A dual-rotor tri-axial turboprop engine

Abstract

The application belongs to the technical field of aero-engines, and discloses a double-rotor three-shaft turboprop engine which comprises a propeller, a speed reducer, an air inlet device, a gas generator and a power turbine; the speed reducer is connected with the air inlet end of the gas generator through the air inlet device, the exhaust end of the gas generator is connected with the power turbine, the gas generator comprises a compressor, a combustion chamber and a gas turbine which are sequentially arranged along the exhaust direction, the power shaft of the power turbine is connected with the input end of the speed reducer at one end, and the output end of the speed reducer is connected with the propeller. The double-rotor three-shaft turboprop engine adopts a unit body structure design, multiple units are divided according to functions, and the double-rotor three-shaft turboprop engine is convenient to assemble and disassemble and is convenient to maintain in the field, and the requirements of airworthiness, reliability, safety, maintainability and economy are fully considered.

Description

Technical Field

[0001] This invention belongs to the field of aero-engine technology, and specifically relates to a dual-rotor three-shaft turboprop engine. Background Technology

[0002] In the existing technology, there are two main structural layout options for turboprop engines with a power range of 4500kW to 8000kW:

[0003] 1. The turboprop engine CPX38 produced by a certain company adopts a combined compressor and a dual-shaft overall structure layout.

[0004] 2. The turboprop engine TP400-D6 produced by a certain company adopts an all-axial compressor and a three-shaft overall structure layout, including a 5-stage axial low-pressure compressor, a 6-stage axial high-pressure compressor, an annular direct-flow combustion chamber, a 1-stage axial high-pressure turbine, a 1-stage low-pressure turbine, a 3-stage axial power turbine, and an offset reduction gear.

[0005] The above two turboprop engine structural layouts have the following disadvantages:

[0006] 1. The CPX38 engine is an improved design based on the GE38-1B turboshaft engine. The gas generator adopts a single rotor structure. The engine has a large starting inertia, which requires high starter power. The single-shaft compressor has many stages and a complex adjustable mechanism. It requires high aerodynamic efficiency of the compressor under the same power, which places high demands on the compressor performance design and matching.

[0007] 2. The TP400-D6 engine has a power output of 8000kW. It uses a full axial flow compressor with multiple compressor stages and a large rotor span, making compressor efficiency control difficult. It also has a less compact structure, higher weight, and more difficult stiffness control of the load-bearing system.

[0008] In summary, the existing turboprop engines in the power range of 4500kW to 8000kW have a non-compact structural layout, are difficult to disassemble and maintain, have a large starting inertia, and have a slow starting speed. Summary of the Invention

[0009] To address the above problems, this invention provides a dual-rotor, three-shaft turboprop engine, employing the following technical solution:

[0010] A dual-rotor, three-shaft turboprop engine includes a propeller, a reducer, an intake device, a gas generator, and a power turbine; wherein the reducer is connected to the intake end of the gas generator via the intake device, the exhaust end of the gas generator is connected to the power turbine, the gas generator includes a compressor, a combustion chamber, and a gas turbine arranged sequentially along the exhaust direction, one end of the power shaft of the power turbine is connected to the input end of the reducer, and the output end of the reducer is connected to the propeller.

[0011] Furthermore, the reducer is an offset reducer, the air intake device is S-shaped, a rectifier support plate is provided inside the air intake device, and the offset reducer is located in front of the air intake device.

[0012] Furthermore, the gas generator also includes a low-pressure rotor shaft and a high-pressure rotor shaft; the compressor includes a high-pressure compressor, an intermediate casing, and a low-pressure compressor; and the gas turbine includes a gas turbine casing, a high-pressure turbine, and a low-pressure turbine.

[0013] The low-pressure compressor's inlet end is connected to the inlet device, the low-pressure compressor's outlet end is connected to the high-pressure compressor via the intermediate casing, the high-pressure turbine is connected to the high-pressure compressor via the high-pressure rotor shaft, the low-pressure turbine is connected to the low-pressure compressor via the low-pressure rotor shaft, the low-pressure rotor shaft passes through the high-pressure rotor shaft, and the power shaft passes through the low-pressure rotor shaft.

