CN114961868B - Hot end rotor system of micro turbojet engine with core cooling structure - Google Patents

Abstract

The invention discloses a hot end rotor system of a microminiature turbojet engine with a core cooling structure, which comprises a main shaft, a shaft sleeve, a rear bearing, a middle-stage sealing isolation disc, a turbine guider, a turbine rotor, a tail nozzle and a tail cone. When the engine works, high-pressure cooling air flow enters the shaft sleeve, becomes total cooling air flow after a throttling measure, is mixed with lubricating oil sprayed to the rear bearing, and lubricates and cools the rear bearing; the total cooling air flow is divided into a pre-disk cooling air flow and a post-disk cooling air flow, wherein the pre-disk cooling air flow and the post-disk cooling air flow are respectively the main gas flow after the temperature of the front disk surface and the rear disk surface of the turbine rotor is reduced. The invention optimizes and reduces the temperature of the hot end rotor system more effectively, solves the heat transfer problem of the turbine rotor and the main shaft to the bearing fundamentally, improves the cooling of the bearing greatly, allows the adoption of a common industrial-grade bearing, and reduces the cost; meanwhile, the lubrication quantity can be reduced, and the performance of the engine is improved.

Description

A micro turbojet engine hot end rotor system with core cooling structure

technical field

The invention belongs to the technical field of aero-engines, and in particular relates to cooling and structural optimization of a hot-end rotor part of a miniature turbojet engine.

Background technique

Micro turbojet engines have been widely used in unmanned aerial vehicles and various types of loitering bombs, etc. They have the characteristics of high thrust-to-weight ratio, short life, and large consumption. With the batch equipment of such products, performance, cost, and reliability are all affected put forward higher requirements.

The common micro-turbojet engine hot end (turbine end) rotor system is shown in Figure 1. The high-pressure air (about 250°C) used for cooling enters the shaft sleeve through certain throttling measures and mixes part of the lubricating oil. After the rear bearing is lubricated and cooled, it enters the front cavity of the turbine disc and convectively cools the front of the turbine rotor disc; there is a closed cavity between the rear disc of the turbine rotor and the tail vertebra, and the gas inside will be continuously pumped by the high-speed rotating turbine. The disc is thrown out, and the gas (about 800°C) from the outside is continuously entering, forming a certain vortex and causing the turbine disc to be surrounded by high-temperature gas; the end of the main shaft, as well as the turbine nut with a more critical role on the end of the main shaft, are also at high temperature When surrounded by gas, the temperature of the parts here will reach more than 600°C, and the ambient temperature is very harsh. The front contact end surface of the turbine rotor generally adopts the method of milling and sinking to reduce the contact area with the bearing to reduce a certain amount of heat transfer. The temperature of the cylindrical surface in contact with the bearing will reach more than 350°C, and the large temperature difference will cause a large amount of heat to be transferred to the rear bearing; in addition, the axial load of the rotor mainly acts on the rear bearing, and the heat generated by the bearing itself when it rotates at high speed, As a result, the rear bearing severely restricts the life of the engine. As a result, extremely high requirements are put forward for bearing cooling, requiring more lubricating and cooling oil, which is equivalent to increasing the fuel consumption rate of the engine and reducing performance; at the same time, the inner and outer rings of the bearing must also be made of high-temperature-resistant high-nitrogen stainless steel. . In addition, the temperature of the turbine rotor is more than 100°C higher than the temperature of the main shaft, coupled with the centrifugal stress generated by the high speed, these all lead to a tendency for the inner hole of the turbine rotor to increase relative to the main shaft, resulting in a large fit clearance (above 0.02), It may lead to a significant increase in residual unbalance (take a 1kg turbine rotor as an example, a gap of 0.02 may cause an additional additional unbalance of 10g.mm); eventually it may lead to increased vibration or even failure to work properly. The end of the main shaft and the turbine nut bear important preload and torsion transmission functions, and are surrounded by high-temperature gas with a temperature of 600°C to 670°C, which puts forward higher requirements on the high-temperature strength reserve of the material; while the commonly used quenched and tempered steel Endurance strength at high temperatures will be greatly reduced. For example, 18Cr2Ni4WA, 200-hour durability strength as an example, is 696MPa at 450°C and 216MPa at 550°C, which is very sensitive to high temperature; coupled with vibration and other influences, fracture failures are prone to occur; under this design, if the pursuit of high gas temperature ( 800°C or higher), the main shaft and turbine nut must be made of high-cost high-temperature-resistant materials; if general steel is used in consideration of cost, it is necessary to reduce the gas temperature from 800°C to 720°C or even 700°C to have A certain strength margin, but this will lose 5% to 10% of the thrust.

Large gas turbine engines have relatively ample space, and various turbine installations and cooling layouts can solve the above problems through complex designs; while micro engines generally have extremely high speeds (50,000 to hundreds of thousands of revolutions per minute), rotor dynamics and The vibration problem is complex, and the distance between the turbine rotor center and the bearing must be strictly controlled, and the structural space is relatively compact; the above problems are difficult to solve in a micro turbojet engine with extremely limited space, and the cost is high or the margin is low. At present, the cooling scheme of most micro turbojet engines basically focuses on the cooling of the shaft sleeve. The temperature of the outer ring of the bearing can be improved by taking measures on the sleeve, but it is difficult to affect the inner ring of the bearing because of the ceramic balls in the middle; The fundamental solution is to reduce the heat transfer from the hot-end rotor parts to the inner ring of the bearing; however, due to the limited space and structure, the optimization of the rotor is less.

For example, the patent CN108547672 uses a complex special centrifugal impeller to pressurize the bleed air (the pressure must not be lower than the pressure in front of the turbine rotor) and requires a large amount of bleed air. At the same time, the gas temperature is obviously lower, otherwise it will not be able to meet the grease lubrication conditions. ;Furthermore, this scheme opens holes on the turbine rotor disk body, which will have a great impact on the strength of the disk body, and its gas volume cannot be controlled; at the same time, the non-contact seal between the stator part and the rotor part will also waste a lot of energy. The supercharged bleed air increases the load on the turbine and reduces the power of the turbine; in summary, these measures cannot be realized in an extremely compact high-performance tiny turbojet engine.

Such as patent CN108952967, is to slot on the main shaft that cooperates with bearing inner ring, is used for reducing heat transfer area and cooling. Because the pressure difference between the front and back of the bearing is generally small, the cooling effect is limited due to the small amount of bleed air; at the same time, the matching area between the main shaft and the bearing should not be too low, so the effect of reducing the heat transfer of the main shaft is also limited; and it cannot be fundamentally solved The heat transfer problem of the heat source (turbine rotor).

For example, the patent CN104265460 only indirectly cools the outer ring of the bearing through heat exchange to reduce the temperature of the outer ring of the bearing. However, the bearings of micro and small engines generally use ceramic balls, and the heat transfer between the outer ring and the inner ring of the bearing is extremely small, so the temperature is the highest. 1. The inner ring of the bearing that affects the bearing life the most will not have a significant effect.

Contents of the invention

The purpose of the invention is to solve the problems in the prior art that the heat dissipation and cooling effect of the miniature turbojet engine is limited, the high temperature resistance is poor, and the service life is short.

