Duct shaft turbojet engine

Abstract

The invention discloses a culvert shaft turbojet engine which comprises a compressed air rotor, a compressed air stator, a combustion chamber, a culvert shaft, a turbine rotor, a turbine stator, an outer duct rotating structure and an inner duct fixing structure. The air compressor and the turbine rotor are coaxially arranged in the outer duct, the blades are inwards installed on the duct blade disc, heat of the outer duct is taken away by circulating air, the outward guide blades are installed on the blade disc of the turbine stator, air also circulates in the inner duct of the blade disc, and heat of the inner duct of the stator is taken away. The heat dissipation effect of the inner and outer ducts is achieved.

Description

TECHNICAL FIELD

[0001] The present application relates to the field of aero-engines, in particular to a double-ducted heat-dissipating turbofan engine. BACKGROUND

[0002] At present, the single-stage turbojet engine used by the unmanned aerial vehicle does not set a heat dissipation system. The power of the single-stage turbine and the centrifugal single-stage compressor is very small. The main reason is that the heat dissipation efficiency is low. The more the number of turbine stages is, the higher the pressure and temperature are, and the power is also improved. At present, the heat dissipation system of the high-power multi-stage turbine engine is to punch holes in the turbine blades, and the compressed air flows through the holes to take away the heat, thereby achieving the cooling effect and increasing the service life of the turbine blades.

[0003] If the single-stage turbojet engine adopts the heat dissipation mode of a large engine, the structure and process of the turbine blades will lose the overall forming technology of the turbine blade and the blade disc. Therefore, the blade and the blade disc are in a mortise and tenon structure, and a single blade is punched, because the laser cannot punch a bend. Although the overall forming technology of the turbine blade has a very low power, the strength and reliability are better than those of the bulk turbine structure. The mortise and tenon structure has advantages and disadvantages. The overall structural strength is inevitably weakened when the blade is installed in the mortise and tenon structure, so the structural strength needs to be increased. The weight is also inevitably increased. Therefore, the single-stage turbojet engine does not set a heat dissipation function. SUMMARY

[0004] The present application provides a turbofan engine, and provides the following technical solutions. The present application comprises a compressor rotor blade, a compressor rotor disc, a compressor stator guide vane, a compressor stator disc, a combustion chamber, a combustion cylinder, a rotating column, a shaft, a turbine stator guide vane, a turbine stator disc, a turbine rotor blade, and a turbine rotor disc. The compressor rotor blade is arranged and installed in the inner circle leaf groove of the compressor rotor disc in sequence. The inner diameter size can be set to multiple rotor blades. The blade root is provided with a fixed screw corresponding to the fixed hole of the rotor disc leaf groove, and the screw is fastened. The number of compressor rotor stages can be set to multiple stages. The flange hole, the screw, and the nut are used for the butt joint between the stages.

[0005] The compressor stator guide vane is arranged and installed in the outer circle leaf groove of the compressor stator disc in sequence. The outer diameter size can be set to multiple compressor guide vanes. The blade root is provided with a fixed screw corresponding to the fixed hole of the compressor stator disc leaf groove, and the screw is fastened. The number of compressor stator stages is the same as that of the rotor stages. The flange hole, the screw, and the nut are used for the butt joint between the stages.

[0006] The combustion chamber is an inner and outer ring, the support plate is fixedly connected with the inner and outer ring, the combustion cylinder is sleeved in the cavity of the inner and outer ring, the annular combustion chamber is provided with a plurality of cavities and the combustion cylinder, the flange holes of the inner ring correspond to the flange holes of the turbine stator and the compressor stator, the screw is penetrated, and the nut is fastened, the bearing groove of the outer ring is matched with the bearing groove of the shaft, the rotating column is arranged in the outer ring of the combustion chamber and the shaft to be abutted and rolled, a plurality of rotating columns are arranged in annular arrangement, and the retaining ring is arranged to separate the rotating columns.

[0007] The turbine rotor blades are sequentially and circularly arranged in the inner circle blade groove of the turbine rotor blade disc, a plurality of turbine rotor blades can be arranged according to the size of the inner diameter, the root of the turbine blade is provided with a fixed screw hole corresponding to the fixed hole of the turbine rotor blade disc, the screw is fastened, and a plurality of stages of turbine rotors can be arranged.

