Aero-engine with shaft-driven fan configuration and control method thereof
Through the design of an aero-engine with a shaft-driven fan configuration, the independent control of the main air inlet and the additional air inlet, combined with a clutch transmission and a flow regulator, switching between high flight power and low fuel consumption modes is achieved, solving the problem of insufficient adaptability of traditional engines to working conditions within a wide speed range and improving the engine's performance and adaptability.
Patent Information
- Application Number
- CN202510734141.4
- Authority / Receiving Office
- CN · China
- Patent Type
- Applications(China)
- Current Assignee / Owner
- Filing Date
- 2025-06-04
- Publication Date
- 2025-09-23
AI Technical Summary
Traditional fixed-configuration aircraft engines perform well under specific flight conditions, but they are difficult to meet the requirements of high flight power and low fuel consumption at the same time within a wide operating range, showing obvious limitations.
An aircraft engine with a shaft-driven fan configuration is adopted. Through the design of the main air inlet and the additional air inlet, combined with a clutch transmission and a flow regulator, switching between different operating modes is achieved. By utilizing independent control of the gas turbine engine and the propulsion fan, the bypass ratio and exhaust system are optimized, and a multi-bypass propulsion system is constructed to achieve switching between high flight power and low fuel consumption modes.
Under different flight conditions, aircraft engines can flexibly switch working modes, improve propulsion efficiency and power, reduce fuel consumption, enhance cycle regulation capabilities, adapt to complex and changing flight environments, and improve overall performance and reliability.
Smart Images
Figure CN120684309A_ABST
Abstract
Description
Technical Field
[0001] The invention relates to the technical field of aircraft power systems and discloses an aero-engine with a shaft-driven fan configuration and a control method thereof. Background Art
[0002] Traditional fixed-configuration aircraft engines typically only have a single operating mode, performing well only under specific flight conditions while failing under other conditions. Future development aims to maintain consistent performance across a wider operating range, but traditional fixed-configuration aircraft engines are unable to simultaneously meet the dual demands of high power and low fuel consumption, demonstrating increasingly significant limitations. Summary of the Invention
[0003] The purpose of the present invention is to provide an aircraft engine with a shaft-driven fan configuration and a control method thereof, which can adopt different working modes in a targeted manner to maximize the satisfaction of the engine performance requirements of different tasks.
[0004] In order to achieve the above technical effects, the technical solution adopted by the present invention is:
[0005] An aircraft engine with a shaft-driven fan configuration, the aircraft engine comprising a main air inlet and an additional air inlet, the main air inlet being an annular structure, the additional air inlet being located on the inner side of an inner ring of the main air inlet; a gas turbine engine being disposed in the main air inlet, a propulsion fan being disposed in the additional air inlet, the propulsion fan being driven and connected to the rotating shaft of the gas turbine engine via a clutch transmission; a nozzle assembly being disposed at the tail of the main air inlet, and an additional exhaust assembly being disposed at the tail of the additional air inlet.
[0006] Furthermore, an inlet flow regulator is provided at the inlet end of the additional air inlet duct, for controlling the opening of the inlet end of the additional air inlet duct.
[0007] Furthermore, an air intake cone is provided at the inlet end of the additional air intake duct, one end of the inlet flow regulator is hinged to the air intake cone, and the other end is located at a position away from the incoming flow direction.
[0008] Furthermore, an outlet flow regulator is provided in the additional exhaust component for controlling the opening of the outlet end of the additional exhaust component.
[0009] Furthermore, the outlet flow regulator is a common regulating component located between the nozzle assembly and the additional exhaust assembly, so that the nozzle assembly and the additional exhaust assembly are both single-sided exhaust nozzle configurations.
[0010] Furthermore, the main air inlet duct is configured as an S-shaped flow path at an upstream position of the gas turbine engine.