In order to achieve the purpose of the present invention, the present invention provides a micro-miniature turbojet engine hot end rotor system with a core cooling structure, including a main shaft, a shaft sleeve, a rear bearing, an intermediate stage sealing isolation disc, a turbine guide, and a turbine rotor , tail nozzle, and tail cone; the rear bearing, the intermediate stage sealing isolation plate, and the turbine rotor are sequentially inserted and installed on the main shaft, and are fixed by the turbine nut; the rear bearing is installed on the shaft sleeve, and is pre-tensioned by the spring from the front end of the main shaft The force and the aerodynamic axial force in the working process make the rear bearing and the end face of the bushing close; the turbine guide is fixed on the bushing through threads or fasteners; the tail cone is integrated with the tail nozzle in the form of welding through the support plate , the tail nozzle is connected to the turbine guide;

When the engine is working, the high-pressure cooling airflow enters the shaft sleeve, and after throttling measures, it becomes the total cooling airflow, which is mixed with the lubricating oil sprayed to the rear bearing to lubricate and cool the rear bearing; the total cooling airflow is divided into the front cooling airflow and the cooling airflow after the disk, the cooling airflow before the disk and the cooling airflow after the disk respectively cool the front disk surface and the rear disk surface of the turbine rotor, and then merge into the main gas flow.

Further, the middle stage seals the isolation disc, the shaft sleeve, and the rear bearing to form a bleed air splitting cavity; the front end of the turbine rotor is provided with a turbine disc front cavity, and the rear end of the turbine rotor is provided with a disc rear caudal cavity;

After the total cooling air arrives at the rear bearing and enters the bleed air distribution chamber, it is divided into the cooling air flow before the disk and the cooling air after the disk; the cooling air flow before the disk enters the front chamber of the turbine disk through the grating teeth of the intermediate sealing isolation disk, and affects the front of the turbine rotor. After cooling the disk surface, it flows out from the second disk front channel between the turbine guider and the turbine rotor, and merges into the main gas flow; the cooling airflow behind the disk passes through the main shaft, the intermediate sealing isolation disk and the turbine rotor. The flow channel of the turbine rotor enters the cavity behind the disc between the turbine rotor and the tail cone, and after the heat insulation and cooling of the rear disc surface of the turbine rotor and the turbine nut, it flows out from the seventh channel behind the disc between the turbine rotor and the tail cone , into the main gas flow.

Further, the main shaft is a cylindrical shaft body, the rear end of the main shaft is provided with a front matching cylindrical surface and a rear matching cylindrical surface, and a channel cylindrical surface is arranged in the middle of the two matching cylindrical surfaces; the end of the front matching cylindrical surface close to the channel cylindrical surface is provided with a The front end surface of the channel; the end of the rear mating cylindrical surface close to the cylindrical surface of the channel is provided with the rear end surface of the channel; the cylindrical surface of the channel is provided with radial holes distributed in the circumferential direction, and the bottom of the radial holes is connected with an axial hole;

The intermediate seal isolation disc has a stepped ring structure, and the ring structure of the intermediate seal isolation disc close to the rear bearing has a smaller diameter, and the ring structure away from the rear bearing has a larger diameter; the ring structure with a smaller diameter is close to the end of the rear bearing There is a front mating end surface, and a front mating inner cylindrical surface is arranged on the inner side, and a rear mating end surface is arranged at the end away from the rear bearing; a rear mating inner cylindrical surface is arranged on the inner side of the ring structure with a larger diameter, and the maximum diameter of the ring structure with a larger diameter is There is a sealing structure on the outside of the center; there is a circle of circumferentially distributed radial air-introduction holes on the middle-level sealing isolation plate, leading to the rear mating inner cylindrical surface;

The front mating cylindrical surface forms a centering fit with the rear bearing and the front mating inner cylindrical surface; the intermediate sealing isolation disc is used to isolate and reduce the heat transfer from the turbine rotor to the rear bearing; the intermediate sealing isolation disc has a more The large heat dissipation area and cooling channel can disperse the heat transferred from the turbine rotor to enhance the cooling effect, and the sealing structure is used to achieve the effect of gas sealing;

The blade and disk structure of the turbine rotor is an axial flow turbine structure. There is an inner hole in the center of the turbine rotor, and a ring structure protrudes outward around the inner hole. The outer surface of the ring structure is the outer matching cylindrical surface of the front boss; the ring The front part of the center of the shape structure is provided with a number of front bosses with three faces. The surface of the hole is the inner chamfer of the front bosses, and the front boss grooves are arranged between several front bosses; the front boss grooves of the turbine rotor are used to reduce the contact area with the intermediate sealing isolation plate, and reduce the The stage seal tightly isolates the heat transfer of the disc and the rear bearing; at the same time, it acts as a part of the cooling airflow channel behind the disc.

Furthermore, the bleed air splitting cavity is formed by the sealing structure above the sealing isolation plate of the intermediate stage, the protruding ring at the rear of the shaft sleeve, and the rear bearing; the sealing structure and the protruding ring at the rear of the shaft sleeve form the first A channel in front of the disk; the cooling air flow in front of the disk passes through the first channel in front of the disk, enters the front cavity of the turbine disk, flows out through the second channel in front of the disk between the turbine guider and the turbine rotor, and merges into the main gas flow.

Further, the post-display cooling airflow sequentially passes through the first post-display channel, the second post-display channel, the third post-display channel, the fourth post-display channel, the fifth post-display channel, and the sixth post-display channel to form a complex airflow channel, Isolate the direct heat transfer of the turbine rotor to the main shaft, cool the turbine rotor and the rear of the main shaft, and also reduce the heat transfer from the turbine rotor to the main shaft and the intermediate sealing isolation disc; then enter the disc between the turbine rotor and the tail vertebra The rear tail vertebra cavity, after cooling the rear disc surface of the turbine rotor and the turbine nut, flows out from the seventh disc rear channel between the turbine rotor and the tail vertebra, and merges into the main gas flow.

Further, the channel behind the first disc is formed by a circle of circumferentially distributed radial air-introducing holes opened on the middle-level sealing isolation disc, and the radial holes are distributed along the circumferential direction;

The channel behind the second disc is surrounded by the outer chamfer of the front boss on the turbine rotor, the rear mating end surface of the intermediate seal isolation disc and the rear mating inner cylindrical surface after installation, and is a circumferential annular channel with a triangular cross-section. ;

The rear channel of the third disc is formed after the rear mating end surface on the intermediate stage seal isolation disc is installed and fitted with the front end surface on the turbine rotor, and the top of the front boss channel on the turbine rotor is covered by the rear mating end face;

The fourth channel behind the disk is formed by the clearance space between the cylindrical surface of the channel on the main shaft and the inner hole of the turbine rotor; the diameter of the cylindrical surface is smaller than the diameter of the inner hole, and the gap between the two forms a circumferential annular passage with a rectangular cross section ;In order to ensure the reliability and smoothness of the airflow, the front end of the channel partially overlaps with the channel behind the third plate, or overlaps with the inner chamfer of the front boss of the turbine rotor;

The channel behind the fifth disc is formed by a circle of circumferentially distributed radial holes on the main shaft, and the radial holes are evenly distributed along the circumferential direction;

The sixth rear channel is an axial hole on the main shaft that communicates with the radial hole.

