Blisk cooling structure of micro turbojet engine
By employing additive manufacturing and rotary supercharging cooling design in the integral bladed disk of the micro turbojet engine, the problem of turbine cooling difficulties has been solved, achieving efficient cooling and performance improvement.
Patent Information
- Authority / Receiving Office
- CN · China
- Patent Type
- Applications(China)
- Current Assignee / Owner
- Filing Date
- 2025-12-17
- Publication Date
- 2026-04-03
AI Technical Summary
Existing micro turbojet engines face bottlenecks in performance improvement due to difficulties in turbine cooling and limited turbine inlet temperature, especially since the integral bladed disk structure makes it difficult to achieve complex hollow cooling channels.
The integral bladed disk is manufactured using additive manufacturing technology. The design integrates airflow channels, sealing structures, and pressurizing blades to form a highly integrated cooling flow path. It utilizes the rotational energy of the turbine rotor to pressurize and cool the airflow, achieving efficient cooling through a complex hollow cooling channel.
It significantly increases turbine blade temperature, reduces mixing losses, improves engine performance, and increases thrust-to-weight ratio by more than 15%.
Smart Images
Figure CN121781978A_ABST
Abstract
Description
Technical Field
[0001] This invention belongs to the field of micro aero-turbojet engine technology, specifically relating to an integral bladed disk cooling structure for a micro turbojet engine. Background Technology
[0002] Miniature turbojet engines (miniature turbojet engines) have broad application prospects in military and civilian fields such as unmanned aerial vehicles (UAVs), target drones, and loitering munitions due to their high thrust-to-weight ratio and compact structure. One of the key ways to improve the performance of miniature turbojet engines is to increase the turbine inlet temperature, which directly affects the engine's thermal efficiency and output power. However, turbine components, especially turbine rotor blades, operate in high-temperature and high-pressure combustion environments for extended periods, posing severe challenges to the material properties of these components.
[0003] In medium and large aero engines, a separate structure for the turbine disk and blades is typically adopted. The disk can be made of powder metallurgy high-temperature alloy, while the blades can be cast single-crystal hollow blades supplemented by complex film cooling technology. This "disk-blade separation" design provides the space and structural basis for implementing efficient cooling.
[0004] However, in micro turbojet engines, due to their extremely compact structural dimensions and manufacturing costs, an "integral bladed disk" design is commonly adopted, where the turbine disk and blades are manufactured as a single unit. While this design simplifies the structure, reduces weight, and avoids stress concentration at tenon joints, it also introduces new challenges: traditional casting processes struggle to achieve the complex hollow cooling channels within the integral bladed disk, especially in the blade section. Therefore, most existing micro turbojet engine turbine rotors can only be designed as solid structures or have extremely simple cooling designs. This severely limits the improvement of turbine inlet temperature, becoming one of the main bottlenecks restricting further performance enhancements of micro turbojet engines.
[0005] With the rapid development of additive manufacturing (3D printing) technology, especially the maturity of metal additive manufacturing processes such as selective laser melting and electron beam melting, it has become possible to manufacture metal parts with extremely complex internal structures. High-temperature alloy additive manufacturing technology has begun to be applied in the aerospace field, which makes it technologically feasible to realize hollow cooling channels similar to those in large engine blades inside micro integral bladed disks.
[0006] However, manufacturing processes alone are insufficient; a highly efficient cooling system that matches the overall bladed disk structure and is suitable for the size and flow characteristics of a micro-engine is also required. This system needs to address issues such as the introduction, distribution, flow, and pressurization of cooling air, as well as its mixing with the mainstream combustion gas. While achieving effective cooling, it must minimize the negative impact on turbine efficiency, and even turn disadvantages into advantages, thereby improving engine performance. Summary of the Invention
[0007] To address the aforementioned shortcomings of existing technologies, this invention provides an integral bladed disk cooling structure for a micro turbojet engine, aiming to overcome the problem that existing micro turbojet engines suffer from limited turbine inlet temperature and performance bottlenecks due to difficulties in turbine cooling.
