ОСНОВНОЕ УСТРОЙСТВО ДЛЯ ВПРЫСКИВАНИЯ ЖИДКОГО ТОПЛИВА ДЛЯ ОДНОЙ КАМЕРЫ СГОРАНИЯ, ИМЕЮЩЕЙ КАМЕРУ ПРЕДВАРИТЕЛЬНОГО СМЕШИВАНИЯ, ГАЗОВОЙ ТУРБИНЫ С МАЛЫМ УРОВНЕМ ВЫБРОСА ЗАГРЯЗНЯЮЩИХ ОКРУЖАЮЩУЮ СРЕДУ ВЕЩЕСТВ

Основное устройство для впрыскивания жидкого топлива для одной камеры (10) сгорания, имеющей камеру предварительного смешивания, газовой турбины с малым уровнем выброса загрязняющих окружающую среду веществ, содержит группу каналов для впрыскивания жидкого топлива, распределенных внутри камеры (12) предварительного смешивания и группу лопаток, отступающих в радиальном направлении относительно оси симметрии камеры сгорания. Группа лопаток снабжена проходами для охлаждающего воздуха, окружающими каналы для подачи жидкого топлива и охлаждающими их для предотвращения формирования отложений нагара на поверхностях лопаток. Каждая лопатка, по меньшей мере, на одной боковой поверхности снабжена, по меньшей мере, одним каналом для впрыскивания жидкого топлива и, по меньшей мере, одной точкой для впрыскивания охлаждающего воздуха. Изобретение направлено на снижение выброса загрязняющих окружающую среду веществ, а также на повышение устойчивости пламени и уменьшение колебаний давления в камере сгорания ...

ТОПЛИВНЫЙ ИНЖЕКТОР КАМЕРЫ СГОРАНИЯ ГАЗОТУРБИННОГО ДВИГАТЕЛЯ, ГАЗОТУРБИННЫЙ ДВИГАТЕЛЬ И СПОСОБ ЕГО ФУНКЦИОНИРОВАНИЯ

Топливный инжектор камеры сгорания газотурбинного двигателя содержит первый трубопровод подачи топлива, обеспечивающий работу в режиме малого газа, и второй основной трубопровод подачи топлива, обеспечивающий работу в других режимах, включая работу на полную мощность, а также первые впускные отверстия для работы в режиме малого газа и вторые основные впускные отверстия, с которыми два питательных трубопровода соединяются соответствующим образом. Топливный инжектор содержит определенное количество (n1) первых отверстий и определенное количество (n2) вторых отверстий в соотношении n1/n2<1. Впускные отверстия располагаются в виде кольца, а первые отверстия для работы в режиме малого газа занимают сектор вышеупомянутого кольца. При этом инжектор содержит также центральный канал снабжения первичным воздухом, а отверстия впрыска располагаются в виде кольца вокруг данного центрального канала. Изобретение обеспечивает ступенчатое сгорание, т.е. позволяющее создать зону горения в режиме малого газа ...

УСТРОЙСТВО ВПРЫСКИВАНИЯ СМЕСИ ВОЗДУХА С ТОПЛИВОМ, КАМЕРА СГОРАНИЯ И ГАЗОТУРБИННЫЙ ДВИГАТЕЛЬ, СНАБЖЕННЫЕ ТАКИМ УСТРОЙСТВОМ

Устройство впрыскивания смеси воздуха с топливом в камеру сгорания газотурбинного двигателя имеет ось (X) симметрии вращения и содержит расположенные, если смотреть в направлении спереди назад по ходу течения потока газов, скользящий переходный элемент (20), имеющий ось (Y) вращения и связанный с радиальными спиральными элементами (40) при помощи кольцевой чашки (30), и конический аэродинамический корпус (60), отстоящий в осевом направлении от этих радиальных спиральных элементов (40). Скользящий переходный элемент (20) содержит переднюю по потоку сходящуюся коническую стенку (21), продолжающуюся цилиндрической стенкой, имеющей ось (X), и задним по потоку фланцем (23), проходящим в радиальном направлении на заднем по потоку конце цилиндрической стенки и снабженным отверстиями подвода воздуха под давлением, называемыми также отверстиями (22) продувки. Передняя по потоку сходящаяся стенка (21) скользящего переходного элемента (20) снабжена, по меньшей мере, одним рядом дополнительных отверстий ...

СПОСОБ ЗАЖИГАНИЯ ДЛЯ КАМЕРЫ СГОРАНИЯ ГАЗОТУРБИННОГО ДВИГАТЕЛЯ

Способ зажигания для камеры сгорания газотурбинного двигателя, питаемой топливом через форсунки и имеющей свечу зажигания, содержит первоначальную фазу, во время которой в камеру впрыскивают топливо с постоянным расходом одновременно с активизацией свечи зажигания, и, - при отсутствии воспламенения в камере в конце первоначальной фазы, - вторую фазу. Во время второй фазы резко увеличивают расход впрыскиваемого топлива на 20-30%. За второй фазой следует фаза постепенного увеличения расхода топлива, которая является менее интенсивной и менее быстрой, чем вторая фаза. Изобретение направлено на повышение надежности воспламенения в камере сгорания. 6 з.п. ф-лы, 3 ил.

