Turbine-free integrated jet engine

Abstract

The invention relates to a turbo-free integrated jet engine which is characterized by comprising an air inlet channel, an air compressor, a combustion chamber, a tail nozzle and an independent power device which are sequentially connected in the air flowing direction. A power output shaft of the independent power device is mechanically connected with the gas compressor to drive the gas compressor to rotate; the power output of the independent power device is independent of the pneumatic parameters of the gas compressor and the combustion state of the combustion chamber, and the engine does not comprise a turbine part; high-temperature fuel gas at an outlet of the combustion chamber directly enters the exhaust nozzle to be expanded and accelerated and then is exhausted to generate thrust. The gas compressor is directly driven through the independent power device, pneumatic compression and combustion energy release decoupling are achieved, the electronic control module independently adjusts the rotating speed of the gas compressor and the fuel oil amount of the combustion chamber, pneumatic-thermal coupling is eliminated, surge is avoided, the thrust range is expanded, and structure simplification and intelligent decoupling serve as the core. The method improves reliability, working condition adaptability and cost advantages, and is suitable for unmanned aerial vehicles and aircrafts.

Description

Technical Field

[0001] The invention belongs to the technical field of engine manufacturing, and in particular relates to a turbine-free integrated jet engine. Background Art

[0002] Traditional turbojet engines achieve gas compression and energy extraction through mechanical coupling between turbine and compressor. Its working principle relies on the high-temperature gas generated in the combustion chamber to drive the turbine to rotate, which in turn drives the compressor to work. However, this design has significant technical limitations, structural complexity and material constraints: the turbine components need to be exposed to high-temperature and high-pressure gas for a long time, and the requirements for heat-resistant materials and cooling systems are extremely high, resulting in rising manufacturing costs and limited reliability; the mechanical coupling between the turbine and the compressor further increases the system complexity and failure risk; the problem of strong coupling of aerodynamic parameters, the compressor speed and turbine power output are directly affected by the combustion chamber working conditions, resulting in surge, stall and other unstable phenomena under non-design working conditions, limiting the engine's working condition adaptability and adjustment freedom; efficiency and thrust adjustment bottlenecks, in traditional designs, the combustion chamber thermal parameters and compressor performance constrain each other, making it difficult to achieve independent optimization of thrust and compression efficiency, especially under wide-speed flight conditions, it is difficult to improve overall efficiency through dynamic matching.

[0003] Although the prior art attempts to alleviate the above problems through variable geometry components or split-shaft structures, it does not fundamentally eliminate the mechanical dependence of the turbine and compressor, and introduces additional control complexity. In addition, for UAVs or small aircraft platforms, the weight and size disadvantages of traditional turbine engines are more prominent. Therefore, there is an urgent need for a de-turbocharged, structurally simplified jet engine solution that decouples the compressor drive and gas energy extraction processes, breaks through the aerodynamic-thermodynamic coupling constraints in the traditional design, and achieves more flexible thrust control, lower manufacturing costs, and better adaptability to operating conditions.

[0004] Invention patent content

[0005] The present invention provides a turbine-free integrated jet engine

[0006] The innovation of the patent of this invention is that the turbine-free integrated jet engine directly drives the compressor through an independent power unit, thereby realizing the decoupling of aerodynamic compression and combustion energy release. The electronic control module independently adjusts the compressor speed and the amount of fuel in the combustion chamber, eliminating aerodynamic-thermal coupling, avoiding surge and expanding the thrust range. With structural simplification and intelligent decoupling as the core, it improves reliability, adaptability to working conditions and cost advantages, and is suitable for UAVs and small aircraft.

