CN115539212A - Speed reduction device and aircraft engine - Google Patents

Speed reduction device and aircraft engine Download PDF

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Publication number
CN115539212A
CN115539212A CN202211290000.0A CN202211290000A CN115539212A CN 115539212 A CN115539212 A CN 115539212A CN 202211290000 A CN202211290000 A CN 202211290000A CN 115539212 A CN115539212 A CN 115539212A
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CN
China
Prior art keywords
gear
engine
power turbine
turbine rotor
rotor
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Pending
Application number
CN202211290000.0A
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Chinese (zh)
Inventor
雷寰兴
王涛
唐志清
游志伟
董红涛
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Hunan Aviation Powerplant Research Institute AECC
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Hunan Aviation Powerplant Research Institute AECC
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Application filed by Hunan Aviation Powerplant Research Institute AECC filed Critical Hunan Aviation Powerplant Research Institute AECC
Priority to CN202211290000.0A priority Critical patent/CN115539212A/en
Publication of CN115539212A publication Critical patent/CN115539212A/en
Pending legal-status Critical Current

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    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F16—ENGINEERING ELEMENTS AND UNITS; GENERAL MEASURES FOR PRODUCING AND MAINTAINING EFFECTIVE FUNCTIONING OF MACHINES OR INSTALLATIONS; THERMAL INSULATION IN GENERAL
    • F16H—GEARING
    • F16H57/00—General details of gearing
    • F16H57/02—Gearboxes; Mounting gearing therein
    • F16H57/021—Shaft support structures, e.g. partition walls, bearing eyes, casing walls or covers with bearings
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F16—ENGINEERING ELEMENTS AND UNITS; GENERAL MEASURES FOR PRODUCING AND MAINTAINING EFFECTIVE FUNCTIONING OF MACHINES OR INSTALLATIONS; THERMAL INSULATION IN GENERAL
    • F16H—GEARING
    • F16H57/00—General details of gearing
    • F16H57/02—Gearboxes; Mounting gearing therein
    • F16H57/023—Mounting or installation of gears or shafts in the gearboxes, e.g. methods or means for assembly
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F16—ENGINEERING ELEMENTS AND UNITS; GENERAL MEASURES FOR PRODUCING AND MAINTAINING EFFECTIVE FUNCTIONING OF MACHINES OR INSTALLATIONS; THERMAL INSULATION IN GENERAL
    • F16H—GEARING
    • F16H57/00—General details of gearing
    • F16H57/04—Features relating to lubrication or cooling or heating

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  • Engineering & Computer Science (AREA)
  • General Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • General Details Of Gearings (AREA)

Abstract

The invention discloses a speed reducing device and an aircraft engine, wherein the speed reducing device comprises a shell assembly, an accessory transmission speed reducing assembly and a power turbine speed reducing assembly; the shell assembly is provided with a first rotor hole and a second rotor hole which are concentrically arranged; the input end of the accessory transmission speed reducing component is connected with a gas turbine rotor of the engine, and the output end of the accessory transmission speed reducing component is suitable for being respectively connected with accessories of the engine; the input end of the power turbine speed reducing assembly is connected with the power turbine rotor of the engine, the output end of the power turbine speed reducing assembly is an output gear shaft, and the axis of the output gear shaft and the power turbine rotor of the engine are concentrically arranged and are suitable for being connected with a rotor wing or a propeller of the airplane. The invention improves the speed reducer of the aircraft engine, integrates the accessory transmission speed reducer and the power turbine speed reducer into a whole, and the axis of the output gear shaft of the power turbine speed reducer is concentric with the power turbine rotor and the gas turbine rotor, thereby reducing the volume of the speed reducer and the overall weight.