Furthermore, the first channel behind the plate and the third channel behind the plate have the function of centrifugal ventilation at the same time, preventing fuel or lubricating oil from entering the channels behind the plate, and causing flow channel blockage caused by glue in the cavity; correspondingly, consider centrifugal ventilation The windage effect of the device controls the pressure difference between the air-entraining shunt cavity and the caudal cavity after the disc, so that the pressure in the air-entraining shunt cavity is more than 20% higher than the pressure in the caudal cavity after the disc;

Control the flow of the cooling airflow before the disk and the cooling airflow after the disk, so that the pressure in the bleed air distribution cavity is greater than the pressure in the front cavity of the turbine disk and the pressure in the caudal cavity after the disk; at the same time, the pressure in the front cavity of the turbine disk is higher than that behind the turbine guide The pressure of the external gas flow, the pressure in the caudal cavity behind the disk is higher than the pressure of the external gas flow behind the turbine rotor.

Furthermore, the channel in front of the second disc is composed of the rear disc of the turbine guide and the front disc of the turbine rotor, and the channel behind the seventh disc is composed of the rear disc of the turbine and the front end of the tail cone; The width of the channel and the seventh channel behind the disc makes it difficult for external gas to flow back into the front cavity of the turbine disc and the tail vertebra cavity behind the disc, while ensuring sufficient safety margin; in addition, except for the channel between the tail vertebra and the turbine rotor, The rest is sealed and airtight.

The axial gap distance between the rear disk of the turbine guide and the front disk of the turbine rotor shall not exceed 2mm, and shall not be less than 0.8mm in consideration of the safety margin;

The axial gap between the rear disk of the turbine and the front end of the tail cone shall not exceed 3mm, but shall not be less than 1.5mm in consideration of the safety margin;

In order to further reduce the gap value of the channel behind the seventh disk under the premise of ensuring a safety margin, it is necessary to make the maximum outer diameter of the tail vertebra lower than the inner cylindrical surface of the trailing edge of the turbine disc, so that the gap value of the channel behind the seventh disc Reach below 1.5mm, when space permits, it can even reach 0 or negative value.

Aiming at reducing the gap value of the channel behind the seventh disk, a further optimization idea is given: the tail vertebra is folded to form a cylindrical surface of the tail vertebra, and the rim of the rear disk of the turbine rotor is optimized to meet the strength requirements. Under the premise, reduce the radial wall thickness at this place to form the inner cylindrical surface of the turbine disk rim; ensure that the radial dimension of the tail cone cylindrical surface is lower than the radial dimension of the rightmost end of the inner cylindrical surface of the turbine disk rim.

Further, the rear bearing and the main shaft are centered at the first mating surface, the middle-stage sealing isolation disk and the main shaft are centered at the second mating surface, and the turbine rotor and the intermediate-stage sealing isolation disk are at the third mating surface Realize centering at the center, and realize good centering of the turbine rotor and the rear bearing under the working heat state;

The outer mating cylindrical surface of the front boss of the turbine rotor is in a mating relationship with the rear mating end surface on the intermediate sealing isolation disc. Under the working heat state, the temperature of the turbine rotor at the third mating surface is significantly higher than that of the intermediate sealing isolation disc. Even considering the centrifugal stress of the two, at this time, there is a tendency to tighten between the outer matching cylindrical surface of the front boss and the rear matching end surface; thus, in the hot state, the centering effect at the third matching surface is strengthened compared with that during assembly ;

In the working state, at the second mating surface, the temperature of the main shaft and the sealing isolation plate of the intermediate stage is the same, and as a small disk surface, the expansion of the inner cylindrical surface of the front fit under the centrifugal stress is negligible, so the main shaft and the intermediate stage are in the working state The centering effect between the sealed isolation discs remains unchanged compared with the cold state; while the second mating surface and the first mating surface are actually the same processing surface on the main shaft; thus ensuring that the turbine rotor and the rear bearing are working Controllable eccentricity in hot state.

Further, the airflow through the middle passage cools the rear disk surface of the turbine rotor, and the cooling airflow passes through the middle passage formed by opening the radial hole in front of the bearing on the main shaft and the axial hole facing the tailbone into the tailbone cavity behind the disk, and the turbine rotor Cool down; or cool down the rear disk surface of the turbine rotor through the air flow of the external channel, use the pipeline to lead the cooling air from the engine or other high-pressure air generation facilities, enter the tail vertebrae cavity after the disk through the external channel opened on the tail vertebra, and cool the turbine rotor Cool down.

Compared with the prior art, the remarkable progress of the present invention lies in: 1) the hot-end rotor system has been more optimized and more effectively cooled; Improved the cooling of the bearings; allows the use of ordinary industrial-grade bearings to reduce costs; at the same time, it can reduce the amount of lubrication and improve engine performance; 2) The cooling of the main shaft has been redesigned, so that the main shaft uses cheap Easy-to-process materials are possible, reducing costs and improving functional expansion; or on the basis of using the same material for the main shaft, the gas temperature of the turbine system can be further increased, and the gas temperature before the turbine has a greater impact on engine performance. Improved performance; further, even high-strength stainless steel materials that are not resistant to high temperatures can be used as the main shaft material to achieve a good balance between cost and corrosion; 3) without changing the rotor dynamics and vibration characteristics, without increasing assembly In the case of complexity, at the cost of the simplest structure, the problem of thermal centering of the turbine rotor and bearing on the micro turbojet engine is optimized and solved.

In order to more clearly illustrate the functional characteristics and structural parameters of the present invention, further description will be given below in conjunction with the accompanying drawings and specific embodiments.

Description of drawings

The accompanying drawings described here are used to provide a further understanding of the present invention and constitute a part of the application. The schematic embodiments of the present invention and their descriptions are used to explain the present invention and do not constitute improper limitations to the present invention. In the attached picture:

Fig. 1 is the schematic diagram of the rotor system of the hot end (turbine end) of the miniature turbojet engine in the prior art;

Fig. 2 is the hot end (turbine end) rotor system structure and the schematic diagram of cooling air flow of the present invention;

Fig. 3 is a schematic diagram of a three-dimensional partial section of an intermediate-level sealing isolation disk of the present invention;

Fig. 4 is a three-dimensional schematic diagram of a turbine rotor of the present invention;

Fig. 5 is a schematic diagram of a three-dimensional partial section of the rear part of the main shaft of the present invention;

6 is a detailed enlarged schematic diagram of the cooling air flow path of the present invention;

Fig. 7 is an enlarged schematic diagram of the cooling air flowing into the fuel gas of the present invention;

8 is a schematic diagram of the temperature distribution achieved by the hot end (turbine end) rotor system of the present invention;

Fig. 9 is a schematic diagram of the centering structure of the hot end rotor of the present invention;

Fig. 10 is a schematic diagram of bleed air in other embodiments of the present invention;

