• 专利附录
  • 专利说明
  • 权利要求
  • 类似技术
  • 同族专利
  • 引用文献
  • 法律状态
  • 优先权
    • 专利摘要:
    Gas turbine engine comprising a shaft (20) on which the compressors (4) and (5), the turbines (6) and (7) are arranged in series. The high-pressure combustion chamber (11) is arranged between the compressor (5) and the turbine (6). The bypass passage (16) for a portion of the working fluid connects the outlet of the compressor (4) to the inlet of the turbine (7) via the low pressure combustion chamber (12). The redistribution of the working medium through the channel (16) takes place unhindered.
    公开号:CH708180B1
    申请号:CH01486/14
    申请日:2013-03-26
    公开日:2018-04-13
    发明作者:Iosifovich Belous Vladimir
    申请人:Iosifovich Belous Vladimir；
    IPC主号:
    专利说明:

    Description: The invention relates to a gas turbine engine according to the
    Preamble of claim 1.
    The invention can be used for gas turbine engines with continuous combustion in a high-speed gas stream according to the open scheme with high-energy gas turbine fuels. It can be used in traffic systems, for example in air and energy systems, and also as a drive in gas compressor units.
    A power plant for the generation of electrical energy with a coal gasification under pressure according to the patent document US 4 199 933 is known. In this plant hot gas, which requires further combustion, flows simultaneously from a gas generator onto a high-pressure turbine and also past it in a low pressure combustor and further into a low pressure turbine with a subsequent exit into the atmosphere. Such a post-combustion process for heating gas is only suitable for a low-heating gas, which is produced in a gas generator, for example, by coal gasification under pressure.
    Also known is the patent document US 5 103 630. Fuel is pre-oxygen deficient in a high pressure combustor with a subsequent passage through a turbine stage and further with gas post-combustion in a low pressure combustor with subsequent work in the turbine stage and exit to the atmosphere burned.
    This combustion method of fuels is only suitable for low-heating fuels. For a full post-combustion of fuel with high-energy fuels, many pressure levels of additional air supply are required until the mass fuel consumption is less than 5% of the total consumption of the working body at the exit from the gas turbine engine, or a water supply into the combustion chambers is required.
    A system for fuel combustion in a gas turbine engine according to patent document GB 2 288 640 is also known. This gas turbine engine is provided with four successive generators and four successive turbines. It is offered to introduce twice less air into the combustion chamber of this gas turbine engine than is necessary for a full combustion of the fuel. And further, downstream, it is offered to carry out a step-by-step post-combustion of the fuel when a working body is working in the turbine stages. It should be noted here that when ordinary high-energy gas turbine fuels are used, the gas temperature before the first turbine stage becomes impermissibly high, since the unburned half of the fuel cannot lower the temperature of the combustion products to an acceptable value.
    It is an object of the invention to provide a gas turbine engine which achieves a maximum complete stage combustion of high-energy fuels and atmospheric oxygen. It is envisaged that after the first high-pressure turbine stage, ordinary combustion products of high-energy fuels, in which, for example, the oxygen excess number is approximately 3, should be mixed with an additional parallel bypass flow of a fuel partially burned in a low-pressure combustion chamber due to lack of oxygen. After this mixture and a post-combustion reaction of the unburned fuel, the temperature of the working body flow increases again by a required value, for example to 1200 K. After working in the second turbine stage, the same mixing and post-combustion procedure is again offered before the third turbine stage.