Description

Speed reduction device and aircraft engine
Technical Field
The invention relates to the technical field, in particular to a speed reducing device and an aircraft engine.
Background
The speed reducer of the aviation gas turbine engine generally comprises an accessory transmission speed reducer and a power turbine speed reducer, wherein the accessory transmission speed reducer is used for extracting power from a gas turbine rotor of the engine and transmitting the power to engine accessories (comprising a lubricating oil pump, a fuel pump, a motor and the like) according to a certain rotating speed and a certain steering direction, and the power turbine speed reducer is used for transmitting the power of a power turbine of the engine to an airplane according to a certain rotating speed and a certain steering direction so as to drive a rotor wing or a propeller of the airplane to work.
At present, for low-power aviation gas turbine shaft engines, there are two common structural forms of accessory drive reducers and power turbine reducers: the first is that the accessory transmission reducer and the power turbine reducer are designed separately and are respectively an independent gear box; and secondly, the accessory transmission reducer and the power turbine reducer share a gearbox casing, and are integrated into an integrated design (an integrated gearbox) which is an integral gearbox in appearance. However, these two forms of accessory drive reducer drive train are independent of the power turbine reducer drive train, which draws power from the gas turbine rotor, and the power turbine reducer drive train draws power from the power turbine. The output gear shaft of the power turbine reducer is typically eccentrically disposed, i.e., the power turbine rotor is not concentric with the output gear shaft of the power turbine reducer. Therefore, the reduction gear of the prior aviation gas turbine engine has larger integral volume, is not beneficial to the design of an aircraft nacelle with limited size and is also not beneficial to the weight reduction design of a gear box.
Disclosure of Invention
The invention mainly aims to provide a speed reducer and an aircraft engine, aiming at reducing the volume of the speed reducer, so as to be beneficial to the design of an aircraft nacelle with limited size and the weight reduction design of a gear box.
To achieve the above object, the present invention provides a reduction gear, comprising:
a casing assembly formed with first and second concentrically disposed rotor bores, the first rotor bore adapted for insertion of an engine gas turbine rotor and the second rotor bore adapted for insertion of an engine power turbine rotor;
the accessory transmission speed reducing component is arranged in the shell component, the input end of the accessory transmission speed reducing component is connected with the gas turbine rotor of the engine, and the output end of the accessory transmission speed reducing component is suitable for being respectively connected with accessories of the engine; and
the power turbine speed reduction assembly is arranged in the shell assembly, the input end of the power turbine speed reduction assembly is connected with the power turbine rotor of the engine, the output end of the power turbine speed reduction assembly is an output gear shaft, and the axis of the output gear shaft is concentrically arranged with the power turbine rotor of the engine and is suitable for being connected with a rotor wing or a propeller of the aircraft.
Optionally, the accessory transmission reduction assembly includes a first gear, a second gear, a third gear, a fourth gear, a fifth gear, a sixth gear, a seventh gear, and an eighth gear, where the first gear is a cylindrical gear, the first gear is sleeved on the engine gas turbine rotor and engaged with the second gear, the second gear is connected with the third gear through a spline, the third gear is engaged with the fifth gear and the sixth gear, the fifth gear is an idler gear, the sixth gear is adapted to drive a lubricating oil pump, the fourth gear is engaged with the fifth gear, the fourth gear is adapted to drive a fuel pump, the sixth gear is engaged with the seventh gear, the seventh gear is an idler gear, the seventh gear is engaged with the eighth gear, and the eighth gear is adapted to drive a motor.
Optionally, the engine gas turbine rotor to second gear ratio is 2.12;
the transmission ratio of the engine gas turbine rotor to the fourth gear is 4.75;
the engine gas turbine rotor to fifth gear ratio is 4.34;
the engine gas turbine rotor to sixth gear ratio is 4.34;
the transmission ratio of the engine gas turbine rotor to the seventh gear is 2.12;
the engine gas turbine rotor to the eighth gear ratio is 4.34.
Optionally, the power turbine speed reduction assembly includes a ninth gear, a tenth gear, an eleventh gear and a twelfth gear, the ninth gear is connected to the engine power turbine rotor through a spline, the ninth gear is meshed with the tenth gear, the tenth gear rotates coaxially with the eleventh gear, the eleventh gear is meshed with the twelfth gear, and a gear shaft of the twelfth gear constitutes the output gear shaft.