The reference signs in the figure are: main shaft 1, front matching cylindrical surface 101, channel front end surface 102, channel cylindrical surface 103, channel rear end surface 104, rear matching cylindrical surface 105, radial hole 106, axial hole 107, Shaft sleeve 2, circular ring 201, rear bearing 3, intermediate sealing isolation disc 4, outer toothed disc 401, radial air-introduction hole 402, front mating inner cylindrical surface 403, rear mating end face 404, front mating end face 405, rear mating end face 405, rear Cooperate with inner cylindrical surface 406, turbine guider 5, turbine guider rear disc surface 501, turbine rotor 6, front boss outer chamfer 601, front end face 602, front boss inner chamfer 603, front boss channel 604, inner Hole 605, outer matching cylindrical surface 606 of front boss, rear turbine disk surface 608, inner cylindrical surface of turbine disk rim 609, tail nozzle 7, tail vertebra 8, tail vertebra front end surface 801, tail vertebra cylindrical surface 802, total cooling airflow 9 , pre-pan cooling airflow 910, first pre-pan channel 911, second pre-pan channel 912, post-pan cooling airflow 920, first post-pan channel 921, second post-pan channel 922, third post-pan channel 923, fourth The rear channel 924, the fifth channel 925, the sixth channel 926, the seventh channel 927, the middle channel airflow 930, the outer channel airflow 940, the turbine nut 10, the bleed air distribution chamber H1, the turbine disk front chamber H2, postdisc caudal cavity H3, the first mating surface 3-1, the second mating surface 4-1, and the third mating surface 4-6.

Detailed ways

The technical solutions in the embodiments of the present invention will be clearly and completely described below in conjunction with the accompanying drawings in the embodiments of the present invention. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of them; based on The embodiments of the present invention and all other embodiments obtained by persons of ordinary skill in the art without making creative efforts belong to the protection scope of the present invention.

As shown in Figure 1, Figure 1 is a common rotor system at the hot end (turbine end) of a micro turbojet engine. The high-pressure air (about 250°C) used for cooling enters the shaft sleeve through certain throttling measures, and is mixed with Part of the lubricating oil, after lubricating and cooling the rear bearing, enters the front cavity of the turbine disc to convectively cool the front of the turbine rotor disc; there is a cavity between the rear disc of the turbine rotor and the tail cone, and the gas inside will be continuously The high-speed rotating turbine disk is thrown out, and the gas (about 800°C) from the outside is constantly entering, forming a certain vortex and causing the turbine disk to be surrounded by high-temperature gas; the most end of the main shaft, and the turbine nut on the end of the main shaft is more important , is also surrounded by high-temperature gas, where the temperature of the parts will reach more than 600 ° C, and the ambient temperature is very harsh.

Example

In this embodiment, for the convenience of description, with reference to the heading of the engine, the left side of the picture is defined as the front, and the right side is defined as the rear; the outside in the text is relative to the inside in the same structure;

As shown in Figure 2, the present invention is a double-optimized cooling and structure hot-end rotor system, including a main shaft 1, a bushing 2, a rear bearing 3, an intermediate stage sealing isolation disc 4, a turbine guide 5, and a turbine rotor 6. Exhaust nozzle 7, tail cone 8, turbine nut 10; among them, the rear bearing 3, the intermediate stage sealing isolation plate 4, and the turbine rotor 6 are sequentially inserted and installed on the main shaft 1, and are compressed and fixed by the turbine nut 10; the rear bearing 3 is installed on the shaft sleeve 2, and the rear bearing 3 is closely attached to the end surface of the shaft sleeve 2 through the spring preload force from the front end of the main shaft and the aerodynamic axial force during the working process; the turbine guide 5 is fixed on the shaft sleeve 2 through threads or fasteners on the shaft sleeve 2; the tail cone 8 is integrated with the tail nozzle 7 in the form of welding through the support plate; and the tail nozzle 7 and the turbine guide 5 are generally directly fixed by fasteners, and can also be used with other components (such as machine parts). Cassette) fixed mode remains relatively fixed with the position of turbine guide 5.

Specifically, when the engine is working, the high-pressure cooling airflow enters the bushing 2, and after certain throttling measures, it becomes the total cooling airflow 9, which is mixed with the lubricating oil sprayed to the rear bearing to lubricate and cool the rear bearing 3; In the middle of the rear bearing 3 and the turbine rotor 6, there is an intermediate stage sealing isolation disc 4, and the intermediate stage sealing isolation disc 4 plays the roles of heat insulation, bleed air cooling and sealing.

Specifically, the middle-stage sealing isolation disc 4, the shaft sleeve 2, and the rear bearing 3 surround the bleed air distribution chamber H1; the total cooling airflow 9 for lubricating and cooling the rear bearing 3 reaches the bleed air distribution chamber H1, Divided into two streams, respectively, the front cooling airflow 910 and the rear cooling airflow 920; the front cooling airflow 910 enters the turbine disk front chamber H2 through the intermediate sealing isolation disk 4, and mainly cools the front disk surface of the turbine rotor 6 After cooling down, it flows out from the front channel 912 between the turbine guide 5 and the turbine rotor 6, and merges into the main gas flow; the cooling air flow 920 behind the disk passes through the main shaft 1, the intermediate sealing isolation disk 4 and the turbine rotor 6 The complex flow path formed by the interaction enters the post-disc caudal cavity H3 between the turbine rotor 6 and the tail vertebra 8, and mainly heats and cools the rear part of the main shaft 1, the rear disc surface of the turbine rotor 6, and the turbine nut 10. It flows out from the back-disk channel 927 between the turbine rotor 6 and the tail cone 8, and merges into the main gas flow.

As shown in Figure 3, the middle-level sealing isolation disc 4 is a stepped annular structure, including a front matching inner cylindrical surface 403, a rear matching end surface 404, a front matching end surface 405, a rear matching inner cylindrical surface 406, and a rear matching inner cylindrical surface. There is a circle of circumferentially distributed radial air-introduction holes 402 above the 406; at the maximum diameter of the intermediate sealing isolation disc 4, there is an outer toothed disc 401. The volume and weight of the middle-stage sealing isolation disc 4 is larger than that of the front end of the original turbine rotor, which can disperse the heat transferred from the turbine rotor 6, and convect heat through its relatively complex profile and cooling channels, thereby strengthening the Cooling effect, and its outer toothed disc structure can achieve gas sealing effect.

As shown in FIG. 4 , the structure of blades and discs of the turbine rotor 6 is similar to that of a general axial flow turbine, but a cylindrical boss is designed in front of the turbine rotor to realize part of the functional requirements of the present invention. The turbine rotor 6 includes an outer chamfer 601 of the front boss, a front end face 602 , an inner chamfer 603 of the front boss, a channel 604 of the front boss, an inner hole 605 , and an outer matching cylindrical surface 606 of the front boss. The function of the front boss channel 604 of the turbine rotor is to reduce the contact area with the intermediate-stage sealing isolation disk 4 and reduce the heat transfer to the intermediate-stage sealing isolation disk 4 and the rear bearing 3; part of the way.

