the relative positional relationship and movement of each component in a certain specific posture (as shown in the figure). If the specific posture changes, the directional indication will also change accordingly.

[0027] Combination Figure 1 , Figure 2As shown, a direct-drive medium-high bypass ratio aero-engine includes a fan casing 3, an inner inlet casing 19, an intermediate casing 6 located in the middle of the engine, and a fan structure 1 located directly in front of the engine, and a turbine support 12 located at the tail of the engine. The fan structure includes a ring motor, fan blades 25, a fan wheel 26, and a main fan mounting ring 24. The ring motor includes a stator 21, coil windings 22, and a rotor 23. The stator 21 of the ring motor is mounted on the fan casing 3. The main fan mounting ring 24 is fixedly connected to the inner hole of the ring motor rotor 23. The blade crowns of the fan blades 25 are connected to the main fan mounting ring 24. The blade roots of the fan blades 25 are connected to the fan wheel 26. The fan wheel 26 is rotatably mounted on a connecting component inside the inner inlet casing 19. Specifically, the fan wheel 26 is mounted on the connecting component inside the inner inlet casing 19 via a bearing E20.

[0028] Combination Figure 1 As shown, it also includes a flow divider ring 2 disposed behind the fan structure 1. The outer ring of the flow divider ring 2 is connected to the fan casing 3 through the fan support 4. The flow divider ring 2 divides the flow channel behind the fan into an inner channel and an outer channel. The inner channel is formed by the inner ring of the flow divider ring 2 and the outer wall of the inner inlet casing 19. The outer channel is formed by the outer ring of the flow divider ring 2 and the inner wall of the fan casing 3.

[0029] Combination Figure 1 As shown, an inner intake deflector 27 is provided in the inner ring of the flow divider ring 2. Under the control of the corresponding mechanical actuation system, the inner intake deflector 27 is closed (in state a in the figure) or open (in state b in the figure) depending on the different states of the engine. When the aircraft lands, the inner intake deflector 27 is opened to connect the outer bypass duct and the inner bypass duct.

[0030] Combination Figure 1 As shown, the aero-engine also includes a low-pressure rotor system, which includes a low-pressure shaft 14, a low-pressure turbine 11, and a booster compressor 5. The low-pressure shaft 14 is located inside the engine, with its rear end rotatably mounted on a connecting component inside the turbine support 12, and its front end rotatably mounted on a connecting component inside the intermediate casing 6. Specifically, the rear end of the low-pressure shaft 14 is mounted on the connecting component inside the turbine support 12 via bearing A13, and the front end is mounted on the connecting component inside the intermediate casing 6 via bearing B18. The booster compressor 5 is located behind the inner duct and at the front end of the intermediate casing 6, and is connected to the front end of the low-pressure shaft 14. The low-pressure turbine 11 is located near the tail of the engine and is connected to the rear end of the low-pressure shaft 14.

[0031] Combination Figure 1As shown, the aero-engine also includes a high-pressure rotor system, which includes a high-pressure turbine 10, a high-pressure compressor 7, and a high-pressure shaft 16. The high-pressure compressor 7 is located at the rear of the intermediate casing 6, and the high-pressure turbine 10 is located at the front of the low-pressure turbine 11. A combustion chamber 8 and a high-pressure turbine guide 9 are disposed between the high-pressure compressor 7 and the high-pressure turbine 10, with the high-pressure turbine guide 9 located at the rear end of the combustion chamber 8. The front end of the high-pressure shaft 16 is rotatably mounted on a connecting component inside the intermediate casing 6, and the rear end is rotatably mounted on a connecting component of the high-pressure turbine guide 9. Specifically, the front end of the high-pressure shaft 16 is mounted on the connecting component inside the intermediate casing 6 via bearing D17, and the rear end is mounted on the connecting component of the high-pressure turbine guide 9 via bearing C15. The high-pressure compressor 7 is connected to the front end of the high-pressure shaft 16; the high-pressure turbine 10 is connected to the rear end of the high-pressure shaft 16.

