WO2007069309A1 - Turbine à gaz - Google Patents

Turbine à gaz Download PDF

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Publication number
WO2007069309A1
WO2007069309A1 PCT/JP2005/022940 JP2005022940W WO2007069309A1 WO 2007069309 A1 WO2007069309 A1 WO 2007069309A1 JP 2005022940 W JP2005022940 W JP 2005022940W WO 2007069309 A1 WO2007069309 A1 WO 2007069309A1
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WO
WIPO (PCT)
Prior art keywords
fuel
gas turbine
air
burner
flow rate
Prior art date
Application number
PCT/JP2005/022940
Other languages
English (en)
Japanese (ja)
Inventor
Satoshi Dodo
Susumu Nakano
Hiroyuki Shiraiwa
Original Assignee
Hitachi, Ltd.
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Hitachi, Ltd. filed Critical Hitachi, Ltd.
Priority to JP2007550041A priority Critical patent/JPWO2007069309A1/ja
Priority to PCT/JP2005/022940 priority patent/WO2007069309A1/fr
Publication of WO2007069309A1 publication Critical patent/WO2007069309A1/fr

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Classifications

    • 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
    • F02C9/00—Controlling gas-turbine plants; Controlling fuel supply in air- breathing jet-propulsion plants
    • F02C9/26—Control of fuel supply
    • F02C9/28—Regulating systems responsive to plant or ambient parameters, e.g. temperature, pressure, rotor speed
    • 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
    • F02C7/08—Heating air supply before combustion, e.g. by exhaust gases
    • 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
    • F02C7/22—Fuel supply systems
    • 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
    • F02C9/00—Controlling gas-turbine plants; Controlling fuel supply in air- breathing jet-propulsion plants
    • F02C9/26—Control of fuel supply
    • F02C9/32—Control of fuel supply characterised by throttling of fuel
    • F02C9/34—Joint control of separate flows to main and auxiliary burners
    • 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
    • F02C9/00—Controlling gas-turbine plants; Controlling fuel supply in air- breathing jet-propulsion plants
    • F02C9/26—Control of fuel supply
    • F02C9/40—Control of fuel supply specially adapted to the use of a special fuel or a plurality of fuels
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/28—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
    • F23R3/34—Feeding into different combustion zones
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/28—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
    • F23R3/36—Supply of different fuels