As shown in Figure 5, the rear half of the main shaft 1 includes a front matching cylindrical surface 101, a rear matching cylindrical surface 105, a channel cylindrical surface 103 in the middle of the two matching surfaces, a front end surface 102 of the channel, a rear end surface 104 of the channel, and a circumferential The distributed radial holes 106, the axial holes 107 communicating with the radial holes 106, the front mating cylindrical surface 101, the rear bearing 3 and the front mating inner cylindrical surface 403 form a centering fit.

As shown in Figure 6, Figure 6 is a detailed enlarged schematic diagram of the cooling air flow path of the present invention, and the detailed implementation of the cooling air flow path is as follows: the outer toothed plate 401 on the middle stage sealing isolation plate 4, the rear part of the shaft sleeve 2 The protruding ring 201 and the rear bearing 3 surround the bleed air distribution chamber H1, and the total cooling airflow 9 is divided into two airflows, the front cooling airflow 910 and the rear cooling airflow 920, in the bleed air distribution chamber H1. The outer grate plate 401 and the protruding ring 201 at the rear of the shaft sleeve 2 form the first channel 911 in front of the plate. This channel is a typical seal structure of the grate tooth. The unilateral sealing gap of the protruding circular ring 201 is between 0.05 mm and 0.25 mm. The main function of this sealing structure is to ensure that the bleed air distribution chamber H1 reaches a relatively high pressure, to ensure that the cooling air maintains a pressure gradient, and can flow stably in accordance with the design direction; at the same time, it can prevent the cooling airflow in front of the disk leaking from the channel 911 in front of the first disk The flow rate of 910 is throttled to reduce the performance loss of the engine. The pre-disc cooling airflow 910 passes through the first disc front channel 911, enters the turbine disc front chamber H2, and then flows out through the second disc front channel 912 between the turbine guide 5 and the turbine rotor 6, and then merges into the main gas flow. From the perspective of realization, the sealing of the channel 911 in front of the first plate is not only limited to the grate tooth sealing, but also can use sealing methods such as floating rings, expansion rings, honeycombs, and brushes.

The post-disc cooling airflow 920 enters the post-disc caudal cavity H3 between the turbine rotor 6 and the caudal vertebra 8 through the complex flow channel formed by the interaction of the main shaft 1 , the intermediate sealing isolation disc 4 and the turbine rotor 6 .

Specifically, the first rear-disk channel 921 through which the post-disk cooling airflow 920 passes is formed by a circle of circumferentially distributed radial air-introducing holes 402 on the middle-stage sealing isolation disc 4 , and the radial holes are distributed along the circumferential direction.

Specifically, the first rear channel 922 through which the cooling airflow 920 passes through is formed by the outer chamfer 601 of the front boss on the turbine rotor 6, the rear mating end surface 404 on the intermediate sealing isolation disc 4, and the rear mating inner cylindrical surface. 406 is enclosed after installation and is a circumferential circular channel with a triangular cross-section; the outer chamfer 601 of the front boss is chamfered in the size of C1~C3, which is selected according to the actual situation. Too short will affect the controllability of the airflow connection. If it is too large, it will affect the length of the mating surface; the section can also be in other shapes such as a rectangle; it is also possible to process a circumferential ring groove outward through the intersection of the rear mating end surface 404 and the rear mating inner cylindrical surface 406 on the middle stage sealing isolation disc 4 , communicate with the radial air-introduction hole 402 to achieve the same effect as the second disc rear channel 922, and at this time, the outer chamfer 601 of the front boss may not be required for the size of the flow-through purpose.

Specifically, the third rear-disk channel 923 through which the cooling airflow 920 passes is formed by the turbine rotor 6 after the rear mating end surface 404 on the isolation disk 4 of the intermediate stage is fitted and bonded to the front end surface 602 on the turbine rotor 6 . The channel formed after the top of the front boss channel 604 is covered by the rear mating end surface 404 .

Specifically, the fourth back-disk channel 924 through which the post-disk cooling airflow 920 passes is formed by the gap space between the cylindrical surface 103 of the channel on the main shaft 1 and the inner hole 605 of the turbine rotor; the diameter of the cylindrical surface 103 is smaller than that of the inner hole 605 Diameter, the gap between the two forms a circumferential annular channel with a rectangular cross-section; the axial length of the annular fourth disc rear channel 924 can be longer than the axial length of the turbine rotor, and the schematic diagram is only for illustration; in order to ensure the reliability of the airflow To be unblocked, the front boundary 102 of the annular cavity should partially overlap with the third disk rear channel 923 (the depth of the front boss channel 604 ) of the turbine rotor, or partially overlap with the front boss inner chamfer 603 of the turbine rotor 6 .

Specifically, the fifth rear channel 925 through which the cooling airflow 920 passes is formed by a circle of circumferentially distributed radial holes 106 on the main shaft 4, and the radial holes are preferably uniformly distributed along the circumferential direction; Stress is more complicated, the number of holes is controlled at 2 to 8, and the diameter of the holes does not exceed 30% of the diameter of the matching cylindrical surface 105 behind the main shaft.

Specifically, the sixth rear-disk passage 926 through which the after-disk cooling airflow 920 passes is the axial hole 107 on the main shaft 1 that communicates with the radial hole 106 . The post-disc cooling airflow 920 sequentially passes through the above-mentioned complex airflow channels formed by the above-mentioned post-disk channels 921-926, which isolates the direct heat transfer from the turbine rotor to the main shaft, and cools the turbine rotor 6 and the main shaft 1. The temperature rises slightly, and then enters the post-disc caudal cavity H3 between the turbine rotor 6 and the tail vertebra 8, after cooling the rear disk surface of the turbine rotor 6 and the turbine nut 10, and then from between the turbine rotor 6 and the tail vertebra 8 The seventh after-market channel 927 in between flows out and merges into the main gas flow.

Specifically, the leakage flow rate of the cooling airflow 920 behind the disk can be throttled and controlled through the machining depth of the turbine rotor front boss groove 604, or the matching gap between the groove cylindrical surface 103 and the turbine rotor inner hole 605; The leakage of the after-disc cooling airflow 920 is controlled at a low level to reduce engine performance loss.

Specifically, the radial first disc channel 921 and the radial first disc channel 923 also have the function of centrifugal ventilation, which can prevent fuel or lubricating oil from entering the internal channels 921-926 and causing glue formation in the cavity. Problems such as flow channel blockage. Correspondingly, considering the windage effect of the centrifugal fan, it is necessary to control the pressure difference between the bleed air shunt chamber H1 and the retrocaudal cavity H3 so that the pressure in the bleed air shunt chamber (H1) is 20% higher than the pressure in the retrocaudal chamber H3. %; Correspondingly, the radial air-introduction hole 402 should be radially outward; the hole can be radial, or can be tilted forward at a certain angle, but the forward tilt should not exceed 50°.

Specifically, the setting of the second disc rear channel 922 and the fourth disc rear channel 924 in the inner and outer double-ring chambers can ensure smooth air flow, and at the same time, it is unnecessary to consider the circumferential positioning between the intermediate stage sealing isolation disc 4 and the turbine rotor 6, reducing the increased assembly complexity. The fourth disk rear channel 924 can also be extended to the front end of the turbine nut, and a radial groove is opened on a certain part at the end face to realize airflow extraction.