Optionally, the engine power turbine rotor to tenth gear ratio is 3.30;
the transmission ratio of the engine power turbine rotor to the twelfth gear is 6.49.
Optionally, the center distance between the ninth gear and the tenth gear is the same as the center distance between the eleventh gear and the twelfth gear.
Optionally, the number of teeth of the ninth gear is 65, and the modulus is 1.75mm;
the number of teeth of the tenth gear is 33, and the modulus is 1.75mm;
the number of teeth of the eleventh gear is 89, and the modulus is 1.50mm;
the number of teeth of the twelfth gear is 27, and the modulus is 1.50mm.
Optionally, the casing assembly includes a gearbox casing, a support frame and a bearing seat, and the support frame and the bearing seat are respectively fixed to the gearbox casing by screws; the support frame is used for supporting the accessory drive reduction assembly and the power turbine reduction assembly; the bearing seat is provided with a first bearing hole and a second bearing hole, the first bearing hole is used for mounting a bearing at one end of the twelfth gear, and the second bearing hole is used for mounting a bearing at one end of the tenth gear.
Optionally, the radial cross section of the support frame is arranged in a shape like a Chinese character 'gong', the support frame comprises a support body arranged in a ring shape and a plurality of wing plates respectively arranged outwards along the radial direction of the support body, and the plurality of wing plates are provided with reinforcing ribs.
To achieve the above object, the present invention also proposes an aircraft engine including a reduction gear as described above, the reduction gear including:
a casing assembly formed with first and second concentrically disposed rotor bores, the first rotor bore adapted for insertion of an engine gas turbine rotor and the second rotor bore adapted for insertion of an engine power turbine rotor;
the accessory transmission speed reducing component is arranged in the shell component, the input end of the accessory transmission speed reducing component is connected with the engine gas turbine rotor, and the output end of the accessory transmission speed reducing component is suitable for being respectively connected with accessories of an engine; and
the power turbine speed reduction assembly is arranged in the shell assembly, the input end of the power turbine speed reduction assembly is connected with the power turbine rotor of the engine, the output end of the power turbine speed reduction assembly is an output gear shaft, and the axis of the output gear shaft is concentrically arranged with the power turbine rotor of the engine and is suitable for being connected with a rotor wing or a propeller of the aircraft.
In the technical scheme of the invention, the speed reducing device comprises a machine shell assembly, an accessory transmission speed reducing assembly and a power turbine speed reducing assembly; the casing assembly is formed with a first rotor bore and a second rotor bore arranged concentrically, the first rotor bore being adapted for insertion of an engine gas turbine rotor and the second rotor bore being adapted for insertion of an engine power turbine rotor; the accessory transmission speed reducing component is arranged in the shell component, the input end of the accessory transmission speed reducing component is connected with a gas turbine rotor of the engine, and the output end of the accessory transmission speed reducing component is suitable for being respectively connected with accessories of the engine; the power turbine speed reducing assembly is arranged in the shell assembly, the input end of the power turbine speed reducing assembly is connected with the power turbine rotor of the engine, the output end of the power turbine speed reducing assembly is an output gear shaft, and the axis of the output gear shaft and the power turbine rotor of the engine are concentrically arranged and are suitable for being connected with a rotor wing or a propeller of the aircraft.
On one hand, the speed reducer can extract power from a gas turbine rotor of an aircraft engine and drive engine accessories such as a lubricating oil pump, a fuel pump and a motor; on the other hand, the speed reducing device simultaneously realizes the power output of the power turbine rotor of the engine, namely, the driving of a rotor or a propeller of the airplane.