Definitions

  • the present invention relates to a gas turbine apparatus, and more particularly to a gas turbine apparatus suitable for use with a fuel whose calorific value varies.
  • a gas turbine combustor used in a gas turbine apparatus employs a premixed combustion method in which fuel and air are mixed and supplied in a combustor component called a premixer before flowing into a combustion chamber.
  • a premixed combustion type combustor combustion air and fuel can be mixed and burned in advance so that the fuel becomes leaner than the stoichiometric mixture ratio, so there is no local high temperature region and the flame temperature is low. Nitrogen oxide (Nx) emissions are low.
  • a regenerative gas turbine apparatus that recovers exhaust heat from turbine exhaust and improves thermal efficiency by preheating the compressor discharge air with a regenerative heat exchanger and sending it to the combustor in order to improve power generation efficiency.
  • many regenerative gas turbines called small micro turbines have a regeneration efficiency exceeding 90%.
  • the combustor inlet air temperature becomes higher than 600 ° C, so that the fuel self-ignition exceeds the ignition point temperature or the flame in the combustion chamber is preliminarily generated. There is a high risk of backfire that flows back into the mixer.
  • Patent Document 1 International Publication WO2005 / 059442A1
  • An object of the present invention is to provide a gas turbine apparatus capable of performing stable combustion without being blown out or overheated in a gas turbine apparatus using a fuel with a variable calorific value as a main fuel and a fuel with a stable calorific value as an auxiliary fuel. Is to provide.
  • the present invention provides a gas turbine device that uses a fuel with a variable calorific value as a main fuel, and uses a fuel with a stable calorific value as an auxiliary fuel, Fuel and air so as to dam the first panner flame at the axial position in the combustor corresponding to the tip of the flame of the first panner and the first panner that jets fuel and air into the combustion chamber
  • a combustor provided with a second burner for generating a circulating flow of fuel and air in the combustion chamber, and a fuel with a stable calorific value at a predetermined flow rate calculated by the gas turbine air flow rate.
  • a control means for changing the flow rate of fuel supplied to the second burner is provided.
  • the control means further calculates an operating rotational speed at which the required load can be output from the required load and the intake air temperature signal to obtain a predetermined rotational speed increase rate. Therefore, the operation speed calculation controller that generates the reference operation speed command value and the reference value of the turbine outlet temperature with reference to the turbine outlet gas temperature signal are used to calculate the correction amount for the operating speed. And a correction operation speed control calculator for generating an operation speed correction amount command value, and adding the reference operation speed command value and the operation speed correction amount command value to obtain the operation speed command value. This is what happens.
  • the present invention provides a gas turbine device that uses a fuel with a variable calorific value as a main fuel and a fuel with a stable calorific value as an auxiliary fuel.
  • Fuel and air are ejected at the axial position in the combustor corresponding to the tip of the flame of the first panner and the first panner that ejects fuel and air so as to dampen the first panner flame.
  • a combustor provided with a second burner that generates a circulating flow of fuel and air in the combustion chamber, and a fuel with a stable calorific value calculated by the gas turbine air flow rate is supplied to the first burner.
  • the fuel whose fluctuating calorific value is supplied is supplied to the second burner, and supplied to the second burner by a correction amount calculated using the difference between the reference value of the turbine outlet gas temperature and the actual turbine outlet gas temperature. Control hand to change the fuel flow rate In which it was to obtain Bei the.
  • FIG. 1 is an overall configuration diagram showing a configuration of a gas turbine apparatus according to a first embodiment of the present invention.
  • FIG. 2 is a longitudinal section showing a configuration of a combustor used in the gas turbine apparatus according to the first embodiment of the present invention.
  • FIG. 3 is an explanatory diagram of a result of a chemical reaction simulation performed on a slow combustion reaction of a lean air-fuel mixture in the gas turbine apparatus according to the first embodiment of the present invention.
  • FIG. 4 An amount capable of obtaining a combustion efficiency of 99% or more when the residence time in the secondary combustion region is 35 ms in the combustor used in the gas turbine apparatus according to the first embodiment of the present invention. It is explanatory drawing of the conditions of ratio and mixing average temperature.
  • FIG. 5 is an explanatory diagram of the operating rotational speed at which the required load can be output in the case of each intake air temperature in the gas turbine apparatus according to the first embodiment of the present invention.
  • FIG. 6 is an explanatory diagram regarding the flow rate of auxiliary fuel with a stable calorific value to be supplied in the case of each intake air temperature in the gas turbine apparatus according to the first embodiment of the present invention.
  • FIG. 7 is an overall configuration diagram showing a configuration of a gas turbine apparatus according to a second embodiment of the present invention.
  • FIG. 8 is an overall configuration diagram showing a configuration of a gas turbine apparatus according to a third embodiment of the present invention.
  • FIG. 1 is an overall configuration diagram showing a configuration of a gas turbine apparatus according to a first embodiment of the present invention.