[0014] Furthermore, the gas turbine is connected to the power turbine via the gas turbine casing, and the exhaust end of the power turbine is provided with an exhaust casing;

[0015] The intake device is equipped with a first bearing, a second bearing, and a third bearing; the intermediate casing is equipped with a fourth bearing and a fifth bearing; a sixth bearing is provided between the outlet end of the high-pressure compressor and the high-pressure turbine; the gas turbine casing is equipped with a seventh bearing and an eighth bearing; and the exhaust casing is equipped with a ninth bearing. The high-pressure rotor shaft is supported on the fifth bearing and the sixth bearing; the low-pressure rotor shaft is supported on the third bearing, the fourth bearing, and the seventh bearing; and the power shaft is supported on the first bearing, the second bearing, the eighth bearing, and the ninth bearing.

[0016] Furthermore, the combustion chamber includes an inner casing, an outer casing, a flame tube, dual fuel line nozzles, and a fuel main;

[0017] The outer casing is connected at its front end to the diffuser casing of the high-pressure compressor, and at its rear end to the casing of the high-pressure turbine. The inner casing is fixed to the diffuser casing of the high-pressure compressor. The head of the flame tube is fixedly connected to the outer casing. The head of the flame tube is also provided with a two-stage axial vortex generator with opposite rotation directions and a dual-fuel nozzle. The centerline of the two-stage axial vortex generator forms an angle with the central axis of the dual-rotor three-shaft turboprop engine. The dual-fuel nozzle is connected to the fuel main.

[0018] Furthermore, the head of the flame tube is also equipped with an air intake cap.

[0019] Furthermore, the high-pressure turbine includes high-pressure turbine guide vanes, a high-pressure turbine rotor, a high-pressure front sealing disc, and a high-pressure rear sealing disc, and the low-pressure turbine includes low-pressure turbine guide vanes, a low-pressure turbine rotor, and a low-pressure front sealing disc.

[0020] The gas turbine is connected to the outer casing and the power turbine via the gas turbine casing. The high-pressure turbine guide vane, the high-pressure turbine rotor, the low-pressure turbine guide vane, and the low-pressure turbine rotor are sequentially arranged inside the gas turbine casing along the exhaust direction. The high-pressure turbine rotor is connected to the high-pressure rotor shaft. The high-pressure front sealing plate is disposed between the high-pressure turbine guide vane and the high-pressure turbine rotor. The high-pressure rear sealing plate is disposed between the high-pressure turbine rotor and the low-pressure turbine guide vane. The low-pressure front sealing plate is disposed between the low-pressure turbine guide vane and the low-pressure turbine rotor.

[0021] Furthermore, the power turbine also includes a power turbine casing, a transition section integrated guide vane, a turbine first-stage guide vane, a turbine first-stage rotor, a turbine second-stage guide vane, a turbine second-stage rotor, a turbine third-stage guide vane, and a turbine third-stage rotor.

[0022] The transition section integrated guide vane is connected to the gas turbine casing and the power turbine casing at both ends, respectively. The turbine first-stage guide vane, the turbine first-stage rotor, the turbine second-stage guide vane, the turbine second-stage rotor, the turbine third-stage guide vane, and the turbine third-stage rotor are sequentially arranged inside the power turbine casing along the exhaust direction.

[0023] Furthermore, the transition section integrated guide vane and the turbine first-stage guide vane are integrally formed.

[0024] Furthermore, it also includes an accessory drive device, which is connected to the compressor drive device.

[0025] Furthermore, the accessory drive mechanism includes an accessory housing, a central drive assembly, and an accessory drive assembly;

[0026] The accessory housing and the intermediate housing are connected. The accessory transmission assembly is disposed in the accessory housing, and the central transmission assembly is disposed in the intermediate housing. The input end of the central transmission assembly is connected to the high-pressure rotor shaft, and the output end of the central transmission assembly is connected to the accessory transmission assembly.

[0027] Furthermore, the accessory housing is provided with an internal oil passage, which is integrally formed with the accessory housing.

[0028] The beneficial effects of this invention are:

[0029] 1. The dual-rotor three-shaft turboprop engine of the present invention adopts a modular structure design, which is divided into multiple units according to function, making it convenient for assembly, disassembly and field maintenance, and fully considering the requirements of airworthiness, reliability, safety, maintainability and economy.

[0030] 2. The dual-rotor three-shaft turboprop engine of the present invention adopts power output at the front of the shaft and places the reducer at the cold end position, which improves the reliability, safety and life of the high load transmission components, reduces maintenance costs and demonstrates good economic efficiency.

