CN107842442B - Aircraft engine - Google Patents

Aircraft engine Download PDF

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Publication number
CN107842442B
CN107842442B CN201711102132.5A CN201711102132A CN107842442B CN 107842442 B CN107842442 B CN 107842442B CN 201711102132 A CN201711102132 A CN 201711102132A CN 107842442 B CN107842442 B CN 107842442B
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China
Prior art keywords
compressor
engine
motor
combustion chamber
magnetoelectric
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CN201711102132.5A
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Chinese (zh)
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CN107842442A (en
Inventor
邱名
范召林
肖中云
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Computational Aerodynamics Institute of China Aerodynamics Research and Development Center
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Computational Aerodynamics Institute of China Aerodynamics Research and Development Center
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Classifications

    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02KJET-PROPULSION PLANTS
    • F02K7/00Plants in which the working fluid is used in a jet only, i.e. the plants not having a turbine or other engine driving a compressor or a ducted fan; Control thereof
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02KJET-PROPULSION PLANTS
    • F02K1/00Plants characterised by the form or arrangement of the jet pipe or nozzle; Jet pipes or nozzles peculiar thereto
    • F02K1/78Other construction of jet pipes
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23RGENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/28Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23RGENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/42Continuous combustion chambers using liquid or gaseous fuel characterised by the arrangement or form of the flame tubes or combustion chambers
    • HELECTRICITY
    • H02GENERATION; CONVERSION OR DISTRIBUTION OF ELECTRIC POWER
    • H02KDYNAMO-ELECTRIC MACHINES
    • H02K44/00Machines in which the dynamo-electric interaction between a plasma or flow of conductive liquid or of fluid-borne conductive or magnetic particles and a coil system or magnetic field converts energy of mass flow into electrical energy or vice versa
    • H02K44/08Magnetohydrodynamic [MHD] generators

Abstract

The invention discloses an aircraft engine, which comprises an air inlet channel, an air compressor, a combustion chamber and a spray pipe, wherein air flow enters the air compressor through the air inlet channel, enters the combustion chamber after being compressed, and is discharged through the spray pipe after being combusted; the biggest innovation point of the invention is that a turbine and a corresponding cooling system are cancelled in the engine, plasma thermal jet is used for generating electricity, and the motor is used for driving the air compressor, thereby solving the problem that the total temperature in front of the turbine is limited at present.

