JPS61140722A - Gas turbine combustor - Google Patents

Gas turbine combustor

Info

Publication number
JPS61140722A
JPS61140722A JP26170284A JP26170284A JPS61140722A JP S61140722 A JPS61140722 A JP S61140722A JP 26170284 A JP26170284 A JP 26170284A JP 26170284 A JP26170284 A JP 26170284A JP S61140722 A JPS61140722 A JP S61140722A
Authority
JP
Japan
Prior art keywords
combustion
fuel
air
main
pilot
Prior art date
Legal status (The legal status 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 status listed.)
Pending
Application number
JP26170284A
Other languages
Japanese (ja)
Inventor
Masahiko Yamada
正彦 山田
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Toshiba Corp
Original Assignee
Toshiba Corp
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 Toshiba Corp filed Critical Toshiba Corp
Priority to JP26170284A priority Critical patent/JPS61140722A/en
Publication of JPS61140722A publication Critical patent/JPS61140722A/en
Pending legal-status Critical Current

Links

Classifications

    • 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
    • F23R3/346—Feeding into different combustion zones for staged combustion
    • 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
    • F23R3/02—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
    • F23R3/04—Air inlet arrangements
    • F23R3/06—Arrangement of apertures along the flame tube

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

Abstract

PURPOSE:To provide a gas turbine combustor having less decreasing in output in response to limitation of generation of nitrogen oxide by a method wherein mixture is divided into three sections of high, medium and low at a side wall of a combustor device and some through-pass holes are made so as to enable the mixtures to be supplied either to a pilot combustion area or main combustion area and the mixtures are supplied to each of the combustion areas. CONSTITUTION:Air for use in making mixture is mixed with fuel fed through the through-pass holes 21a, 21b and 21c before it is flowed into the passage ports 22a, 22b and 22c, passed through the passage ports 22a, 22b and 22c into the pilot combustion chamber 2 and the main combustion chamber 3. When it is approximately in a non-loaded condition, only the pilot fuel is supplied and a dispersion combustion under a high efficiency of combustion is carried out. Under a low loaded condition, all the main fuels are supplied only to the outer passage port 22a and only the air is injected from the other passage ports 22b and 22c. Due to this fact, even under such a flow rate of main fuel in which the fuel is mixed with the air for entire premixing operation to cause the excessive lean condition of the mixture, a proper concentration of the mixture can be attained, ignited and ignited under a lean premixed condition, so that a combustion with low nitrogen oxide under a high efficiency of combustion is carried out.

Description

【発明の詳細な説明】 〔発明の技術分野〕 本発明は、一つの燃焼室にパイロット燃焼領域とメイン
燃焼領域とに分区されたそのメイン燃焼領域に希薄燃料
としての混合気を混入し、窒素酸化物（以下ＮＯｘと記
す）の生成を抑制するガスタービン燃焼装置の改良に関
する。Detailed Description of the Invention [Technical Field of the Invention] The present invention has a single combustion chamber divided into a pilot combustion region and a main combustion region, and a mixture as a lean fuel is mixed into the main combustion region. The present invention relates to an improvement in a gas turbine combustion device that suppresses the production of oxides (hereinafter referred to as NOx).

〔発明の技術的背景〕[Technical background of the invention]

近時、ガスタービン設備は、大気汚染防止の見地から、
低ＮＯｘの開発がなされており、その一つとして燃焼ガ
スに希薄燃料（混合気）を混入し、局所的に高温領域を
取除く手法が講じられている。Recently, gas turbine equipment has been developed from the viewpoint of preventing air pollution.
Efforts have been made to reduce NOx, and one method is to mix lean fuel (mixture) into combustion gas to locally remove high temperature regions.

ところが、この手法は定格運転時において、かなりの成
果をおさめているものの、一度、ガスタービン設備が部
分負荷に入ると、混合気の燃料濃度が可燃限界以下にな
り、十分な燃焼がしなくなるおそれがある。However, although this method has achieved considerable success during rated operation, once the gas turbine equipment enters partial load, there is a risk that the fuel concentration in the mixture will fall below the flammable limit and sufficient combustion will not occur. There is.