[0011] To achieve the above technical effects, the present invention further provides a control method for an aircraft engine with a shaft-driven fan configuration. The control method is based on the aircraft engine with the shaft-driven fan configuration, and includes:
[0012] Constructing a simulation model of the aircraft engine, and using the simulation model to conduct simulation analysis under multiple typical operating points to obtain the flight power and fuel consumption rate of the aircraft engine under different combinations of gas turbine engine fuel flow rate, nozzle throat area, and propulsion fan speed;
[0013] The fuel flow rate, nozzle throat area, propulsion fan speed and fuel consumption rate of the gas turbine engine at the design point are used to perform dimensionless processing on the fuel flow rate, nozzle throat area, propulsion fan speed and fuel consumption rate of the gas turbine engine at each typical operating point, and the dimensionless fuel flow rate, dimensionless nozzle throat area, dimensionless propulsion fan speed and dimensionless fuel consumption rate corresponding to the typical operating point are obtained;
[0014] Taking the flight power requirements of a typical operating point as constraints, based on the dimensionless fuel flow rate, dimensionless nozzle throat area, dimensionless speed of the propulsion fan, and dimensionless fuel consumption rate, weight coefficients corresponding to the dimensionless fuel flow rate, dimensionless nozzle throat area, dimensionless speed of the propulsion fan, and dimensionless fuel consumption rate are given respectively, and an objective function for aircraft engine parameter optimization is constructed; and combined with a simulation model of the aircraft engine, the fuel flow rate, nozzle throat area, and propulsion fan speed of the gas turbine engine corresponding to the minimum value of the objective function are determined as the optimal adjustment parameters for the current operating point, so that the gas turbine engine drives the propulsion fan to operate through the clutch transmission.
[0015] Furthermore, based on the dimensionless fuel flow rate, dimensionless nozzle throat area, dimensionless propulsion fan speed and dimensionless fuel consumption rate, the objective function of aircraft engine parameter optimization is constructed as follows: Where J is the objective function value, w f is the fuel flow rate of the gas turbine engine at one of the typical operating points, A8 is the nozzle throat area corresponding to the typical operating point, N f is the propulsion fan speed corresponding to the typical operating point, SFC is the fuel consumption rate corresponding to the typical operating point, a1 is the fuel flow weight coefficient of the gas turbine engine under the typical operating point, a2 is the nozzle throat area weight coefficient under the typical operating point, a3 is the propulsion fan speed weight coefficient under the typical operating point, a4 is the fuel consumption rate weight coefficient under the typical operating point, is the fuel flow rate of the gas turbine engine at the design point, A 8,design is the nozzle throat area at the design point, SFC designis the fuel consumption rate at the design point, N f,design is the propulsion fan speed at the design point;
[0016] Furthermore, under the condition that the flight power requirement is a constraint, in order to achieve the optimal fuel consumption rate, the w f / w f,design The weight coefficient is not greater than 1, the A8 / A 8,design The weight coefficient is not greater than 1, the N f / N f,design The weight coefficient is not greater than 1, the SFC / SFC design The weight coefficient is not less than 5.
[0017] Compared with the prior art, the present invention has the following beneficial effects:
[0018] 1. In a low fuel consumption mode in which the aircraft's fuel consumption is lower than a preset fuel consumption value, the aircraft engine of the present invention can have both the inlet flow regulator and the outlet flow regulator open, utilizing the gas turbine engine's rotating shaft to drive the propulsion fan, thereby increasing the bypass ratio of the entire propulsion system, improving propulsion efficiency, and reducing fuel consumption. In a high flight power mode in which the aircraft requires power greater than a preset power value, both the inlet flow regulator and the outlet flow regulator are closed, the clutch transmission disengages the fan and the engine's rotating shaft, and power is generated solely by the gas turbine engine. The engine operates in a low bypass ratio mode, thereby improving flight power.