    With such an engine construction, the degree of excess oxygen in the working body of an engine is continuously reduced from one stage to the other. For example, after the fourth turbine stage, the oxygen excess number can be 1, which means that almost all of the atmospheric oxygen burns. Then this flow of the working body is to be mixed with an additional stoichiometric bypass flow of the working body with the appropriate pressure and temperature from a corresponding low-pressure combustion chamber. As a result, the current temperature of the working body remains at a high level even before the last fifth turbine stage. As a result of the heat being supplied to the working body five times, the motor efficiency is approximately 80%. The temperature of exhaust gases, for example, rises to 1000 K. The engine output per unit of weight increases. The use of corresponding compressor stages with corresponding mass air consumption is provided for the generation of parallel air flows with corresponding pressure. The mass consumption of the additional parallel bypass flows of the working body can, for example, be a few percent of the main flow of the working body. This creates the prerequisites for an unimpeded redistribution of momentary workload between the main flow and the additional flow. This makes the combustion process in the combustion chambers, in contrast to known gas turbine engines, with intermediate heating of the working body during the combustion of high-energy fuels reliable and free of vibrations. 1 shows three possible variants of the specific arrangement of a low-pressure combustion chamber. All three variants have the prerequisites for the unrestricted redistribution of torque expenditure of the working body between the main flow of the working body and the additional bypass flow, which avoids the possibility of vibration combustion occurring in the combustion chambers of an engine. It is planned to use corresponding petroleum products - liquid gas and aviation turbine fuels, liquid gas, natural or shale gas - as high-energy fuels. The heat of combustion of such fuels is over 43,000 kJ / kg.
    The invention will be explained with reference to the accompanying drawings of embodiments of the invention. It shows:
    Fig. 1 shows the diagram of a gas turbine engine for generating power with different variants of the diversion of part of the working body in the turbine stages under appropriate pressure and
    Fig. 2 shows the scheme of a subsonic turbo aircraft engine with intermediate heating of the working body.
    1 is equipped with compressors 1, 2, 3, 4, 5 and turbines 6, 7, 8, 9, 10. A high pressure combustion chamber 12 is arranged between the high pressure compressor 5 and the high pressure turbine 6.
    The input of a low-pressure combustion chamber 11 is connected to the output of the compressor 4 and its output to the turbine stage 7. The outlet of a low-pressure combustion chamber 11 and the outlet of a compressor 3 are connected together to the inlet of the turbine stage 8, and the inlet of this chamber 11 is connected to the outlet of the turbine stage 7. The outlet of a low-pressure combustion chamber 14 is connected to the inlet of the turbine 9 and its inlet to the outlet of the compressor 2 and to the outlet of the turbine 8. The inlet of a low-pressure combustion chamber 15 is connected to the outlet of the compressor 1 and its outlet to the inlet of the turbine stage 10 with the appropriate pressure. Bypass air lines 16, 17, 18, 19 can be designed in the form of ring channels or can be divided into several parallel channels with their own low-pressure combustion chambers. A motor shaft 20 is connected to a power generator 21. In the high-pressure combustion chamber 11 there are injection nozzles for supplying high-energy gas turbine fuels. Each of the low pressure combustion chambers 12, 13, 14 and 15 has fuel injection injectors for burning the fuels in these chambers. The combustion chambers 11, 12 and 15 are also provided with flame lighters.
    The engine works in the following way:
    At the outlet from the high-pressure combustion chamber 11, hot gases have a temperature after the engine has started, which is not dangerous for continuous continuous operation of the engine. However, the excess air figure - approx. 3 - is high. Therefore, the air flows into the low pressure combustion chamber 12 through a bypass air line 16, for example about 4% in relation to the main flow, which runs continuously through the high pressure compressor 5. Then it is possible to ensure the combustion of high-energy fuels in the combustion chamber 12 under oxygen deficiency conditions. After the exit from the low-pressure combustion chamber 12, the fuel that is not completely burned mixes with the working body consumed in the high-pressure turbine 6, which has a good excess of oxygen. Then the fuel is afterburned in the afterburning passage 12A. A working body with a further increased temperature runs to the entrance of the turbine 7. This has a technical effect that cannot be ascertained in engines that use a staged post-combustion of low-heating fuels under low-oxygen conditions. Namely, the low-pressure combustion chamber 12 increases the temperature of the working body behind the high-pressure turbine stage 6, that is to say, which in this respect functions as a combustion chamber which is consequently connected to the high-pressure combustion chamber 11. At the