A combined compressor turbojet engine

Abstract

The present application belongs to the field of engine technology, and provides a combined compressor type turbojet engine, which comprises a compressor, and the air inlet end of the turbojet engine is provided with an axial flow compressor, the axial flow compressor is connected with the shaft of the turbojet engine, and is located upstream of the compressor. The present application can improve the compression ratio of the engine, thereby improving the power and reducing the fuel consumption.

Description

Technical Field

[0001] This invention belongs to the field of engine technology and provides a combined compressor turbojet engine. Background Technology

[0002] In the fields of target drones, target missiles, and small and medium-sized loitering aircraft, low-cost, short-life micro-turbojet engines are the mainstream power plants. These engines adopt a classic architecture of a single-stage compressor, a combustion chamber, and a single-stage axial turbine, which features a simple structure, convenient assembly, and strong operational stability, and can meet the power requirements of basic flight missions.

[0003] Currently, the core performance of these micro-turbojet engines is limited by the design capabilities of the compressor. The pressure ratio of a single-stage compressor can only reach 4 to 5, resulting in significant bottlenecks in compression efficiency and boosting capacity. This directly leads to a fixed limitation on the overall engine performance: the engine's fuel consumption rate has long remained around 1.35 kg / h / daN, and the thrust-to-weight ratio is only around 7, making it impossible to achieve performance breakthroughs. When the aircraft performs medium- to long-range loitering missions, the poor fuel economy and insufficient power output of the engine generally result in insufficient fuel reserves and short range, making it difficult to meet the requirements of long-range operations.

[0004] In summary, existing micro-turbojet engines, due to their excessively low compressor pressure ratio, cannot achieve the dual goals of improving power performance and reducing fuel consumption, thus becoming a core technical problem restricting the long-range operational capabilities of target drones, target missiles, and loitering aircraft. Summary of the Invention

[0005] To address the aforementioned technical problems, this invention provides a combined compressor turbojet engine that can increase the engine's compression ratio, thereby increasing power and reducing fuel consumption.

[0006] The technical solution of the present invention includes a compressor. An axial compressor is provided at the air intake end of the turbojet engine. The axial compressor is connected to the shaft of the turbojet engine and is located upstream of the compressor.

[0007] Furthermore, the axial compressor includes an axial compressor rotor and an axial compressor stator; the axial compressor rotor is coaxially arranged with the compressor, and the axial compressor rotor and the shaft are interference-fitted; one end of the axial compressor stator is fixedly connected to the casing at the air inlet of the turbojet engine, and the other end is rotatably connected to the shaft through a pivot bearing.

[0008] Furthermore, it also includes a generator, with a cavity on the left side of the compressor rotor-shaft connection for housing the generator; the generator rotor is connected to the shaft, and the generator stator is connected to the compressor rotor.

[0009] Furthermore, the turbojet engine has two support bearings on its shaft, located at opposite ends of the shaft.

[0010] Furthermore, the compressor is a mixed-flow compressor.

[0011] Furthermore, the stator of the axial compressor is equipped with stator blades, which are used to rectify the airflow discharged from the rotor of the axial compressor and adjust the airflow direction.

[0012] The technical solution provided by this invention has the following advantages compared with the prior art: The most distinctive feature of this invention is that When a combined-compressor turbojet engine is running, the outside airflow first enters the axial compressor at the intake end for primary compression, and then flows into the compressor located downstream of the axial compressor for secondary compression. The high-pressure gas after two stages of compression enters the combustion chamber and mixes with fuel for combustion. The high-temperature, high-pressure gas drives the turbine to rotate, and the turbine synchronously drives the axial compressor and the main compressor to work continuously via the engine shaft. Finally, the gas is ejected at high speed to generate thrust. By adopting a combined architecture of axial compressor and main compressor in series, the overall pressure ratio is increased to 6 to 10, breaking through the pressure ratio limit of 4 to 5 of traditional single-stage compressors, significantly reducing engine fuel consumption and solving the problem of insufficient fuel for target drones and target missiles during medium- and long-range loitering flights.