[0008] To achieve the above objectives, the technical solution adopted by the present invention is as follows: a micro turbojet engine integral bladed disk cooling structure is provided, which includes a rotating shaft, on which a centrifugal impeller, a turbine rotor and a bearing housing for supporting the rotating shaft are coaxially arranged, and an airflow channel is provided between the bearing housing and the rotating shaft; The turbine rotor is an integral bladed disk structure, which includes a turbine disk and turbine blades. The turbine disk has an annular cavity inside, and the turbine blades have hollow cooling channels that communicate with the annular cavity inside. A sealing sleeve is provided between the bearing chamber and the turbine rotor. A set of sealing grates is provided on the inner wall of the sealing sleeve. A protruding part matching the sealing grates is provided on the front end face of the turbine disk. The two together constitute a sealing structure. The sealing sleeve and the front side of the turbine disk form a sealing cavity that communicates with one end of the airflow channel. The other end of the airflow channel is located at the outlet of the centrifugal impeller. The sealing sleeve, bearing chamber, and turbine guide located in front of the turbine rotor are fixed by bolts through a stop fit; The high-pressure gas flow path after the centrifugal impeller is configured as follows: the gas flow passes sequentially through the airflow channel of the bearing chamber, the sealing cavity, the annular cavity and the hollow cooling channel, and then is discharged from the trailing edge of the turbine blade to the main flow channel of the engine.
[0009] The basic principle of the integral bladed disk cooling structure for a micro turbojet engine of the present invention is as follows: by setting a special airflow channel in the bearing chamber, the high-pressure cold air after the centrifugal impeller is directly led to the sealing cavity on the front side of the turbine rotor, and then distributed to the hollow cooling channels of each blade through the annular cavity inside the turbine disk, forming a highly integrated, well-defined, sealed and controllable cooling flow path, with the cooling path minimized, reducing flow resistance and pressure loss; the sealing cavity ensures efficient utilization of high-pressure cold air, avoids cold air leakage to non-cooled areas, and improves cooling efficiency and economy.
[0010] Furthermore, the turbine rotor is formed using additive manufacturing processes. Additive manufacturing processes (such as SLM and EBM) can achieve integrated molding of complex internal cavities and flow channels that are difficult to complete using traditional casting or machining, supporting more optimized cooling channel topology designs, such as conformal cooling channels and porous structures, thereby further improving cooling performance.
[0011] Furthermore, multiple pressurizing blades are evenly distributed circumferentially on the front end face of the turbine disk. When the turbine rotor rotates, these pressurizing blades compress the gas flowing through the front end face of the turbine disk into the annular cavity. By setting circumferentially distributed pressurizing blades with a certain installation angle on the front side of the turbine disk, and utilizing the centrifugal effect and aerodynamic compression during turbine rotor rotation, the gas within the sealed cavity is pressurized and then forced into the annular cavity. This actively increases the cooling gas pressure, ensuring it can overcome subsequent flow resistance and fully fill the cooling channel; it also enhances the kinetic energy of the cooling gas flow, facilitating velocity matching with the mainstream at the blade trailing edge and reducing mixing losses; and by utilizing the rotor's own rotational energy, no external pressurization device is required, resulting in a simple and efficient structure.
[0012] Furthermore, the hollow cooling channel is a curved, rotating channel. The inlet of the hollow cooling channel is located at the root of the turbine blade and communicates with the annular cavity, while the outlet is located at the trailing edge of the turbine blade. This design significantly increases the airflow path and residence time inside the blade, enhancing convective heat transfer. Simultaneously, the hollow cooling channel design guides the airflow to impact the high-heat areas of the blade's inner wall. This greatly improves heat transfer efficiency, achieving a uniform reduction in blade wall temperature to adapt to the internal temperature gradient and stress distribution of the blade. Furthermore, the compact structure of the hollow cooling channel maximizes the heat transfer area within a limited space.
[0013] Furthermore, the hollow cooling channel has multiple outlets, which are vertically spaced and evenly arranged along the trailing edge of the turbine blades of the turbine rotor. This multi-outlet design allows for more uniform and thorough coverage and spraying of the cooling air along the trailing edge, forming a continuous or discrete air film protective layer. This improves the local cooling effect at the trailing edge, prevents hot spots, and enhances the mixing uniformity of the cooling air with the main flow, further reducing aerodynamic losses.
[0014] Furthermore, the turbine rotor is manufactured using a nickel-based superalloy. Nickel-based superalloys possess excellent high-temperature strength, creep resistance, corrosion resistance, and oxidation resistance, making them ideal materials for hot-end components of aero-engines, ensuring the long-term reliable operation of the turbine rotor under high-temperature and high-stress environments.
