A combustion and lubricating oil afterburning system at the outlet of the combustion chamber for cooling the bearings of an aeroengine

Abstract

The present invention belongs to the technical field of the research, development and optimization of small aero-engines, and particularly relates to a combustion augmentation system for the outlet of a fuel / oil combustion chamber for cooling an aero-engine bearing; the high-pressure gas provided by a compressor passes the fuel / oil mixture for cold-state bearing cooling through the cavity formed by a front bearing, an aero-engine rotating shaft and an aero-engine shaft cavity, and a rear bearing respectively, and then is discharged into a position near the outlet of the combustion chamber through a drain pipe on the aero-engine shaft cavity for combustion augmentation, which not only increases the outlet temperature of the combustion chamber, thereby improving the efficiency of the aero-engine, but also utilizes the fuel / oil for cooling for combustion augmentation, reducing the fuel consumption rate of the engine. It improves the efficiency of the prototype aero-engine and reduces the fuel consumption rate. The modification cost of the improved structure is low, which can significantly reduce the fuel consumption rate and improve the propulsion efficiency of the aero-engine.

Description

Technical Field

[0001] The present invention belongs to the technical field of small aircraft engine research and development and optimization, and in particular relates to a combustion chamber outlet supplementary combustion system for cooling aircraft engine bearings and lubricating oil. Background Art

[0002] Aircraft engines are known as the jewel in the crown of industry. They involve materials, energy, chemistry, mechanics, and other disciplines, and are very complex. All countries with conditions attach great importance to the research and development and expansion of aircraft engines, not only in military technology, but also in civilian and commercial fields. In recent years, the development of ultra-small aircraft engines has received more and more attention from all aspects. As people's requirements for resources, efficiency, entertainment, services, and safety are getting higher and higher, the demand for drone exploration, drone delivery, small flying cars, target drones, teaching aids, and high-end toys is increasing, and the market is also growing. As the power of small aircraft, small aircraft engines are the most important technical core. Compared with power devices such as electric motors and small rockets, they have unique advantages and functions. Their efficiency, cost, safety and stability and other indicators have the most important impact on development and application.

[0003] Jet aircraft engine technology has been greatly developed since World War II due to the promotion of military use and civil transportation. Military use mainly focuses on performance and stability, and cost factors are secondary, while civil transportation mainly considers large aircraft engine equipment and rarely involves ultra-small multi-purpose aircraft engines. Therefore, under the condition of cost constraints, ultra-small multi-purpose jet aircraft engines still face serious problems such as insufficient thrust, low efficiency, poor safety and stability, high fuel consumption and high pollution. How to improve the performance and stability of ultra-small multi-purpose aircraft engines at a lower cost has become an urgent problem that needs to be solved.

[0004] The ultra-small turbojet engines commonly seen on the market currently have a high speed, and the bearings are generally cooled by a mixture of fuel and lubricating oil in a certain proportion. Unlike large aircraft engines that have cooling circulation accessory systems and sealing systems such as oil-gas separators, lubricating oil pumps, heat exchangers, and lubricating oil return pumps, ultra-small turbojet engines have a simple structure, low cost, and a short safe and stable working time, so the cooled blended mixture is directly discharged from the engine, resulting in a high fuel consumption rate and serious pollution. At the same time, from the perspective of the combustion chamber, the material selection of ultra-small aircraft engines is relatively simple. For example, thin-walled materials such as commonly used industrial stainless steel are usually used as combustion chamber materials, and the combustion chamber structure is usually only simply optimized, so the oil and gas ratio is low, the flow loss is high, the combustion efficiency is low, and the combustion chamber outlet temperature is low, resulting in insufficient thrust, low propulsion efficiency, and a small thrust-to-weight ratio for the aircraft engine. If the fuel consumption rate can be reduced and the reasonable combustion chamber outlet temperature of the aircraft engine can be increased while saving costs to the maximum extent, thereby improving the efficiency of the engine, it will have greater market competitiveness and development space.

[0005] Therefore, based on the above technical problems, it is necessary to design a new fuel and lubricating oil combustion chamber outlet supplementary combustion system for aircraft engine bearing cooling. Summary of the invention

[0006] The purpose of the present invention is to provide a combustion chamber outlet supplementary combustion system for cooling the bearings of an aircraft engine.