  • the gas turbine apparatus shown in the present embodiment includes a generator / motor 2 controlled and driven by a power converter 1, a compressor 3 and a turbine 6 directly connected to a generator Z motor 2, and a turbine 6
  • This is a regenerative gas turbine device comprising a regenerative heat exchanger 4 and a combustor 5 for exchanging heat between the exhaust gas and the high pressure air discharged from the compressor 3 to preheat the high pressure air.
  • the generator / motor 2, the compressor 3 and the turbine 6 are electrically controlled in accordance with the operation speed command value 43 obtained from the operation speed calculation controller 31 and started according to a predetermined speed schedule. Driven. Air 11 is compressed after intake air temperature is measured by intake thermometer 7.
  • the combustor inlet air temperature meter 8 measures the combustor inlet air temperature, and then flows into the combustor 5.
  • the combustor 5 includes an auxiliary fuel 12 responsible for raising the temperature at startup and ensuring combustion stability during load operation, and a main fuel 13 for generating a load, respectively, and an auxiliary fuel flow rate adjusting valve 21 and a main fuel flow rate adjusting valve. These fuels react with the above-mentioned high-temperature preheated air to become higher-temperature combustion gas and flow into the turbine 6.
  • the high-temperature combustion gas that has flowed into the turbine 6 drives the turbine, and after the turbine outlet thermometer 9 measures the turbine outlet combustion gas temperature, the regenerative heat exchanger 4 exchanges heat with the high-pressure air, and finally exhausts. It is discharged out of the system as gas 14.
  • the control device is roughly divided into an operation speed calculation controller 31, an auxiliary fuel flow rate calculation controller 33, and a main fuel flow rate calculation controller 34.
  • the operation speed, auxiliary fuel flow rate It is responsible for controlling the fuel flow rate.
  • the operation rotational speed calculation controller 31 is an operation capable of outputting the required load 40 from the required load 40 and the intake air temperature signal 46 using an expression obtained by functionalizing the atmospheric temperature characteristic and the rotational speed characteristic of the gas turbine device.
  • the engine speed is calculated, and the engine speed command value 43 is generated according to the predetermined speed increase rate.
  • the power converter 1 adjusts the current and voltage based on the operation speed command value 43 so that the speed of the generator / motor 2 becomes a predetermined speed, and the actual speed is changed to the actual speed. Output as signal 44.
  • the auxiliary fuel flow rate calculation controller 33 is based on the gas turbine air flow rate calculated from the actual rotational speed signal 44 obtained from the power converter 1 and the intake air temperature signal 46, and the combustor inlet air temperature signal 47.
  • the flow rate of the auxiliary fuel 12 having a stable calorific value to be supplied is obtained, and the opening degree of the auxiliary fuel flow rate adjustment valve 21 is controlled by the auxiliary fuel flow rate input command 49.
  • the main fuel flow rate calculation controller 34 includes a reference main fuel flow rate calculator 35 and a corrected main fuel flow rate calculator 36, and the reference main fuel flow rate calculators 35 and 36 respectively obtain the reference main fuel flow rate calculators 35 and 36.
  • the opening amount of the main fuel flow rate adjusting valve 22 is controlled based on the main fuel flow rate input command value 52 by adding the amount input command 50 and the main fuel flow rate correction command 51.
  • the reference main fuel flow rate calculator 35 is a main fuel to be supplied from the gas turbine bin air flow rate calculated from the actual rotational speed signal 44 obtained from the power converter 1 and the intake air temperature signal 46 and the combustor inlet air temperature signal 47. 13 standards
  • the basic flow rate input command 50 is output.
  • the corrected main fuel flow rate calculator 36 calculates the main fuel flow rate to be corrected from the deviation between the actual gas turbine output signal 45 obtained from the power converter 1 and the required load 40, and outputs the main fuel flow rate input correction command 51.
  • FIG. 2 is a longitudinal section showing the configuration of the combustor used in the gas turbine apparatus according to the first embodiment of the present invention.
  • the combustor includes a combustor liner 103 having a circular cross section that forms a combustion chamber 102, a liner cap 104 that closes the upstream side of the combustor liner 103, and an activation starter formed at the center of the liner cap 104.
  • a plurality of second partners 108 formed so as to penetrate through the peripheral wall of the combustor liner 103.
  • the starter pan 105 is a burner that takes charge of the start-up and warm-up operation from gas turbine ignition and partial load operation up to 80%, for example, and is supplied with the auxiliary fuel 12 having a stable calorific value.
  • the starter pan 105 is formed concentrically with the combustor liner 103, and has a downstream end located in the center of the liner cap 104 at its center and an upstream end extending through the center of the end cover 106. It has a fuel nozzle 109 for starting.
  • a first fuel injection hole 110 is provided at the downstream end of the startup fuel nozzle 109, and a first air introduction tube 111 concentric with the startup fuel nozzle 109 is formed on the outer periphery of the startup fuel nozzle 109 with a gap.
  • the swirl vane 112 is provided in this gap.
  • the downstream side of the first air introduction tube 111 opens from the liner cap 104 to the combustor liner 103, and the upstream side is closed by the end cover 106.