[0015] The present invention provides an integral bladed disk cooling structure for a micro turbojet engine, the advantages of which are as follows: 1. A micro turbojet engine integral bladed disk cooling structure that integrates all cooling structures inside the rotor, eliminating the need for additional complex piping and external devices, maintaining the advantages of a compact micro engine structure, while also achieving high overall structural strength.
[0016] 2. A micro turbojet engine integral bladed disk cooling structure, wherein the turbine rotor is an integral bladed disk manufactured using high-temperature alloy additive manufacturing, with hollow cooling channels designed internally; circumferentially distributed pressurizing blades are provided on the front side of the turbine disk; sealing grates are provided inside the sealing sleeve, forming a sealing structure with the protruding part on the front side of the turbine disk. During operation, high-pressure cold air drawn from behind the centrifugal impeller flows sequentially through the airflow channel and the sealing cavity, is pressed and pressurized by the pressurizing blades of the turbine disk, then enters the annular cavity inside the turbine disk, and flows radially into the hollow cooling channels of each turbine blade under the action of centrifugal force, finally exiting from the trailing edge of the turbine blades. This structure not only achieves effective cooling of the turbine rotor (especially the turbine blades), allowing for an increase in the turbine blade inlet temperature, but also minimizes mixing losses with the mainstream when the pressurized cooling air is ejected from the trailing edge, and its momentum can be converted into additional turbine output work, significantly improving the overall performance of the micro turbojet engine. Attached Figure Description
[0017] Figure 1 This is a schematic diagram of the overall structure of an integral bladed disk cooling structure for a micro turbojet engine.
[0018] Figure 2 This is a partial structural diagram of the integral bladed disk cooling structure of a micro turbojet engine.
[0019] Figure 3 This is a schematic diagram of the turbine rotor.
[0020] Figure 4 This is a schematic diagram of the turbine blade structure.
[0021] Among them, 1. centrifugal impeller; 2. shaft; 3. bearing chamber; 4. sealing sleeve; 41. sealing grate; 5. turbine guide; 6. turbine rotor; 61. turbine disk; 62. turbine blade; 63. booster blade; 7. sealing cavity; 8. airflow channel; 9. annular cavity; 10. hollow cooling channel. Detailed Implementation
[0022] The specific embodiments of the present invention are described below to enable those skilled in the art to understand the present invention. However, it should be understood that the present invention is not limited to the scope of the specific embodiments. For those skilled in the art, various changes are obvious as long as they are within the spirit and scope of the present invention as defined and determined by the appended claims. All inventions utilizing the concept of the present invention are protected.
[0023] like Figures 1-2As shown, a micro turbojet engine has a core rotor component comprising a centrifugal impeller 1, a shaft 2, and a turbine rotor 6. The core rotor component is supported within a bearing housing 3 by bearings at two fulcrums. A sealing sleeve 4 and a turbine guide 5 are sequentially connected to the rear end of the bearing housing 3. The three components are positioned by a precision stop and fastened with bolts, ensuring the coaxiality and stability of the structure.
[0024] The present invention provides a micro turbojet engine integral bladed disk cooling structure, the core improvement of which lies in the turbine rotor 6 and its related cooling system. The bearing chamber 3 and the shaft 2 are provided with airflow channels 8.
[0025] like Figures 1-4 As shown, the turbine rotor 6 is an integral bladed disk structure, comprising a turbine disk 61 and turbine blades 62. The turbine rotor 6 is an integral bladed disk formed using laser selective melting technology with a nickel-based superalloy (such as Inconel 718). During manufacturing, an annular cavity 9 surrounding the shaft is formed inside the turbine disk 61. Simultaneously, a complex hollow cooling channel 10 is formed inside each turbine blade 62. This channel communicates with the annular cavity 9 of the disk at the blade root, rotates multiple times inside the blade to increase the heat exchange area, and finally opens an outlet at the blade trailing edge.
[0026] A sealing sleeve 4 is provided between the bearing chamber 3 and the turbine rotor 6. A set of sealing grates 41 is provided on the inner wall of the sealing sleeve 4. A protruding part matching the sealing grates 41 is provided on the front end face of the turbine disk 61. The two together constitute a sealing structure. The sealing sleeve 4 and the front side of the turbine disk 61 form a sealing cavity 7 that communicates with one end of the airflow channel 8. The other end of the airflow channel 8 is located at the outlet of the centrifugal impeller 1.