[0007] In order to solve the above technical problems, the present invention provides a combustion chamber outlet supplementary combustion system for cooling fuel and lubricating oil for aircraft engine bearings, comprising:

[0008] Aircraft engine shafts, front bearings, aeroengine shaft cavities, rear bearings and combustion chambers;

[0009] The aero-engine shaft cavity acts on the aero-engine rotating shaft through the front bearing and the rear bearing to bear force and limit position;

[0010] The combustion chamber is sleeved outside the aeroengine shaft cavity;

[0011] The engine shaft is a high-speed rotor component;

[0012] The aero-engine shaft cavity and combustion chamber are stator parts.

[0013] Furthermore, the aircraft engine shaft comprises: a rotor shaft, a bearing inner ring positioning ring, a sealing ring, a positioning adjustment ring and a clamping nut;

[0014] The compression nut is arranged at one end of the rotor shaft;

[0015] The bearing inner ring positioning ring, the sealing ring and the positioning and adjusting ring are sleeved on the rotor shaft in sequence;

[0016] The positioning and adjusting ring is arranged close to the compression nut;

[0017] The bearing inner rings of the front bearing and the rear bearing are sleeved on the rotor shaft. The bearing inner ring of the front bearing is arranged on one side of the bearing inner ring of the rear bearing, and the bearing inner ring of the rear bearing is arranged on one side of the bearing inner ring positioning ring.

[0018] Further, the aero-engine shaft cavity includes: a mounting and positioning edge and a shaft cavity;

[0019] The inner cavity of the mounting and positioning edge is matched with the bearing outer ring of the front bearing;

[0020] The rear inner cavity of the shaft cavity is matched with the bearing outer ring of the rear bearing.

[0021] Further, a plurality of rows of oil pipes are circumferentially and equidistantly arranged on the outer wall of the shaft cavity;

[0022] The length direction of the row of oil pipes is arranged along the radial direction of the shaft cavity.

[0023] Further, a combustion chamber evaporation pipe is arranged in the combustion chamber;

[0024] The combustion chamber evaporation pipe adopts a threaded pipe shape;

[0025] The axis direction of the threaded pipe is arranged at a certain angle with the axis direction of the aero-engine rotating shaft.

[0026] Further, the front bearing serves as an axial positioning and anchoring position;

[0027] The mounting and positioning edge is positioned by other stator structures of the aero-engine.

[0028] Further, the relative positions of the row of oil pipes and the sealing ring are adjusted by the bearing inner ring positioning ring and the positioning and adjusting ring.

[0029] Further, the cooling fuel / lubricating oil and high-pressure air mixture of the rear bearing converges through the chamber formed between the rear bearing, the aero-engine shaft cavity and the sealing ring, and is discharged into the combustion chamber from the row of oil pipes.

[0030] Further, the row of oil pipes is installed in the threaded hole of the aero-engine shaft cavity through the air inlet hole of the inner ring of the combustion chamber. By adjusting the degree of screw-in of the row of oil pipes in the shaft cavity, it is ensured that the outlet of the row of oil pipes and the inner wall surface of the inner ring of the combustion chamber are on the same plane.

[0031] Further, the center of the outlet of the row of oil pipes is concentric with the center of the air inlet hole of the inner wall surface of the inner ring of the combustion chamber.

[0032] The beneficial effects of the present invention are as follows: The present invention includes an aeroengine rotating shaft, a front bearing, an aeroengine shaft cavity, a rear bearing, and a combustion chamber. The aeroengine shaft cavity acts on the aeroengine rotating shaft through the front bearing and the rear bearing for load bearing and positioning. The combustion chamber is sleeved outside the aeroengine shaft cavity. The engine rotating shaft is a high-speed rotating part, and the aeroengine shaft cavity and the combustion chamber are stator parts. The high-pressure gas provided by the compressor cools the cold-state bearings by passing the fuel / oil mixture through the cavities formed by the front bearing, the aeroengine rotating shaft, and the aeroengine shaft cavity, and then through the rear bearing, and is discharged into the position near the outlet of the combustion chamber through the drain pipe on the aeroengine shaft cavity for afterburning. This not only increases the outlet temperature of the combustion chamber, thereby improving the efficiency of the aeroengine, but also utilizes the fuel / oil mixture for cooling for afterburning, reducing the fuel consumption rate of the engine. It improves the efficiency of the prototype aeroengine and reduces the fuel consumption rate. The modification cost of the improved structure is low, which can significantly reduce the fuel consumption rate and improve the propulsion efficiency of the aeroengine.