かかる欠点を解消する一手段として、一つの燃焼器に二
つの燃料供給系統を設け、一系統はメイン燃焼領域を作
り出すために使用され、残の系統はパイロット燃焼領域
を作り出すために使用されるものがある。One way to overcome this drawback is to provide one combustor with two fuel supply systems, one used to create the main combustion zone and the remaining system used to create the pilot combustion zone. There is.

ここに、第４図は上記従来のガスタービン燃焼装置の概
略構成を示し、燃焼器（１）にはパイロット燃焼室（２
）とメイン燃焼室（３）とが一つの容器内に区分されて
いる。パイロット燃焼室（２）の先端は、第１の燃料供
給系ｆｃ（４）から送られてくる燃料を噴霧・噴口する
燃料噴射装置（５）が設けである。この燃料噴射装置（
５）の周囲は、燃料に旋回流を与えて好適に燃焼せしめ
るスワラ−（６）が環状列に周設されている。一方、メ
イン燃焼室（３）Ｋは、その側壁に予混合室（７）が連
設されており、この予混合室（７）は第２の燃料供給系
統（８）に結ばれていて、送られてくる燃料に過分の空
気を加えて混合気（９）を作り出している。したがって
、パイロット燃焼室（２）のガス温度が高くとも、予混
合室（力からの混合気（９）によって適度に温度が下げ
られ、これによってＮＯｘの生成を極力抑えている。な
お、符号０〔はパイロット燃焼室（２）の空気孔であり
、また符号（１１）もメイン燃焼室（３）の空気孔であ
る。Here, FIG. 4 shows a schematic configuration of the conventional gas turbine combustion device, in which the combustor (1) has a pilot combustion chamber (2).
) and the main combustion chamber (3) are divided into one container. The tip of the pilot combustion chamber (2) is provided with a fuel injection device (5) that sprays and injects fuel sent from the first fuel supply system fc (4). This fuel injector (
5) is surrounded by an annular array of swirlers (6) that give a swirling flow to the fuel and cause it to burn properly. On the other hand, the main combustion chamber (3) K has a premixing chamber (7) connected to its side wall, and this premixing chamber (7) is connected to a second fuel supply system (8). An air-fuel mixture (9) is created by adding an excess amount of air to the incoming fuel. Therefore, even if the gas temperature in the pilot combustion chamber (2) is high, the temperature is appropriately lowered by the air-fuel mixture (9) from the premixing chamber (force), thereby suppressing the generation of NOx as much as possible. [ is an air hole in the pilot combustion chamber (2), and symbol (11) is also an air hole in the main combustion chamber (3).

〔背景技術の問題点〕[Problems with background technology]

ところで、上記二系統の燃料供給ラインを有していても
、ガスタービン設備の負荷変動は滋しく、その運転が部
分負荷領域に入ると、たちまちメイン燃焼室（３）の燃
料が可燃限界以下に落ち入ることがある。このため、燃
料の可燃限界以下になると、例えば第５図乃至第６図に
示されるように、第２の燃料供給系統（８）が用いられ
、燃料の不足分をバックアップしている。すなわち、第
５図は負荷変化に対するメイン燃焼室（３）に送給され
る燃料量およびパイロット燃焼室（２）に送給される燃
料量のそれぞれの分担分を図式化したものであって、こ
の場合、燃焼不良があると、メイン燃焼室（２）Ｋ送給
される第２の燃料供給系統（８）からの燃料が増分され
、部分負荷時の出力低下を補っている（第６図参照）。By the way, even with the above-mentioned two systems of fuel supply lines, load fluctuations in gas turbine equipment are slow, and when the operation enters the partial load region, the fuel in the main combustion chamber (3) immediately drops below the flammable limit. Sometimes I feel depressed. Therefore, when the fuel becomes below the flammable limit, the second fuel supply system (8) is used to back up the insufficient amount of fuel, as shown in FIGS. 5 and 6, for example. That is, FIG. 5 is a diagram illustrating the respective portions of the amount of fuel fed to the main combustion chamber (3) and the amount of fuel fed to the pilot combustion chamber (2) with respect to load changes, In this case, if there is a combustion failure, the fuel from the second fuel supply system (8) that is fed to the main combustion chamber (2) is increased to compensate for the decrease in output at partial load (Figure 6). reference).