[0027] Preferably, but not limited to, multiple pressurizing blades 63 are evenly distributed circumferentially on the front end face of the turbine disk 61. When the turbine rotor 6 rotates, the multiple pressurizing blades 63 compress the gas flowing through the front end face of the turbine disk 61 into the annular cavity 9. By providing circumferentially distributed pressurizing blades 63 with a certain installation angle on the front side of the turbine disk 61, the centrifugal effect and aerodynamic compression of the turbine rotor 6 during rotation are used to pressurize the gas in the sealed cavity 7 and then compress it into the annular cavity 9. This actively increases the cooling gas pressure, ensuring it can overcome the subsequent flow channel resistance and fully fill the cooling channel; it enhances the kinetic energy of the cooling gas flow, which is beneficial for achieving velocity matching with the mainstream at the blade trailing edge and reducing mixing losses; and by utilizing the rotor's own rotational energy, no external pressurization device is required, resulting in a simple and efficient structure.
[0028] When the engine is running, a portion of the high-pressure, relatively low-temperature air (referred to as "high-pressure cold air") drawn from the outlet of the centrifugal impeller 1 is directed to the front and rear ends of the airflow passage 8 to cool and lubricate the high-speed bearings. After completing its bearing cooling task, this portion of gas flows into the sealed cavity 7, which is surrounded by the sealing sleeve 4 and the front side of the turbine disk 61. As the turbine rotor 6 rotates at a high speed of tens of thousands to hundreds of thousands of revolutions per minute, the booster blades 63 at the front end of the turbine disk 61 also rotate at high speed. These booster blades 63 act like the impellers of a miniature centrifugal compressor, performing work on the gas within the sealed cavity 7, increasing its pressure and velocity. Under the centrifugal force and the pumping action of the booster blades 63, the gas is forced into the annular cavity 9 within the turbine disk 61. The gas accumulates within the annular cavity 9. Due to the powerful centrifugal force generated by the high-speed rotation of the turbine disk 61, the gas experiences an outward volume force within the cavity. This force drives the gas from the outer circumference of the annular cavity 9, through multiple radial flow channels or directly through the inlet at the blade root, radially into the hollow cooling channel 10 of each turbine blade 62. As the cooling gas flows through the rotating channels inside the blade, it undergoes intense convective heat transfer with the blade wall surface heated by the high-temperature combustion gas, thereby carrying away heat from the blade metal and maintaining the blade temperature within the safe range allowed by the material. The heated cooling gas remains at a high pressure and is ejected at high speed from the outlet at the trailing edge of the turbine blade 62.
[0029] Through the prior pressurization of the supercharger blades 63 and the design of the flow channel, the velocity of the airflow at the outlet is adjusted to be very close to the velocity and direction of the mainstream combustion gas at the turbine rotor 6 outlet. These two airflows mix smoothly behind the trailing edge, avoiding strong shearing and turbulence caused by sudden velocity changes, thus reducing the aerodynamic losses of mixing to an extremely low level. More importantly, the cooling gas is ejected with high momentum, generating a reaction force on the turbine blades 62. This momentum is partially recovered in the turbine rotor 6 and converted into additional turbine output power, improving the efficiency of the turbine rotor 6.
[0030] The metal temperature of turbine blades 62 can be reduced by several hundred degrees, allowing designers to raise the gas temperature in front of the turbine by 200°C or even higher. According to the Brayton cycle principle, this directly and significantly improves the engine's thermal efficiency. Simultaneously, the additional work contributed by the cooling gas and the extremely low mixing losses further enhance the efficiency of turbine rotor 6. Overall, the engine's thrust and thrust-to-weight ratio are expected to achieve a performance improvement of over 15%.
[0031] Preferably, but not limited to, the hollow cooling channel 10 is a curved rotary channel. The inlet of the hollow cooling channel 10 is located at the root of the turbine blade 62 and communicates with the annular cavity 9, while the outlet of the hollow cooling channel 10 is located at the trailing edge of the turbine blade 62. This design significantly increases the airflow path and residence time inside the blade, enhancing convective heat transfer. Simultaneously, the hollow cooling channel 10 design guides the airflow to impact the high-heat areas of the blade's inner wall. This greatly improves heat transfer efficiency, achieving a uniform reduction in blade wall temperature to adapt to the internal temperature gradient and stress distribution of the blade. Furthermore, the hollow cooling channel 10 has a compact structure, maximizing the heat transfer area within a limited space.