Meanwhile, the axis of the output gear shaft of the power turbine speed reduction assembly is concentric with the power turbine rotor of the engine, and the power turbine rotor of the engine is concentric with the gas turbine rotor of the engine, namely, the axis of the output gear shaft is concentric with the first rotor hole and the second rotor hole, so that the space occupied by the power output shaft is reduced, the area of a gear box is reduced, and the accessory transmission speed reduction assembly and the power turbine speed reduction assembly are integrated into a whole, so that the whole volume of the speed reduction device is reduced, the design of an aircraft nacelle when the size is limited is facilitated, and the weight reduction design of the gear box is facilitated. And due to the symmetry of the layout of the reduction device, when the structure is applied to a turboprop engine, the structure does not generate adverse disturbance to airflow around the propeller.
Drawings
In order to more clearly illustrate the embodiments or technical solutions of the present invention, the drawings used in the embodiments or technical solutions of the prior art will be briefly described below, it is obvious that the drawings in the following description are only some embodiments of the present invention, and for those skilled in the art, other drawings can be obtained according to the structures shown in the drawings without creative efforts.
FIG. 1 is a schematic view of a transmission chain structure of a reduction gear according to an embodiment of the present invention;
FIG. 2 isbase:Sub>A cross-sectional view taken at A-A of FIG. 1;
FIG. 3 is a schematic front view of a support frame according to an embodiment of the reduction gear transmission of the present invention;
FIG. 4 is a schematic view of a rear side of a support frame in an embodiment of the reduction gear transmission of the present invention;
fig. 5 is a sectional view of a support frame in an embodiment of the reduction gear transmission of the present invention.
The reference numbers indicate:
10. a housing assembly; 20. an accessory drive reduction assembly; 30. a power turbine speed reduction assembly; 101. an engine gas turbine rotor; 102. an engine power turbine rotor; 21. a first gear; 22. a second gear; 23. a third gear; 24. a fourth gear; 25. a fifth gear; 26. a sixth gear; 27. a seventh gear; 28. an eighth gear; 31. a ninth gear; 32. a tenth gear; 33. an eleventh gear; 34. a twelfth gear; 11. a gearbox casing; 12. a support frame; 13. a bearing seat; 13a, a first bearing hole; 13b, a second bearing hole; 131. a support body; 132. a wing plate; 133. and (5) reinforcing ribs.
The implementation, functional features and advantages of the objects of the present invention will be further explained with reference to the accompanying drawings.
Detailed Description
The technical solutions in the embodiments of the present invention will be clearly and completely described below with reference to the drawings in the embodiments of the present invention, and it is obvious that the described embodiments are only a part of the embodiments of the present invention, and not all of the embodiments. All other embodiments, which can be derived by a person skilled in the art from the embodiments given herein without making any creative effort, shall fall within the protection scope of the present invention.
It should be noted that, if directional indications (such as up, down, left, right, front, and back … …) are involved in the embodiment of the present invention, the directional indications are only used to explain the relative position relationship between the components, the motion situation, and the like in a specific posture (as shown in the drawing), and if the specific posture is changed, the directional indications are changed accordingly.
In addition, if there is a description of "first", "second", etc. in an embodiment of the present invention, the description of "first", "second", etc. is for descriptive purposes only and is not to be construed as indicating or implying relative importance or implicitly indicating the number of technical features indicated. Thus, a feature defined as "first" or "second" may explicitly or implicitly include at least one of the feature. In addition, if appearing throughout the text, "and/or" is meant to include three juxtaposed aspects, taking "A and/or B" as an example, including either the A aspect, or the B aspect, or both A and B satisfied aspects. In addition, technical solutions between various embodiments may be combined with each other, but must be realized by a person skilled in the art, and when the technical solutions are contradictory or cannot be realized, such a combination should not be considered to exist, and is not within the protection scope of the present invention.
In order to maintain the fuel, lubricant, and electricity required to operate an aircraft gas turbine engine, the engine is typically equipped with accessories such as a fuel pump, a lubricant pump, and a generator, which are typically mounted on an accessory drive reducer, and the power required to operate the accessories is extracted from the gas turbine rotor.
Because of the high output speed of the engine, a reduction device, i.e., a power turbine reducer, is typically required between the engine and the propeller (or rotor of the helicopter) to transfer the power of the power turbine rotor of the engine at a low speed.