Specifically, it is necessary to control the flow rates of the pre-disc cooling airflow 910 and the post-disc cooling airflow 920 within a certain range to ensure that the pressure in the bleed air splitting chamber H1 is greater than the pressure in the turbine disc anterior cavity H2 and the post-disk caudal cavity H3 ;Further, ensure that the pressure in the front chamber H2 of the turbine disk is higher than the pressure of the external gas flow behind the turbine guide, and ensure that the pressure in the caudal cavity H3 after the disk is higher than the pressure of the external gas flow behind the turbine rotor; finally ensure that all cooling air flows Able to flow stably in the designed direction.

As shown in Figure 7, in order to prevent excessive backflow of gas, it is necessary to ensure that the second disc front channel 912 and the seventh disc rear channel 927 are as small as possible so that the external gas is as difficult as possible to flow into the turbine disc front cavity H2 and the rear disc caudal cavity H3; while ensuring sufficient safety margin. In addition, except for the seventh back-disc channel 927 between the tail vertebra 8 and the turbine rotor 6, the remaining parts are closed and airtight.

Specifically, for the second disk front channel 912, on the premise of sufficient structural margin, the axial clearance distance between the turbine guider rear disk 501 and the turbine rotor front disk 607 is designed to be sufficiently small, not exceeding 2mm.

Specifically, for the seventh disc rear channel 927, it is also necessary to reduce the axial clearance distance between the turbine rear disc surface 608 and the front end surface 801 of the tail cone as much as possible, and at the same time avoid the high-speed rotating turbine rear disc surface 608 and the front end of the tail vertebra When the surface 801 collides, a safe distance needs to be maintained; for this reason, further optimization is as follows: under the premise of ensuring that the gas flow field does not deteriorate, the tail vertebra is folded to form the tail vertebra cylindrical surface 802; through the turbine rotor 6 Optimizing the rim of the rear disk surface, and reducing the radial wall thickness at the rim on the premise of satisfying the strength, forms the inner cylindrical surface 609 of the turbine disk rim. Ensure that the radial dimension of the tail cone cylindrical surface 802 is lower than the radial dimension of the rightmost end of the inner cylindrical surface 609 of the turbine disk rim, and there is sufficient structural margin; on this premise, the inner cylindrical surface 609 of the turbine disk rim can also have A certain slope; at the same time, the cylindrical surface 802 of the caudal vertebra can also be a non-cylindrical surface, or even not folded. In this way, the front end surface 801 of the tail cone can be close enough to the rear disk surface 608 of the turbine, reaching less than 2mm or even 0mm; further, under the premise of sufficient space, the front end surface 801 of the tail cone can extend forward to overlap with the turbine rotor in the axial direction to achieve negative Value; the above sealing measures, coupled with the ejection effect of high-speed gas, can greatly reduce the gas entering the tail cone, reduce the flow demand for the cooling airflow 920 behind the disc, reduce the bleed air loss, which is equivalent to improving the engine performance.

As shown in Figure 8, Figure 8 is a schematic diagram of the temperature distribution achieved by the hot end (turbine end) rotor system of the present invention. Through the above cooling measures, combined with conventional heat insulation and cooling of the shaft sleeve, the temperature of the hot end rotor system can be improved. Cooling optimization will have the desired effect. The post-disc caudal cavity H3 in the coccyx 8 has changed from a high-temperature region to a medium-low temperature region. Even considering the heating of the cooling airflow and the infiltration of a very small amount of gas, the temperature of the airflow in the coccyx 8 will not exceed 450°C. The temperature in the central area of the disk body of the turbine rotor 6 has dropped by more than 100°C. The rear of the main shaft 1 is reduced due to the passage of cooling airflow and the decrease in heat transfer caused by the temperature drop of the adjacent turbine rotor 6 itself. The rear of the main shaft 1 The temperature must not exceed 450°C, and the drop rate exceeds 200°C; cheap and easy-to-process materials can be used. Due to the existence of the intermediate sealing isolation disc 4, the heat transfer from the hot end parts to the bearing is further reduced. The temperatures are all lower than 300°C, fundamentally solving the problem of heat transfer to the bearing 2 . At the same time, it can be seen that the temperature at the joint between the intermediate sealing isolation disk 4 and the main shaft 1 is basically close, and there is no large temperature difference. During the calculation of the present invention, it is assumed that the temperature of the cooling air cited is: a maximum temperature of 250° C. that can be achieved by a general micro-turbojet engine through compressor compression and temperature rise, without considering the temperature drop of the oil-gas mixture after being mixed with oil and the temperature of the mixture. The improvement of the convective heat transfer coefficient. Therefore, the actual bearing temperature can be kept below 200°C. In this case, industrial-grade high-speed bearings of common materials supplied in large quantities can be directly used; there is no need to use high-nitrogen stainless steel bearings specially used for current micro-turbojet engines, which greatly reduces costs.

Specifically, the present invention can be used in grease lubricated bearings as well as oil lubricated bearings.

As shown in Fig. 1, Fig. 3, Fig. 4 and Fig. 9, the intermediate sealing isolation disc 4 is designed with stepped holes, and has a large overlapping area with the turbine rotor 6 in the axial direction, shortening the total axial length. Realize their respective functions without affecting the strength of the turbine disk; compared with the original hot end rotor system of the same type and specification, the quality of the rotor system, the distance from the center of gravity to the bearing is basically the same, so it has no impact on rotor dynamics and vibration. There is an additional impact that can be easily retrofitted to existing rotors. The setting of the second disc rear channel 922 and the fourth disc rear channel 924 in the inner and outer double ring chambers can ensure smooth air flow, and at the same time, the intermediate stage seals the isolation disc 4, the main shaft 1, and the turbine rotor 6 without circumferential positioning, reducing the Assembly complexity.

As shown in Fig. 9, in view of the problem that the centering of the general hot-end rotor parts deteriorates under the thermal state, the solution of this embodiment is as follows: the turbine rotor 6 and the sealing isolation disc 4 in the middle stage are centered at the mating surface 4-6, The centering of the intermediate sealing isolation disc 4 and the main shaft 1 is realized at the second mating surface 4-1, and the centering of the rear bearing 3 and the main shaft 1 is realized at the first mating surface 3-1. The ultimate goal is to realize the turbine rotor 6 Good centering with rear bearing 3 in working (hot) condition. The outer mating cylindrical surface 606 of the front boss of the turbine rotor 6 is in a mating relationship with the rear mating end face 404 on the intermediate sealing isolation disc 4. Under the hot working state, the temperature of the turbine rotor at the third mating surface 4-6 is significantly higher than The intermediate stage seals the isolation plate tightly. At this time, the outer matching cylindrical surface 606 of the front boss and the rear matching end surface 404 tend to be tightened. The centering effect of the third matching surface 4-6 in the working state is stronger than that during assembly. of.