しかしながら、この方法でも中間負荷時に混合気が希薄
になり、パイロット燃焼室（２）からの燃焼ガスに混入
しても十分燃焼しきれず、燃焼効率が悪くなっている。However, even with this method, the air-fuel mixture becomes lean during intermediate loads, and even if mixed with the combustion gas from the pilot combustion chamber (2), it cannot be fully combusted, resulting in poor combustion efficiency.

この欠点を補うため、第７図に示す手法もある。第７図
も負荷変化に対するメイン燃焼室（３）に送給される燃
料量およびパイロット燃焼室（２）に送給される燃料量
のそれぞれの分担分を図式化したものである。この方法
では、第８図に示されるように、燃焼効率の悪い中間負
荷時、パイロット燃料だけで負荷に対応し、ある段階を
越えたらパイロット燃料をステップ状に減らしてその分
メイン燃料を注入する。するとメイン部分の当量比が燃
焼に充分な程度高くなり高い燃焼効率を保つことができ
る。しかしこの方法では中間負荷の大部分が通常の一段
燃焼器と同じことになり、パイロット部での局所当量比
が１前後になって燃焼ガスが高温化し、ＮＯｘが多量発
生してしまう。In order to compensate for this drawback, there is also a method shown in FIG. FIG. 7 also diagrammatically shows the respective portions of the amount of fuel fed to the main combustion chamber (3) and the amount of fuel fed to the pilot combustion chamber (2) with respect to load changes. In this method, as shown in Figure 8, during intermediate loads where combustion efficiency is poor, pilot fuel alone is used to cope with the load, and once a certain stage is exceeded, pilot fuel is reduced in steps and main fuel is injected accordingly. . Then, the equivalence ratio of the main part becomes high enough for combustion, and high combustion efficiency can be maintained. However, in this method, most of the intermediate load is the same as in a normal single-stage combustor, and the local equivalence ratio in the pilot section becomes around 1, which increases the temperature of the combustion gas and generates a large amount of NOx.

〔発明の目的〕[Purpose of the invention]

本発明は、上記の事情に照してなされたものであって、
全負荷領域においてＮＯｘの生成が抑制され、またＮＯ
ｘ生成の抑制に基づく出力低下のないようにするガスタ
ービン燃焼装置を提供するものである。The present invention has been made in light of the above circumstances, and includes:
NOx generation is suppressed in the entire load range, and NOx
An object of the present invention is to provide a gas turbine combustion device that prevents output from decreasing due to suppression of x generation.

〔発明の概要〕[Summary of the invention]

上記目的達成のため、本発明はガスタービン燃焼装置の
側壁に混合気を犬、中、小に分区してパイロット燃焼領
域またはメイン燃焼領域に送給できるよう透口を穿設し
たものであって、これら透口にはそれぞれ通路口が連設
されることを特徴とし、これら通路口を通じて各燃・焼
領域に混合気を送給するものである。In order to achieve the above object, the present invention has a through hole bored in the side wall of a gas turbine combustion device so that the air-fuel mixture can be divided into dog, medium, and small and fed to the pilot combustion area or the main combustion area. Each of these through holes is characterized in that a passage port is connected to each other, and the air-fuel mixture is supplied to each combustion region through these passage ports.

〔発明の実施例〕[Embodiments of the invention]

以下添付図を参照して本発明の一実施例を説明する。 An embodiment of the present invention will be described below with reference to the accompanying drawings.

第１図において、符号（１）は燃焼器を示し、との燃焼
器（１）の先端は、燃料噴射装置（５）を有している。In FIG. 1, reference numeral (1) indicates a combustor, and the tip of the combustor (1) has a fuel injection device (5).