At present, for low-power aviation gas turbine shaft engines, there are two common structural forms of accessory drive reducers and power turbine reducers: the first is that the accessory transmission reducer and the power turbine reducer are designed separately and are respectively an independent gear box; and secondly, the accessory transmission reducer and the power turbine reducer share a gearbox casing, and are integrated into an integrated design (an integrated gearbox) which is an integral gearbox in appearance. However, these two forms of accessory drive reducer drive train are independent of the power turbine reducer drive train, which draws power from the gas turbine rotor, and the power turbine reducer drive train draws power from the power turbine. The output gear shaft of the power turbine reducer is typically eccentrically disposed, i.e., the power turbine rotor is not concentric with the output gear shaft of the power turbine reducer.
For the condition that a power output gear shaft of the power turbine speed reducer is eccentrically arranged, namely a power turbine rotor is not concentric with the output gear shaft of the power turbine speed reducer, the volume of the integrated gear box is larger; due to the asymmetry of the eccentric arrangement, the structure, when applied to a turboprop engine, can adversely affect the airflow around the propeller.
Therefore, the invention provides a speed reducing device which is suitable for an aircraft engine, in particular to a 200 kW-grade aircraft gas turbine shaft engine, and the speed reducing device is not limited in the process.
Referring to fig. 1 and 2, in one embodiment of the present invention, the reduction apparatus includes a housing assembly 10, an accessory drive reduction assembly 20, and a power turbine reduction assembly 30; casing assembly 10 is formed with first and second concentrically disposed rotor bores, the first rotor bore being adapted for insertion of an engine gas turbine rotor 101 and the second rotor bore being adapted for insertion of an engine power turbine rotor 102; the accessory transmission speed reducing component 20 is arranged in the machine shell component 10, the input end of the accessory transmission speed reducing component 20 is connected with the engine gas turbine rotor 101, and the output end of the accessory transmission speed reducing component 20 is suitable for being respectively connected with accessories such as a lubricating oil pump, a fuel pump and a motor of an engine; the power turbine speed reduction assembly 30 is disposed within the enclosure assembly 10, an input end of the power turbine speed reduction assembly 30 is coupled to the engine power turbine rotor 102, an output end of the power turbine speed reduction assembly 30 is an output gear shaft, and an axis of the output gear shaft is concentrically disposed with the engine power turbine rotor 102 and is adapted to be coupled to an aircraft rotor or propeller.
In this embodiment, the casing assembly 10 may include a gearbox casing 11, a support frame 12, and a bearing seat 13, wherein the support frame 12 and the gearbox casing 11 may be fixed by screws, and the bearing seat 13 and the gearbox casing 11 may also be fixed by screws. Wherein support frame 12 is configured to support accessory drive reduction assembly 20 and power turbine reduction assembly 30; the bearing housing 13 is provided with a first bearing hole 13a and a second bearing hole 13b, the first bearing hole 13a being for bearing mounting of one end of the twelfth gear 34, and the second bearing hole 13b being for bearing mounting of one end of the tenth gear 32.
In this embodiment, the accessory drive reduction assembly 20 and the power turbine reduction assembly 30 may each include a plurality of gears, and the drive train of the accessory drive reduction assembly 20 and the drive train of the power turbine reduction assembly 30 are independent and may both be fixed-axis gear trains, and there is no drive relationship between the two. The specific configurations of the accessory drive reduction assembly 20 and the power turbine reduction assembly 30 are not limited herein.
It will be appreciated that the reduction gear of the invention is an integrated gearbox, which on the one hand is capable of extracting power from the aircraft engine gas turbine rotor 101 to drive the operation of engine accessories such as the lube pump, fuel pump, electric motor, etc.; on the other hand, the speed reduction device simultaneously achieves the power output of the power turbine rotor 102 of the engine, namely, the driving of the rotor or propeller of the airplane.