[0032] In this embodiment, the fan blade 25, fan disc 26, and fan main mounting ring 24 are integrally formed. The forming method can be casting. The integral forming structure ensures that both ends of the fan blade 25 are rigidly connected to the fan disc 26 and the fan main mounting ring 24, respectively, reducing the degree of freedom of movement of the fan blade 25 and greatly reducing the vibration of the fan blade 25 caused by aerodynamic forces.

[0033] In this invention, when the aircraft lands, the ring motor of the fan structure drives the blades to rotate in the opposite direction to generate reverse thrust. At the same time, the inner intake deflector 27 opens (as shown in state b in the figure), and the airflow is introduced into the engine inner duct from the rear of the engine outer bypass duct through the inner intake deflector 27, ensuring the engine's operational stability, replacing the traditional thrust reverser mechanism in the aircraft engine nacelle, and reducing the overall weight of the engine.

[0034] In this embodiment, bearings A13, B18, C15, D17, and E20 are in the form of rolling element bearings, such as ball bearings or roller bearings.

[0035] The working principle of this invention is as follows: When the engine is working (generating positive thrust), the ring motor on the fan structure 1 is driven by the electrical control system on the aircraft / engine, which in turn drives the fan blades 25 to rotate forward, drawing in outside air into the engine. The airflow is compressed as it flows through the fan blades 25, increasing its pressure. After the fan blades 25, the airflow is divided into inner and outer bypass streams by the splitter ring 2. The outer bypass airflow enters the outer bypass duct and flows backward around components such as the booster compressor 5, high-pressure compressor 7, combustion chamber 8, high-pressure turbine 10, and low-pressure turbine 11. At this time, the inner bypass intake deflector 27 on the inner ring of the splitter ring is closed (in state a in the figure). The inner bypass airflow is sequentially pressurized by the booster compressor 5 and high-pressure compressor 7, and then enters the combustion chamber 8, where it mixes with fuel injected by the fuel system and is ignited to achieve the predetermined high temperature. The high-temperature, high-pressure gas flow expands and does work after passing through the high-pressure turbine 10 and low-pressure turbine 11, while simultaneously driving the high-pressure compressor 7 and booster compressor 5 to continuously draw in and pressurize air. Finally, the airflow from the inner duct and the airflow from the outer bypass duct are discharged from the engine at high speed through mixed exhaust or separate exhaust, depending on the actual situation, generating counter-thrust. When the aircraft lands, the fan directly driven by the ring motor rotates in the opposite direction to generate counter-thrust. At the same time, the inner duct inlet deflector 27 on the inner ring of the splitter ring 2 opens (in state b in the figure). The airflow is introduced into the engine inner duct from the rear of the outer bypass duct through the inner duct inlet deflector 27, ensuring the engine's operational stability, simplifying the design of the engine nacelle, and replacing the traditional thrust reverser mechanism in the aircraft engine nacelle.

[0036] The above description is merely a preferred embodiment of the present invention and does not limit the patent scope of the present invention. Any equivalent structural transformations made using the contents of the present invention's specification and drawings under the inventive concept of the present invention, or direct / indirect applications in other related technical fields, are included within the patent protection scope of the present invention.