[0031] 3. The power turbine of the present invention adopts a simple support scheme with two support points at the front and rear, which ensures the rigidity of the power turbine, facilitates the maintenance of the rotor-stator clearance, thereby slowing down the performance degradation of the engine and reducing the cost over the entire life cycle.

[0032] 4. The dual-rotor three-shaft turboprop engine of the present invention has a compact structure, high integration, few external pipelines, and is easy to disassemble and maintain.

[0033] 5. The dual-rotor three-shaft turboprop engine of the present invention has a smaller starting inertia than the single-rotor gas generator, and its starting speed is faster under the same starting power.

[0034] 6. The high and low pressure compressors of the present invention operate at their respective optimal speeds, have a large surge margin, low aerodynamic efficiency requirements to achieve the target power, and can maintain a low fuel consumption rate even at non-design points.

[0035] 7. The dual-rotor three-shaft turboprop engine of the present invention enables the engine to achieve a higher power level under lower turbine inlet temperature conditions, has a longer hot end life, and has greater development and growth potential within the same frame size, covering dual-rotor three-shaft turboprop engines with power levels of 4500kW to 8000kW.

[0036] Other features and advantages of the invention will be set forth in the following description, and will be apparent in part from the description, or may be learned by practicing the invention. The objects and other advantages of the invention may be realized and obtained by means of the structures pointed out in the description and the drawings. Attached Figure Description

[0037] 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 some embodiments of the present invention. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0038] Figure 1 A schematic diagram of the overall structure of a dual-rotor, three-shaft turboprop engine according to an embodiment of the present invention is shown.

[0039] Figure 2 An installation schematic diagram of a gas generator according to an embodiment of the present invention is shown;

[0040] Figure 3 A schematic diagram of a dual-rotor support scheme according to an embodiment of the present invention is shown;

[0041] Figure 4 A schematic diagram of the air intake device according to an embodiment of the present invention is shown;

[0042] Figure 5 A schematic diagram of a compressor structure according to an embodiment of the present invention is shown;

[0043] Figure 6 A schematic diagram of a combustion chamber structure according to an embodiment of the present invention is shown;

[0044] Figure 7 A schematic diagram of a gas turbine structure according to an embodiment of the present invention is shown;

[0045] Figure 8 A schematic diagram of a power turbine structure according to an embodiment of the present invention is shown;

[0046] Figure 9 A schematic diagram of the transmission device of the accessory transmission device according to an embodiment of the present invention is shown;

[0047] Figure 10 A schematic diagram of the internal oil passage arrangement of the accessory housing according to an embodiment of the present invention is shown.

[0048] In the diagram: 1. Propeller; 2. Gearbox; 3. Intake system; 4. Gas generator; 5. Power turbine; 6. First bearing; 7. Second bearing; 8. Third bearing; 9. Fourth bearing; 10. Fifth bearing; 11. Sixth bearing; 12. Seventh bearing; 13. Eighth bearing; 14. Ninth bearing; 31. Rectifier support plate; 41. Compressor; 42. Combustion chamber; 43. Gas turbine; 44. High-pressure rotor shaft; 45. Low-pressure rotor shaft; 46. Power shaft; 411. Low-pressure compressor; 412. Intermediate casing; 413. High-pressure compressor; 4131. High-pressure first-stage axial compressor; 4132. High-pressure first-stage centrifugal compressor; 421. Inner casing; 422. Outer casing; 423. Flame tube; 424. Dual-oil nozzle; 425. Gas turbine. 426. Oil main pipe; 431. Diffuser casing; 432. Gas turbine casing; 433. High-pressure turbine; 4321. High-pressure turbine guide vane; 4322. High-pressure turbine rotor; 4323. High-pressure front sealing plate; 4324. High-pressure rear sealing plate; 4331. Low-pressure turbine guide vane; 4332. Low-pressure turbine rotor; 4333. Low-pressure front sealing plate; 51. Power turbine casing; 52. Exhaust casing; 53. Transition section integrated guide vane; 54. Turbine first-stage rotor; 55. Turbine second-stage guide vane; 56. Turbine second-stage rotor; 57. Turbine third-stage guide vane; 58. Turbine third-stage rotor; 59. Bearing housing; 61. Accessory drive assembly; 62. Central drive assembly; 63. Accessory casing; 64. Internal oil passage. Detailed Implementation

[0049] To make the objectives, technical solutions, and advantages of the embodiments of the present invention clearer, 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 some embodiments of the present invention, not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.