УСТРОЙСТВО СГОРАНИЯ С ИМПУЛЬСНЫМ РАЗДЕЛЕНИЕМ ТОПЛИВА

Дано описание управляющего блока устройства сгорания и устройства сгорания, например, газовой турбины, который на основе по меньшей мере одного рабочего параметра определяет, находится ли устройство сгорания в заданной рабочей фазе. В ответ на это генерируется управляющий сигнал, предназначенный для установки соотношения по меньшей мере двух различных входных потоков топлива на заданное значение (psc1, psc3) для заданного времени (dt) в случае, если устройство сгорания находится в заданной рабочей фазе. Технический результат - уменьшение опасности перегрева камеры сгорания и уменьшение выброса вредных веществ в широком рабочем диапазоне. 4 н. и 4 з.п. ф-лы, 3 ил.

ТОПЛИВНЫЙ ИНЖЕКТОР С ЦЕНТРАЛЬНЫМ КОРПУСОМ В СБОРЕ

Раскрыт центральный корпус в сборе (700) для топливного инжектора (600) с прямым впрыском обедненной смеси. Центральный корпус в сборе определяет первичный канал (721) для жидкости, проход (880) для жидкости, первый основной канал (875) для жидкости, второй основной канал (879) для жидкости и распылитель в сборе (850). Первичный канал для жидкости подает жидкое топливо в проход для жидкости. Проход для жидкости выровнен с первичным каналом для жидкости и распределяет топливо на первый основной канал для жидкости и впоследствии на второй основной канал для жидкости. Распылитель в сборе находится в связи по текучей среде со вторым основным каналом для жидкости и обеспечивает распыление топлива для сжигания. 9 з.п. ф-лы, 13 ил.

УСТРОЙСТВО ДЛЯ ПОДАЧИ ТОПЛИВА В ОСНОВНУЮ КАМЕРУ СГОРАНИЯ

Изобретение относится к устройствам регулирования подачи топлива в основную камеру сгорания ГТД в топливной форсунке. Устройство для подачи топлива в основную камеру сгорания ГТД содержит малую топливную форсунку, а также форсунку, выполненную в виде внешнего корпуса (гильзы) малой форсунки, внешний и внутренний завихрители воздуха, при этом форсунка (гильза) выполнена высокорасходной и удерживается в неподвижном положении пружиной, а лопатки внешнего завихрителя выполнены поворотными. Изобретение позволяет расширить диапазон устойчивого горения и обеспечить стабильность процесса горения при увеличении расхода топлива. 7 ил.

ТОПЛИВНЫЙ ИНЖЕКТОР С ГАЗОРАСПРЕДЕЛЕНИЕМ ЧЕРЕЗ МНОЖЕСТВО ТРУБОК

Настоящее изобретение в целом относится к газотурбинным двигателям и к топливному инжектору с контуром газораспределения через множество трубок. Раскрыт топливный инжектор (600) для газотурбинного двигателя. Топливный инжектор (600) содержит фланец в сборе (610) с распределительным блоком (612), три трубки для основного газа и распылительное устройство (630). Распределительный блок (612) равномерно распределяет основное газообразное топливо по трем трубкам для основного газа. Распылительное устройство (630) содержит корпус (640) инжектора с первичным проходом для газа, имеющим кольцеобразную форму. Три трубки для основного газа подключены параллельно между распределительным блоком (612) и первичным проходом для газа. Три трубки для газа находятся в связи по текучей среде с первичным проходом для газа и предоставляют основное газообразное топливо от одного и того же источника основного газообразного топлива. Выравнивание термического расширения в каждой трубке для основного газа может предотвратить ...

Топливная форсунка с радиальным и осевым завихрителями для газовой турбины и газовая турбина

Варианты выполнения изобретения относятся к топливным форсункам газовых турбин, имеющим радиальный и осевой завихрители, и к газовым турбинам, в которых применяются указанные форсунки. Данное изобретение направлено на решение проблемы, заключающейся в обеспечении надежности работы топливной форсунки и решаемой путем обеспечения стабильности пламени и снижения выбросов NOx. Согласно первым вариантам выполнения, топливная форсунка содержит радиальный завихритель и осевой завихритель, причем радиальный завихритель выполнен с возможностью закручивания первого потока первой смеси топлива и окислителя, а осевой завихритель выполнен с возможностью закручивания второго потока второй смеси топлива и окислителя. Первый поток может подаваться по центральному каналу, а второй поток может подаваться по кольцевому каналу, окружающему центральный канал. С радиальным завихрителем форсунки взаимосвязана первая зона рециркуляции, а с осевым завихрителем форсунки взаимосвязана вторая зона рециркуляции, по ...