[0033] Other features and advantages of the present invention will be described in the following specification, and some of them will become obvious from the specification or be understood by implementing the present invention. The objectives and other advantages of the present invention are achieved and obtained by the structures specifically pointed out in the specification and the drawings.

[0034] To make the above objectives, features, and advantages of the present invention more obvious and understandable, the following specifically provides preferred embodiments and, in conjunction with the accompanying drawings, details are described as follows. Brief Description of the Drawings

[0035] To more clearly illustrate the specific embodiments of the present invention or the technical solutions in the prior art, the following will briefly introduce the drawings required for use in the description of the specific embodiments or the prior art. Obviously, the following drawings are some embodiments of the present invention. For those of ordinary skill in the art, without creative efforts, other drawings can also be obtained based on these drawings.

[0036] Figure 1 It is a schematic structural diagram of a fuel / oil afterburning system at the outlet of the combustion chamber for cooling the bearings of an aeroengine according to the present invention;

[0037] Figure 2 It is a schematic structural diagram of the aeroengine rotating shaft according to the present invention;

[0038] Figure 3 It is a schematic structural diagram of the aeroengine shaft cavity according to the present invention.

[0039] In the figure:

[0040] 1 is the rotating shaft of an aeroengine; 1.1 is the rotor shaft; 1.2 is the bearing inner ring positioning ring; 1.3 is the sealing ring; 1.4 is the positioning and adjusting ring; 1.5 is the compression nut; 2 is the front bearing; 3 is the aeroengine shaft cavity; 3.1 is the installation and positioning edge; 3.2 is the shaft cavity; 3.3 is the drain pipe; 4 is the rear bearing; 5 is the combustion chamber; 5.1 is the combustion chamber evaporation pipe; Specific embodiments

[0041] To make the objectives, technical solutions, and advantages of the embodiments of the present invention clearer, the technical solutions of the present invention will be clearly and completely described below with reference to the accompanying drawings. Apparently, the described embodiments are some, but not all, of the embodiments of the present invention. All other embodiments obtained by those of ordinary skill in the art based on the embodiments of the present invention without creative efforts shall fall within the protection scope of the present invention.

[0042] As Figures 1 to 3 As shown, this embodiment provides a combustion and lubricating oil combustion chamber outlet supplementary combustion system for cooling the bearings of an aeroengine, including: the aeroengine rotating shaft 1, the front bearing 2, the aeroengine shaft cavity 3, the rear bearing 4, and the combustion chamber 5; the rear bearing 4 can be a roller bearing, and the entire shaft allows a small axial expansion; the aeroengine shaft cavity 3 acts on the aeroengine rotating shaft 1 through the front bearing 2 and the rear bearing 4 for force bearing and positioning; the combustion chamber 5 is sleeved outside the aeroengine shaft cavity 3; the engine rotating shaft is a high-speed rotor component; the aeroengine shaft cavity 3 and the combustion chamber 5 are stator components; it realizes that the high-pressure gas provided by the compressor passes the fuel / oil mixture for cooling the cold bearings through the cavity formed by the front bearing 2, the aeroengine rotating shaft 1, and the aeroengine shaft cavity 3, and the rear bearing 4, and then is discharged into the position near the outlet of the combustion chamber 5 through the drain pipe 3.3 on the aeroengine shaft cavity 3 for supplementary combustion. This not only increases the outlet temperature of the combustion chamber 5, thereby improving the efficiency of the aeroengine, but also uses the fuel / oil for cooling for supplementary combustion, reducing the fuel consumption rate of the engine. It improves the efficiency of the prototype aeroengine and reduces the fuel consumption rate. The modification cost of the improved structure is low, which can significantly reduce the fuel consumption rate and improve the propulsion efficiency of the aeroengine.