[0050] It should be noted that the terms "first," "second," etc., used in this application are used to distinguish similar objects and are not necessarily used to describe a specific order or sequence. It should be understood that such data can be interchanged where appropriate for the embodiments of this application described herein. In this application, the terms "upper," "lower," "left," "right," "front," "rear," "top," "bottom," "inner," "outer," "middle," "vertical," "horizontal," "lateral," "longitudinal," etc., indicate the orientation or positional relationship based on the orientation or positional relationship shown in the accompanying drawings.

[0051] This invention provides a dual-rotor, three-shaft turboprop engine with a simple and compact structure, large airflow, and a power range of 4500kW to 8000kW, designed to meet the requirements of fixed-wing aircraft (regional turboprop airliners, military and civilian short- and medium-range transport aircraft, and special-purpose aircraft such as carrier-based fixed-wing early warning aircraft) for high power-to-weight ratio and lightweight power plants. With adaptive modifications, the dual-rotor, three-shaft turboprop engine structure of this invention can also be used in turboshaft and turbofan engines.

[0052] like Figure 1 As shown, a dual-rotor tri-shaft turboprop engine includes a propeller 1, a reducer 2, an intake device 3, a gas generator 4, a power turbine 5, an accessory transmission device, a fuel system, a control system, a lubricating oil system, an air system, a starting system, an electrical system, and a health management system.

[0053] In this configuration, the reducer 2 is connected to the intake end of the gas generator 4 via the intake device 3, and the exhaust end of the gas generator 4 is connected to the power turbine 5. Figure 2 As shown, the gas generator 4 includes a compressor 41, a combustion chamber 42 and a gas turbine 43 arranged sequentially along the exhaust direction. One end of the power shaft 46 of the power turbine 5 is connected to the input end of the reducer 2, the output end of the reducer 2 is connected to the propeller 1, and the accessory transmission device is connected to the compressor 41 in a transmission connection.

[0054] The fuel system, control system, lubricating oil system, air system, starting system, electrical system, and health management system are not improvements of this invention. The corresponding existing technologies can be applied to the dual-rotor three-shaft turboprop engine of this invention, so they will not be described in detail.

[0055] The dual-rotor three-shaft turboprop engine of this invention adopts a modular structure design. According to its function, the whole engine is divided into two main units: a reduction gear unit 2, an air intake device unit 3, a gas generator unit 4, a power turbine unit 5, and an accessory transmission device unit. This facilitates assembly, disassembly, and field maintenance, and fully considers requirements such as airworthiness, reliability, safety, maintainability, and economy.

[0056] The dual-rotor three-shaft turboprop engine of this invention adopts power output at the front of the shaft and places the reducer 2 at the cold end position, which improves the reliability, safety and life of the high load transmission components, reduces maintenance costs and has good economic benefits.

[0057] Specifically, gas generator 4 has a dual-rotor structure, such as... Figure 3 As shown, it includes a low-pressure rotor shaft 45 and a high-pressure rotor shaft 44, wherein, as Figure 5As shown, the compressor 41 includes a high-pressure compressor 413, an intermediate casing 412, and a low-pressure compressor 411. The inlet end of the low-pressure compressor 411 is connected to the inlet device 3, and the outlet end of the low-pressure compressor 411 is connected to the high-pressure compressor 413 through the intermediate casing 412. The high-pressure compressor 413 and the low-pressure compressor 411 are connected in terms of pneumatic flow path and mechanical connection through the intermediate casing 412. For example, the low-pressure compressor 411 can be a 4-stage axial flow compressor 41, and the high-pressure compressor 413 can be a combination compressor of a high-pressure 1-stage axial flow compressor 4131 and a high-pressure 1-stage centrifugal compressor 4132.

[0058] like Figure 7 As shown, the gas turbine 43 includes a gas turbine casing 431, a high-pressure turbine 432 and a low-pressure turbine 433. The high-pressure turbine 432 is connected to the high-pressure compressor 413 via a high-pressure rotor shaft 44, and the low-pressure turbine 433 is connected to the low-pressure compressor 411 via a low-pressure rotor shaft 45. The low-pressure rotor shaft 45 passes through the center of the high-pressure rotor shaft 44, and the power shaft 46 of the power turbine 5 passes through the center of the low-pressure rotor shaft 45.

[0059] In one embodiment, the high-pressure rotor adopts a 1-1-0 support scheme, the low-pressure rotor adopts a 1-1-1 three-point support scheme, and the rotor of the power turbine 5 adopts a 2-1-1 support scheme.