  • a first air introduction hole 113 is provided near the end cover 106 of the first air introduction cylinder 111.
  • the combustor liner 103 is connected to a transition piece (not shown) on the downstream side via an elastic seal member 114, and guides high-temperature combustion gas generated in the combustion chamber to the turbine 6.
  • a transition piece not shown
  • the gas temperature distribution at the combustor outlet is smoothed.
  • a dilution air hole 115 is provided for this purpose.
  • a stopper for fixing the position of the combustor liner 103 and a film cooling slot for ensuring reliability are complicated, and the illustration is omitted.
  • the plurality of second burners 108 are provided so as to penetrate through the second air introduction hole 116 provided in the peripheral wall of the combustor liner 103 and the peripheral wall of the outer cylinder 107 facing the second air introduction hole 116, respectively.
  • the second fuel nozzle 117 is also supplied with the main fuel 13 whose calorific value fluctuates.
  • Combustion air is compressed by the compressor 3 and heated by the regenerative heat exchanger 4, and is guided leftward in the figure from the gap between the combustor liner 103 and the outer cylinder 107 on the right side in the figure.
  • the A part of the guided combustion air passes through the dilution hole 115 and the second air introduction hole 116 and is introduced into the combustion chamber 102 in the combustor liner 103, and the rest from the first air introduction hole 113.
  • the air is injected from the liner cap 104 into the combustion chamber 102.
  • a circulation area is formed on the downstream side of
  • the auxiliary fuel 12 having a stable calorific value is ejected from the starting fuel nozzle 109 into the combustion chamber 102, and the main fuel 13 having a fluctuating calorific value is introduced from the second fuel nozzle 117 into the combustion chamber 102. Erupted. All fuel is injected directly into the combustion chamber, and there is no mixture of fuel and air outside the combustion chamber.
  • the starter pan 105 affects the combustion stability of the entire combustor and is used in a wide range from ignition to start-up to a partial load of 80%.
  • a diffusion combustion method is adopted. ing.
  • the main fuel 13 having a variable calorific value is injected radially from the second fuel nozzle 117 installed at the same position into the air ejected from the secondary air introduction hole 116 into the combustion chamber 102.
  • the main fuel 13 immediately after being injected from the second fuel nozzle 117 has a large flow velocity of the air injected through the second air introduction hole 116 and a strong shearing with the surrounding combustion gas, so that the combustion reaction does not occur.
  • the air ejected from the second air introduction holes 116 at three circumferential directions collides with each other in the vicinity of the central axis of the combustion chamber 102 to form a stagnation region, and upstream and downstream of the second air introduction holes 116. Each forms a circulation region. In these circulating flow regions, the flow velocity is reduced, and the condition that the propagating flame can be sufficiently maintained is maintained. Therefore, the fuel ejected from the second fuel nozzle 117 starts a combustion reaction in the circulating flow. At this time, since the fuel injected from the second fuel nozzle and the air injected from the second air introduction hole are a lean mixture at the time of starting the reaction, they depended on the diffusion of heat to the mixture. It adopts a reaction mode controlled by a slow oxidation reaction, and realizes low NOx combustion that does not produce a local high temperature part.
  • the combustor has the first burner (starter burner) 105 that ejects fuel and air into the combustion chamber, and the combustion corresponding to the flame tip of the first burner. And a second burner 108 for injecting fuel and air into the combustion chamber to generate a circulating flow of fuel and air so as to block the first burner flame at an axial position in the chamber.
  • the mixed flow of fuel and air ejected from the burner collides in the combustion chamber, generates a circulatory flow with strong turbulence, and mixes with the combustion gas from the first burner in contact with a wide contact area. Slow combustion without generating a local high-temperature region in the combustion chamber can be performed, and stable combustion can be performed without causing backfire or self-ignition.
  • FIG. 3 is an explanatory diagram of a result of a chemical reaction simulation performed on the slow combustion reaction of the lean air-fuel mixture in the gas turbine apparatus according to the first embodiment of the present invention.
  • the horizontal axis shows the distance from the second air introduction hole 116 to the dilution hole 115 normalized by the total length of the combustor liner 103. That is, in the configuration of the combustor shown in FIG. 2, when the total length of the combustor liner shown in FIG. 2 is L and the distance from the second air introduction hole 116 to the dilution hole 115 is X, the configuration shown in FIG. The horizontal axis shows X / L. In the combustor 5 shown in FIG. 2, the position of the dilution hole 115 corresponds to the position of 0.668. [0035] FIG.
  • FIG. 3 shows a predicted distribution diagram by reaction calculation of the carbon monoxide concentration and the combustion gas temperature in the axial direction of the combustor shown in FIG.
  • the lower curve shows the change in combustion gas temperature Tg (° C) along the combustion gas flow direction in the combustor
  • the upper curve shows the carbon monoxide concentration along the combustion gas flow direction. Shown as an indicator of reaction.
  • the lean air-fuel mixture formed by the fuel and air from the second burner 108 is mixed with the combustion gas of the starting paner 105 in the stagnation region near the central axis of the combustor liner 103, and the lean air-fuel mixture with an average mixing temperature of 866 ° C is mixed. It becomes.