[0019] 2. The aircraft engine with a shaft-driven fan configuration of the present invention can also utilize the speed regulation function of the clutch transmission to achieve regulation of the propulsion fan, so that the aircraft engine has stronger cycle regulation capability and adaptability, and can adopt different working modes in a targeted manner to maximize the satisfaction of the engine performance requirements of different tasks. BRIEF DESCRIPTION OF THE DRAWINGS
[0020] Figure 1 This is a schematic diagram of an aircraft engine with a shaft-driven fan configuration operating in a low fuel consumption mode in an embodiment;
[0021] Figure 2 Schematic diagram of an aircraft engine with a shaft-driven fan configuration operating in a high-flight power mode in an embodiment;
[0022] Among them, 1. Main air inlet duct; 2. Additional air inlet duct; 3. Gas turbine engine; 4. Propulsion fan; 5. Clutch transmission; 6. Nozzle assembly; 7. Additional exhaust assembly; 8. Inlet flow regulator; 9. Inlet cone; 10. Outlet flow regulator; 11. Low-pressure drive shaft. DETAILED DESCRIPTION
[0023] The present invention will be described in further detail below with reference to the embodiments and accompanying drawings. However, this should not be construed as limiting the scope of the present invention to the following embodiments, as all technologies implemented based on the present invention fall within the scope of the present invention.
[0024] Example
[0025] See also Figure 1 、 Figure 2 , an aircraft engine with a shaft-driven fan configuration, the aircraft engine includes a main air inlet 1 and an additional air inlet 2, the main air inlet 1 is a ring-shaped structure, and the additional air inlet 2 is located on the inner side of the inner ring of the main air inlet 1; a gas turbine engine 3 is arranged in the main air inlet 1, and a propulsion fan 4 is arranged in the additional air inlet 2, and the propulsion fan 4 is driven and connected to the rotating shaft of the gas turbine engine 3 through a clutch transmission 5; a nozzle assembly 6 is arranged at the tail of the main air inlet 1, and an additional exhaust assembly 7 is arranged at the tail of the additional air inlet 2.
[0026] In this embodiment, the engine shaft is connected to the input end of the clutch transmission 5 via a low-pressure drive shaft 11, and the output end of the clutch transmission 5 is connected to the propulsion fan 4. In a low-fuel consumption mode where the aircraft's fuel consumption is lower than a preset fuel consumption value, the gas turbine engine 3 shaft drives the propulsion fan 4, increasing the bypass ratio of the entire propulsion system, improving propulsion efficiency, and reducing fuel consumption. In a high-power flight mode where the aircraft requires power greater than a preset power value, the clutch transmission 5 disconnects the fan and engine shafts, generating power solely from the gas turbine engine 3. The engine operates in a low-bypass ratio mode, thereby increasing flight power. The shaft-driven fan configuration of the present invention can also utilize the speed regulation function of the clutch transmission 5 to achieve control of the propulsion fan 4, giving the aircraft engine stronger cycle regulation capabilities and adaptability. Different operating modes can be used in a targeted manner to maximize the performance requirements of different missions.
[0027] In this embodiment, an inlet flow regulator 8 is provided at the inlet end of the additional air inlet duct 2 to control the opening of the inlet end of the additional air inlet duct 2. This allows for regulation of the air flow rate of the additional air inlet duct 2 in low fuel consumption mode, thereby optimizing the engine's air intake efficiency. In low fuel consumption mode, inlet flow regulator 8 dynamically adjusts the air flow rate based on the specific operating conditions and fuel consumption requirements of the aircraft, ensuring high-efficiency engine operation and further improving engine fuel economy.
[0028] In this embodiment, an intake cone 9 is provided at the inlet end of the additional intake duct 2. This cone 9 also provides a degree of rectification and guidance for the airflow entering the additional intake duct 2, allowing it to enter the engine more smoothly and helping to improve the engine's intake efficiency and stability. Furthermore, the hinged connection between the intake cone 9 and the inlet flow regulator 8, with one end hinged to the cone 9 and the other end positioned away from the incoming flow, reduces the flow resistance of the inlet flow regulator 8 to the incoming flow.