[0060] like Figure 1 , Figure 3 , Figure 5 - Figure 8 As shown, specifically, the gas turbine 43 is connected to the power turbine 5 via a gas turbine casing 431, and an exhaust casing 52 is provided at the exhaust end of the power turbine 5. The intake device 3 is equipped with a first bearing 6, a second bearing 7, and a third bearing 8; the intermediate casing 412 is equipped with a fourth bearing 9 and a fifth bearing 10; a sixth bearing 11 is provided between the outlet end of the high-pressure compressor 413 and the high-pressure turbine 432; a seventh bearing 12 and an eighth bearing 13 are provided on the gas turbine casing 431; and a ninth bearing 14 is provided on the exhaust casing 52.

[0061] The high-pressure rotor shaft 44 is supported on the fifth bearing 10 and the sixth bearing 11, the low-pressure rotor shaft 45 is supported on the third bearing 8, the fourth bearing 9 and the seventh bearing 12, and the power shaft 46 of the power turbine 5 is supported on the first bearing 6, the second bearing 7, the eighth bearing 13 and the ninth bearing 14.

[0062] In this embodiment, the front support point of the low-pressure rotor and the front support point of the power turbine 5 shaft are supported by the intake device 3; the front support point of the high-pressure rotor and the middle support point of the low-pressure rotor are supported by the intermediate casing 412; the rear support point of the low-pressure rotor and the front support point of the power turbine 5 disc are supported by the turbine casing; and the rear support point of the power turbine 5 is supported solely by the exhaust casing 52. The power turbine 5 adopts a simple support scheme with two support points at the front and rear, ensuring the rigidity of the power turbine 5, which is beneficial for maintaining the rotor-stator clearance, thereby slowing down the performance degradation of the engine and reducing the cost over the entire life cycle.

[0063] In one embodiment, the reducer 2 is an offset reducer 2, such as... Figure 4 As shown, the intake device 3 is S-shaped, and a rectifier support plate 31 is installed inside the intake device 3. The offset reducer 2 is located in front of the intake device 3. The intake device 3 adopts an integrated design, providing a mounting edge to connect with the reducer 2 and the compressor 41 to realize load transfer. It also provides front support for the rotor of the power turbine 5 and the low-pressure rotor. The intake device 3 is also equipped with a lubricating oil interface, an anti-icing interface, an intake air temperature sensor interface, and a torque sensor interface.

[0064] In this embodiment, the gearbox 2 is offset, which facilitates the installation of propeller 1 accessories and aircraft accessories, improving maintainability. This embodiment of the dual-rotor three-shaft turboprop engine allows for the replacement of the field gearbox 2 and the power turbine 5, improving the maintainability and reliability of the dual-rotor three-shaft turboprop engine and reducing maintenance costs.

[0065] Both the integrated air intake device 3 and the rectifier support plate 31 use lubricating oil anti-icing, which ensures the aircraft's anti-icing capability in all weather and across the entire flight envelope, and improves overall safety. The lubricating oil supply and return are built-in cast pipelines, which reduces the number of external pipelines and improves reliability. The integral cast air intake device 3 simplifies the number of parts and improves engine maintainability.

[0066] In one embodiment, such as Figure 6 As shown, the combustion chamber 42 adopts an inclined annular direct current structure, including an inner casing 421, an outer casing 422, a flame tube 423, dual fuel line nozzles 424, a fuel main pipe 425, and an ignition nozzle (not shown in the figure). The front end of the outer casing 422 is bolted to the diffuser casing 426 of the high-pressure compressor 413, and the rear end of the outer casing 422 is connected to the casing of the high-pressure turbine 432. The inner casing 421 is fixed on the diffuser casing 426 of the high-pressure compressor 413. The flame tube 423 adopts a positioning support structure with a fixed front end and a free rear end. That is, the head of the flame tube 423 is fixedly connected to the outer casing 422, and the tail end of the flame tube 423 is located at the air inlet of the high-pressure turbine 432. Both the inner and outer rings of the flame tube 423 are machined from integral forgings.

[0067] The combustion chamber 423 head is equipped with a dual-stage axial vortex generator with opposite rotation directions and a dual-fuel nozzle. The dual-fuel nozzle is matched with the dual-stage axial vortex generator and is connected to the fuel main 425. The centerline of the dual-stage axial vortex generator has an angle with the central axis of the dual-rotor three-shaft turboprop engine to reduce the axial length of the flame tube 423. A compact structural layout is adopted to reduce the cooling area of ​​the flame tube 423. A film cooling structure and thermal barrier coating technology are used to reduce the wall temperature and temperature gradient of the flame tube 423 and improve the service life of the combustion chamber 42.