燃料噴射装置（５）は制御弁（８ａ）を経て第１の燃料
供給系統（８）に結ばれており、この第１の燃料供給系
統（８）からの燃料をスワラ−（６）Ｋよって噴霧・旋
回せしめている。また、燃焼器（１）の側壁には、ヘッ
ダ翰が設けてあり、このヘッダ（２■は第２の燃料供給
系統（４）を経て制御弁（４ａ）　Ｋ結ばれている。The fuel injection device (5) is connected to a first fuel supply system (8) via a control valve (8a), and the fuel from the first fuel supply system (8) is fed through a swirler (6)K. It is sprayed and swirled. Further, a header is provided on the side wall of the combustor (1), and this header (2) is connected to a control valve (4a) K via a second fuel supply system (4).

上記ヘッダ（イ）は、燃焼器（１）の側壁に環状的に配
設されており、その−側は透口（２１ａ）、（２１ｂ）
、（２１ｃ）を有している。これら透口のうち、第１の
透口（２１ａ）は側壁の横断軸に対し、傾め前方に開設
され、第２の透口（２１ｂ）は側壁の横断軸と平行に開
設され、まだ第３の透口（２１Ｃ）は側壁の横断軸に対
し、傾め後方に開設されている。第１の透口（２１ａ）
は第１の通路口（２２ａ）に、第２の透口（２１ｂ）は
第２の通路口（２２ｂ）に、第３の透口（２１ｃ）は第
３の通路口（２２ｃ　）にそれぞれ連通せしめられてお
り、これらの通路口（２２ａ）　、　（２２ｂ）　、　
（２２Ｇ）はパイロット燃焼室（２）、メイン燃焼室（
３）に通じている。なお、符号Ｅは図示しない圧縮機か
ら送られてくる高圧空気を示し、この高圧空気によって
燃焼器（１）の側壁（特に内筒の側壁）を冷している。The header (a) is annularly arranged on the side wall of the combustor (1), and the - side has through holes (21a) and (21b).
, (21c). Among these openings, the first opening (21a) is opened obliquely forward with respect to the transverse axis of the side wall, and the second opening (21b) is opened parallel to the transverse axis of the side wall, and the second opening (21b) is opened parallel to the transverse axis of the side wall. No. 3 through hole (21C) is opened at an inclined rear with respect to the transverse axis of the side wall. First opening (21a)
communicates with the first passage opening (22a), the second passage opening (21b) with the second passage opening (22b), and the third passage opening (21c) with the third passage opening (22c). These passage ports (22a), (22b),
(22G) is the pilot combustion chamber (2), the main combustion chamber (
3). Note that the symbol E indicates high-pressure air sent from a compressor (not shown), and this high-pressure air cools the side wall of the combustor (1) (particularly the side wall of the inner cylinder).

次に作用を説明する。図示しない圧縮機から吐出した圧
縮空気（２）はアニユラ−空気流路（２５）と通路口（
２２ａ）、（２２ｂ）、（２２Ｃ）にはいり、アニユラ
−空気流路（２５）にはいっだ空気は内筒壁面冷却空気
のとスワラ−（６）を通過した旋回空気＋２６１になる
。この旋回空気ＣＩ！６１　ｅ用いて燃料噴射装置（５
）から噴射される燃料（２７）がパイロット燃焼室（２
）で燃焼する。混合気を作るだめの空気は通路口（２２
ａ）、（２２ｂ）、（２２ｃ）に流入する前に透口（２
１ａ）　、（２１ｂ）　、（２１ｃ）からの燃料と混合
し、通路口（２２ａ）、（２２ｂ）、（２２ｃ）を経て
パイロット燃焼室（２）、メイン燃焼室（３）に入る。Next, the effect will be explained. Compressed air (2) discharged from a compressor (not shown) flows through an annular air flow path (25) and a passageway (
22a), (22b), and (22C), and the air that enters the annular air flow path (25) becomes inner cylinder wall surface cooling air and swirling air +261 that has passed through the swirler (6). This swirling air CI! 61 e using fuel injection device (5
) The fuel (27) injected from the pilot combustion chamber (2
) to burn. The air that creates the mixture is passed through the passageway opening (22
a), (22b), and (22c).
1a), (21b), and (21c), and enters the pilot combustion chamber (2) and the main combustion chamber (3) through passage ports (22a), (22b), and (22c).