[0029] An outlet flow regulator 10 is provided in the additional exhaust component 7 for controlling the opening of the outlet end of the additional exhaust component 7. The outlet flow regulator 10 of the additional exhaust component 7 can be closed in the high flight power mode, or in the low fuel consumption mode, the outlet flow regulator 10 and the inlet flow regulator 8 can be matched and adjusted according to the specific working conditions and thrust requirements of the aircraft to adapt to the power requirements of the aircraft in different flight phases, so that the aircraft engine of this embodiment can better adapt to the complex and changeable flight environment and improve the overall performance and reliability of the aircraft.
[0030] In this embodiment, the outlet flow regulator 10 is a common adjustment component located between the nozzle assembly 6 and the additional exhaust assembly 7, so that the nozzle assembly 6 and the additional exhaust assembly 7 are both single-sided exhaust nozzle configurations, which not only simplifies the structure of the exhaust system, but also avoids mutual interference between the main and auxiliary while achieving adjustment; the common adjustment component precisely controls the exhaust flow of the nozzle assembly 6 and the additional exhaust assembly 7. Under different flight phases and power requirements, the common adjustment component can respond quickly and adjust the exhaust flow to ensure that the engine always operates in the best state, thereby achieving optimized management of the engine exhaust system.
[0031] In this embodiment, the main air inlet 1 is configured as an S-shaped flow path at the upstream position of the gas turbine engine 3, which effectively reduces the electromagnetic detectability of the aircraft, so that the aircraft engine of this embodiment has stronger concealment capability while maintaining high performance, providing more comprehensive performance guarantees for modern aircraft.
[0032] To address the technical shortcomings of traditional aircraft engines, which, due to their fixed configuration, only achieve optimal performance under specific flight conditions and are unable to adapt to flight conditions over a wide speed range, this embodiment proposes a novel and flexible shaft-driven fan-configured aircraft engine. By innovatively proposing a combined inlet and engine design, and adopting a structural layout where the propulsion fan is located within an independent duct, dual-mode control of "high flight power and low fuel consumption" is achieved with minimal modification cost. By utilizing low-pressure shaft power diversion and a modular external fan design, an independently controllable multi-duct propulsion system is constructed, breaking through the coupling limitations of traditional engine cycle parameters and significantly improving operating condition adaptability. Furthermore, a breakthrough is achieved in the dynamic coordination of the propulsion fan and inlet geometric parameters, effectively resolving the problem of the traditional fixed-configuration engine's limited operating condition adaptability over a wide speed range. This significantly improves the engine's aerodynamic performance adjustment margin and multi-condition adaptability, effectively overcoming the narrow operating range and insufficient adjustment capability of existing engines, and providing core power support for the expansion of aircraft multi-mission performance.
[0033] Based on the same inventive concept, this embodiment further provides a control method for an aircraft engine with a shaft-driven fan configuration. The control method is based on the aircraft engine with the shaft-driven fan configuration, and includes:
[0034] Step 1: construct a simulation model of the aircraft engine, and use the simulation model to conduct simulation analysis under multiple typical operating points to obtain the flight power and fuel consumption rate of the aircraft engine under different combinations of fuel flow rate of the gas turbine engine 3, nozzle throat area, and propulsion fan 4 speed;
[0035] Step 2: Using the fuel flow rate, nozzle throat area, propulsion fan 4 speed, and fuel consumption rate of the gas turbine engine 3 at the design point, perform dimensionless processing on the fuel flow rate, nozzle throat area, propulsion fan 4 speed, and fuel consumption rate of the gas turbine engine at each typical operating point, thereby obtaining the dimensionless fuel flow rate, dimensionless nozzle throat area, dimensionless speed, and dimensionless fuel consumption rate corresponding to the typical operating point;
[0036] Step 3: Taking the flight power requirements of the typical operating point as constraints, based on the dimensionless fuel flow rate, dimensionless nozzle throat area, dimensionless speed of the propulsion fan 4 and dimensionless fuel consumption rate, weight coefficients corresponding to the dimensionless fuel flow rate, dimensionless nozzle throat area, dimensionless speed of the propulsion fan and dimensionless fuel consumption rate are given respectively, and an objective function for aircraft engine parameter optimization is constructed; and combined with the simulation model of the aircraft engine, the fuel flow rate, nozzle throat area and speed of the propulsion fan 4 corresponding to the minimum value of the objective function are determined as the optimal adjustment parameters for the current operating point, so that the gas turbine engine 3 drives the propulsion fan 4 to work through the clutch transmission.