[0068] The head of the flame tube 423 is also equipped with an intake cap to reduce total pressure loss. Through a two-stage axial vortex and dual-oil-line fuel nozzles, the concentration distribution of fuel in the main combustion zone is controlled, reducing pollution emissions from the combustion chamber 42 and reducing engine smoke under high conditions.

[0069] In one embodiment, such as Figure 7 As shown, the high-pressure turbine 432 and the low-pressure turbine 433 adopt a high-low pressure reversal design. The high-pressure turbine 432 adopts a high-reversal design to obtain a larger outlet airflow pre-swirl angle and reduce the aerodynamic load on the low-pressure guide vanes. The high-pressure turbine 432 includes a high-pressure turbine guide vane 4321, a high-pressure turbine rotor 4322, a high-pressure front sealing disk 4323, and a high-pressure rear sealing disk 4324. The low-pressure turbine 433 includes a low-pressure turbine guide vane 4331, a low-pressure turbine rotor 4332, and a low-pressure front sealing disk 4333.

[0070] Among them, the gas turbine 43 is connected to the outer casing 422 of the combustion chamber 42 and the power turbine 5 through the gas turbine casing 431. The high-pressure turbine guide vane 4321, the high-pressure turbine rotor 4322, the low-pressure turbine guide vane 4331, and the low-pressure turbine rotor 4332 are arranged sequentially inside the gas turbine casing 431 along the exhaust direction. The high-pressure turbine rotor 4322 is connected to the high-pressure rotor shaft 44. The high-pressure front sealing plate 4323 is arranged between the high-pressure turbine guide vane 4321 and the high-pressure turbine rotor 4322. The high-pressure rear sealing plate 4324 is arranged between the high-pressure turbine rotor 4322 and the low-pressure turbine guide vane 4331. The low-pressure front sealing plate 4333 is arranged between the low-pressure turbine guide vane 4331 and the low-pressure turbine rotor 4332.

[0071] The rotors of the high-pressure compressor 413 and the low-pressure compressor 411 are not mechanically connected. The high-pressure turbine rotor 4322 and the low-pressure turbine rotor 4332 are not mechanically connected. The low-pressure compressor 411 has adjustable inlet guide vanes and a venting device. There is an interstage venting function between the high-pressure compressor 413 and the low-pressure compressor 411.

[0072] The high-pressure turbine guide vane 4321, the high-pressure turbine rotor 4322, the low-pressure turbine guide vane 4331, and the low-pressure turbine rotor 4332 all adopt a high-efficiency cooling design to meet the overall high thermodynamic cycle parameter requirements.

[0073] In one embodiment, such as Figure 8 As shown, the power turbine 5 is a three-stage axial flow turbine. The power turbine 5 also includes a power turbine casing 51, a transition section integrated guide vane 53, a turbine first-stage guide vane, a turbine first-stage rotor 54, a turbine second-stage guide vane 55, a turbine second-stage rotor 56, a turbine third-stage guide vane 57, and a turbine third-stage rotor 58. A bearing housing 59 is provided on the exhaust casing 52, and the ninth bearing 14 is installed in the bearing housing 59.

[0074] The transition section integrated guide vane 53 is connected to the gas turbine casing 431 and the power turbine casing 51 at both ends, respectively. The turbine first-stage guide vane, turbine first-stage rotor 54, turbine second-stage guide vane 55, turbine second-stage rotor 56, turbine third-stage guide vane 57 and turbine third-stage rotor 58 are arranged sequentially inside the power turbine casing 51 along the exhaust direction.

[0075] Furthermore, the transition section integrated guide vane 53 and the turbine first-stage guide vane adopt an integrated structure, which effectively shortens the axial length of the engine. The aerodynamics adopts a moderate load coefficient and flow coefficient design, and the rotor blades are crowned. These design features enable the power turbine 5 to obtain a high-efficiency power turbine 5 solution with a wide high-efficiency range.

[0076] In one embodiment, such as Figure 9 and Figure 10 As shown, the accessory drive system includes an accessory drive assembly 61, a central drive assembly 62, an accessory housing 63, a drive starter, an oil-gas separator, a fuel pump regulator, a lubricating oil pump, and an alternator. The accessory drive system also provides a mounting platform for the radiator, speed sensor, and lubricating oil filter.