[0043] In this embodiment, the aeroengine rotating shaft 1 includes: a rotor shaft 1.1, a bearing inner ring positioning ring 1.2, a sealing ring 1.3, a positioning and adjusting ring 1.4, and a compression nut 1.5; the compression nut 1.5 is arranged at one end of the rotor shaft 1.1; the bearing inner ring positioning ring 1.2, the sealing ring 1.3, and the positioning and adjusting ring 1.4 are sequentially sleeved on the rotor shaft 1.1; the positioning and adjusting ring 1.4 is arranged close to the compression nut 1.5; the bearing inner rings of the front bearing 2 and the rear bearing 4 are sleeved on the rotor shaft 1.1, the bearing inner ring of the front bearing 2 is arranged on one side of the bearing inner ring of the rear bearing 4, and the bearing inner ring of the rear bearing 4 is arranged on one side of the bearing inner ring positioning ring 1.2.

[0044] In this embodiment, the aeroengine shaft cavity 3 includes: a mounting and positioning edge 3.1 and a shaft cavity 3.2; the inner cavity of the mounting and positioning edge 3.1 is matched with the bearing outer ring of the front bearing 2; the rear inner cavity of the shaft cavity 3.2 is matched with the bearing outer ring of the rear bearing 4.

[0045] In this embodiment, a number of drain pipes 3.3 are circumferentially and equidistantly arranged on the outer wall of the shaft cavity 3.2; the length direction of the drain pipes 3.3 is arranged along the radial direction of the shaft cavity 3.2; for example, an even number such as 4, 6, 8, etc. are axially evenly distributed, aiming to make the outlet temperature of the combustion chamber 5 more uniform, and they are connected to the shaft cavity 3.2 by means of screw threads. The drain pipes 3.3 can have a screwing structure such as an external hexagon and a positioning and pressing device such as a nut; there is a wrench screwing space between the shaft cavity 3.2 and the inner wall of the combustion chamber 5.

[0046] In this embodiment, a combustion chamber evaporation pipe 5.1 is arranged in the combustion chamber 5; the combustion chamber evaporation pipe 5.1 adopts a form such as a threaded pipe to increase turbulence; the axial direction of the threaded pipe is arranged at a certain angle with the axial direction of the aeroengine rotating shaft 1, and the optimal angle is obtained through experimental calculations and other means to increase the efficiency of fuel atomization.

[0047] In this embodiment, the front bearing 2 serves as an axial positioning and anchoring position, and assembly and correction are both positioned at this point. The mounting and positioning edge 3.1 is precisely positioned with other stator structures of the aeroengine, such as the compressor stator casing, and has concentric constraint and axial position constraint.

[0048] In this embodiment, the relative positions of the drain pipe 3.3 and the sealing ring 1.3 are adjusted by the bearing inner ring positioning ring 1.2 and the positioning and adjusting ring 1.4; the cooling fuel / lubricating oil and high-pressure air mixture of the rear bearing 4 converges through the chamber formed between the rear bearing 4, the aeroengine shaft cavity 3 and the sealing ring 1.3, and is discharged into the combustion chamber 5 from the drain pipe 3.3; the functions of the bearing inner ring positioning ring 1.2 and the positioning and adjusting ring 1.4 are not only for accurate assembly, but also for facilitating the adjustment of the combustion chamber 5 to approach the afterburner near the tail nozzle to prevent the ablation of the turbine guide vane caused by excessive temperature.

[0049] In this embodiment, the drain pipe 3.3 is installed in the threaded hole of the aeroengine shaft cavity 3 through the air intake hole in the inner ring of the combustion chamber 5. By adjusting the degree of screw-in of the drain pipe 3.3 in the shaft cavity 3.2, it is ensured that the outlet of the drain pipe 3.3 and the inner ring wall surface of the combustion chamber 5 are on the same plane. The purpose is to prevent the drain pipe 3.3 extending into the combustion chamber 5 from disturbing the main flow of the combustion chamber 5 and increasing additional flow losses, and to isolate the contact and flame transmission between the flame in the combustion chamber 5 and the drain pipe 3.3 through the air intake function of the combustion chamber 5 air intake hole.