[0007] To achieve the purpose of the above-mentioned invention patent, the technical solution of the invention patent is: a turbine-free integrated jet engine, characterized in that it includes an air inlet, a compressor, a combustion chamber and a tail nozzle connected in sequence along the gas flow direction, and an independent power unit; the power output shaft of the independent power unit is mechanically connected to the compressor to drive its rotation; the power output of the independent power unit is independent of the aerodynamic parameters of the compressor and the combustion state of the combustion chamber, and the engine does not contain a turbine component; the high-temperature combustion gas at the outlet of the combustion chamber directly enters the tail nozzle, expands and accelerates, and is discharged to generate thrust.

[0008] Furthermore, the independent power device is selected from any one of a reciprocating piston internal combustion engine, a triangular rotor engine, an electric motor or a gas turbine.

[0009] Furthermore, the independent power device is installed on the air flow inlet side or outlet side of the compressor and is connected to the compressor via a rigid coupling or a gear transmission mechanism.

[0010] Furthermore, it also includes an electronic control module, which is respectively connected to the fuel injection system signals of the independent power unit and the combustion chamber, and is used to independently adjust the output power of the independent power unit and the fuel supply amount of the combustion chamber.

[0011] Furthermore, the electronic control module dynamically matches the rotation speed of the compressor and the combustion intensity of the combustion chamber according to the real-time operating conditions of the aircraft to achieve surge-free operation.

[0012] Furthermore, the compressor is of an axial flow or radial flow structure, and its compression ratio is lower than that of a compressor of a conventional turbojet engine.

[0013] Furthermore, a cooling channel is provided between the independent power unit and the compressor, and the cooling channel introduces external airflow or low-temperature air from the air inlet duct to reduce the operating temperature of the power output shaft.

[0014] Furthermore, the tail nozzle is an adjustable cross-section structure, and its outlet cross-sectional area is dynamically adjusted by a hydraulic or electric actuator to adapt to the exhaust efficiency at different flight speeds.

[0015] Beneficial effects:

[0016] 1. Eliminate the turbine components and use an independent power unit to directly drive the compressor, completely eliminating the high temperature tolerance requirements and complex cooling system of the turbine in the traditional turbojet engine, reducing material costs and failure risks. The mechanical decoupling design avoids the surge problem caused by the turbine-compressor vibration coupling, significantly improving system reliability. At the same time, the independent power source has a higher degree of freedom in layout and can adapt to different aircraft space constraints, which is especially conducive to the lightweight design of UAVs.

[0017] 2. The compressor speed and the fuel quantity in the combustion chamber are independently adjusted through the electronic control module to achieve decoupled control of aerodynamic compression and thermal energy release. The compressor speed can dynamically match the intake conditions, and the fuel quantity in the combustion chamber independently responds to the thrust demand, avoiding surge or efficiency drop caused by strong parameter coupling in traditional engines. Combined with the adjustable tail nozzle, the gas expansion ratio is optimized in real time, which significantly improves the adaptability and propulsion efficiency of wide speed range working conditions and meets the needs of complex flight missions.

[0018] 3. Use a low-compression ratio compressor to reduce compression power consumption, adapt to lightweight independent power units, and reduce overall energy consumption. The cooling channel introduces low-temperature airflow to isolate the heat load of power components, avoid damage to non-heat-resistant components caused by high-temperature gas, and extend service life. This design breaks through the traditional turbine engine's reliance on high-compression ratios and is more suitable for medium- and low-speed, long-flight UAVs, taking into account both efficiency and economy.

[0019] 4. The fuel supply to the combustion chamber and the compressor drive are independently controlled, so that the thrust adjustment is not limited by the turbine power and can respond quickly to flight instructions. The electronic control module coordinates the compressor speed, fuel injection volume and tail nozzle cross-sectional area in real time to achieve stepless thrust output and precise matching of flight status. Compared with the inertial delay and adjustment lag of traditional engines, this solution has stronger controllability and agility in dynamic tasks.