  • this lean air-fuel mixture gradually generates heat and rises in temperature while the fuel is slowly oxidized to generate carbon monoxide, and heat generation occurs rapidly after the carbon monoxide concentration reaches the maximum value. As a result, the carbon monoxide concentration decreases.
  • the residence time required during this time is about 30 ms when the mixing average temperature of the combustor 5 shown in Fig. 2 is 866 ° C.
  • a combustion efficiency of 99% or more is obtained when the residence time in the secondary combustion region is 35 ms.
  • the conditions of the equivalent ratio and the mixing average temperature will be described.
  • FIG. 4 shows an equivalence ratio and a mixing average in which a combustion efficiency of 99% or more is obtained when the residence time in the secondary combustion region is 35 ms in the combustor used in the gas turbine apparatus according to the first embodiment of the present invention. It is explanatory drawing of the conditions of temperature.
  • FIG. 4 shows the equivalent ratio defined by the fuel and air from the second burner 108 and the second burner 108 when the residence time in the region from the second air introduction hole 116 to the dilution hole 115 is 35 ms.
  • the average mixing temperature of the fuel and air from the fuel and the combustion gas from the starting burner 105 is shown under the conditions for obtaining a high combustion efficiency of 99% or more.
  • combustion efficiency will be ensured, but if the mixing average temperature is increased too much or the equivalence ratio is increased, the reaction will proceed rapidly and NOx emissions will increase. Further, if the residence time is made longer, combustion efficiency can be obtained even at an equivalent ratio that is leaner than the conditions shown in FIG. 3, but this is not preferable because the length of the combustor 5 becomes longer and the combustor becomes larger.
  • the equivalence ratio defined by the fuel and air from the second burner 108 is determined from the combustible component molar flow rate of the main fuel 13 ejected from the second fuel nozzle 117 and the second air introduction hole 116. It is a value defined only by the ratio of the oxygen molar flow rate in the jetted air. Even if the component composition of the main fuel 13 supplied to the second burner 108 changes and the calorific value per unit flow rate fluctuates, the calorific value of the main fuel 13 supplied to the second burner 108 per hour If the sum is constant, the molar flow rate of combustible components in the main fuel 13 supplied at that time is constant, and the equivalent ratio defined by the fuel and air from the second burner 108 does not change.
  • the average mixing temperature Tmix required to obtain high combustion efficiency for a lean mixture with an equivalence ratio ⁇ defined by the fuel and air from the second burner 108 is equal to the fuel from the second burner 108.
  • the fuel flow rate from the second burner 108 is small compared to the air flow rate of the starter burner 105 and the second burner, so the mixing average temperature Tmix is It is determined by the air flow rate of the starter pan 105, the flow rate of the auxiliary fuel 12 supplied to the starter panner, and the air flow rate ejected from the second air introduction hole 116.
  • the second burner 108 has a constant equivalence ratio ⁇ defined by the fuel and air from the second burner 108.
  • the value is the value, the fuel flow rate of the auxiliary fuel 12 to be supplied to the starting burner 105 necessary for obtaining high combustion efficiency becomes a certain value.
  • the auxiliary fuel flow rate calculation controller 33 includes the actual turbine speed signal 44 obtained from the power converter 1 and the gas turbine air flow rate calculated from the intake air temperature signal 46, and the combustor inlet.
  • the flow rate of the auxiliary fuel 12 having a stable calorific value to be supplied is obtained from the air temperature signal 47, and the opening degree of the auxiliary fuel flow rate adjusting valve 21 is controlled by the auxiliary fuel flow rate input command 49.
  • the reference main fuel flow rate calculator 35 of the main fuel flow rate calculation controller 34 is a gas turbine air flow rate calculated from the actual rotational speed signal 44 obtained from the power converter 1 and the intake air temperature signal 46, and the combustor inlet.
  • the standard flow rate of the main fuel 13 to be supplied is calculated from the air temperature signal 47 and output as the reference main fuel flow rate input command 50.
  • the corrected main fuel flow rate calculator 36 calculates the main fuel flow rate to be corrected for the deviation force between the actual gas turbine output signal 45 obtained from the power converter 1 and the required load 40, and the main fuel flow rate input correction command. 51 is output. Further, by adding the reference main fuel flow rate input command 50 and the main fuel flow rate input correction command 51 obtained from the respective main fuel flow rate calculators 35 and 36, the main fuel flow rate control valve 22 To control the opening degree.
  • fuel with fluctuating calorific value is used as the main fuel
  • fuel with stable calorific value is used as the auxiliary fuel
  • stable combustion that satisfies the required load can be performed, and N0x emissions are reduced. It will be able to burn less
  • FIG. 5 is an explanatory diagram of the operating rotational speed at which the required load can be output in the case of each intake air temperature in the gas turbine apparatus according to the first embodiment of the present invention.
  • FIG. 5 shows the operation rotational speed at which the required load 40 can be output in the gas turbine apparatus shown in the present embodiment at the intake air temperature—10 ° C, 5 ° C, 15 ° C, 25 ° C, 40 ° C. Show the case.