Specifically, in the working state, the temperature of the main shaft 1 at the second mating surface 4-1 is basically the same as that of the intermediate sealing isolation disc 4, and as a small disc body, the front mating inner cylindrical surface 403 under centrifugal stress The expansion is negligible, so in the working state, the centering effect between the main shaft 1 and the intermediate sealing isolation plate 4 remains basically unchanged compared with the cold state; while the second mating surface 4-1 and the first mating surface 3- 1 is actually the same processing surface on the main shaft 1, thus, it can ensure that the eccentricity of the turbine rotor 6 and the rear bearing 3 is controllable under working conditions.

As shown in Figure 10, the present invention can also directly open a radial hole at the position in front of the bearing on the main shaft 1 and an axial hole facing the tail vertebra 8 to form the middle channel airflow 930 as shown in Figure 10, and enter the rear tail of the disc. The vertebral cavity H3 achieves a similar cooling effect. Pipelines can also be used to lead cooling air from the engine or other high-pressure air generation facilities, such as the external channel airflow 940 shown in Figure 10, into the caudal cavity H3 behind the disc, and the cooling effect can also be achieved.

Specifically, the seal between the turbine rotor 6 and the tail cone 8 can also adopt a more complicated seal structure such as a toothed tooth, which can achieve better results; but the cost will increase.

Specifically, if the actual engine pressure ratio is lower, or measures are taken to introduce cooling air with lower pressure (but sufficient) and temperature, a better cooling effect will be achieved, and the temperature of the bearing and the main shaft will be lower.

It should be noted that in this article, relational terms such as first and second are only used to distinguish one entity or operation from another entity or operation, and do not necessarily require or imply that there is a relationship between these entities or operations. There is no such actual relationship or order between them. Furthermore, the term "comprises", "comprises" or any other variation thereof is intended to cover a non-exclusive inclusion such that a process, method, article, or apparatus comprising a set of elements includes not only those elements, but also includes elements not expressly listed. other elements of or also include elements inherent in such a process, method, article, or device.

Although the embodiments of the present invention have been shown and described, those skilled in the art can understand that various changes, modifications and substitutions can be made to these embodiments without departing from the principle and spirit of the present invention. and modifications, the scope of the invention is defined by the appended claims and their equivalents.

Claims (10)