[0077] The accessory housing 63 and the intermediate housing 412 are connected. The accessory transmission assembly 61 is located inside the accessory housing 63, and the central transmission assembly 62 is located inside the intermediate housing 412. The input end of the central transmission assembly 62 is connected to the high-pressure rotor shaft 44 through gear transmission, and the output end of the central transmission assembly 62 is connected to the accessory transmission assembly 61 through gear transmission. A portion of the power transmitted by the high-pressure rotor shaft 44 of the high-pressure turbine 432 is transmitted to the accessory that needs to extract power through the accessory transmission device.

[0078] In this embodiment, the entire device employs a simple transmission system. The transmission chains of the accessory transmission assembly 61 and the central transmission assembly 62 include three pairs of spiral bevel gears, four cylindrical gears, eight cylindrical roller bearings, and four deep groove ball bearings. Furthermore, the accessory housing 63 is internally equipped with an internal oil passage 64, which is integrally formed with the accessory housing 63. The integrated structure of gears and bearings in the accessory transmission assembly 61 and the central transmission assembly 62 reduces weight and improves reliability and maintainability.

[0079] In this embodiment, a large number of internal pipes are used in the casing to reduce the number of external pipes and improve the reliability of the engine.

[0080] The dual-rotor, three-shaft turboprop engine of this invention has a simple structure and high reliability. It has been implemented in the design, trial production and testing stages, and has passed the test of practical engineering applications. It has achieved the expected design concept and has good results. Under certain conditions, it can be quickly transformed into a practical product or equipment.

[0081] The dual-rotor three-shaft turboprop engine of this invention has a compact structure, high integration, few external pipelines, and is easy to disassemble and maintain.

[0082] The dual-rotor three-shaft turboprop engine of this invention has a smaller starting inertia than the single-rotor gas generator 4, and a faster starting speed under the same starting power.

[0083] The high and low pressure compressors of the dual-rotor three-shaft turboprop engine in this embodiment of the invention operate at their respective optimal speeds, have a large surge margin, low aerodynamic efficiency requirements to achieve the target power, and can maintain a low fuel consumption rate even at non-design points.

[0084] The twin-rotor three-shaft turboprop engine of this invention enables the engine to achieve a higher power level under lower turbine inlet temperature conditions, has a longer hot end life, and has greater development and growth potential within the same frame size, covering twin-rotor three-shaft turboprop engines with power levels of 4500kW to 8000kW.

[0085] Although the present invention has been described in detail with reference to the foregoing embodiments, those skilled in the art should understand that modifications can still be made to the technical solutions described in the foregoing embodiments, or equivalent substitutions can be made to some of the technical features; and these modifications or substitutions do not cause the essence of the corresponding technical solutions to deviate from the spirit and scope of the technical solutions of the embodiments of the present invention.

Claims

1. A dual-rotor, three-shaft turboprop engine, characterized in that, It includes a propeller, a reducer, an intake system, a gas generator, and a power turbine; The reducer is connected to the intake end of the gas generator via the intake device, the exhaust end of the gas generator is connected to the power turbine, the gas generator includes a compressor, a combustion chamber and a gas turbine arranged sequentially along the exhaust direction, one end of the power shaft of the power turbine is connected to the input end of the reducer, and the output end of the reducer is connected to the propeller. The gas generator further includes a low-pressure rotor shaft and a high-pressure rotor shaft; the compressor includes a high-pressure compressor, an intermediate casing, and a low-pressure compressor; the gas turbine includes a gas turbine casing, a high-pressure turbine, and a low-pressure turbine; wherein, the inlet end of the low-pressure compressor is connected to the inlet device, the exhaust end of the low-pressure compressor is connected to the high-pressure compressor through the intermediate casing, the high-pressure turbine is connected to the high-pressure compressor through the high-pressure rotor shaft, the low-pressure turbine is connected to the low-pressure compressor through the low-pressure rotor shaft, the low-pressure rotor shaft passes through the high-pressure rotor shaft, and the power shaft passes through the low-pressure rotor shaft. The combustion chamber includes an inner casing, an outer casing, a flame tube, dual fuel nozzles, and a fuel main. The front end of the outer casing is connected to the diffuser casing of the high-pressure compressor, and the rear end of the outer casing is connected to the casing of the high-pressure turbine. The inner casing is fixed to the diffuser casing of the high-pressure compressor. The head of the flame tube is fixedly connected to the outer casing. The head of the flame tube is also equipped with a two-stage axial vortex generator with opposite rotation directions and dual fuel nozzles. The centerline of the two-stage axial vortex generator forms an angle with the central axis of the dual-rotor three-shaft turboprop engine. The dual fuel nozzles are connected to the fuel main.