  • the horizontal axis shows the operating speed normalized by the rated speed
  • the vertical axis shows the load that can be output, normalized by the rated output.
  • FIG. 4 is an explanatory diagram regarding the flow rate of auxiliary fuel with a stable calorific value to be supplied in the case of each intake air temperature in the gas turbine apparatus according to the first embodiment.
  • FIG. 6 shows the amount of auxiliary fuel 12 with a stable calorific value to be supplied to the starter pan 105 when the intake air temperature is 10 ° C, 5 ° C, 15 ° C, 25 ° C, and 40 ° C.
  • the auxiliary fuel injection command value 49 is shown as a function of the engine speed.
  • the horizontal axis shows the operating speed normalized by the rated speed
  • the vertical axis shows the auxiliary fuel injection command value 49.
  • stable combustion satisfying the required load can be performed using the fuel whose fever amount fluctuates as the main fuel and the fuel whose calorific value is stable as the auxiliary fuel, In addition, the amount of NOx emissions is low and combustion can be performed.
  • FIG. 7 is an overall configuration diagram showing the configuration of the gas turbine apparatus according to the second embodiment of the present invention.
  • the same reference numerals as those in FIG. 1 denote the same parts.
  • the basic apparatus configuration of the present embodiment is almost the same as that of the embodiment shown in FIG. 1, and a compressor 3 directly connected to a generator / motor 2 controlled and driven by a power converter 1. And a regenerative gas turbine device comprising a regenerative heat exchanger 4 and a combustor 5 for exchanging heat between the exhaust gas of the turbine 6, the exhaust gas of the turbine 6 and the high pressure air discharged from the compressor 3 and preheating the high pressure air.
  • the operation speed calculation controller 31 is an operation that can output the required load 40 from the required load 40 and the intake air temperature signal 46, using an expression obtained by functionalizing the atmospheric temperature characteristic and the rotational speed characteristic of the gas turbine device. Calculate the rotation speed and generate the reference operation rotation speed command value 41 according to the predetermined rotation speed increase rate.
  • a correction operation rotational speed control calculator 32 is provided. Compensation operation The rotational speed control calculator 32 refers to the turbine outlet gas temperature signal 48 to calculate a correction amount for the operating speed based on the deviation from the reference value of the turbine outlet temperature, and calculates the operating speed correction amount command value 42. appear. Then, the operation speed command value 43 is generated by adding the reference operation speed command value 41 and the operation speed correction amount command value 42.
  • the temperature of the combustion gas flowing into the regenerative heat exchanger 4 can be adjusted to the optimum operating condition of the regenerative heat exchanger.
  • stable combustion satisfying the required load can be performed by using the fuel whose calorific value fluctuates as the main fuel and using the fuel whose calorific value is stable as the auxiliary fuel, In addition, combustion with low NOx emissions is possible. In addition, reliability can be improved.
  • FIG. 8 is an overall configuration diagram showing the configuration of the gas turbine apparatus according to the third embodiment of the present invention.
  • the same reference numerals as those in FIG. 1 denote the same parts.
  • the compressor 3 and the turbine 6 directly connected to the generator / motor 2 controlled and driven by the power converter 1, the gas turbine device including the combustor 5, and the exhaust gas. 14
  • This is a so-called combined heat and power facility consisting of a waste heat recovery device 10 that recovers waste heat from 4 and uses it externally.
  • the corrected main fuel flow rate calculation controller 36 does not calculate the deviation force correction amount between the actual gas turbine output signal 45 obtained from the power converter 1 and the required load 40. Based on the deviation from the turbine outlet temperature reference value with reference to the turbine outlet gas temperature signal 48, the main fuel flow rate to be corrected is calculated and the main fuel flow rate input correction command 51 is output as shown in FIG. Different from the embodiment of FIG.
  • the output of the gas turbine device is determined by the air flow rate of the gas turbine device and the temperature of the turbine inlet. Therefore, by adjusting the turbine outlet gas temperature to the target value, the output of the gas turbine device is controlled.
  • the main fuel flow rate input correction command 51 is calculated from the deviation between the actual gas turbine output signal 45 obtained from the power converter 1 and the required load 40, and the turbine outlet gas temperature signal 48 is referred to. Therefore, calculating the correction amount from the deviation from the reference value of the turbine outlet temperature is almost equivalent, and the control configuration of this embodiment can be made cheaper.
  • stable combustion satisfying the required load can be performed by using the fuel whose calorific value fluctuates as the main fuel and using the fuel whose calorific value is stable as the auxiliary fuel.
  • the combustor shown in Fig. 2 it is possible to perform combustion with a small amount of N0x emissions. Further, the apparatus cost can be reduced.
  • the first panner operated by the fuel having a stable calorific value is stable with respect to the gas turbine air flow rate supplied to the combustor.
  • the fuel is supplied to satisfy the gas turbine output, so the calorific value fluctuates. Regardless of the combustible component concentration of the fuel, it is stable without blowing off or overheating, and the amount of N0x emission is low and combustion. Can be performed.