[0050] In this embodiment, the center of the outlet of the drain pipe 3.3 is concentric with the center of the air intake hole of the inner ring wall surface of the combustion chamber 5; to make the fluid discharged from the drain pipe 3.3 uniform, and after knowing the air flow temperature of the air intake hole of the combustion chamber 5, the inner hole diameter of the drain pipe 3.3 is determined by calculation. The purpose is to cause flame extinction and prevent the flame from spreading in the drain pipe 3.3, to ensure that the flame of the combustion chamber 5 does not spread to the chamber formed between the rear bearing 4, the aeroengine shaft cavity 3 and the sealing ring 1.3, and to reduce the flow resistance of the drain pipe 3.3 through optimized calculation.

[0051] In this embodiment, before the outlet of the combustion chamber 5, there is a structure similar to that of the combustion chamber 5 that increases vortex mixing and cooling, such as a thin-walled right-angled edge structure. This structure can form a low-temperature and low-pressure vortex region before the gas flows out of the outlet of the combustion chamber 5 to prevent the ablation of the turbine guide vane due to excessive temperature, but it reduces the outlet temperature of the combustion chamber 5 and has a large flow loss. Obviously, afterburning at this position can offset part of the excessive temperature reduction, and the formation of the vortex helps the secondary mixing of fuel and air, which is beneficial to reducing the reaction time and increasing the combustion heat release efficiency.

[0052] In summary, the present invention relates to an aero-engine rotating shaft 1, a front bearing 2, an aero-engine shaft cavity 3, a rear bearing 4, and a combustion chamber 5. The aero-engine shaft cavity 3 acts on the aero-engine rotating shaft 1 through the front bearing 2 and the rear bearing 4 for load bearing and positioning. The combustion chamber 5 is sleeved outside the aero-engine shaft cavity 3. The engine rotating shaft is a high-speed rotating part, and the aero-engine shaft cavity 3 and the combustion chamber 5 are stator parts. It realizes that the high-pressure gas provided by the compressor cools the fuel / oil mixture for the cold bearings and then discharges it into the position near the outlet of the combustion chamber 5 through the cavity formed by the front bearing 2, the aero-engine rotating shaft 1, and the aero-engine shaft cavity 3, and the rear bearing 4, and then through the drain pipe 3.3 on the aero-engine shaft cavity 3 for supplementary combustion. This not only increases the outlet temperature of the combustion chamber 5, thereby improving the efficiency of the aero-engine, but also uses the fuel / oil mixture for cooling for supplementary combustion, reducing the fuel consumption rate of the engine. It improves the efficiency of the prototype aero-engine and reduces the fuel consumption rate. The modification cost of the improved structure is low, which can significantly reduce the fuel consumption rate and improve the propulsion efficiency of the aero-engine.

[0053] All the components selected in this application (components without specific structures described) are common standard components or components known to those skilled in the art. Their structures and principles can be known by those skilled in the art through technical manuals or obtained through conventional experimental methods.

[0054] In the description of the embodiments of the present invention, unless otherwise clearly specified and limited, the terms "installation", "connection", and "connection" should be understood in a broad sense. For example, it can be a fixed connection, a detachable connection, or an integral connection; it can be a mechanical connection or an electrical connection; it can be directly connected or indirectly connected through an intermediate medium, and it can be the communication inside two components. For those of ordinary skill in the art, the specific meanings of the above terms in the present invention can be understood according to specific situations.

[0055] In the description of the present invention, it should be noted that the orientation or positional relationship indicated by the terms "center", "upper", "lower", "left", "right", "vertical", "horizontal", "inner", "outer", etc. is based on the orientation or positional relationship shown in the drawings. It is only for the convenience of describing the present invention and simplifying the description, rather than indicating or implying that the device or component referred to must have a specific orientation, be constructed and operated in a specific orientation, so it cannot be understood as a limitation to the present invention. In addition, the terms "first", "second", and "third" are only used for descriptive purposes and cannot be understood as indicating or implying relative importance.

[0056] In several embodiments provided by this application, it should be understood that the disclosed systems, devices, and methods can be implemented in other ways. The device embodiments described above are merely illustrative. For example, the division of the units is only a logical function division. In actual implementation, there may be other division methods. For another example, multiple units or components can be combined or integrated into another system, or some features can be ignored or not executed. Another point is that the displayed or discussed coupling, direct coupling, or communication connection to each other can be through some communication interfaces. The indirect coupling or communication connection of the devices or units can be in electrical, mechanical, or other forms.