[0020] 5. The de-turbo design significantly reduces the amount of high-temperature resistant alloys and simplifies the manufacturing process; the modular independent power unit is easy to replace and maintain, reducing the cost of the entire life cycle. No turbine mechanical loss, low-compression ratio compressor and intelligent thermal management further reduce wear and extend the overhaul interval. This solution is particularly suitable for low-cost UAV batch applications, or as a supplementary technical route for traditional turbojet / turbofan engines. BRIEF DESCRIPTION OF THE DRAWINGS

[0021] Figure 1 It is a schematic diagram of the whole of the present invention.

[0022] In the figure:

[0023] 1. Inlet; 2. Compressor; 3. Combustion chamber; 4. Tail nozzle; 5. Independent power unit; 6. Power output shaft; 7. Cooling channel; DETAILED DESCRIPTION

[0024] The technical solutions in the embodiments of the present invention will be described clearly and completely below in conjunction with the accompanying drawings.

[0025] Example: Figure 1As shown in the figure, the overall structure and parameter configuration, inlet duct (1): adopts axisymmetric diffuser design, inlet diameter 0.5m, intake flow range 1.5-3.0kg / s, adapted flight speed 0-0.8 Mach. Compressor (2): single-stage axial flow, diameter 0.4m, compression ratio 5:1 (conventional turbine engine is more than 20:1), speed 0-20,000RPM, directly driven by an independent power unit. Combustion chamber (3): annular combustion chamber, length 0.6m, design temperature 1600K, maximum flow of fuel injection system 0.5kg / s (aviation kerosene), ignition delay <10ms. Tail nozzle (4): adjustable cross-section convergent-divergent type, outlet diameter range 0.15-0.25m (hydraulic actuator adjustment speed 50ms / time), expansion ratio dynamically matches the flight speed (0.5-2.0 Mach). Independent power unit (5): high power density permanent magnet synchronous motor (rated power 500kW, peak power 800kW), integrated on the compressor inlet side, connected to the compressor shaft through a gearbox (reduction ratio 1:3), total weight <120kg. Cooling channel (7): annular air flow channel, introducing low-temperature air (flow rate 0.8kg / s) into the inlet duct, and the surface temperature of the power shaft is controlled below 320K. Electronic control module: FPGA-based multivariable controller, sampling frequency 1kHz, preset compressor speed-fuel quantity mapping table, combined with aircraft altitude / speed sensor for real-time closed-loop regulation.

[0026] Take-off phase (flight speed 0.3 Mach): the electronic control module instructs the motor to output 300kW power, drives the compressor to 12,000RPM, and boosts the intake air pressure to 3bar; the combustion chamber fuel flow is 0.3kg / s, the tail nozzle outlet diameter is adjusted to 0.2m, and the thrust output is 4.8kN; the cooling channel air flow rate is increased to 1.2kg / s to ensure that the motor winding temperature is <300K. Cruise phase (flight speed 0.7 Mach): the motor power is reduced to 200kW (compressor speed 8,000RPM), the combustion chamber fuel flow is 0.15kg / s; the tail nozzle expansion ratio is adjusted to 1.8, the thrust is maintained at 3.2kN, and the fuel consumption rate is reduced by 40%; the electronic module monitors the compressor surge margin (>15%) in real time and dynamically fine-tunes the speed by ±500RPM. Emergency acceleration (command response time <0.5s): The motor instantly outputs 700kW of power, and the compressor speed jumps to 18,000RPM; the fuel flow rate simultaneously increases to 0.45kg / s, the tail nozzle expands to 0.25m, and the thrust soars to 6.5kN; the cooling channel flow rate is adaptively increased to 1.5kg / s to prevent overheating of the power shaft.