ГОРЕЛКА МНОГОКОНУСНОГО ТИПА ПРЕДВАРИТЕЛЬНОГО СМЕШИВАНИЯ ДЛЯ ГАЗОВОЙ ТУРБИНЫ

Горелка предварительного смешивания многоконусного типа для газовой турбины содержит множество кожухов, расположенных вокруг центральной оси горелки и являющихся частями виртуального аксиально продолжающегося общего конуса , открытого в направлении вниз по потоку. Указанные части смещены перпендикулярно оси горелки для образования тангенциальной щели между каждой парой смежных кожухов. Виртуальный общий конус имеет угол конусности, изменяющийся в осевом направлении. Изобретение направлено на повышение стабильности горения. 8 з.п. ф-лы, 6 ил.

КОРПУС ТОПЛИВНОЙ ФОРСУНКИ ДЛЯ ГАЗОТУРБИННОГО ДВИГАТЕЛЯ (ВАРИАНТЫ) И СПОСОБ ЕГО ИЗГОТОВЛЕНИЯ

Корпус топливной форсунки для газотурбинного двигателя, выполненный из детали из твердого материала, содержит входную и выходную торцевые части. В корпусе между входной и выходной торцевыми частями выполнено, по крайней мере, одно пазовое средство, которое по всей длине герметизировано, по крайней мере, одним закрывающим средством с образованием, по крайней мере, одного канала подачи топлива от упомянутой входной торцевой части к упомянутой выходной торцевой части. Входная торцевая часть выполнена с возможностью соединения с подсоединенным к топливному инжектору переходником топливопровода и подачи топлива через упомянутый корпус топливной форсунки. При изготовлении корпуса топливной форсунки сначала рассверливают противоположные торцы детали из твердого материала. Затем выполняют, по крайней мере, одно пазовое средство по длине упомянутой детали между входными и выходными частями. Герметизируют пазовые средства посредством, по крайней мере, одного закрывающего устройства и образуют, по ...

ЗАВИХРИТЕЛЬ, ГОРЕЛКА И СИСТЕМА СГОРАНИЯ ДЛЯ ГАЗОТУРБИННОГО ДВИГАТЕЛЯ

Изобретение относится к завихрителю, предназначенному для использования в системе сгорания газотурбинного двигателя (10), содержащему множество простирающихся в основном радиально внутрь каналов, циклически разнесенных по окружности в шахматном порядке, причем каждый канал имеет радиально внешний входной конец, радиально внутренний выходной конец, первую и вторую простирающиеся в основном радиально внутрь боковые поверхности, а также поверхность основания и верхнюю поверхность. Так, во время эксплуатации завихрителя топливо и воздух движутся по каналам от их входных концов к их выходном концам таким образом, что создают рядом с выходными концами завихряющуюся топливовоздушную смесь, при этом по меньшей мере одна поверхность по меньшей мере одного канала содержит по меньшей мере одно отверстие для впрыска газообразного топлива. Также поверхность завихрителя, имеющая отверстие для впрыска газообразного топлива, содержит по меньшей мере одно глухое отверстие, радиально окружающее отверстие ...

ТОПЛИВНЫЙ ИНЖЕКТОР И ТОПЛИВНАЯ СИСТЕМА ДЛЯ ДВИГАТЕЛЯ ВНУТРЕННЕГО СГОРАНИЯ

Топливный инжектор (26) для двигателя (10) внутреннего сгорания содержит инжекторную головку (28), содержащую сопло (44), устройство (48) предварительного смешивания и распределитель (70), приспособленный для распределения нескольких разных типов топлива в разные наборы проходов (58, 60, 66) для снабжения топливом в устройстве (48) предварительного смешивания. Узел (103) пилотного впрыска топливного инжектора (26) соединен с устройством (48) предварительного смешивания и содержит первый тракт (110) снабжения топливом для первого топлива и второй тракт (111) снабжения топливом для второго топлива. Несколько наборов проходов (58, 60, 66) для снабжения топливом расположены в топливном инжекторе (26), при этом наборы (58, 60, 66) проходов для снабжения топливом выборочно соединяются с несколькими источниками (20, 22, 24) разного топлива, и как одни, так и другие имеют такие расположение и размер, чтобы соответствовать широкому диапазону расхода топлива для обеспечения работы двигателя (10) ...

ГОРЕЛКА И СПОСОБ ЭКСПЛУАТАЦИИ ГОРЕЛКИ

Горелка содержит выходное отверстие горелки с, по меньшей мере, двумя секторами. Каждому сектору соответствует, по меньшей мере, одна топливная форсунка, причем имеются, по меньшей мере, два раздельных ведущих к топливным форсункам разных секторов трубопровода для подачи топлива, и имеется устройство регулирования потока массы топлива, текущего через соответствующий трубопровод для подачи топлива. В качестве устройства регулирования потока массы топлива, текущего через соответствующий трубопровод для подачи топлива, используются расположенные в соответствующем трубопроводе для подачи топлива регулируемые вентили. Вентили регулируются раздельно таким образом, что в режиме полной нагрузки предусмотрена равномерная подача топлива во все сектора так, что возникает однородное распределение температуры. В режиме частичной нагрузки за счет раздельного управления подачей топлива в отдельные сектора горелки в камере сгорания можно создавать более горячие и более холодные зоны, причем более горячие ...