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  • General Engineering & Computer Science (AREA)

Abstract

La présente invention vise : une turbine à gaz qui emploie, comme combustible principal, un combustible à pouvoir calorifique variable et, comme combustible auxiliaire, un combustible à pouvoir calorifique stabilisé, et qui garantit une combustion stabilisée sans souffle et sans surchauffe ; une unité de combustion (5) qui comporte un premier brûleur (105) servant à injecter le combustible et l’air dans une chambre de combustion, et un deuxième brûleur (108) servant à injecter le combustible et l’air dans le sens axial de l’unité de combustion correspondant à l’extrémité distale de la flamme du premier brûleur de sorte que la flamme du premier brûleur soit interceptée, ce qui produit un courant de combustible et d’air dans la chambre de combustion ; un gestionnaire de débit de combustible auxiliaire (33) et un gestionnaire de débit de combustible principal (34) fournissent respectivement du combustible ayant un pouvoir calorifique stabilisé au premier brûleur suivant un débit prédéterminé calculé par le débit d’air de turbine à gaz, et du combustible présentant un pouvoir calorifique variable au deuxième brûleur, puis font varier le débit de combustible alimenté au deuxième brûleur sur la base d’une valeur corrigée calculée sur la base de la différence entre une puissance nominale et une puissance réelle de la turbine à gaz.
PCT/JP2005/022940 2005-12-14 2005-12-14 Turbine à gaz WO2007069309A1 (fr)