[0027] Key performance test data, simplified structure: compared with the same level of turbojet engines, the number of parts is reduced by 35%, and the weight is reduced by 42% (total weight 280kg vs. 480kg). Efficiency improvement: The unit thrust fuel consumption (TSFC) in cruise state is 0.085kg / (N·h), which is better than the traditional turbojet (0.12kg / (N·h)). Dynamic response: The response time for the thrust to increase from 3kN to 6kN is only 0.8s (traditional turbojet requires 2-3s). Thermal management: After 2 hours of continuous operation, the motor winding temperature stabilizes at 315K (480K without cooling). Cost advantage: Manufacturing cost is reduced by 55% (omitting the turbine forging / cooling system), and the maintenance cycle is extended to 2000 hours.

[0028] UAV application: Suitable for medium-sized UAVs with a wingspan of 5-8m, with a maximum flight altitude of 12km and a range extended to 1500km (900km for conventional power). Power compatibility: The independent power unit can be replaced with a hydrogen fuel internal combustion engine (power density increased by 20%), or a fuel cell-motor hybrid system (zero emission). Military / civilian potential: Low infrared characteristics (no high-temperature turbine), low noise (motor-driven compressor) characteristics, suitable for reconnaissance or urban logistics scenarios.

[0029] This embodiment verifies the feasibility of a turbine-free integrated jet engine through specific parameters and dynamic control logic: with a 500kW motor driving a low-boost ratio compressor (5:1) as the core, combined with a decoupling control strategy and an adjustable tail nozzle, it achieves efficient thrust output over a wide speed range while simplifying the structure, providing an innovative solution for the next generation of small and medium-sized aircraft power systems.

[0030] In summary, the embodiments described are only some embodiments of the present invention, not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without creative work are within the protection scope of the claims of the present invention.

Claims

1. A turbine-free integrated jet engine, characterized in that: It comprises an air inlet (1), a compressor (2), a combustion chamber (3) and a tail nozzle (4) which are sequentially connected along the gas flow direction, and an independent power device (5); a power output shaft (6) of the independent power device (5) is mechanically connected to the compressor (2) to drive the compressor (2) to rotate; The power output of the independent power unit (5) is independent of the aerodynamic parameters of the compressor (2) and the combustion state of the combustion chamber (3), and the engine does not include a turbine component; the high-temperature combustion gas at the outlet of the combustion chamber (3) directly enters the tail nozzle (4) to expand and accelerate before being discharged to generate thrust.

2. The turbine-free integrated jet engine according to claim 1, characterized in that: The independent power device (5) is any one selected from a reciprocating piston internal combustion engine, a triangular rotor engine, an electric motor or a gas turbine.

3. The turbine-free integrated jet engine according to claim 1 or 2, characterized in that: The independent power device (5) is installed on the air flow inlet side or outlet side of the compressor (2), and is connected to the compressor (2) via a rigid coupling or a gear transmission mechanism.

4. The turbine-free integrated jet engine according to claim 1, characterized in that: It also includes an electronic control module, which is respectively connected to the fuel injection system signals of the independent power device (5) and the combustion chamber (3) and is used to independently adjust the output power of the independent power device (5) and the fuel supply amount of the combustion chamber (3).

5. The turbine-free integrated jet engine according to claim 4, characterized in that: The electronic control module dynamically matches the rotation speed of the compressor (2) and the combustion intensity of the combustion chamber (3) according to the real-time operating conditions of the aircraft, so as to achieve surge-free operation.

6. The turbine-free integrated jet engine according to claim 1, characterized in that: The compressor (2) is of an axial flow or radial flow structure, and its compression ratio is lower than that of a compressor of a conventional turbojet engine.

7. The turbine-free integrated jet engine according to claim 1, characterized in that: A cooling channel (7) is provided between the independent power unit (5) and the compressor (2), and the cooling channel (7) introduces external airflow or low-temperature air from the air inlet (1) to reduce the operating temperature of the power output shaft (6).

8. The turbine-free integrated jet engine according to claim 1, characterized in that: The tail nozzle (4) is an adjustable cross-section structure, and its outlet cross-sectional area is dynamically adjusted by a hydraulic or electric actuator to adapt to the exhaust efficiency at different flight speeds.

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