СПОСОБ СЖИГАНИЯ С НИЗКИМ УРОВНЕМ ЗВУКОВЫХ ЭФФЕКТОВ (ВАРИАНТЫ)

Изобретение предназначено для использования в энергетике. Способ уменьшения пульсаций давления в камере сгорания газотурбинного двигателя, возникающих в результате горения в ней топлива и воздуха, предусматривает сжигание смеси топлива с воздухом в камере сгорания вниз по технологической цепочке от плоскости выходного сечения топливной форсунки так, чтобы зоны рециркуляции, образуемые посредством топливной форсунки, отстояли от плоскости выходного сечения, а продукты сгорания были отделены от топлива и воздуха в зоне смешения при всех режимах работы двигателя. Изобретение обеспечивает уменьшение звуковых эффектов и пульсаций. 2 с.п. ф-лы, 3 ил.

СЖИГАЮЩЕЕ УСТРОЙСТВО ГАЗОТУРБИННОЙ УСТАНОВКИ

Изобретение относится к сжигающему устройству газотурбинной установки. Сжигающее устройство газотурбинной установки содержит пилотную горелку, пилотный клапан регулирования подачи топлива, который регулирует расход топлива, подаваемого в пилотную горелку, основную горелку для горения предварительно приготовленной смеси, расположенную на внешней периферийной стороне пилотной горелки, множество основных клапанов регулирования подачи топлива, которые регулируют расходы топлива, индивидуальным образом подаваемого во множество секторов горелки, на которые разделена основная горелка в окружном направлении, и контроллер, выполненный с возможностью управления пилотным клапаном регулирования подачи топлива и множеством основных клапанов регулирования подачи топлива, при этом контроллер выполнен с возможностью управления множеством основных клапанов регулирования подачи топлива таким образом, что, когда топливо подлежит подаче во все из множества секторов горелки, возникает различие в расходе топлива ...

ГАЗОВАЯ ТУРБИНА ДЛЯ СЖИГАНИЯ ГОРЮЧЕГО ГАЗА

Газовая турбина для сжигания горючего газа имеет систему трубопроводов, выполненную таким образом, что часть горючего газа отводится, направляется через каталитическую ступень предварительного формирования для преобразования содержащегося в горючем газе углеводорода в спирт и/или альдегид и затем подводится к горючему газу для снижения его температуры воспламенения. Такое осуществление изобретения приводит к тому, что в ступени предварительного формирования из горючего газа получают сравнительно легковоспламеняемые вещества спирт и/или альдегид. Смешанный с этими веществами горючий газ поэтому воспламеняется при значительно более низкой температуре воспламенения, чем горючий газ без предварительно сформированных компонентов. 4 з.п.ф-лы, 1 ил.

СПОСОБ СЖИГАНИЯ ТОПЛИВА В СЖАТОМ ВОЗДУХЕ

Способ сжигания топлива в сжатом воздухе относится к сжиганию жидкого или газообразного топлива в сжатом воздухе в газовой турбине. Сжатый воздух подают в перемещающемся вдоль оси потоке. При этом вначале от потока отделяют множество частичных потоков. К каждому из потоков по отдельности подводят часть топлива и сжигают во вдающемся в потокопилотном пламени. Остальное топливо подводят в различных местах к потоку. Распределяют в потоке неоднородно. Образующееся за счет неоднородного распределения топливо в потоке имеет локальный максимум на каждом пилотном пламени. Далее топливо поджигают на пилотных пламенах и сжигают. Такое осуществление способа способствует стабилизации горения и обеспечению того, что все топливо сжигается. 4 з.п.ф-лы, 6 ил.

МАЛОЭМИССИОННАЯ КАМЕРА СГОРАНИЯ ГАЗОТУРБИННОГО ДВИГАТЕЛЯ

... 1. Камера сгорания газотурбинного двигателя с выходной электрической мощностью 100-130 МВт, двухопорным жестким ротором, выполненным с консольной барабанной секцией компрессора и сварными барабанно-дисковыми секциями компрессора и турбины, содержащая 20-22 модульных элемента, расположенных по окружности ротора снаружи по отношению к последним ступеням компрессора, причем каждый модульный элемент камеры сгорания содержит силовой корпус, установленные внутри него по меньшей мере одно фронтовое устройство с концентрически расположенными и снабженными выходными насадками цилиндрическими внутренней пилотной горелкой и внешней основной горелкой предварительного смешения, по меньшей мере один электрический поджигатель, а также примыкающую к выходному насадку основной горелки жаровую трубу с пламенным и газоотводным участками, причем входной торец пламенного участка жаровой трубы имеет больший диаметр по отношению к выходному диаметру указанного насадка и соединен с последним посредством кольцевого ...