[0057] The units described as separate components may or may not be physically separated. The components displayed as units may or may not be physical units, that is, they may be located in one place, or they may be distributed to multiple network units. Some or all of the units can be selected according to actual needs to achieve the purpose of the solution of this embodiment.

[0058] In addition, in each embodiment of the present invention, the functional units can be integrated into one processing unit, or each unit can exist physically alone, or two or more units can be integrated into one unit.

[0059] Based on the above inspiration from the ideal embodiments of the present invention, through the above description, relevant staff can completely make various changes and modifications without departing from the technical idea of this invention. The technical scope of this invention is not limited to the content in the specification, and its technical scope must be determined according to the scope of the claims.

Claims

1. A combustion and lubricating oil afterburning system at the outlet of a combustion chamber for cooling an aeroengine bearing, characterized in that, Comprising: An aeroengine rotating shaft, a front bearing, an aeroengine shaft cavity, a rear bearing, and a combustion chamber; The aeroengine shaft cavity acts on the aeroengine rotating shaft through the front bearing and the rear bearing for load bearing and positioning; The combustion chamber is sleeved outside the aeroengine shaft cavity; The engine rotating shaft is a high-speed rotor component; The aeroengine shaft cavity and the combustion chamber are stator components; The aeroengine shaft cavity includes: a mounting and positioning edge and a shaft cavity; The inner cavity of the mounting and positioning edge is fitted with the outer bearing ring of the front bearing; The rear inner cavity of the shaft cavity is fitted with the outer bearing ring of the rear bearing; A number of rows of oil pipes are circumferentially and equidistantly arranged on the outer wall of the shaft cavity; The length direction of the oil pipes is arranged along the radial direction of the shaft cavity; The center of the oil pipe outlet is concentric with the center of the air inlet hole on the inner ring wall surface of the combustion chamber.

2. The fuel and lubricating oil combustion chamber outlet afterburning system for aeroengine bearing cooling according to claim 1, characterized in that The aeroengine rotating shaft includes: a rotor shaft, a bearing inner ring positioning ring, a sealing ring, a positioning and adjusting ring, and a compression nut; The compression nut is arranged at one end of the rotor shaft; The bearing inner ring positioning ring, the sealing ring, and the positioning and adjusting ring are sequentially sleeved on the rotor shaft; The positioning and adjusting ring is arranged close to the compression nut; The inner bearing rings of the front bearing and the rear bearing are sleeved on the rotor shaft. The inner bearing ring of the front bearing is arranged on one side of the inner bearing ring of the rear bearing, and the inner bearing ring of the rear bearing is arranged on one side of the bearing inner ring positioning ring.

3. The fuel and lubricating oil combustion chamber outlet afterburning system for aeroengine bearing cooling according to claim 2, characterized in that A combustion chamber evaporation pipe is arranged in the combustion chamber; The combustion chamber evaporation pipe adopts a threaded pipe shape; The axis direction of the threaded pipe is arranged at a certain angle with the axis direction of the aeroengine rotating shaft.

4. The fuel and lubricating oil combustion chamber outlet afterburning system for aeroengine bearing cooling according to claim 3, characterized in that The front bearing serves as an axial positioning and anchoring position; The mounting and positioning edge is positioned through other stator structures of the aeroengine.

5. The fuel and lubricating oil combustion chamber outlet afterburning system for aeroengine bearing cooling according to claim 4, characterized in that The relative positions of the oil pipes and the sealing ring are adjusted through the bearing inner ring positioning ring and the positioning and adjusting ring.

6. The fuel and lubricating oil combustion chamber outlet afterburning system for aeroengine bearing cooling according to claim 5, characterized in that The cooling fuel / lubricating oil and high-pressure air mixture of the rear bearing converges through the chamber formed between the rear bearing, the aeroengine shaft cavity, and the sealing ring, and is discharged into the combustion chamber from the oil pipes.

7. The fuel and lubricating oil combustion chamber outlet afterburning system for aeroengine bearing cooling according to claim 6, characterized in that The oil pipes are installed in the threaded holes of the aeroengine shaft cavity through the air inlet holes of the inner ring of the combustion chamber. By adjusting the degree of screw-in of the oil pipes in the shaft cavity, it is ensured that the outlets of the oil pipes and the inner ring wall surface of the combustion chamber are on the same plane.

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