Meanwhile, the axis of the output gear shaft of the power turbine speed reducing assembly 30 is concentrically arranged with the engine power turbine rotor 102, the engine power turbine rotor 102 is concentrically arranged with the engine gas turbine rotor 101, namely, the axis of the output gear shaft is concentrically arranged with the first rotor hole and the second rotor hole, so that the space occupied by a power output shaft is reduced, the area of a gear box is reduced, and the accessory transmission speed reducing assembly 20 and the power turbine speed reducing assembly 30 are integrated into a whole, so that the whole size of the speed reducing device is reduced, the design of an aircraft nacelle when the size is limited is facilitated, and the weight reduction design of the gear box is facilitated. And due to the symmetry of the layout of the reduction device, when the structure is applied to a turboprop engine, the structure does not generate adverse disturbance to airflow around the propeller.
In order to further reduce the volume of the reduction device and improve the performance of the reduction device, referring to fig. 1 and 2, in an embodiment, the accessory drive reduction assembly 20 may include a first gear 21, a second gear 22, a third gear 23, a fourth gear 24, a fifth gear 25, a sixth gear 26, a seventh gear 27, and an eighth gear 28, the first gear 21 is a cylindrical gear, the first gear 21 is sleeved on the engine gas turbine rotor 101 and is meshed with the second gear 22, the second gear 22 is connected with the third gear 23 through a spline, the third gear 23 is respectively meshed with the fifth gear 25 and the sixth gear 26, the fifth gear 25 is an idle gear, the sixth gear 26 is adapted to drive a lubricating oil pump, the fourth gear 24 is meshed with the fifth gear 25, the fourth gear 24 is adapted to drive a fuel pump, the sixth gear 26 is meshed with the seventh gear 27, the seventh gear 27 is an idle gear, the seventh gear 27 is meshed with the eighth gear 28, and the eighth gear 28 is adapted to drive an electric motor.
Wherein the transmission ratio of the engine gas turbine rotor 101 to the second gear 22 may preferably be 2.12; the gear ratio of the engine gas turbine rotor 101 to the fourth gear 24 may preferably be 4.75; the gear ratio of the engine gas turbine rotor 101 to the fifth gear 25 may preferably be 4.34; the gear ratio of the engine gas turbine rotor 101 to the sixth gear 26 may preferably be 4.34; the gear ratio of the engine gas turbine rotor 101 to the seventh gear 27 may preferably be 2.12; the gear ratio of the engine gas turbine rotor 101 to the eighth gear 28 may preferably be 4.34.
In this embodiment, the accessory drive reduction assembly 20 extracts power from the engine gas turbine rotor 101 through the spur gear and transfers the power to the other gears at the desired speed and direction by dividing the gear ratio of each gear pair of the drive train.
In order to further reduce the volume of the speed reducer and improve the performance of the speed reducer, referring to fig. 1 and 2, in an embodiment, the power turbine speed reducer assembly 30 may include a ninth gear 31, a tenth gear 32, an eleventh gear 33 and a twelfth gear 34, the ninth gear 31 is connected with the engine power turbine rotor 102 through a spline, the ninth gear 31 is meshed with the tenth gear 32, the tenth gear 32 rotates coaxially with the eleventh gear 33, the eleventh gear 33 is meshed with the twelfth gear 34, and a gear shaft of the twelfth gear 34 constitutes an output gear shaft.
Wherein the gear ratio of the engine power turbine rotor 102 to the tenth gear 32 is 3.30; the gear ratio of the engine power turbine rotor 102 to the twelfth gear 34 is 6.49.
In order to realize the concentricity of the axis of the output gear shaft with the engine gas turbine rotor 101 and the engine power turbine rotor 102, so as to further reduce the volume of the integrated speed reduction device, in the present embodiment, the center distance between the ninth gear 31 and the tenth gear 32 and the center distance between the eleventh gear 33 and the twelfth gear 34 may be set to be the same.
Further, to achieve better performance, ease of installation, and at the same time reduce the size of the reduction gear as much as possible, the gear module and the coefficient of variation shown in table 1 below may be selected:
TABLE 1 Power turbo reducer Gear parameters
Gear wheel Number of teeth Modulus (mm) Angle of pressure Helix angle Coefficient of variation
Ninth gear 31 65 1.75 22.5 12.57 -0.1226
Tenth gear 32 33 1.75 22.5 12.57 0.1226
Eleventh gear 33 89 1.5 22.5 8 -0.2126
Twelfth gear 34 27 1.5 22.5 8 0.2126
In this embodiment, the gears may be supported by bearings, the drive train of the accessory drive reduction assembly 20 shares a gearbox housing 11 and a support frame 12 with the drive train of the power turbine reduction assembly 30, and the support frame 12 may be secured to the gearbox housing 11 by a plurality of screws, so that the entire reduction assembly is a single, self-contained gearbox that may be maintained separately.
In order to ensure that the tenth gear 32 and the twelfth gear 34 are coaxially assembled, the bearing seat 13 is designed, and the bearing seat 13 can be connected with the gearbox casing 11 through screws. Referring to fig. 1 and 2, during assembly, the twelfth gear 34 and the bearing housing 13 may be assembled first, and then the tenth gear 32 and the support frame 12 may be assembled.