Claims

1. A direct drive medium bypass ratio aero-engine comprising a fan case (3), an inner bypass inlet inner case (19), an intermediate case (6) arranged in the middle of the engine, and a fan structure (1) arranged at the front of the engine, a turbine support (12) arranged at the tail of the engine, characterized in that: The fan structure comprises a ring motor, fan blades (25), a fan wheel disc (26), and a fan main mounting ring (24); The ring motor comprises a stator (21), a coil winding (22), and a rotor (23); The stator (21) of the ring motor is mounted on the fan casing (3); The fan main mounting ring (24) is fixedly connected in the inner hole of the rotor (23) of the ring motor; The blade crown of the fan blade (25) is connected with the fan main mounting ring (24), and the blade root of the fan blade (25) is connected with the fan wheel disc (26) to form a bridge type mounting structure; The fan blade (25), the fan wheel disc (26), and the fan main mounting ring (24) are integrally formed; The fan wheel disc (26) is rotatably mounted on the connecting component inside the inner duct inlet inner casing (19); When the ring motor is running, the magnetic field generated by the stator (21) exciting the coil winding (22) drives the rotor (23) to rotate, and also provides the rotor (23) with a magnetic suspension force, so as to avoid the friction resistance and vibration caused by the mechanical contact between the motor rotor system and the motor stator; The fan structure (1) further comprises a shunt ring (2) arranged at the rear of the fan structure (1), and the outer ring of the shunt ring (2) is connected with the fan casing (3) through a fan support (4); the shunt ring (2) divides the flow channel behind the fan into an inner duct and an outer duct; the inner duct is formed by the inner ring of the shunt ring (2) and the outer wall of the inner duct inlet inner casing (19); and the outer duct is formed by the outer ring of the shunt ring (2) and the inner wall of the fan casing (3); An inner duct inlet air baffle (27) is arranged on the inner ring of the shunt ring (2), and when the airplane lands, the inner duct inlet air baffle (27) is opened to communicate the outer duct with the inner duct, and the fan structure ring motor drives the blades to rotate reversely to generate reverse cruise anti-thrust.

2. A direct drive intermediate bypass ratio aero-engine as claimed in claim 1, characterized in that, The fan structure (1) further comprises a low pressure rotor system, which comprises a low pressure shaft (14), a low pressure turbine (11), and a booster stage compressor (5); the low pressure shaft (14) is arranged inside the engine, the rear end of the low pressure shaft (14) is rotatably mounted on the connecting component inside the turbine support (12), and the front end is rotatably mounted on the connecting component inside the intermediate casing (6); The booster stage compressor (5) is arranged at a position behind the inner duct and at the front end of the intermediate casing (6), and is connected with the front end of the low pressure shaft (14); the low pressure turbine (11) is arranged at a position close to the tail of the engine and is connected with the rear end of the low pressure shaft (14).

3. A direct drive intermediate bypass ratio aero-engine as claimed in claim 2, characterised in that, The fan structure (1) further comprises a high pressure rotor system, which comprises a high pressure turbine (10), a high pressure compressor (7), and a high pressure shaft (16); The high pressure compressor (7) is located at the rear of the intermediate casing (6), and the high pressure turbine (10) is located at the front end of the low pressure turbine (11); a combustion chamber (8) and a high pressure turbine guide vane (9) are arranged between the high pressure compressor (7) and the high pressure turbine (10), and the high pressure turbine guide vane (9) is located at the rear end of the combustion chamber (8); The front end of the high pressure shaft (16) is rotatably mounted on the connecting component inside the intermediate casing (6), and the rear end is rotatably mounted on the connecting component of the high pressure turbine guide vane (9). The high pressure compressor (7) is connected to the front end of the high pressure shaft (16) and the high pressure turbine (10) is connected to the rear end of the high pressure shaft (16).

4. A direct drive intermediate bypass ratio aero-engine as claimed in claim 2, wherein, The rear end of the low pressure shaft (14) is mounted by bearing A (13) to a connection member inside the turbine support (12) and the front end is mounted by bearing B (18) to a connection member inside the intermediate casing (6).

5. A direct drive intermediate bypass ratio aero-engine as claimed in claim 3, wherein, The front end of the high pressure shaft (16) is mounted by bearing D (17) to a connection member inside the intermediate casing (6) and the rear end is mounted by bearing C (15) to a connection member of the high pressure turbine guide vane (9).

6. A direct drive intermediate bypass ratio aero-engine as claimed in claim 1, wherein, The fan wheel disc (26) is mounted by bearing E (20) to a connection member inside the inner hub inlet inner casing (19).

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