[0008] The turbine stator guide vane is sequentially and circularly arranged in the outer circle blade groove of the turbine stator blade disc, a plurality of guide vanes can be arranged according to the size of the outer diameter, the root of the guide vane is provided with a fixed screw hole corresponding to the fixed hole of the turbine stator blade disc, the screw is fastened, and the number of stages of the turbine stator is the same as that of the turbine rotor. The flange holes of the turbine stator and the turbine rotor are abutted, the screw is penetrated, and the nut is fastened.

[0009] Further, the flange holes of the combustion chamber are respectively abutted to the flange holes of the compressor stator blade disc and the turbine stator blade disc, the screw is penetrated, and the nut is fastened.

[0010] Further, the flange holes of the shaft are respectively abutted to the flange holes of the compressor rotor and the turbine rotor blade disc, the screw is penetrated, and the nut is fastened.

[0011] Further, the flange holes of the turbine rotor blade disc correspond to the flange holes of the heat dissipation blade disc, the screw is penetrated, and the nut is fastened.

[0012] Further, the flange holes of the turbine stator blade disc correspond to the flange holes of the diffuser cover, the screw is penetrated, and the nut is fastened.

[0013] Further, the flange holes of the turbine stator blade disc correspond to the flange holes of the diffuser cover, the screw is penetrated, and the nut is fastened.

[0014] Further, the flange holes of the turbine stator blade disc correspond to the flange holes of the diffuser cover, the screw is penetrated, and the nut is fastened.

[0015] In terms of heat dissipation and materials technology, both are decisive factors in turbine blade lifespan. Although we currently lack major breakthroughs in materials technology, we employ a dual-duct cooling system in the heat dissipation field. Compared to the single-stage turbojet engines used in existing UAVs, this solution offers high heat dissipation efficiency, significantly extending the turbine blade's lifespan. Attached Figure Description

[0016] To more clearly illustrate the embodiments of the present invention or the existing technical solutions, the accompanying drawings used in the description are briefly introduced.

[0017] Figure 1 This is a structural diagram of the present invention.

[0018] Figure 2 This is a three-view structural diagram of the compressor rotor installation of the present invention.

[0019] Figure 3 This is a three-view structural diagram of the compressor stator installation of the present invention.

[0020] Figure 4 This is a structural diagram of the combustion chamber and rotating column installation of the present invention.

[0021] Figure 5 This is a three-view structural diagram of the combustion chamber and combustion cylinder installation of the present invention.

[0022] Figure 6 This is a three-view structural diagram of the combustion chamber and culvert shaft installation of the present invention.

[0023] Figure 7 This is a three-view structural diagram of the turbine stator installation of the present invention.

[0024] Figure 8 This is a three-view structural diagram of the turbine rotor installation of the present invention.

[0025] Figure 9 This is a structural diagram of the heat dissipation impeller of the present invention.

[0026] Figure 10 This is a structural diagram of the diffuser cover of the present invention.

[0027] Figure 11 This is a structural diagram of the hanging bracket cover of the present invention.

[0028] Figure 12 This is a structural diagram of the starting gear ring of the present invention.

[0029] The figure includes: compressor rotor blades (1), compressor rotor disk (2), fastening screws (201), flange holes (202), steps (203), blade slots (204), compressor stator guide vanes (3), fastening screws (301), flange holes (302), steps (303), compressor stator disk (4), combustion chamber (5), flange holes (501), cage (502), steps (503), oil seal ring (504), combustion chamber (6), inner ring (601), outer ring (602), support plate (603), heat dissipation holes (604), oil nozzles (605), rotor column (7), culvert (8), flange holes (8). 01) Turbine stator guide vane (9) Fastening screw (901) Flange hole (902) Step (903) Turbine stator blade (10) Turbine rotor blade (11) Fastening screw (1101) Flange hole (1102) Step (1103) Turbine rotor blade (12) Cooling blade (13) Flange hole (1301) Cooling blade (14) Diffuser (15) Exhaust hole (1501) Flange hole (1502) Starter motor (16) Hanger cover (17) Flange hole (1701) Inlet hole (1702) Flange hole (1703) Starter gear ring (18) Flange hole (1801). Detailed Implementation

[0030] The technical solutions of the present invention will be clearly and completely described below with reference to the accompanying drawings of the embodiments of the present invention. Obviously, the described embodiments are only some embodiments of the present invention, and not all embodiments.