Referring to fig. 3 to 5, in this embodiment, the support frame 12 is a skeleton structure, and is essentially a degraded gearbox casing 11, which effectively reduces the weight of the entire gearbox. Support frame 12 may be comprised of a support body 131 and a plurality of radially extending wings 132, the position of each wing 132 being designed according to the fulcrum position. The support body 131 is annularly disposed, and the plurality of wings 132 may be respectively disposed outward in a radial direction of the support body 131. To improve the stiffness of the rear support frame 12, the radial cross-section of the entire support frame 12 may be "bowed" as shown in FIG. 3. The number of the wings 132 may be preferably 7, but is not limited thereto.
In order to reduce the weight of support frame 12, the present invention uses a thin-walled profiled support frame 12, which is machined using a forged aluminum alloy blank, and in order to ensure the rigidity and strength of support frame 12, a rib 133 may be provided on a flange 132 of support frame 12, as shown in fig. 4. In order to ensure that the positions of the bearing points on the support frame 12 correspond to the positions of the bearing points on the gearbox casing 11 one to one, the support frame 12 can be positioned by positioning pins when being mounted on the gearbox casing 11. It should be noted that the bearing bracket 12 and the gearbox casing 11 should be machined with strict control over the positional tolerance of the bearing mounting holes with respect to the positioning pins. The invention adopts the framework type supporting frame 12 with the reinforcing ribs 133, and has the advantages of light weight, simple structure and convenient assembly.
The drive sprocket of the accessory drive speed reduction assembly 20 of the present invention can be a spur gear, which does not have axial force (neglecting drive error) when working, and can be supported by a deep groove ball bearing, and the axial positioning can depend on a lock nut and a shoulder structure on a gear shaft.
The transmission sprocket of the power turbine speed reduction assembly 30 of the present invention may be a helical gear, which has high transmission power and relatively low rotation speed, and has axial force and large radial force during operation, and may be supported by both ball bearing and cylindrical roller bearing.
In addition, the support bearings of the input drive gear of the power turbine reduction assembly 30 may be cooled and lubricated using an under-ring oil supply, and all the meshing points of the gear pairs and the remaining support bearings may be cooled and lubricated using direct injection of lubricating oil. The support bearings of the accessory drive reduction assembly 20 may be cooled and lubricated by an oil mist, and the remaining gears may be cooled and lubricated by an oil mist, except for the gear pair having the highest rotational speed, which is cooled and lubricated by a spray of oil. The shaft extension end of the gear shaft of the accessory drive reduction assembly 20 can be sealed using graphite dynamic seals, and the static seals can be O-ring seals.
It should be noted that the speed reducer provided by the invention has been subjected to a performance debugging test and a 60-hour endurance test along with the whole aircraft gas turbine engine, a predetermined function is realized in the test, and no abnormality is found after the test. And the simulation work is carried out on the speed reducing device, and the simulation result shows that:
1. compared with the existing eccentric output structure (namely the output gear shaft of the power turbine speed reducer is not concentric with the power turbine rotor), the windward area of the gear box is reduced by 7 percent, and is reduced to 166452mm2 from 178981mm 2.
2. Compared with the traditional case, the invention adopts the framework type supporting frame 12 with the reinforcing ribs 133, so that the weight of the rear case is reduced by 60 percent and is reduced from 2750g to 1100g.
The invention also proposes an aeroengine comprising a reduction gear, the specific structure of which is as described above with reference to the above embodiments, and since the aeroengine proposed by the invention comprises all the solutions of all the embodiments of the reduction gear described above, it has at least the same technical effects as the reduction gear, and it is not set forth herein.
The above description is only an alternative embodiment of the present invention, and not intended to limit the scope of the present invention, and all modifications and equivalents of the present invention, which are made by the contents of the present specification and the accompanying drawings, or directly/indirectly applied to other related technical fields, are included in the scope of the present invention.