この実施例は、メイン燃料が定格流量のときの図であり
。This example is a diagram when the main fuel is at the rated flow rate.

このうち透口（２１ａ）から噴出する燃料は貫通距離が
大きく、最も内側の通路口（２２ａ）−ｊで達し、ここ
で混合気になる。角度と径の違う透口（２１ｂ）。Among these, the fuel ejected from the through port (21a) has a long penetration distance and reaches the innermost passage port (22a)-j, where it becomes an air-fuel mixture. Openings (21b) with different angles and diameters.

（２１Ｃ）から噴出する燃料は、それぞれの貫通距離の
違いにより、２つの通路口（２２ｂ）、、　（２２ｃ）
に達するようになっている。このようなしくみでそれぞ
れの通路口（２２ａ）、（２２ｂ）、（２２ｃ）で適度
な燃料濃度の予混合気になるようになっている。メイン
燃料が定格より小さいときの燃料ジェットのようすを示
したのが第２図である、燃料流量が少量のときはどのジ
ェットにも貫通力が足りなく、最も外側の通路口（２２
ａ）に全部の燃料が流入する。燃料流量が中間量のとき
は貫通距離の大きいジェットが中間の通路口（２２ｂ）
まで達してここで予混合気を形成し、貫通距離の小さい
ジェットは外側の通路口（２２ａ）で予混合気を形成す
る。定格流量時は前述したとおりである。(21C) The fuel ejected from the two passage ports (22b), (22c) due to the difference in their penetration distances.
It is designed to reach. With this mechanism, a premixture with an appropriate fuel concentration is created at each of the passage ports (22a), (22b), and (22c). Figure 2 shows the state of the fuel jet when the main fuel is smaller than the rated value.
All fuel flows into a). When the fuel flow rate is intermediate, the jet with a large penetration distance is located at the intermediate passage port (22b).
The jet reaches up to the outer passageway opening (22a) and forms a premixture there, and the jet having a small penetration distance forms a premixture at the outer passageway opening (22a). At rated flow rate, it is as described above.

以上のように動作する本実施例がもたらす効果を説明す
る。このような構造で動作させると無負荷状態に近いと
きはパイロット燃料のみを供給し、高燃焼効率で拡散燃
焼をする。負荷が小さいときはすべてのメイン燃料が外
側の通路口（２２ａ）にだけ供給され、他の通路口（２
２ｂ）　、（２２ｃ）からは空気だけが噴出する。この
ため全体の予混合気用空気に混合すれば希薄すぎて着火
しないようなメイン燃料流量でも適度の濃度になり、着
火し、希薄予混合燃焼するので高燃焼効率の低ＮＯｘ燃
焼をする。The effects brought about by this embodiment that operates as described above will be explained. When operated with this structure, only pilot fuel is supplied when there is almost no load, and diffusion combustion is performed with high combustion efficiency. When the load is small, all the main fuel is supplied only to the outer passage port (22a), and the other passage ports (22a) are supplied only to the outer passage port (22a).
Only air blows out from 2b) and (22c). For this reason, even if the main fuel flow rate is too lean to ignite when mixed with the entire premix air, it becomes an appropriate concentration, ignites, and performs lean premix combustion, resulting in high combustion efficiency and low NOx combustion.

負荷が中間量の場合は尾部側の透口（２１ｃ）だけ燃料
が供給されず空気だけが噴出し、他の透口（２１ａ）。When the load is an intermediate amount, fuel is not supplied to the tail side opening (21c) and only air is ejected, and the other opening (21a) is not supplied with fuel.

（２１ｂ）からは適度の燃料濃度の予混合気が噴出し、
希薄予混合燃焼して高燃焼効率の低ＮＯｘ燃焼をする。From (21b), a premixture with an appropriate fuel concentration is ejected,
Performs lean premix combustion to achieve high combustion efficiency and low NOx combustion.