2. The dual-rotor, three-shaft turboprop engine according to claim 1, characterized in that, The reducer is an offset reducer, the air intake device is S-shaped, a rectifier support plate is provided inside the air intake device, and the offset reducer is located in front of the air intake device.

3. The dual-rotor, three-shaft turboprop engine according to claim 1, characterized in that, The gas turbine is connected to the power turbine via the gas turbine casing, and the exhaust end of the power turbine is provided with an exhaust casing. The intake device is equipped with a first bearing, a second bearing, and a third bearing; the intermediate casing is equipped with a fourth bearing and a fifth bearing; a sixth bearing is provided between the outlet end of the high-pressure compressor and the high-pressure turbine; the gas turbine casing is equipped with a seventh bearing and an eighth bearing; and the exhaust casing is equipped with a ninth bearing. The high-pressure rotor shaft is supported on the fifth bearing and the sixth bearing; the low-pressure rotor shaft is supported on the third bearing, the fourth bearing, and the seventh bearing; and the power shaft is supported on the first bearing, the second bearing, the eighth bearing, and the ninth bearing.

4. The dual-rotor, three-shaft turboprop engine according to claim 1, characterized in that, The head of the flame tube is also equipped with an air intake cap.

5. The dual-rotor, three-shaft turboprop engine according to claim 1, characterized in that, The high-pressure turbine includes high-pressure turbine guide vanes, a high-pressure turbine rotor, a high-pressure front sealing disc, and a high-pressure rear sealing disc; the low-pressure turbine includes low-pressure turbine guide vanes, a low-pressure turbine rotor, and a low-pressure front sealing disc. The gas turbine is connected to the outer casing and the power turbine via the gas turbine casing. The high-pressure turbine guide vane, the high-pressure turbine rotor, the low-pressure turbine guide vane, and the low-pressure turbine rotor are sequentially arranged inside the gas turbine casing along the exhaust direction. The high-pressure turbine rotor is connected to the high-pressure rotor shaft. The high-pressure front sealing plate is disposed between the high-pressure turbine guide vane and the high-pressure turbine rotor. The high-pressure rear sealing plate is disposed between the high-pressure turbine rotor and the low-pressure turbine guide vane. The low-pressure front sealing plate is disposed between the low-pressure turbine guide vane and the low-pressure turbine rotor.

6. The dual-rotor, three-shaft turboprop engine according to any one of claims 1-5, characterized in that, The power turbine also includes a power turbine casing, a transition section integrated guide vane, a turbine first-stage guide vane, a turbine first-stage rotor, a turbine second-stage guide vane, a turbine second-stage rotor, a turbine third-stage guide vane, and a turbine third-stage rotor; The transition section integrated guide vane is connected to the gas turbine casing and the power turbine casing at both ends, respectively. The turbine first-stage guide vane, the turbine first-stage rotor, the turbine second-stage guide vane, the turbine second-stage rotor, the turbine third-stage guide vane, and the turbine third-stage rotor are sequentially arranged inside the power turbine casing along the exhaust direction.

7. The dual-rotor, three-shaft turboprop engine according to claim 6, characterized in that, The transition section integrated guide vane and the turbine first-stage guide vane are integrally formed.

8. The dual-rotor, three-shaft turboprop engine according to any one of claims 1-5, characterized in that, It also includes an accessory drive device, which is connected to the compressor drive device.

9. The dual-rotor, three-shaft turboprop engine according to claim 8, characterized in that, The accessory drive system includes an accessory housing, a central drive assembly, and an accessory drive assembly; The accessory housing and the intermediate housing are connected. The accessory transmission assembly is disposed in the accessory housing, and the central transmission assembly is disposed in the intermediate housing. The input end of the central transmission assembly is connected to the high-pressure rotor shaft, and the output end of the central transmission assembly is connected to the accessory transmission assembly.

10. The dual-rotor, three-shaft turboprop engine according to claim 9, characterized in that, The accessory housing has an internal oil passage, which is integrally formed with the accessory housing.

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