КАМЕРА СГОРАНИЯ ГАЗОТУРБИННОГО ДВИГАТЕЛЯ

Камера сгорания газотурбинного двигателя, содержащая, по крайней мере, одно запальное устройство и жаровую трубу, в которой расположены по окружности ряды отверстий для подачи воздуха, в одном из рядов выполнено отверстие для установки запального устройства, при этом отверстия этого ряда имеют одинаковую площадь, отличающаяся тем, что вокруг запального устройства образован кольцевой канал для подачи воздуха, площадь которого больше площади отверстия для подачи воздуха, площадь которого больше площади отверстия для подачи воздуха, в котором установлено запальное устройство, в 1,1-1,4 раза.

КАМЕРА СГОРАНИЯ С НИЗКИМ УРОВНЕМ ЗАГРЯЗНЕНИЯ И СПОСОБ УПРАВЛЕНИЯ СГОРАНИЕМ ДЛЯ НЕЕ

Изобретение относится к камере сгорания с низким уровнем загрязнения и способу управления сгоранием для нее. Камера сгорания с низким уровнем загрязнения содержит головную часть камеры сгорания, содержащую ступень сжигания основной смеси и ступень предварительного сжигания, ступень сжигания основной смеси содержит канал ступени сжигания основной смеси и завихритель ступени сжигания основной смеси, расположенный в канале ступени сжигания основной смеси, при этом ступень сжигания основной смеси дополнительно содержит предварительную пленочную пластину, расположенную в канале ступени сжигания основной смеси, и предварительная пленочная пластина радиально разделена на предварительную пленочную пластину внешнего слоя и предварительную пленочную пластину внутреннего слоя, и при этом положения и направления впрыска точек впрыска топлива ступени сжигания основной смеси конфигурируются, чтобы регулировать топливо ступени сжигания основной смеси, которое должно быть впрыснуто в канал ступени сжигания ...

УСТРОЙСТВО ДЛЯ РАСПЫЛА ТОПЛИВА В КАМЕРЕ СГОРАНИЯ

... 1. Устройство для распыла топлива в камере сгорания газотурбинного двигателя, содержащее топливную форсунку и размещенный вокруг форсунки воздушный осесимметричный тракт, разделенный продольно кольцевым элементом на наружный и внутренний каналы, причем наружный канал ограничен патрубком, включающим сужающийся и расширяющийся участки, а внутренний канал включает цилиндрический и сужающийся участки, ограниченные кольцевым элементом с острой кромкой на выходе, при этом наружный канал на входе имеет лопаточный завихритель, расположенный между патрубком и кольцевым элементом, а внутренний канал - кольцевой завихритель, расположенный в кольцевом элементе после топливной форсунки, кроме того, устройство включает фланцевую втулку, размещенную над патрубком наружного канала с образованием дополнительного кольцевого канала и скрепленную с этим патрубком концевыми частями, отличающееся тем, что патрубок на выходе дополнительно снабжен цилиндрическим участком и кольцевым козырьком, где внутренняя поверхность ...

ГОРЕЛКА, В ЧАСТНОСТИ ДЛЯ ГАЗОВОЙ ТУРБИНЫ

Горелка, в частности для газовой турбины, содержит каталитическую камеру сгорания. Камера сгорания имеет в направлении течения топлива в основном цилиндрическую протяженность и содержит на обращенной к топливу стенке каталитически активное покрытие для окисления топлива. Камера сгорания выполнена с возможностью подведения топлива, содержащего главный поток топлива, предварительно сформированный частичный поток топлива и воздуха. Для предварительного формирования предусмотрена обтекаемая частичным потоком топлива каталитическая ступень предварительного формирования, разлагающая топливо, по крайней мере частично, на легко воспламеняющиеся вещества, в частности на спирты, альдегиды или водород. Для ввода предварительно сформированного, частичного потока топлива, смешанного при необходимости с воздухом, в стенке камеры сгорания предусмотрены отверстия. За счет каталитически индуцированного сжигания топлива достигается особенно низкое содержание окислов азота в отходящем газе горелки. Одновременно ...

КАМЕРА СГОРАНИЯ ДЛЯ ГАЗОВОЙ ТУРБИНЫ

Камера сгорания для газовой турбины является протекаемой вдоль оси текущим из компрессорной части к турбинной части потоком сжатого воздуха, который имеет завихрение относительно оси. При этом на входе камеры сгорания предусмотрен кольцевой канал с впускной частью для отделения частичного потока из потока, которая содержит средства для устранения от частичного потока его завихрения и сообщается с каналами охлаждения для охлаждения камеры сгорания, а также с пилотными горелками для стабилизации горения в камере сгорания. Такое выполнение камеры сгорания приводит к уменьшению термодинамических потерь, и позволяет использовать отделение частичного потока как для целей охлаждения, так и для целей стабилизации. 6 з.п.ф-лы, 4 ил.