[0031] Please see Figure 1 As shown in the embodiment of the present invention, the culverted turbojet engine includes compressor rotor blades (1), compressor rotor disk (2), compressor stator guide vanes (3), compressor stator disk (4), combustion chamber (5), combustion cylinder (6), rotor (7), culvert (8), turbine stator guide vanes (9), turbine stator disk (10), turbine rotor blades (11), and turbine rotor disk (12). Figure 2 As shown, the compressor rotor blades (1) are arranged and installed in sequence on the inner ring blade groove (204) of the compressor rotor disk (2). The inner diameter can be set with multiple rotor blades (1). The blade root is provided with screws (201) for fastening, corresponding to the fixing holes of the blade groove of the rotor disk (2), and the screws are fastened. The compressor rotor stage can be set with multiple stages. The flange holes (202) between the two stages are connected, the screws pass through, and the nuts are fastened. The compressor rotor disk (2) is provided with an annular step (203) and the rotor disk (2) is integrally formed and fastened. The annular step between the two stages of rotor disk (2) is merged to form the circular rail groove of the compressor stator guide vane.

[0032] Figure 3 As shown, the compressor stator guide vanes (3) are arranged and installed sequentially on the outer ring blade groove (304) of the compressor stator blade disk (4). Multiple compressor guide vanes (3) can be set in the outer diameter. The blade root is provided with screws (301) for fastening, corresponding to the fixing holes of the blade groove (304) of the compressor stator blade disk (4), and the screws are fastened. The number of stages of the compressor stator is the same as the number of stages of the rotor. The flange holes (302) between the two stages are connected, the screws pass through, and the nuts are fastened. The compressor stator blade disk (4) is provided with an annular step (303) and is integrally formed with the stator blade disk (4) for fastening. The annular steps between the two stages of the stator blade disk (4) are combined to form the circular rail groove of the compressor rotor blade.

[0033] Figure 4 , Figure 5 , Figure 6 As shown, the combustion chamber (5) is a cavity formed by an inner ring (601) and an outer ring (602) with multiple support plates (603) separating and fixing the inner ring (601) and the outer ring (602). The combustion cylinder (6) is installed inside the cavity, and the number of annular combustion cylinders (6) is the same as the number of cavities. The flange holes (501) of the inner ring correspond to the turbine stator flange holes (902) and the compressor stator flange holes (302) respectively. The screws pass through and the nuts are tightened. The bearing groove step (503) of the outer ring is fitted with the bearing groove (802) of the duct shaft (8). The rotating column (7) is placed on the outer ring of the combustion chamber (5) and rolls with the duct shaft (8). Multiple rotating columns (7) are arranged in a ring, and the rotating columns (7) are separated by a retainer (502) to provide rolling support for rotation. The installation is as follows. Figure 6 As shown, the gap between the combustion chamber (5) and the two ends of the culvert (8) is sealed with an oil seal ring (504) to prevent the lubricating oil of the rotating column (7) from leaking out.

[0034] Figure 8 As shown, the turbine rotor blades (11) are arranged in a circular pattern and installed in the inner ring blade groove (1104) of the turbine rotor disk (12). The inner diameter can be set with multiple turbine rotor blades (11). The root of the turbine blade is provided with screw holes (1101) for fastening. The screws are fastened to the corresponding fixing holes of the blade groove of the turbine rotor disk (12). The turbine rotor can be set with multiple stages. The connection between them is the flange hole (1102). The screw passes through and the nut is fastened. The turbine rotor disk (12) is provided with an annular step (1103) and the rotor disk (12) is integrally formed and fastened. The annular step between the two stages of rotor disk (12) is merged to form the circular rail groove of the turbine rotor blade.