[0013] Other advantages, objectives and features of the present invention will become apparent in part from the following description, and in part from those skilled in the art through study and practice of the invention. Attached Figure Description

[0014] 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 these drawings without creative effort.

[0015] Figure 1 This is a structural illustration of an embodiment of the present invention. Figure 1 .

[0016] Figure 2 This is a structural illustration of an embodiment of the present invention. Figure 2 .

[0017] Figure 3 This is a structural illustration of an embodiment of the present invention. Figure 3 .

[0018] Figure 4 This is a structural illustration of an embodiment of the present invention. Figure 4 .

[0019] Figure label: 1. Axial compressor rotor; 2. Axial compressor stator; 3. Shaft; 4. Generator; 5. Support bearing; 6. Compressor; 7. Mixed-flow compressor. Detailed Implementation

[0020] The following detailed description of a specific embodiment of the present invention is provided in conjunction with the accompanying drawings. However, it should be understood that the scope of protection of the present invention is not limited to the specific embodiment.

[0021] In the description of this invention, it should be understood that the terms "center," "longitudinal," "lateral," "length," "width," "thickness," "upper," "lower," "front," "rear," "left," "right," "vertical," "horizontal," "top," "bottom," "inner," "outer," "axial," "radial," and "circumferential" indicate the orientation or positional relationship based on the orientation or positional relationship shown in the accompanying drawings. They are only for the convenience of describing the technical solution of this invention and simplifying the description, and do not indicate or imply that the device or element referred to must have a specific orientation, or be constructed and operated in a specific orientation. Therefore, they should not be construed as limitations on this invention.

[0022] In the description of the embodiments of the present invention, unless otherwise stated, "a plurality of" means two or more.

[0023] like Figure 1 As shown, the present invention provides a combined compressor 6 type turbojet engine, including compressor 6, an axial flow compressor is provided at the air intake end of the turbojet engine, the axial flow compressor is connected to the shaft 3 of the turbojet engine and is located upstream of compressor 6.

[0024] In the embodiments provided by the present invention, the axial compressor includes an axial compressor rotor 1 and an axial compressor stator 2; the axial compressor rotor 1 is coaxially arranged with the compressor, and the axial compressor rotor 1 and the shaft 3 are interference-fitted; one end of the axial compressor stator 2 is fixedly connected to the casing at the air inlet of the turbojet engine, and the other end is rotatably connected to the shaft 3 through a pivot bearing 5.

[0025] The axial compressor rotor 1 is interference-fitted with the engine shaft 3 and rotates coaxially to achieve primary compression of the airflow. One end of the axial compressor stator 2 is fixed to the engine inlet casing, and the other end is rotatably connected to the shaft 3 via a pivot bearing 5. The stator remains stationary to cooperate with the rotor in completing compression and provides rotational support for the rotor. The installation and transmission structure of the axial compressor is clearly defined. The coaxial interference fit of the rotor ensures transmission accuracy, and the double-end fixing of the stator enhances structural stability, ensuring that the axial compressor stably achieves a primary pressure ratio of 1.5 to 2.

[0026] like Figure 2As shown, in the embodiment provided by the present invention, a generator 4 is also included. A cavity is provided on the left side of the part where the rotor of the compressor is connected to the shaft 3. The cavity is used to place the generator 4. The rotor of the generator 4 is connected to the shaft 3, and the stator of the generator 4 is connected to the rotor of the compressor 6.

[0027] When the engine shaft 3 rotates, it drives the rotor of the generator 4 to rotate synchronously. The stator of the generator 4 remains relatively fixed to the rotor of the compressor. The relative motion between the rotor and the stator generates electrical energy through electromagnetic induction. The generator 4 is housed in the cavity on the left side where the compressor rotor connects to the shaft 3, and it continuously generates electricity as the engine runs. This integration of power generation without increasing the engine length utilizes the internal space effectively, optimizes the rotor structure, reduces the complexity of rotor dynamics design, and enhances the engine's functionality and compactness.

[0028] like Figure 3 As shown, in the embodiment provided by the present invention, the shaft 3 of the turbojet engine has two support bearings 5, which are located at both ends of the shaft 3 respectively.