Claims (10)

1. A reduction gear, comprising:
a casing assembly formed with first and second concentrically disposed rotor bores, the first rotor bore adapted for insertion of an engine gas turbine rotor and the second rotor bore adapted for insertion of an engine power turbine rotor;
the accessory transmission speed reducing component is arranged in the shell component, the input end of the accessory transmission speed reducing component is connected with the engine gas turbine rotor, and the output end of the accessory transmission speed reducing component is suitable for being respectively connected with accessories of an engine; and
the power turbine speed reduction assembly is arranged in the shell assembly, the input end of the power turbine speed reduction assembly is connected with the power turbine rotor of the engine, the output end of the power turbine speed reduction assembly is an output gear shaft, and the axis of the output gear shaft is concentrically arranged with the power turbine rotor of the engine and is suitable for being connected with a rotor wing or a propeller of the aircraft.
2. The reduction device according to claim 1, wherein the accessory drive reduction assembly includes a first gear, a second gear, a third gear, a fourth gear, a fifth gear, a sixth gear, a seventh gear and an eighth gear, the first gear is a cylindrical gear, the first gear is sleeved on the engine gas turbine rotor and is meshed with the second gear, the second gear and the third gear are connected through splines, the third gear is respectively meshed with the fifth gear and the sixth gear, the fifth gear is an idle gear, the sixth gear is adapted to drive a lubricating oil pump, the fourth gear is meshed with the fifth gear, the fourth gear is adapted to drive a fuel pump, the sixth gear is meshed with the seventh gear, the seventh gear is an idle gear, the seventh gear is meshed with the eighth gear, and the eighth gear is adapted to drive an electric motor.
3. The reduction unit of claim 2, wherein the engine gas turbine rotor has a gear ratio to the second gear of 2.12;
the gear ratio of the engine gas turbine rotor to the fourth gear is 4.75;
the gear ratio of the engine gas turbine rotor to the fifth gear is 4.34;
the gear ratio of the engine gas turbine rotor to the sixth gear is 4.34;
the transmission ratio of the engine gas turbine rotor to the seventh gear is 2.12;
the engine gas turbine rotor to eighth gear ratio is 4.34.
4. The reduction unit of claim 1, wherein the power turbine reduction assembly includes a ninth gear splined to the engine power turbine rotor, a tenth gear in mesh with the tenth gear, the tenth gear rotating coaxially with the eleventh gear, the eleventh gear in mesh with the twelfth gear, and a gear shaft of the twelfth gear constituting the output gear shaft.
5. The reduction unit of claim 4, wherein the transmission ratio of the engine power turbine rotor to the tenth gear is 3.30;
the transmission ratio of the engine power turbine rotor to the twelfth gear is 6.49.
6. The reduction gear according to claim 4, wherein a center distance between the ninth gear and the tenth gear is the same as a center distance between the eleventh gear and the twelfth gear.
7. A reduction unit according to claim 6, characterised in that the number of teeth of the ninth gear is 65, the module being 1.75mm;
the number of teeth of the tenth gear is 33, and the modulus is 1.75mm;
the number of teeth of the eleventh gear is 89, and the modulus is 1.50mm;
the number of teeth of the twelfth gear is 27, and the modulus is 1.50mm.
8. The reduction gear according to claim 4, wherein the casing assembly includes a gearbox casing, a support bracket, and a bearing seat, the support bracket and the bearing seat are respectively fixed to the gearbox casing by screws; the support frame is used for supporting the accessory drive reduction assembly and the power turbine reduction assembly; the bearing seat is provided with a first bearing hole and a second bearing hole, the first bearing hole is used for mounting a bearing at one end of the twelfth gear, and the second bearing hole is used for mounting a bearing at one end of the tenth gear.
9. The reduction apparatus according to claim 8, wherein the support frame has a radial cross section in a "bow" shape, and the support frame includes a support body having an annular shape and a plurality of wing plates respectively provided radially outward of the support body, and the plurality of wing plates are provided with reinforcing ribs.
10. An aircraft engine comprising a reduction unit according to any one of claims 1 to 9.
CN202211290000.0A 2022-10-20 2022-10-20 Speed reduction device and aircraft engine Pending CN115539212A (en)

Priority Applications (1)

Application Number Priority Date Filing Date Title
CN202211290000.0A CN115539212A (en) 2022-10-20 2022-10-20 Speed reduction device and aircraft engine

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
CN202211290000.0A CN115539212A (en) 2022-10-20 2022-10-20 Speed reduction device and aircraft engine

Publications (1)

Publication Number Publication Date
CN115539212A true CN115539212A (en) 2022-12-30

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Family Applications (1)

Application Number Title Priority Date Filing Date
CN202211290000.0A Pending CN115539212A (en) 2022-10-20 2022-10-20 Speed reduction device and aircraft engine

Country Status (1)

Country Link
CN (1) CN115539212A (en)

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