[0037] In this embodiment, based on the fuel flow rate, nozzle throat area, propulsion fan 4 speed and fuel consumption rate of the gas turbine engine 3, the objective function of the aircraft engine parameter optimization is constructed as follows: Where J is the objective function value, w f is the fuel flow rate of the gas turbine engine at one of the typical operating points, A8 is the nozzle throat area corresponding to the typical operating point, N f is the speed of the propulsion fan 4 corresponding to the typical operating point, SFC is the fuel consumption rate corresponding to the typical operating point, a1 is the fuel flow weight coefficient of the gas turbine engine 3 under the typical operating point, a2 is the nozzle throat area weight coefficient under the typical operating point, a3 is the speed weight coefficient of the propulsion fan 4 under the typical operating point, a4 is the fuel consumption rate weight coefficient under the typical operating point, is the fuel flow rate of gas turbine engine 3 at the design point, A 8,design is the nozzle throat area at the design point, SFC design is the fuel consumption rate at the design point, N f,design is the speed of propulsion fan 4 at the design point.
[0038] As in some other embodiments, under the condition that the flight power requirement is a constraint, in order to achieve the best fuel consumption rate, the w f / w f,design The weight coefficient is not greater than 1, the A8 / A 8,design The weight coefficient is not greater than 1, the N f / N f,design The weight coefficient is not greater than 1, the SFC / SFC design The weight coefficient is not less than 5.
[0039] In this embodiment, the fuel flow rate, nozzle throat area, propulsion fan 4 speed, and fuel consumption rate of the gas turbine engine 3 are selected as the core control variables of the objective function. This aims to address the poor operating adaptability and efficiency bottlenecks caused by the strong coupling of aerodynamic parameters in traditional engines through multi-parameter collaborative optimization, thereby breaking through the narrow adjustment range of cycle parameters in traditional engine control modes. While ensuring flight power constraints, this design significantly reduces operating point matching losses and improves aircraft propulsion efficiency through comprehensive optimization of the control parameters of fuel flow rate, nozzle throat area, propulsion fan 4 speed, and fuel consumption rate. Compared to the mechanical adjustment scheme of traditional variable duct engines, the aircraft engine in this embodiment uses decoupled control between the main inlet 1 and the additional inlet 2, reducing structural modifications. This achieves optimal operating point configuration while maintaining core engine stability, effectively resolving the industry's current challenges of achieving both power and fuel efficiency and the high complexity of component matching.
[0040] The above are only preferred embodiments of the present invention and are not intended to limit the present invention. Any modifications, equivalent substitutions and improvements made within the spirit and principles of the present invention should be included in the scope of protection of the present invention.
Claims
1. An aircraft engine with a shaft-driven fan configuration, characterized in that: The aircraft engine includes a main air inlet and an additional air inlet. The main air inlet is an annular structure, and the additional air inlet is located on the inner side of the inner ring of the main air inlet. A gas turbine engine is arranged in the main air inlet, and a propulsion fan is arranged in the additional air inlet. The propulsion fan is connected to the rotating shaft of the gas turbine engine through a clutch transmission. A nozzle assembly is arranged at the tail of the main air inlet, and an additional exhaust assembly is arranged at the tail of the additional air inlet.
2. The aero-engine of shaft-driven fan configuration according to claim 1, characterized in that: An inlet flow regulator is provided at the inlet end of the additional air inlet duct, which is used to control the opening degree of the inlet end of the additional air inlet duct.
3. The aero-engine of shaft-driven fan configuration according to claim 2, characterized in that: An air intake cone is provided at the inlet end of the additional air intake duct, one end of the inlet flow regulator is hinged to the air intake cone, and the other end is located at a position away from the incoming flow direction.