Description

Aircraft engine
Technical Field
The invention belongs to the field of engines, and particularly relates to an aircraft engine which is suitable for an aircraft below Mach 3.
Background
The aircraft engine is a national high-tech strategic industry, and affects the politics and outturn of the country. To increase the thrust and thrust-to-weight ratio of an aircraft engine, it is desirable that the total temperature in front of the turbine be as high as possible. However, the current aircraft engine is limited by materials and cooling technology, and the total temperature before the turbine cannot be too high (the total temperature before the turbine of the fourth generation aircraft engine can reach 2000K).
Due to the influence of high-temperature gas, the current aircraft engine needs to cool the turbine blades, so that a cold air system is designed to bleed air from the air compressor and blow the air out of small holes of the turbine blades; this not only increases the system complexity of the aircraft engine, but also increases flow losses, reducing the thermal efficiency and thrust of the aircraft engine. Meanwhile, the turbine blade of the aero-engine works in a high-temperature and high-pressure environment for a long time, so that the turbine blade is high in production cost and short in service life, and the economy of the aero-engine is restricted.
The turbine is a rotating part and requires bearings and a lubrication system. In order to avoid the overhigh temperature of the turbine bearing, bleed air is needed to cool the parts near the turbine bearing and prevent the high-temperature gas leaked between the stages of the turbine from flowing backwards. In practical use, the turbine blade holes are often blocked, and high-temperature gas flows backwards to the backward flow bearing.
The compressor and the turbine are connected on the same shaft, and the rotating speed and the power are the same, so that the matching problem of the compressor and the turbine of the aircraft engine is caused. The multi-shaft design can improve the interstage matching of the compressor, but the high-pressure rotating component is positioned between the low-pressure rotating components, and the multi-rotating shaft design can be realized only by adopting a spatial shaft; in addition, a combustion chamber is arranged between the compressor and the turbine, and the rotating shaft needs to penetrate through the middle of the combustion chamber, so that the structure is extremely complex, and the design of multiple rotating shafts is difficult. Thus, current aircraft engines typically have only 1-2 spools, and only the 3-spool engine developed by luo.
In order to improve the thrust and thrust-weight ratio of the aero-engine, improve the performance matching of parts, and avoid the limitation of high-temperature materials and cooling technologies, the development of a high-temperature gas aero-engine which is simple in structure and has no turbine blades is needed.
Disclosure of Invention
The invention aims to provide an aircraft engine without turbine blades, which avoids the limitations of high-temperature materials and cooling technologies and realizes higher thrust and higher thrust-weight ratio.
In order to achieve the purpose, the invention adopts the following technical scheme:
the aero-engine comprises an air inlet channel, a compressor, a combustion chamber and a spray pipe, wherein air flow enters the compressor through the air inlet channel, enters the combustion chamber after being compressed, and is discharged through the spray pipe after being combusted, a turbine is not arranged in the aero-engine, and the compressor is driven by a motor to work.
In the technical scheme, the air compressor has a high pressure ratio, the combustion chamber adopts isobaric combustion, and the front section of the spray pipe injects oil to perform isothermal combustion.
In the technical scheme, the motor is arranged in a hub of the air compressor and comprises a coil winding and a magnet, and the coil winding and the air compressor wheel disc are integrated.
In the technical scheme, the compressor is provided with a plurality of stages, a plurality of motors are arranged in a hub of the compressor, and each motor drives at least one stage of the compressor.
In the above technical scheme, the nozzle is provided with a thermal jet magnetoelectric device, and the thermal jet magnetoelectric device comprises a positive electrode, a negative electrode and a magnet.
In the technical scheme, the spray pipe is a circular rotary square spray pipe, the square section comprises an expanding wall and a flat wall, the inner wall surface of the expanding wall is provided with an electrode, and the outer wall surface of the flat wall is provided with a magnet.
In the technical scheme, the spray pipe is provided with a plurality of electrodes and magnets, and the output ends of the electrodes are connected to the motor.
In the technical scheme, the combustion chamber is of a streamline rotary structure, and the rotary radius of the combustion chamber is gradually reduced from one end of the gas compressor to the end of the spray pipe.
In the technical scheme, the front end of the spray pipe is provided with the plurality of rings of oil nozzles, isothermal combustion is carried out at the front end of the spray pipe to replace traditional afterburning, and higher heat efficiency is achieved.
In the above technical solution, the engine has a working flow:
the method comprises the following steps: the external power supply drives the air compressor to work, and simultaneously drives the thermal jet flow magnetoelectric device to work, and after the air compressor works normally, fuel oil is sprayed into the combustion chamber and ignited;
step two: after successful ignition, high-speed charged plasma airflow is ejected after combustion, and the airflow generates current through a thermal jet magnetoelectric device;
step three: after the engine works stably, external power supply is cut off, the heat jet magnetoelectric device supplies power to the motor, the motor drives the compressor to work, and the engine enters an idling working state.
Step four: in the taking-off process, oil supply is increased, the rotating speed is increased, the engine enters a climbing state, the oil supply is reduced and each component system is adjusted after the engine enters the specified height and speed, and the engine enters a cruising state.
In the technical scheme, the ionization catalyst in the fuel ionizes gas into plasma at high temperature, and when plasma gas flows through the middle of a magnet of the thermal jet magnetoelectric device at high speed, the plasma cuts magnetic lines of force in a magnetic field to generate induced electromotive force, so that a potential difference is formed between two electrodes to generate stable direct current.
In the technical scheme, the catalyst is used for reducing the ionization temperature of the airflow, and the ionization temperature of the gas is reduced from about 6000K to about 3000K through the catalyst.
In summary, due to the adoption of the technical scheme, the invention has the beneficial effects that:
the biggest innovation point of the invention is that a turbine and a corresponding cooling system are cancelled in the engine, thereby solving the problem that the total temperature in front of the turbine is limited, removing the thrust loss caused by the cooling system and greatly increasing the thrust of the engine; the plasma thermal jet is used for generating electricity, the compressor is driven by a motor, the electric energy can be freely distributed in the motor of each compression part, multiple rotating shafts are easy to realize, and the matching problem of a compression system is solved; a transmission shaft between the original turbine and the compressor is removed, the structural complexity is greatly reduced, the complexity of the problem of rotor dynamics is weakened, and the influence of the critical rotating speed on the engine is weakened; isothermal combustion after the combustion chamber is adopted to replace the traditional afterburning, so that the heat efficiency is greatly increased.
Drawings
The invention will now be described, by way of example, with reference to the accompanying drawings, in which:
FIG. 1 is a schematic structural view of the present invention;
FIG. 2 is a schematic structural view of the present invention;
FIG. 3 is a schematic view of the convergent-divergent wall configuration of the nozzle of the present invention;
FIG. 4 is a schematic view of a straight wall construction of the nozzle of the present invention;
wherein: 1 is an air inlet channel, 2 is an air compressor, 3 is a combustion chamber, 4 is a spray pipe, 5 is a motor, 6 is a thermal jet magnetoelectric device, 7 is an electrode, and 8 is a magnet.
Detailed Description
All of the features disclosed in this specification, or all of the steps in any method or process so disclosed, may be combined in any combination, except combinations of features and/or steps that are mutually exclusive.
As shown in figures 1 and 2, the engine comprises six core components including an air inlet, an air compressor, a combustion chamber, a spray pipe, a motor and a thermal jet magnetoelectric device, changes the structure of the traditional engine, removes a turbine and uses electric energy to provide compression work.
The invention has the greatest improvement after removing the turbine blades, namely, the air compressor is improved, the rotor of the air compressor is driven to rotate coaxially with the turbine before, and the air compressor is driven by an independent motor. The invention arranges a plurality of motors in the wheel hub of the air compressor, each motor drives at least one stage of air compressor rotor to rotate, the motor of the invention is composed of coil winding and magnet, and the power supply of the motor is powered by a thermal jet magnetoelectric device.
In order to work with the motor, the thermal jet magnetoelectric device is introduced into the spray pipe, as shown in fig. 3 and 4, two electrodes are arranged on the inner wall surface of the expansion wall of the spray pipe (of course, a plurality of electrodes can be arranged on the whole spray pipe, only two electrodes are marked in the drawing for convenient understanding), and a magnet is arranged on the outer wall surface of the straight wall (of course, a plurality of magnets can be arranged on the whole spray pipe, only two electrodes are marked in the drawing for convenient understanding). The magnetic field is provided by a superconducting electromagnet or other permanent magnets with strong magnetism, so that the electric energy which can be continuously obtained by the plasma thermal jet magnetoelectric device of the engine under the working condition can be output to the motor.
The coil winding of the motor and the wheel disc of the air compressor are of an integrated structure, and the coil winding and the wheel disc of the air compressor are combined with an electromagnet (or a permanent magnet) to form a direct current motor, when direct current voltage from a thermal jet flow magnetoelectric device is input into a coil, the motor can start to work and drive the blades of the air compressor to rotate, and the work of the air compressor is realized.
The key point of the invention is how to realize the electric energy supply of the plasma thermal jet magnetoelectric device, so that the invention adds a catalyst into the fuel oil, wherein the catalyst is an ionization catalyst and generally contains an alkali metal substance (such as potassium salt). When fuel is burned to generate high temperature, the catalyst ionizes air into plasma at high temperature. Because a large amount of plasma gas flow is accelerated in the spray pipe and then sprayed out of the spray pipe, and a large amount of charges are carried in the plasma, when the plasma passes through the magnetic fields of the two electrodes, magnetic induction lines are cut, and particles with positive and negative charges can drift towards different electrodes under the action of the magnetic fields, so that a potential difference is formed between the two electrodes.
On the basis of all the above, the invention cancels the traditional turbine blade and cooling gas circuit, which makes the outlet temperature of the combustion chamber higher, the heat efficiency of the engine higher and the unit flow thrust larger. The traditional annular structure of the combustion chamber is changed into the arc-shaped structure of the scheme, the radian of the combustion chamber from one end of the gas compressor to one end of the spray pipe is gradually increased, and the fusion design of the combustion chamber and the spray pipe is realized (as shown in figure 2). This structural improvement allows the burned gas to expand rapidly and makes the engine length short.
The specific working process of the whole engine is as follows:
in a take-off state, the compressor is driven by an external power supply of the airport, and the superconducting electromagnet works; after the compressor works normally, fuel oil (potassium salt or other ionization catalyst is added in the fuel oil) is sprayed into the combustion chamber and ignited;
after successful ignition, the external power supply is maintained for several seconds until the engine stably works; after the engine works stably, an external power supply is cut off, the thermal jet magnetoelectric device supplies power to the gas compressor and the electromagnet, and the engine enters an idling state;
after receiving a takeoff command, increasing the oil supply of the engine, and enabling the engine to enter a maximum climbing state and have maximum thrust;
after a specified altitude and speed, the fuel supply is reduced and the individual component systems are adjusted, and the engine enters cruise conditions.
The ground starting of the engine can adopt three schemes of plug-in starting, battery starting and micro gas turbine starting. When the plug-in starting is adopted, the motor is externally connected with an airport power supply, oil is sprayed and combusted after the driving compressor works, and the power supply is disconnected after the starting is finished; when the battery is adopted for starting, the aircraft needs to be provided with the battery, the power is supplied to the motor during starting, and the battery is charged by the engine after starting; when the micro gas turbine is adopted for starting, the starting process is consistent with that of the existing aircraft engine.
The invention is not limited to the foregoing embodiments. The invention extends to any novel feature or any novel combination of features disclosed in this specification and any novel method or process steps or any novel combination of features disclosed.