same time, the redistribution of momentary expenditure of the working body between the above-mentioned chambers takes place analogously to the combustion chambers operating in parallel. This eliminates self-vibrations when burning fuels in some combustion chambers. The application of the technical effect mentioned in gas turbine engines, in which high-energy fuels are used, makes it possible to increase their efficiency to 80%. In the low-pressure combustion chamber 13, the fuel combustion takes place in a medium in which a substantial part of oxygen has already been burned. The air that enters through an air line 17 promotes the stabilization of the combustion process. Fresh air flows into the entrance of the low-pressure combustion chamber 14 through an air line 18. Together with the incoming fuel, it forms a central part of the torch (flame) inside the chamber 14. The main part of the working body runs closer to the walls of the combustion chamber 14. The working body consumed in the turbine stage 9 contains almost no unburned oxygen more. Therefore, stoichiometric combustion of fuel and air entering through an air line 19 is carried out in the low-pressure combustion chamber 15. The air consumption for the air line 19 is determined accordingly. The nominal operation of the engine takes place at a standard temperature of gases at the entrance to the turbine stages. The power is reduced due to the lowering of the temperature of the working body only before the last turbine stage 10, then before the penultimate stage and so on, at a stable rotational speed of a shaft 20. Several other embodiment variants of the invention are also possible. The low-pressure combustion chambers 13 and 14 can, for example, be designed and connected analogously to the combustion chamber 12. The combustion chamber 15 can also operate in conditions of lack of atmospheric oxygen. Instead of the last turbine stage 10 and an electricity generator 21, a free turbine stage with its own load can be used.
    A two-shaft subsonic aircraft engine has a multi-stage high-pressure compressor 23 and a single-stage high-pressure turbine 24 on a shaft 22. A multi-stage low-pressure compressor 26 with a fan wheel 27 is arranged on another shaft 25. Five low-pressure turbine stages 28 are arranged on the same shaft. The engine also has a high-pressure combustion chamber 29, a low-pressure combustion chamber 30 and a bypass air duct 31. At the entrance and exit of an annular duct 31, automatic blades 32 and 33 are fastened accordingly. Numerous blades 32 and 33 are evenly distributed over the circumference of the cross section of the channel 31, fastened in accordance with the diagram, and have the possibility of passing an air flow in only one direction - from the compressor to the turbine. When the air flow attempts to return, these wings automatically close channel 31.
    Both combustion chambers 29 and 30 are provided with injection nozzles for the supply of the flight petroleum and with flame lighters. The engine is designed and engineered for maximum calculated operation on a flight at 11,000 meters and at a flight speed of Mach 0.8, which differs from a customary development or design of subsonic engines in which the maximum operation is based on ground takeoff conditions is determined.
    The engine is operated in the following way:
    When departing from an airport, a high-pressure combustion chamber 29 is first started. Wings 32 and 33 prevent movement of the working body in the direction from the turbine to the compressor. The shaft revolutions 25 are greatly reduced. The low pressure combustor 30 is then started and the revolutions of a low pressure compressor 26 are increased. Now all of the air is not taken out by a high pressure compressor 23. The wings 32 and 33 open automatically, a certain part of the air runs through the bypass channel 31, whereby the combustion process in the low-pressure combustion chamber 30 is stabilized. The temperature of the heating gases at the exit from the combustion chambers 29 and 30 is not maintained at the highest level on departure. As a result, the turbine blades are not overheated, the stated revolutions of the compressors 26 and 23 are reduced. This makes it possible to increase the flight efficiency during departure, to utilize the metal strength of the turbine blades to a greater extent and to design the engine with a high bypass. As the altitude increases, the temperature and air density at the engine entrance decrease. The temperature of the hot gases at the outlet of the combustion chambers 29 and 30 is increased, thereby regulating the fuel consumption, since the temperature of the compressor air provided for turbine cooling also drops. And it is only at a height of 11,000 meters at a temperature at the engine entrance, such as 244 K, that the engine is brought into maximum calculated operation. This makes it possible to develop an aircraft engine with a high reserve of power during flight and thus to increase flight safety. It is envisaged to reduce engine operation by reducing the temperature of the gases at the outlet of the low-pressure combustion chambers 30. It is also provided that the constant temperature of the working body in front of the high-pressure turbine stage 24 is kept unchangeable over a wide thrust range. This is intended to ensure reliable burning in the combustion chamber 30. Fuel savings are achieved by increasing the heat and flight efficiency.