4. The aircraft engine of the shaft-driven fan configuration according to claim 1, characterized in that: An outlet flow regulator is provided in the additional exhaust component for controlling the opening of the outlet end of the additional exhaust component.
5. The aircraft engine of the shaft-driven fan configuration according to claim 4, characterized in that: The outlet flow regulator is a common regulating component located between the nozzle assembly and the additional exhaust assembly, so that both the nozzle assembly and the additional exhaust assembly are of a single-sided exhaust nozzle configuration.
6. The aircraft engine of the shaft-driven fan configuration according to claim 1, characterized in that: The main air inlet duct is configured as an S-shaped flow path at an upstream position of the gas turbine engine.
7. A control method for an aircraft engine with a shaft-driven fan configuration, the control method being based on the aircraft engine with a shaft-driven fan configuration according to claim 1, characterized in that: include: Constructing a simulation model of the aircraft engine, and using the simulation model to conduct simulation analysis under multiple typical operating points to obtain the flight power and fuel consumption rate of the aircraft engine under different combinations of gas turbine engine fuel flow rate, nozzle throat area, and propulsion fan speed; The fuel flow rate, nozzle throat area, propulsion fan speed and fuel consumption rate of the gas turbine engine at the design point are used to perform dimensionless processing on the fuel flow rate, nozzle throat area, propulsion fan speed and fuel consumption rate of the gas turbine engine at each typical operating point, and the dimensionless fuel flow rate, dimensionless nozzle throat area, dimensionless propulsion fan speed and dimensionless fuel consumption rate corresponding to the typical operating point are obtained; Taking the flight power requirements of a typical operating point as constraints, based on the dimensionless fuel flow rate, dimensionless nozzle throat area, dimensionless speed of the propulsion fan, and dimensionless fuel consumption rate, weight coefficients corresponding to the dimensionless fuel flow rate, dimensionless nozzle throat area, dimensionless speed of the propulsion fan, and dimensionless fuel consumption rate are given respectively, and an objective function for aircraft engine parameter optimization is constructed; and combined with a simulation model of the aircraft engine, the fuel flow rate, nozzle throat area, and propulsion fan speed of the gas turbine engine corresponding to the minimum value of the objective function are determined as the optimal adjustment parameters for the current operating point, so that the gas turbine engine drives the propulsion fan to operate through the clutch transmission.
8. The method for controlling an aircraft engine with a shaft-driven fan configuration according to claim 7, wherein: Based on the dimensionless fuel flow rate, dimensionless nozzle throat area, dimensionless speed of propulsion fan and dimensionless fuel consumption rate, the objective function of aircraft engine parameter optimization is constructed as follows: Where J is the objective function value, w f is the fuel flow rate of the gas turbine engine at one of the typical operating points, A8 is the nozzle throat area corresponding to the typical operating point, N f is the propulsion fan speed corresponding to the typical operating point, SFC is the fuel consumption rate corresponding to the typical operating point, a1 is the fuel flow weight coefficient of the gas turbine engine under the typical operating point, a2 is the nozzle throat area weight coefficient under the typical operating point, a3 is the propulsion fan speed weight coefficient under the typical operating point, a4 is the fuel consumption rate weight coefficient under the typical operating point, is the fuel flow rate of the gas turbine engine at the design point, A 8,design is the nozzle throat area at the design point, SFC design is the fuel consumption rate at the design point, N f,design is the propulsion fan speed at the design point.
9. The method for controlling an aircraft engine with a shaft-driven fan configuration according to claim 7, characterized in that: Under the condition that the flight power requirement is a constraint, in order to achieve the optimal fuel consumption rate, the w f / w f,design The weight coefficient is not greater than 1, the A8 / A 8,design The weight coefficient is not greater than 1, the N f / N f,design The weight coefficient is not greater than 1, the SFC / SFC design The weight coefficient is not less than 5.