Система подачи жидкого топлива газотурбинной установки

Изобретение относится к области энергомашиностроения, конкретно к газотурбостроению, в частности к системе подачи жидкого топлива к горелкам камеры сгорания и может быть использовано в составе двухтопливной энергетической газотурбинной установки. Система подачи жидкого топлива содержит магистраль подачи жидкого топлива 1 к горелкам камеры сгорания ГТУ, линию возврата жидкого топлива 2 от горелок камеры сгорания газовой турбины. Магистраль подачи жидкого топлива 1 содержит последовательно, друг за другом, установленный сдвоенный фильтр жидкого топлива 3, мембранные баки 7, насос жидкого топлива 4, клапан рециркуляции жидкого топлива 17, аварийный запорный клапан 5, регулирующий клапан 6. На линии возврата жидкого топлива 2 последовательно друг за другом установлены аварийный запорный клапан 8 и регулирующий клапан 9. Магистраль подачи жидкого топлива 1 связана с линией возврата жидкого топлива 2 через линию рециркуляции жидкого топлива 26, включающую регулятор сопротивления 25, которая с ...

КАМЕРА СГОРАНИЯ С ОПТИМАЛЬНЫМ ЧИСЛОМ ФОРСУНОК

Камера сгорания газотурбинного двигателя с оптимальным числом форсунок содержит корпус, жаровую трубу и фронтовое устройство с завихрителями воздуха и форсунками подачи жидкого или газообразного топлива. Фронтовое устройство камеры выполнено с оптимальным числом форсунок - три штуки на 100 см2 площади поперечного (миделевого) сечения жаровой трубы. Изобретение позволяет снизить выброс оксидов азота при минимальном недожоге топлива. 3 ил.

Кольцевая камера сгорания

Изобретение относится к стационарным газотурбинным двигателям ГТД, используемым в наземных условиях в судостроении, на газоперекачивающих станциях в качестве привода насоса и для пиковых энергетических установок в качестве привода электрогенератора. Кольцевая камера сгорания состоит из корпуса наружного (1) и внутреннего (2), диффузора (3), топливоподводящего коллектора (4), жаровой трубы (5), включающей наружное (6) и внутреннее (7) кольцо, состоящее из трех секций с двумя рядами отверстий для подвода воздуха, наружный (8) и внутренний (9) кожухи, состоящие из двух секций, внутренние (10) и наружные (11) карманы, уплотнительные кольца (12), фиксаторы (13) для подвешивания жаровой трубы (5), топливные форсунки (14). Фронтовое устройство (15) жаровой трубы (5) содержит головку кольцевую (16) с наружным (17) и внутренним (18) топливными коллекторами, штуцером (19) подвода газа во внутреннюю полость наружного (17) коллектора, закрепленным при помощи проходника (20) и вкладыша (21), установленного ...

УСТРОЙСТВО ГОРЕЛКИ

Использование: в энергетике и относится к горелкам для текучих видов топлива. Сущность: устройство горелки с практически концентрично расположенными кольцевыми каналами для подачи различных рабочих сред /В, С, Е, F/ и внешней практически конически сужающейся системой кольцевых каналов для подвода воздуха, причем предусмотрено множество выпускных форсунок для подмешивания тонко распределенной газообразной среды /В/ или жидкой среды /С/ к воздушному потоку /А/, протекающему в системе кольцевых каналов для подвода воздуха, причем далее со стороны притока выше выпускных форсунок в систему кольцевых каналов для подвода воздуха впадает дополнительный кольцевой канал. Дополнительный кольцевой канал служит для подвода горючего газа с низкой удельной теплотой сгорания, например каменноугольного газа, в то время как устройство горелки согласно уровню техники может эксплуатироваться как горелка с предварительным смешиванием на газе и/или нефти. При всех применяемых видах топлива и видах эксплуатации ...

КАМЕРА СГОРАНИЯ ГТД И ФОРСУНОЧНЫЙ МОДУЛЬ

Камера сгорания ГТД содержит корпус, жаровую трубу, имеющую внешнюю и внутреннюю стенки и плиту кольцевой формы с установленными на ней форсуночными модулями и основной топливный коллектор, соединенный с плитой, полость которого соединена топливными каналами с форсуночными модулями, внешний и внутренний корпусы, внешний и внутренний кожуха, установленные с зазором относительно внешнего и внутреннего корпусов. Число форсуночных модулей выполнено кратным четырем. Форсуночные модули установлены в два ряда: внешний и внутренний. Дополнительно выполнено два топливных коллектора внешний и внутренний. Полость внешнего коллектора соединена топливными каналами с каждым форсуночным модулем через один внешнего ряда форсуночных модулей. Полость внутреннего коллектора соединена с каждым форсуночным модулем через один внутреннего ряда. Основной топливный коллектор соединен с остальными форсуночными модулями обеих рядов. Между плитой и внешней и внутренней стенками жаровой трубы установлены соответственно ...