Claims (4)

1. The utility model provides an aeroengine, includes intake duct, compressor, combustion chamber and spray tube, and the air current gets into the compressor through the intake duct, reentries the combustion chamber after the compression, discharges through the spray tube after the burning, its characterized in that not be provided with the turbine in the engine, by the work of motor drive compressor, the combustion chamber adopts isobaric combustion, spray tube anterior segment oil spout carries out isothermal combustion, the motor sets up in the wheel hub of compressor, the compressor has a plurality of grades, is provided with a plurality of motor in the wheel hub of compressor, each motor drive one-level compressor, the motor includes coil winding and motor magnet, coil winding and compressor rim plate are as an organic whole, be provided with thermal jet magnetoelectric device on the spray tube, thermal jet magnetoelectric device includes positive and negative electrode and magnet, the spray tube is circular square spray tube of turning, and square section is including expanding wall and straight wall, the inner wall surface of the expansion wall is provided with a positive electrode and a negative electrode, the outer wall surface of the straight wall is provided with a magnet, and the output ends of the positive electrode and the negative electrode are connected to the motor.
2. An aircraft engine according to claim 1, wherein the combustion chamber is of streamlined revolution, the radius of revolution of the combustion chamber decreasing progressively from the compressor end to the nozzle end.
3. An aircraft engine according to any one of claims 1 to 2, characterised in that the engine has a working flow comprising:
the method comprises the following steps: the external power supply drives the air compressor to work, and simultaneously drives the thermal jet flow magnetoelectric device to work, and after the air compressor works normally, fuel oil is sprayed into the combustion chamber and ignited;
step two: after successful ignition, the combustion chamber ejects high-speed charged plasma airflow, and the airflow generates current through a thermal jet magnetoelectric device;
step three: after the engine stably works, external power supply is cut off, the thermal jet magnetoelectric device supplies power to the motor, the motor drives the compressor to work, and the engine enters an idling working state;
step four: in the taking-off process, oil supply is increased, the rotating speed is increased, the engine enters a climbing state, the oil supply is reduced and each component system is adjusted after the engine enters the specified height and speed, and the engine enters a cruising state.
4. An aircraft engine according to claim 3, characterized in that the ionization catalyst in the fuel ionizes the gas at high temperature into plasma, and when the plasma gas flow passes through the middle of the magnet of the hot jet magnetoelectric device at high speed, the plasma cuts magnetic lines of force in the magnetic field to generate induced electromotive force, so that a potential difference is formed between the positive and negative electrodes to generate stable direct current.
CN201711102132.5A 2017-11-10 2017-11-10 Aircraft engine Active CN107842442B (en)

Priority Applications (1)

Application Number Priority Date Filing Date Title
CN201711102132.5A CN107842442B (en) 2017-11-10 2017-11-10 Aircraft engine

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
CN201711102132.5A CN107842442B (en) 2017-11-10 2017-11-10 Aircraft engine

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CN107842442B true CN107842442B (en) 2020-09-18

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Publication number Priority date Publication date Assignee Title
CN110608108A (en) * 2018-06-14 2019-12-24 哈尔滨工业大学 Non-turbine jet engine integrated with solid oxide fuel cell

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CA2526591A1 (en) * 2003-04-28 2005-02-10 Marius A. Paul Turbo rocket with real carnot cycle
CN101649781A (en) * 2008-08-11 2010-02-17 刘佳骏 Jet engine
CN103441641A (en) * 2013-09-02 2013-12-11 董国光 Detonation wave ejection loop type magnetic fluid power generation system

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Effective date of registration: 20210827

Address after: 621052 No.6, south section of 2nd Ring Road, Fucheng District, Mianyang City, Sichuan Province

Patentee after: COMPUTATIONAL AERODYNAMICS INSTITUTE OF CHINA AERODYNAMICS RESEARCH AND DEVELOPMENT CENTER

Address before: 621052 No.6, south section of 2nd Ring Road, Fucheng District, Mianyang City, Sichuan Province

Patentee before: COMPUTATIONAL AERODYNAMICS INSTITUTE OF CHINA AERODYNAMICS RESEARCH AND DEVELOPMENT CENTER

Patentee before: Qiu Ming