定格負荷の時は全部の透口（２１ａ）、（２１ｂ）、（
２１ｃ）から予混合気が噴出して同様に高燃焼効率の低
ＮＯｘ燃焼を行なう。以上のように本発明の予混合ダク
トを用いると全負荷範囲に渡って高燃焼効率の低ＮＯｘ
燃焼を達成することのできる燃焼器を提供することがで
きる。At rated load, all through holes (21a), (21b), (
The premixture is ejected from 21c) to similarly perform low NOx combustion with high combustion efficiency. As described above, the use of the premixing duct of the present invention provides high combustion efficiency and low NOx over the entire load range.
A combustor capable of achieving combustion can be provided.

本発明は前述し−だ実施例に限定されることなく、趣旨
を変更しない範囲で種々変形して実施することができる
。その例として第３図に示すように通路口（イ）を多重
アニユラ−にせず、一つのアニユラ−にすることも考え
られる。混合の不均一を利用して予混合気の濃度に差を
つけるものである。メイン燃料が少量のときは燃料ジェ
ットの貫通距離が小さく、全体に混合しない。そのため
、尾部側の予混合気噴出孔（１４Ｃ）には燃料が達せず
、前述した多重アニユラ−の方法と同じ効果がある。多
重アニユラ一方式より濃度に差をつける効果は小さいが
通常の予混合方式よりはるかに中間負荷時の燃焼効率は
改善される。また、多重アニユラ一方式より構造が更に
シンプルになる利点がある。The present invention is not limited to the above-mentioned embodiments, but can be implemented with various modifications without changing the spirit. As an example, as shown in FIG. 3, it is possible to use a single annular for the passageway opening (a) instead of multiple annulars. This method takes advantage of non-uniform mixing to differentiate the concentration of the premixture. When the main fuel is small, the penetration distance of the fuel jet is small and it does not mix completely. Therefore, the fuel does not reach the premixture injection hole (14C) on the tail side, and has the same effect as the multiple annular method described above. Although the effect of making a difference in concentration is smaller than that of the multiple annular single system, the combustion efficiency at intermediate loads is much improved compared to the normal premix system. Further, there is an advantage that the structure is simpler than that of the multi-annulus single type.

〔発明の効果〕〔Effect of the invention〕

以上説明したように、本発明によれば、燃焼器の側壁に
、混合気を犬、中、小に区分してパイロット燃焼領域ま
たはメイン燃焼領域に供給できるよう透口を穿設し、こ
の透口は通路口を経て上記パイロット燃焼領域、メイン
燃焼領域に連通ずるようにしたから、負荷が変化しても
、ＮＯｘの生成は抑制され、またその変化時における出
力低下も防止することが期待できる。As explained above, according to the present invention, a through hole is bored in the side wall of the combustor so that the air-fuel mixture can be divided into small, medium, and small and then supplied to the pilot combustion area or the main combustion area. Since the opening communicates with the pilot combustion area and the main combustion area through the passage opening, it is expected that even if the load changes, the generation of NOx will be suppressed, and a decrease in output will be prevented when the load changes. .

【図面の簡単な説明】[Brief explanation of the drawing]

第１図は本発明の一実施例を示す概略図、第２図は第１
図の部分拡大図、第３図は本発明の他の実施例を示す図
、第４図は従来の実施例を示す概略図、第５図はパイロ
ット燃焼室、メイン燃焼室に供給される燃料量割合を示
す模式図、第６図は第５図の燃料量割合によって燃焼不
良を示す模式図、第７図もパイロット燃焼室、メイン燃
焼室に供給される燃料量割合を示す模式図、第８図は第
７図の燃料量割合によってＮＯｘの発生量を示す模式図
である。 １・・・燃焼器　　　　　　２・・パイロット燃焼室３
・・メイン燃焼室　　　２０・・・ヘッダ２１ａ、２１
ｂ、２１Ｃ透口　　２２ａ　、２２ｂ　、　２２０　　
通路口代理人　弁理士　則　近　憲　佑　（ばか１名）
第５図 負　荷　（Ｚ） 第６図 Ｈ−一勺Ｔ、゛ｒ尭不良−−− 第７図 第８図 −ＮＱ発←FIG. 1 is a schematic diagram showing one embodiment of the present invention, and FIG. 2 is a schematic diagram showing an embodiment of the present invention.
3 is a diagram showing another embodiment of the present invention, FIG. 4 is a schematic diagram showing a conventional embodiment, and FIG. 5 is a fuel supplied to the pilot combustion chamber and the main combustion chamber. Fig. 6 is a schematic diagram showing poor combustion depending on the fuel quantity ratio shown in Fig. 5. Fig. 7 is also a schematic diagram showing the fuel quantity ratio supplied to the pilot combustion chamber and the main combustion chamber. FIG. 8 is a schematic diagram showing the amount of NOx generated depending on the fuel amount ratio shown in FIG. 1... Combustor 2... Pilot combustion chamber 3
...Main combustion chamber 20...Header 21a, 21
b, 21C opening 22a, 22b, 220
Passage agent Patent attorney Kensuke Chika (1 idiot)
Fig. 5 Load (Z) Fig. 6 H-Ichitsu T, ゛r 尭 defective --- Fig. 7 Fig. 8-NQ issue←