1. a micro-miniature turbojet engine hot end rotor system with core cooling structure is characterized in that it comprises a main shaft (1), an axle sleeve (2), a rear bearing (3), an intermediate stage sealing isolation disc (4 ), turbine guider (5), turbine rotor (6), tail nozzle (7), tail cone (8); the rear bearing (3), intermediate stage sealing isolation disc (4), turbine rotor (6 ) are inserted and installed on the main shaft (1) in turn, and fixed by pressing the turbine nut (10); The aerodynamic axial force in the working process makes the rear bearing (3) close to the end face of the bushing (2); the turbine guide (5) is fixed on the bushing (2) by fasteners; the tail cone ( 8) Integrate with the tail nozzle (7) in the form of welding through the support plate, and the tail nozzle (7) is connected to the turbine guide (5); When the engine is working, the high-pressure cooling airflow enters the shaft sleeve (2), and after throttling measures, it becomes the total cooling airflow (9), which is mixed with the lubricating oil sprayed to the rear bearing to lubricate and cool the rear bearing (3) ; The total cooling airflow (9) is divided into the front cooling airflow (910) and the cooling airflow (920) after the disk, and the cooling airflow (910) before the disk and the cooling airflow (920) after the disk are respectively the front disk surface of the turbine rotor (6). After cooling down with the rear panel, it merges into the main gas flow. 2. a kind of micro turbojet engine hot-end rotor system with core cooling structure according to claim 1, is characterized in that, described intermediate stage sealing isolating disc (4), axle sleeve (2), rear The bearing (3) surrounds and forms a bleed air splitting cavity (H1); the front end of the turbine rotor (6) is provided with a turbine disc front cavity (H2), and the rear end of the turbine rotor (6) is provided with a disc caudal cavity ( H3); The total cooling airflow (9) reaches the rear bearing (3) and enters the bleed air distribution cavity (H1) and is divided into a front cooling airflow (910) and a rear cooling airflow (920); the front cooling airflow (910) passes through The intermediate sealing isolation disk (4) enters the turbine disk front cavity (H2), and after cooling the front disk surface of the turbine rotor (6), the second stage between the turbine guider (5) and the turbine rotor (6) The air flow from the front channel (912) of the second disk flows out and merges into the main gas flow; the cooling airflow (920) behind the disk passes through the interaction of the main shaft (1), the intermediate sealing isolation disk (4) and the turbine rotor (6). Each flow passage of the turbine rotor (6) enters the disc caudal cavity (H3) between the turbine rotor (6) and the tail vertebra (8), and heats and cools the rear disc surface of the turbine rotor (6) and the turbine nut (10). After that, it flows out from the seventh disc rear channel (927) between the turbine rotor (6) and the tail cone (8), and merges into the main gas flow. 3. The hot-end rotor system of a micro turbojet engine with a core cooling structure according to claim 1, wherein the main shaft (1) is a cylindrical shaft body, and the rear end of the main shaft (1) is arranged There are a front fitting cylindrical surface (101) and a rear fitting cylindrical surface (105), and a groove cylindrical surface (103) is arranged in the middle of the two fitting cylindrical surfaces; the front fitting cylindrical surface (101) is arranged near the end of the groove cylindrical surface (103) There is a front end surface (102) of the channel; the end of the rear mating cylindrical surface (105) close to the cylindrical surface of the channel (103) is provided with a rear end surface (104) of the channel; the cylindrical surface of the channel (103) is provided with circumferentially distributed radial hole (106), the bottom of the radial hole (106) is connected with an axial hole (107); The intermediate sealing isolation disc (4) has a stepped annular structure, and the diameter of the annular structure of the intermediate sealing isolation disc (4) close to the rear bearing (3) is smaller, and the diameter of the annular structure far away from the rear bearing (3) is smaller. Larger; the end of the annular structure with a smaller diameter near the rear bearing (3) is provided with a front mating end surface (405), the inner side is provided with a front mating inner cylindrical surface (403), and the end far away from the rear bearing (3) is provided with a rear mating end face (404); the inner side of the annular structure with a larger diameter is provided with a rear mating inner cylindrical surface (406), and a sealing structure is arranged on the outer side of the largest diameter of the annular structure with a larger diameter; the middle stage is sealed and isolated There is a circle of circumferentially distributed radial air-introducing holes (402) on the disk (4), leading to the rear mating inner cylindrical surface (406); The front mating cylindrical surface (101) forms a centering fit with the rear bearing (3) and the front mating inner cylindrical surface (403); the intermediate sealing isolation disc (4) is used to isolate and lower the turbine rotor (6) The heat transfer to the rear bearing (3); the middle-stage sealing isolation disc (4) has a larger heat dissipation area and cooling channel than the front end of the turbine rotor (6), which can disperse the heat transferred from the turbine rotor (6) to To strengthen the cooling effect, while the sealing structure is used to achieve the effect of gas sealing; The blade and disc structure of the turbine rotor (6) is an axial flow turbine structure, the center of the turbine rotor (6) is provided with an inner hole (605), and a ring structure is protruded outward around the inner hole (605), so The outer surface of the annular structure is the outer matching cylindrical surface (606) of the front boss; the front center of the ring structure is provided with several front bosses with three faces, and the front boss is far away from the surface of the inner hole (605). It is the outer chamfer of the front boss (601), the front side of the front boss is the front end face (602), the surface of the front boss near the inner hole (605) is the inner chamfer of the front boss (603), several front A front boss channel (604) is arranged between the bosses; the front boss channel (604) is used to reduce the contact area with the intermediate sealing isolation disc (4) and reduce the contact area to the intermediate sealing isolation disc. (4) and the heat transfer of the rear bearing (3); at the same time as a part of the flow path of the cooling airflow (920) behind the disk. 4. The hot-end rotor system of a micro turbojet engine with a core cooling structure according to claim 2, characterized in that, the bleed air split chamber (H1) is sealed by an intermediate stage isolation disc (4) The sealing structure above, the ring (201) protruding from the rear of the shaft sleeve (2), and the rear bearing (3) are formed around; the sealing structure is formed with the ring (201) protruding from the rear of the shaft sleeve (2) The first disk front passage (911); the front cooling airflow (910) passes through the first disk front passage (911), enters the turbine disk front chamber (H2), and then passes through the gap between the turbine guide (5) and the turbine rotor (6) The channel (912) in front of the second disc in the room flows out and merges into the main gas flow. 5. The hot-end rotor system of a micro turbojet engine with a core cooling structure according to claim 4, wherein the cooling air flow (920) after the disk passes through the first passage behind the disk (921), the second The complex airflow channel composed of the second after-display channel (922), the third after-display channel (923), the fourth after-display channel (924), the fifth after-display channel (925), and the sixth after-display channel (926) is The turbine rotor (6) and the rear of the main shaft (1) are cooled, which also reduces the heat transfer from the turbine rotor to the main shaft (1) and the intermediate sealing isolation disc (4); the cooling airflow (920) behind the disc then enters the turbine After cooling the rear disk surface of the turbine rotor (6) and the turbine nut (10) in the post-disc caudal cavity (H3) between the rotor (6) and the tail vertebra (8), the turbine rotor (6) and the turbine nut (10) are then cooled. The seventh post-disc channel (927) between the caudal vertebrae (8) flows out and merges into the main gas flow. 6. a kind of micro-miniature turbojet engine hot end rotor system with core cooling structure according to claim 5, is characterized in that, passage (921) behind the described first disc, is sealed by intermediate stage and isolating disc ( 4) A circle of circumferentially distributed radial air-introducing holes (402) opened on the top is formed; The channel behind the second disc (922) is composed of the outer chamfer (601) of the front boss on the turbine rotor (6), the upper rear mating end surface (404) of the intermediate stage sealing isolation disc (4) and the rear mating inner cylinder The surface (406) is enclosed after installation and is a circumferential annular channel with a triangular cross-section; The passage behind the third disk (923) is installed and bonded to the rear mating end surface (404) on the isolation disk (4) by the intermediate stage and the front end surface (602) on the turbine rotor (6), and the turbine rotor ( 6) The top of the front boss channel (604) on the top is covered by the rear mating end surface (404); The passage behind the fourth disc (924) is formed by the clearance space between the cylindrical surface (103) of the channel on the main shaft (1) and the inner hole (605) of the turbine rotor; the diameter of the cylindrical surface (103) is smaller than that of the inner hole ( 605), the gap between the two constitutes a rectangular section of the circumferential ring channel; in order to ensure the reliability and smoothness of the air flow, the front end of the channel (102) partially overlaps with the third rear channel (923), or The inner chamfer (603) of the front boss of the turbine rotor (6) is partially overlapped; The fifth rear channel (925) is formed by a circle of circumferentially distributed radial holes (106) on the main shaft (1); The sixth rear channel (926) is an axial hole (107) on the main shaft (1) communicating with the radial hole (106). 7. a kind of micro turbojet engine hot end rotor system with core cooling structure according to claim 2, is characterized in that, passage (921) after the first disc and passage (923) after the 3rd disc, simultaneously It has the function of centrifugal ventilation to prevent fuel or lubricating oil from entering the channels behind the discs, causing flow channel blockage caused by glue in the cavities; The air pressure difference in the caudal vertebral cavity (H3) makes the pressure in the bleed air shunt cavity (H1) higher than the pressure in the posterior caudal vertebral cavity (H3) by more than 20%; Controlling the flow of the front cooling airflow (910) and the rear cooling airflow (920), so that the pressure in the bleed air split chamber (H1) is greater than the pressure in the turbine disk front chamber (H2) and the rear caudal cavity (H3); At the same time, the pressure in the front cavity (H2) of the turbine disk is higher than the pressure of the external gas flow behind the turbine guide, and the pressure in the caudal cavity (H3) behind the disk is higher than the pressure of the external gas flow behind the turbine rotor. 8. The hot end rotor system of a micro turbojet engine with a core cooling structure according to claim 5, characterized in that, the second disk front channel (912) is formed by the turbine guide rear disk surface (501) It is composed of the front disk surface (607) of the turbine rotor, and the passage behind the seventh disk (927) is composed of the rear disk surface (608) of the turbine and the front end surface (801) of the tail vertebra; The width of the channel (912) and the seventh channel behind the disk (927) makes it difficult for the external gas to flow back into the front chamber (H2) of the turbine disk and the caudal cavity (H3) after the disk; (6) Outside the passage (927), the remainder is closed and airtight. 9. The hot-end rotor system of a micro turbojet engine with a core cooling structure according to claim 3, wherein the rear bearing (3) and the main shaft (1) are on the first mating surface (3 Centering is realized at -1), and the centering of the intermediate stage sealing isolation disc (4) and the main shaft (1) is realized at the second mating surface (4-1), and the turbine rotor (6) and the middle stage The sealing isolation disc (4) realizes centering at the third mating surface (4-6), and realizes good centering of the turbine rotor (6) and the rear bearing (3) under working heat state; The outer mating cylindrical surface (606) of the front boss of the turbine rotor (6) is in a mating relationship with the rear mating inner cylindrical surface (406) on the intermediate stage sealing isolation plate (4). Under the working heat state, the third mating The temperature of the turbine rotor at the surface (4-6) is significantly higher than that of the intermediate sealing isolation plate. At this time, the outer matching cylindrical surface (606) of the front boss and the inner cylindrical surface (406) of the rear matching are in a tightening trend. The centering effect at the three mating surfaces (4-6) is stronger than that during assembly. 10. The hot-end rotor system of a kind of band core cooling structure according to claim 1, characterized in that, the rear disk surface of the turbine rotor (6) is cooled by the middle channel air flow (930), and the cooling air flow Through the radial hole in front of the bearing on the main shaft (1) and the middle channel formed by the axial hole towards the tail vertebra (8), it enters the tail vertebra cavity (H3) behind the disc to cool the turbine rotor (6); Or cool down the rear disk surface of the turbine rotor (6) through the external channel airflow (940), use pipelines to draw cooling air from the engine or other high-pressure air generation facilities, and enter through the external channel airflow (940) offered on the tail cone (8) The caudal cavity (H3) behind the disc is used to cool the turbine rotor (6).