[0032] Specifically, the hollow cooling channel 10 has multiple outlets, which are vertically spaced and evenly arranged along the trailing edge of the turbine blades 62 of the turbine rotor 6. This multi-outlet design allows for more uniform and thorough coverage and spraying of cooling air along the trailing edge, forming a continuous or discrete air film protective layer. This improves the local cooling effect at the trailing edge, prevents hot spots, and enhances the mixing uniformity of the cooling air with the main flow, further reducing aerodynamic losses.
[0033] In summary, this invention constructs a highly integrated, efficient, and reliable turbine cooling system for a micro turbojet engine through an innovative combination of the internal airflow channel 8 within the bearing housing 3, the internal cooling flow path of the integral bladed disk, and the sealing and pressurizing structure. This system fully utilizes the advantages of additive manufacturing processes to achieve integrated molding of complex cooling channels; through internal flow channel design and the application of pressurizing blades 63, it improves cooling gas pressure and flow efficiency; and finally, by optimizing the cooling gas exhaust method, it reduces mixing losses and is expected to contribute additional thrust. This invention effectively solves the turbine cooling problem of micro turbojet engines, providing a practical technical solution for improving turbine inlet temperature and overall engine performance.
Claims
1. A micro turbojet engine integral bladed disk cooling structure, characterized in that, Includes a rotating shaft (2), on which a centrifugal impeller (1), a turbine rotor (6) and a bearing chamber (3) for supporting the rotating shaft are coaxially arranged, and the bearing chamber (3) and the rotating shaft (2) are provided with an airflow channel (8); The turbine rotor (6) is an integral bladed disk structure. The turbine rotor (6) includes a turbine disk (61) and turbine blades (62). The turbine disk (61) has an annular cavity (9) inside, and the turbine blades (62) have a hollow cooling channel (10) communicating with the annular cavity (9) inside. A sealing sleeve (4) is provided between the bearing chamber (3) and the turbine rotor (6). A set of sealing grates (41) is provided on the inner wall of the sealing sleeve (4). A protruding part matching the sealing grates (41) is provided on the front end face of the turbine disk (61). The two together constitute a sealing structure. The sealing sleeve (4) and the front side of the turbine disk (61) form a sealing cavity (7) that communicates with one end of the airflow channel (8). The other end of the airflow channel (8) is located at the outlet of the centrifugal impeller (1). The sealing sleeve (4), bearing chamber (3) and turbine guide (5) located in front of the turbine rotor (6) are fixed by bolts through a stop fit; The high-pressure gas flow path after the centrifugal impeller (1) is configured as follows: the gas flow path flows sequentially through the airflow channel (8), sealing cavity (7), annular cavity (9) and hollow cooling channel (10) of the bearing chamber (3), and then is discharged from the trailing edge of the turbine blade (62) to the main flow path of the engine.
2. The integral bladed disk cooling structure for a micro turbojet engine according to claim 1, characterized in that, The turbine rotor (6) is formed using an additive manufacturing process.
3. The integral bladed disk cooling structure for a micro turbojet engine according to claim 1, characterized in that, On the front end face of the turbine disk (61), a plurality of pressurizing blades (63) are evenly distributed in the circumferential direction. When the turbine rotor (6) rotates, the plurality of pressurizing blades (63) press the gas flowing through the front end face of the turbine disk (61) into the annular cavity (10).
4. The integral bladed disk cooling structure for a micro turbojet engine according to claim 1, characterized in that, The hollow cooling channel (10) is a curved rotary channel. The inlet of the hollow cooling channel (10) is located at the root of the turbine blade (62) and communicates with the annular cavity (9). The outlet of the hollow cooling channel (10) is located at the trailing edge of the turbine blade (62).
5. The integral bladed disk cooling structure for a micro turbojet engine according to claim 4, characterized in that, The hollow cooling channel (11) has multiple outlets, which are arranged vertically and evenly along the trailing edge of the turbine blades (62) of the turbine rotor (6).
6. The integral bladed disk cooling structure for a micro turbojet engine according to any one of claims 1 to 5, characterized in that, The turbine rotor (6) is made of a nickel-based high-temperature alloy.