[0029] The engine shaft 3 is supported and positioned by two support bearings 5 ​​at both ends. The bearings 5 ​​provide radial and axial constraints for the rotation of the shaft 3, reducing rotational friction and ensuring the smooth operation of the compressor 6, axial compressor, and turbine driven by the shaft 3, thus maintaining the stable operation of the rotor system. The dual-support structure at both ends of the shaft 3 simplifies the support system of the non-generator version of the engine, improves the operating stiffness and stability of the shaft 3, meets the requirements of rotor dynamics design, and reduces the complexity of the engine structure.

[0030] like Figure 4 As shown, in the embodiment provided by this invention, the compressor 6 is a mixed-flow compressor 7. Using the mixed-flow compressor 7 as the main compressor 6, it works in conjunction with the upstream axial-flow compressor to complete a two-stage combined compression of the airflow. The mixed-flow compressor 7 performs a secondary, efficient compression of the airflow that has been initially pressurized by the axial-flow compressor, delivering high-pressure air that meets the pressure requirements to the combustion chamber. The mixed-flow compressor 7 has a smaller diameter, larger flow rate, and higher efficiency, which can reduce the engine's frontal area, optimize engine size while increasing the overall pressure ratio, and further improve thrust and fuel efficiency performance.

[0031] In the embodiments provided by the present invention, the stator 2 of the axial compressor is provided with stator blades, which are used to rectify the airflow discharged from the rotor 1 of the axial compressor and adjust the airflow direction.

[0032] The airflow discharged from the axial compressor rotor 1 enters the stator, where it is guided and rectified by the stator blades. This adjusts the flow angle and direction of the airflow, ensuring it enters the downstream mixed-flow compressor 7 in a manner compatible with the inlet, thus preventing turbulence that could reduce compression efficiency. Optimizing airflow matching through stator blade rectification reduces energy loss and improves the overall compression efficiency of the combined compressor 6, further guaranteeing engine pressure ratio and fuel economy.

[0033] It should be noted that any parts not disclosed or specifically described in this invention are existing technology or conventional configurations, and their specific structures and working principles will not be elaborated further. In this document, the terms "comprising," "including," or any other variations thereof are intended to cover non-exclusive inclusion, such that a process, method, article, or apparatus that comprises a list of elements includes not only those elements but also other elements not expressly listed, or elements inherent to such a process, method, article, or apparatus. Without further limitations, an element defined by the phrase "comprising one..." does not exclude the presence of other identical elements in the process, method, article, or apparatus that includes that element.

[0034] Although embodiments of the present invention have been disclosed above, they are not limited to the applications listed in the specification and embodiments. It can be applied to various fields suitable for the present invention. Other modifications can be readily implemented by those skilled in the art. Therefore, without departing from the general concept defined by the claims and their equivalents, the present invention is not limited to the specific details and examples shown and described herein.

Claims

1. A combined compressor turbojet engine, comprising a compressor, characterized in that, The turbojet engine has an axial compressor at its intake end, which is connected to the engine shaft and located upstream of the compressor.

2. The combined compressor turbojet engine according to claim 1, characterized in that, The axial compressor includes an axial compressor rotor and an axial compressor stator; The axial compressor rotor is coaxially arranged with the compressor, and the axial compressor rotor and shaft are interference-fitted. One end of the axial compressor stator is fixedly connected to the casing at the air inlet of the turbojet engine, and the other end is rotatably connected to the shaft through a pivot bearing.

3. The combined compressor turbojet engine according to claim 1, characterized in that, It also includes a generator, and a cavity is provided on the left side of the part where the compressor rotor is connected to the shaft, the cavity being used to house the generator; The generator's rotor is connected to the shaft, and the generator's stator is connected to the compressor's rotor.

4. A combined compressor turbojet engine according to claim 1, characterized in that, The turbojet engine has two support bearings on its shaft, which are located at both ends of the shaft.

5. A combined compressor turbojet engine according to any one of claims 1 to 4, characterized in that, The compressor is a mixed-flow compressor.

6. The combined compressor turbojet engine according to claim 5, characterized in that, The axial compressor stator is equipped with stator blades, which are used to rectify the airflow discharged from the axial compressor rotor and adjust the airflow direction.

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