    In embodiment variants, spray nozzles can be set up for supplying fuel in the channel 31. A gearbox can also be used in the motor.
    权利要求:
    Claims (1)
    [1]
    claims
    1.Gas turbine engine with two combustion chambers and a diversion channel with at least one low-pressure compressor (4), a high-pressure compressor (5) and a high-pressure turbine (6), a high-pressure combustion chamber (11) arranged between them, a fuel delivery system, a low-pressure combustion chamber (12) with a fuel delivery system and a bypass duct (16) for part of the working medium, which leads from the outlet of the low-pressure compressor (4) past the high-pressure compressor (5), the high-pressure combustion chamber (11), the high-pressure turbine (6), to the inlet of the low-pressure combustion chamber (12), and the outlet of this chamber is connected to the inlet of a low-pressure turbine (7), whereby a flow redistribution between a main flow of the working medium and an additional flow through the bypass duct (16) is achieved, the compressors (4, 5) and the turbines (6 , 7) are arranged on a shaft (20), characterized in that d he mentioned redistribution of the working medium flow between a main flow and an additional diversion can take place without hindrance.
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    同族专利:
    公开号 | 公开日
    CH708180A4|2013-10-03|
    UA103413C2|2013-10-10|
    RU2012115610A|2013-08-10|
    DE112013003321T5|2015-11-26|
    WO2013142941A1|2013-10-03|
    GB201418548D0|2014-12-03|
    GB2515947A|2015-01-07|
    GB2515947B|2020-07-01|
    CA2870615A1|2013-10-03|
    US20150135725A1|2015-05-21|
    引用文献:
    公开号 | 申请日 | 公开日 | 申请人 | 专利标题
    US3677012A|1962-05-31|1972-07-18|Gen Electric|Composite cycle turbomachinery|
    FR2270450B1|1974-03-29|1976-10-08|Snecma|
    US4068471A|1975-06-16|1978-01-17|General Electric Company|Variable cycle engine with split fan section|
    US4054030A|1976-04-29|1977-10-18|General Motors Corporation|Variable cycle gas turbine engine|
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    GB2288640B|1994-04-16|1998-08-12|Rolls Royce Plc|A gas turbine engine|
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    US7584598B2|2005-08-10|2009-09-08|Alstom Technology Ltd.|Method for operating a gas turbine and a gas turbine for implementing the method|GB201518929D0|2015-10-27|2015-12-09|Rolls Royce Plc|Gas turbine engine|
    GB2547674A|2016-02-25|2017-08-30|Rolls Royce Plc|Gas turbine engine|
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    US20190017437A1|2017-07-13|2019-01-17|General Electric Company|Continuous detonation gas turbine engine|
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    CN113323769A|2021-06-07|2021-08-31|北京航空航天大学|Variable cycle engine configuration based on multi-duct intake interstage combustion chamber|
    法律状态:
    2014-12-31| NV| New agent|Representative=s name: ASSOZIATION WEISSRUSSLNNEN IN DER SCHWEIZ, CH |
    2017-03-31| AZW| Rejection (application)|
    2020-05-29| AEN| Modification of the scope of the patent|Free format text: :DIE PATENTANMELDUNG IST AUFGRUND DES WEITERBEHANDLUNGSANTRAGS VOM 31.03.2017 REAKTIVIERT WORDEN |
    2020-10-30| PL| Patent ceased|
    优先权:
    申请号 | 申请日 | 专利标题
    BY20120506|2012-03-30|
    PCT/BY2013/000002|WO2013142941A1|2012-03-30|2013-03-26|Gas-turbine engine|
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