[0035] Figure 7As shown, the turbine stator guide vanes (9) are arranged in a circular pattern and installed on the outer ring of the turbine stator disk (10) in the blade groove (904). The outer diameter can be set with multiple guide vanes (9). The root of the guide vane (9) is provided with screws (901) for fastening. The screws are fastened to the corresponding fixing holes of the turbine stator disk (10). The number of stages of the turbine stator is the same as the number of stages of the turbine rotor. The connection between them is through flange holes (902), through which screws are passed and nuts are fastened. The turbine stator disk (10) is provided with an annular step (903) and is integrally formed with the stator disk (10) for fastening. The annular step between the two stages of the stator disk (10) is merged to form the circular groove of the turbine rotor blade.

[0036] Preferred, Figure 6 As shown, the two ends of the flange hole (501) of the combustion chamber (5) are respectively connected to the flange hole (302) of the compressor stator blade disk and the flange hole (902) of the turbine stator blade disk, with screws passing through and nuts tightening the connection.

[0037] Preferably, the flange hole (801) of the culvert (8) is connected to the flange hole (202) of the compressor rotor and the flange hole (1102) of the turbine rotor disk at both ends, respectively, with screws passing through and nuts fastening the connection.

[0038] Preferred, Figure 9 As shown, the flange hole (1102) of the turbine rotor blade disk (12) corresponds to the flange hole (1301) of the heat dissipation blade disk (13). The screw passes through and the nut is fastened. The heat dissipation blade (14) is integrally formed on the heat dissipation blade disk (13) and fastened. Multiple heat dissipation blades (14) are arranged sequentially on the outer diameter of the heat dissipation blade disk (13).

[0039] Preferred, Figure 10 As shown, the flange hole (902) of the turbine stator blade disk (10) corresponds to the flange hole (1502) of the diffuser shroud (15), the screw passes through, and the nut is fastened to connect. The tip of the diffuser shroud (15) is provided with an air outlet (1501).

[0040] Preferred, Figure 11 As shown, the flange hole (1703) of the mounting cover (17) corresponds to the flange hole (9302) of the compressor rotor disk (2), with screws penetrating and nuts tightening the connection. The mounting cover (17) has an internal cavity through which fuel pipes, lubrication pipes, and ignition coils pass. The mounting cover (17) has multiple air inlets (1702), and the mounting hole (1701) of the mounting cover (17) is located on the aircraft.

[0041] Preferably, the flange hole (202) of the compressor rotor blade disk (2) is merged and corresponds to the flange hole (1801) of the starting gear ring (18), with screws passing through and nuts fastening the connection. The starting gear ring (18) is meshed with the starting motor gear shaft (16), and the mounting hole of the starting motor corresponds to the wall hole of the bracket cover (17), with screws passing through and nuts fastening the connection.

[0042] The working process and principle of this invention are as follows: The starter motor drives the starter gear ring to rotate at high speed. After reaching the air compression speed, fuel is injected and ignited, completing the starting process. The high-speed airflow ejected from the turbine nozzle is pressurized by the diffuser, resulting in a decrease in velocity and an increase in pressure, thus generating a reaction force to drive the engine. The outer bypass duct of the rotor has a cooling impeller that rotates synchronously with the turbine rotor. Cool air is drawn in from the inlet end by the cooling impeller and flows over the surface of the turbine rotor, thereby carrying away a large amount of heat. High-pressure gas is ejected from the middle bypass duct, creating a negative pressure in the middle hole of the diffuser. This causes cool air to flow through the air holes of the mounting cover and through the inner duct of the engine, thereby carrying away the heat generated inside. This achieves a dual-bypass cooling effect.

[0043] Although the present invention has been described in detail with reference to the foregoing embodiments, those skilled in the art can still modify the technical solutions described in the foregoing embodiments or make equivalent substitutions for some of the technical features. Any modifications, equivalent substitutions, improvements, etc., made within the spirit and principles of the present invention should be included within the protection scope of the present invention.