Малоэмиссионная камера сгорания и способ подачи в ней топлива

Малоэмиссионная камера сгорания газотурбинного двигателя содержит как минимум два топливных коллектора: основной и пилотный, кран перераспределения топлива, как минимум два горелочных устройства, каждое из которых снабжено криволинейным каналом, образованным двумя спрофилированными обечайками. Внутренняя поверхность криволинейного канала сопряжена с центральным телом горелочного устройства, на котором установлен аксиальный завихритель и трубки подачи топлива в проточную часть горелочного устройства от основного топливного коллектора. На внешней поверхности криволинейного канала размещен выравнивающий коллектор для подвода топлива через соответствующие трубки из пилотного коллектора в проточную часть горелочного устройства, которое соединено с жаровой трубой, имеющей отверстия для подачи воздуха, внутри которой соосно с ней расположен экран, состоящий из поперечной кольцевой части и продольной цилиндрической части. Экран и жаровая труба образуют проточную полость для воздуха. На входе жаровой ...

СЖИГАЮЩЕЕ УСТРОЙСТВО ГАЗОТУРБИННОЙ УСТАНОВКИ

Изобретение относится к сжигающему устройству газотурбинной установки. Сжигающее устройство газотурбинной установки содержит горелку для горения предварительно приготовленной смеси и камеру сгорания для сжигания топлива и воздуха, подаваемых из горелки для горения предварительно приготовленной смеси, причем горелка для горения предварительно приготовленной смеси содержит топливную форсунку для впрыска топлива, подаваемого из системы подачи топлива, и канал для предварительно приготовленной смеси для смешения топлива, впрыскиваемого из топливной форсунки, и воздуха, подаваемого из воздушного канала, и подачи топливовоздушной смеси в камеру сгорания, при этом топливная форсунка содержит: сужающийся участок, наружный диаметр которого постепенно уменьшается от ближней стороны к дальней стороне топливной форсунки, плоский участок, проходящий от сужающегося участка в направлении дальней стороны топливной форсунки и имеющий постоянный наружный диаметр от ближней стороны к дальней стороне топливной ...

СПОСОБ РЕГУЛИРОВАНИЯ ПОДАЧИ ТОПЛИВА В КАМЕРУ СГОРАНИЯ ГАЗОТУРБИННОЙ УСТАНОВКИ И УСТРОЙСТВО ДЛЯ ЕГО ОСУЩЕСТВЛЕНИЯ

... 1. Способ регулирования подачи топлива в камеру сгорания газотурбинной установки, включающий изменение расхода топлива в зависимости от ее мощности путем дозирования подачи топлива в коллектора дежурных и основных горелок горелочных устройств с предварительным смешением топлива и воздуха, отличающийся тем, что подачу топлива осуществляют в два яруса горелочных устройств, причем на режиме запуска топливо подают в коллектор дежурных горелок наружного яруса, перед выходом на режим малого газа - в коллектор дежурных горелок внутреннего яруса, на режиме малого газа поддерживают близким по величине количество топлива, подаваемого в дежурные горелки наружного и внутреннего ярусов, затем увеличивают подачу топлива в дежурные горелки наружного и внутреннего ярусов, перед выходом на режим холостого хода подают топливо к основным горелкам наружного и внутреннего ярусов, а в диапазоне режимов от холостого хода до номинального увеличивают подачу топлива в основные горелки, при этом уменьшают относительную ...

УСТРОЙСТВО ДЛЯ ЗАКРЕПЛЕНИЯ КАНАЛА ДЛЯ ТЕКУЧЕЙ СРЕДЫ В КОРПУСЕ ТУРБОРЕАКТИВНОГО ДВИГАТЕЛЯ

... 1. Устройство для закрепления канала для текучей среды в отверстии корпуса турбореактивного двигателя, в частности, канала для подачи топлива к кольцу форсунок в форсажной камере, причем устройство содержит средство типа винта и гайки между концом канала и отверстием корпуса, при этом устройство содержит кольцо, ввинченное в отверстие корпуса и имеющее один конец, который упирается в упорное средство, установленное в канале, и гайку, навинченную на конец канала таким образом, что она прижимает кольцо к упорному средству канала для прикрепления его к корпусу. 2. Устройство по п.1, в котором упорное средство канала содержит шайбу, посаженную на конец канала и упирающуюся в выступ указанного канала. 3. Устройство по п.2, в котором шайба имеет центральное отверстие в форме усеченного конуса, входящее в контакт с соответствующей поверхностью в форме усеченного конуса канала. 4. Устройство по п.2, в котором указанная шайба удерживается в обойме, образующей внутренний конец кольца. 5. Устройство ...