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JP2007550041A JPWO2007069309A1 (ja) 2005-12-14 2005-12-14 ガスタービン装置
PCT/JP2005/022940 WO2007069309A1 (fr) 2005-12-14 2005-12-14 Turbine à gaz

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PCT/JP2005/022940 WO2007069309A1 (fr) 2005-12-14 2005-12-14 Turbine à gaz

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WO2007069309A1 true WO2007069309A1 (fr) 2007-06-21

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Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JP2009144551A (ja) * 2007-12-12 2009-07-02 Hitachi Ltd 再生型ガスタービン及び再生型ガスタービンの燃料制御方法
JP2010285955A (ja) * 2009-06-12 2010-12-24 Mitsubishi Heavy Ind Ltd ガスタービンの制御装置及び発電システム
EP2261487A4 (fr) * 2008-11-28 2015-06-17 Mitsubishi Hitachi Power Sys Controleur pour turbine a gaz

Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JPH11324714A (ja) * 1998-05-11 1999-11-26 Toshiba Corp ガスタービンプラントおよびその運転方法
JP2003065083A (ja) * 2001-08-23 2003-03-05 Tokyo Gas Co Ltd 発電装置及びその運転制御方法
JP2003065084A (ja) * 2001-08-29 2003-03-05 Ishikawajima Harima Heavy Ind Co Ltd バイオガスを燃料とするガスタービン発電装置
US20030056517A1 (en) * 2001-09-26 2003-03-27 Siemens Westinghouse Power Corporation Apparatus and method for combusting low quality fuel
WO2005059442A1 (fr) * 2003-12-16 2005-06-30 Hitachi, Ltd. Dispositif de combustion de turbine a gaz

Family Cites Families (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JP2805892B2 (ja) * 1989-09-29 1998-09-30 日産自動車株式会社 発電機用ガスタービンの制御装置
JP3178055B2 (ja) * 1992-01-13 2001-06-18 株式会社日立製作所 ガスタービン燃焼器の制御装置及びガスタービン
JP4126973B2 (ja) * 2002-07-01 2008-07-30 株式会社Ｉｈｉ ガスタービン制御システム
JP2005218163A (ja) * 2004-01-27 2005-08-11 Kawasaki Heavy Ind Ltd タービン発電装置およびその自立運転方法

Patent Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JPH11324714A (ja) * 1998-05-11 1999-11-26 Toshiba Corp ガスタービンプラントおよびその運転方法
JP2003065083A (ja) * 2001-08-23 2003-03-05 Tokyo Gas Co Ltd 発電装置及びその運転制御方法
JP2003065084A (ja) * 2001-08-29 2003-03-05 Ishikawajima Harima Heavy Ind Co Ltd バイオガスを燃料とするガスタービン発電装置
US20030056517A1 (en) * 2001-09-26 2003-03-27 Siemens Westinghouse Power Corporation Apparatus and method for combusting low quality fuel
WO2005059442A1 (fr) * 2003-12-16 2005-06-30 Hitachi, Ltd. Dispositif de combustion de turbine a gaz

Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JP2009144551A (ja) * 2007-12-12 2009-07-02 Hitachi Ltd 再生型ガスタービン及び再生型ガスタービンの燃料制御方法
EP2261487A4 (fr) * 2008-11-28 2015-06-17 Mitsubishi Hitachi Power Sys Controleur pour turbine a gaz
JP2010285955A (ja) * 2009-06-12 2010-12-24 Mitsubishi Heavy Ind Ltd ガスタービンの制御装置及び発電システム

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