Claims

1. A culverted shaft turbojet engine, characterized in that... Including compressor rotor blades (1), compressor rotor blade disk (2), compressor stator guide vanes (3), compressor stator blade disk (4), combustion chamber (5), combustion cylinder (6), rotor column (7), culvert (8), turbine stator guide vanes (9), turbine stator blade disk (10), turbine rotor blades (11), and turbine rotor blade disk (12); The compressor rotor blades (1) are arranged in sequence in the inner ring blade groove of the compressor rotor disk (2). The inner diameter can be set with multiple rotor blades (1). The blade root is provided with fixing screws, which correspond to the fixing holes of the blade groove of the rotor disk (2) and are tightened with screws. The number of compressor rotor stages can be set to multiple stages. The connection between them is a flange hole, with screws passing through and nuts tightening. The compressor stator guide vanes (3) are arranged in sequence on the outer ring blade groove of the compressor stator blade disk (4). Multiple compressor guide vanes (3) can be set in the outer diameter. The root of the blade is provided with fixing screws, which correspond to the fixing holes of the blade groove of the compressor stator blade disk (4) and are tightened with screws. The number of stages of the compressor stator is the same as the number of stages of the rotor. The connection between them is a flange hole, through which screws are passed and nuts are tightened. Combustion chamber (5) consists of inner and outer rings. Support plates are fixedly connected to the inner and outer rings. The combustion cylinder (6) is installed inside the cavity of the inner and outer rings. The annular combustion chamber is provided with multiple cavities and combustion cylinders (6). The flange holes of the inner ring correspond to the flange holes of the turbine stator and the compressor stator, respectively. Screws pass through and nuts are tightened. The bearing groove of the outer ring is fitted with the bearing groove of the duct shaft (8). The rotating column (7) is placed in the combustion chamber (5). The outer ring and the duct shaft (8) are rolled together. Multiple rotating columns (7) are arranged in annularly, and the rotating columns (7) are separated by a retainer. The turbine rotor blades (11) are arranged in a circular pattern and installed in the inner ring blade groove of the turbine rotor disk (12). The inner diameter can be set with multiple turbine rotor blades (11). The root of the turbine blade is provided with a fixing screw hole, which corresponds to the fixing hole of the blade groove of the turbine rotor disk (12). The screw is fastened to the connection. The turbine rotor can be set with multiple stages. The connection between them is a flange hole, through which the screw passes and the nut is tightened. The turbine stator guide vanes (9) are arranged in a circular pattern and installed on the outer ring of the turbine stator disk (10). Multiple guide vanes (9) can be set according to the outer diameter. The root of the guide vanes (9) is provided with fixing screw holes, which correspond to the fixing holes of the turbine stator disk (10). The screws are used to fasten the connection. The number of stages of the turbine stator is the same as the number of stages of the turbine rotor. The connection between them is through flange holes, with screws passing through and nuts tightening.

2. The culverted shaft turbojet engine according to claim 1, characterized in that: The flange holes of the combustion chamber (5) are respectively connected to the flange holes of the compressor stator blade disk and the turbine stator blade disk, with screws passing through and nuts tightening.

3. The culverted shaft turbojet engine according to claim 1, characterized in that: The flange holes of the culvert (8) are respectively connected to the flange holes of the compressor rotor and the turbine rotor blade disk, with screws passing through and nuts tightening.

4. The culverted shaft turbojet engine according to claim 1, characterized in that: The flange hole of the turbine rotor blade disk (12) corresponds to the flange hole of the heat dissipation blade disk (13), the screw passes through, and the nut is fastened to connect. The heat dissipation blade (14) is integrally formed on the heat dissipation blade disk (13) and fastened to connect. Multiple heat dissipation blades (14) are arranged sequentially on the outer diameter of the heat dissipation blade disk (13).

5. The bypass shaft turbojet engine according to claim 1, characterized in that: The turbine stator blade disk (10) corresponds to the flange hole of the diffuser shroud (15), the screw passes through, and the nut is fastened to connect. The pointed part of the diffuser shroud (15) is provided with an air outlet.

6. The bypass shaft turbojet engine according to claim 1, characterized in that: The mounting bracket (17) corresponds to the flange hole of the compressor rotor disk (2), with screws penetrating and nuts tightening the connection. The mounting bracket (17) has an internal cavity through which fuel pipes, lubrication pipes, and ignition coils pass. The mounting bracket (17) has multiple air inlets.

7. The culverted shaft turbojet engine according to claim 1, characterized in that: The flange hole of the compressor rotor disk (2) corresponds to the starting gear ring (18), with screws passing through and nuts tightening the connection. The starting gear ring (18) meshes with the starting motor gear shaft (16). The mounting hole of the starting motor corresponds to the wall hole of the bracket cover (17), with screws passing through and nuts tightening the connection.

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