Claims (2)

【特許請求の範囲】[Claims] （１）燃焼器の先端に燃料噴射装置を設けてパイロット
燃焼領域を作り、また燃焼器の側壁に希薄燃料混合気が
通過する側路を設けてメイン燃焼領域を作り、前記燃料
噴射装置からの燃焼ガスによつてもたらされるメイン燃
焼領域に、前記混合気を混合せしめるガスタービン燃焼
装置において、前記側壁には混合気を大、中、小に分区
して前記パイロット燃焼領域またはメイン燃焼領域に送
給できるよう透口を穿設するとともに、前記透口にそれ
ぞれ通路口が連設されることを特徴とするガスタービン
燃焼装置。(1) A fuel injection device is provided at the tip of the combustor to create a pilot combustion region, and a side passage through which a lean fuel mixture passes is provided on the side wall of the combustor to create a main combustion region, and the fuel injection device is provided with a main combustion region. In a gas turbine combustion device that mixes the air-fuel mixture into a main combustion region brought by combustion gas, the side wall has a wall that divides the air-fuel mixture into large, medium, and small and sends it to the pilot combustion region or the main combustion region. A gas turbine combustion apparatus characterized in that a through hole is provided so that the gas can be supplied to the gas, and a passage port is connected to each of the through holes. （２）透口は、混合気が少ないとき、側壁の横断軸に対
して傾め前方に開設し、混合気が中程度のとき、側壁の
横断軸に平行に、また混合気が多いとき、側壁の横断軸
に対して傾め後方にそれぞれ開設することを特徴とする
特許請求の範囲第１項記載のガスタービン燃焼装置。(2) When the air-fuel mixture is low, the opening is tilted forward with respect to the transverse axis of the side wall, when the air-fuel mixture is medium, the opening is opened parallel to the transverse axis of the side wall, and when the air-fuel mixture is large, the opening is opened at an angle to the front. 2. The gas turbine combustion device according to claim 1, wherein the side walls are provided at rear sides inclined with respect to the transverse axis of the side walls.
JP26170284A 1984-12-13 1984-12-13 Gas turbine combustor Pending JPS61140722A (en)

Priority Applications (1)

Application Number Priority Date Filing Date Title
JP26170284A JPS61140722A (en) 1984-12-13 1984-12-13 Gas turbine combustor

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
JP26170284A JPS61140722A (en) 1984-12-13 1984-12-13 Gas turbine combustor

Publications (1)

Publication Number Publication Date
JPS61140722A true JPS61140722A (en) 1986-06-27

Family

ID=17365522

Family Applications (1)

Application Number Title Priority Date Filing Date
JP26170284A Pending JPS61140722A (en) 1984-12-13 1984-12-13 Gas turbine combustor

Country Status (1)

Country Link
JP (1) JPS61140722A (en)

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP3008391A4 (en) * 2013-06-11 2016-07-06 United Technologies Corp Combustor with axial staging for a gas turbine engine

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP3008391A4 (en) * 2013-06-11 2016-07-06 United Technologies Corp Combustor with axial staging for a gas turbine engine
US11143407B2 (en) 2013-06-11 2021-10-12 Raytheon Technologies Corporation Combustor with axial staging for a gas turbine engine

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