ТОПЛИВНЫЙ ИНЖЕКТОР С КАМЕРОЙ ПРЕДВАРИТЕЛЬНОГО СМЕШИВАНИЯ С ЗАЩИТНЫМ ПОКРЫТИЕМ, НАПЛАВЛЕННЫМ ЛАЗЕРОМ

... 1. Топливный инжектор (12) газотурбинного двигателя (10), содержащийцентральное тело (36), расположенное на продольной оси (50), икамеру (26) предварительного смешивания, расположенную в радиальном направлении внешне относительно центрального тела, и образующую кольцевой канал (16) между ними, проходящий от входного патрубка, соединенного с возможностью передачи потока с компрессором (2) газотурбинного двигателя, к выходному патрубку, соединенному с возможностью передачи потока с камерой сгорания (14), при этом камера предварительного смешивания содержитпервый участок (24), расположенный на входном патрубке и состоящий из нержавеющей стали, ивторой участок (22), расположенный на выходном патрубке, состоящий из жаропрочного сплава на основе никеля и присоединенный к первому участку посредством лазерной наплавки.2. Топливный инжектор по п. 1, в котором второй участок камеры предварительного смешивания имеет длину примерно 12,7-63,5 мм и толщину примерно 3,81-12,7 мм.3. Топливный инжектор ...

ГАЗОТУРБИННЫЙ ДВИГАТЕЛЬ, ФОРСУНКА ФОРСАЖНОЙ КАМЕРЫ (ВАРИАНТЫ) И СПОСОБ МОДЕРНИЗАЦИИ ФОРСАЖНОЙ КАМЕРЫ

Газотурбинный двигатель снабжен центральной ступенью, расположенной в газовом тракте двигателя от его вышерасположенной по направлению основного газового потока части до нижерасположенной части и имеющей нижерасположенное по направлению основного газового потока устройство факела выхлопных газов и направляющее устройство. Газотурбинный двигатель также снабжен группой лопаток, группой топливных форсунок и группой воспламенителей. Направляющее устройство расположено в зоне вышерасположенного по направлению основного газового потока края устройства факела выхлопных газов. Группа лопаток расположена в газовом тракте за пределами центральной ступени. Лопатки содержат простирающиеся сквозь них распылительные направляющие. Топливные форсунки установлены на внутренних концах соответствующих распылительных направляющих. Каждая форсунка имеет вход, выход и проход между входом и выходом. Проход имеет часть, расположенную с возможностью направления потока топлива на первую часть поверхности прохода ...

КАМЕРА СГОРАНИЯ ГАЗОВОЙ ТУРБИНЫ И ГАЗОВАЯ ТУРБИНА

Изобретение относится к камере сгорания газовой турбины и газовой турбине. Камера сгорания газовой турбины принимает сжатый воздух из компрессора, смешивает сжатый воздух с топливом, сжигает смесь для получения газа сгорания и подает газ сгорания в турбину. Камера сгорания включает в себя: внутренний цилиндр, образующий внутри пространство камеры сгорания; внешний цилиндр, охватывающий внутренний цилиндр и образующий цилиндрический внешний круговой канал потока между внутренним и внешним цилиндрами для прохождения сжатого воздуха; и горелку, установленную на конце внешнего цилиндра, который расположен с противоположной стороны относительно стороны турбины, и обращенную в пространство камеры сгорания. Горелка включает в себя цилиндрический основной корпус, включающий в себя полость, распределяющую топливо, и топливные форсунки, расположенные по кругу, если смотреть из пространства камеры сгорания, и соединенные с полостью. Если смотреть из пространства камеры сгорания, в основном корпусе ...

КАМЕРА СГОРАНИЯ ГАЗОВОЙ ТУРБИНЫ

Камера сгорания газовой турбины включает в себя кромку завихрителя, располагающуюся на внешней границе пластины с отверстиями для воздуха с выступанием в сторону полости сгорания, и пружинное уплотнение, установленное на участке сопряжения вкладыша камеры сгорания с пластиной с отверстиями для воздуха и кромкой завихрителя. Пружинное уплотнение имеет отверстие для воздуха, предназначенное для обеспечения прохождения части воздуха для горения, введенного в зазор между внешней границей пластины с отверстиями для воздуха и вкладышем камеры сгорания. Кромка завихрителя имеет отверстие для воздуха, предназначенное для ввода части воздуха для горения, прошедшего через отверстие для воздуха в пружинном уплотнении, в полость сгорания. Вкладыш камеры сгорания имеет отверстие для воздуха в положении напротив отверстия для воздуха в пружинном уплотнении, предназначенное для ввода части воздуха для горения снаружи вкладыша камеры сгорания в зазор между внешней границей пластины с отверстиями для воздуха ...

КОЛЬЦЕВАЯ КАМЕРА СГОРАНИЯ ГАЗОТУРБИННОГО ДВИГАТЕЛЯ И СПОСОБ ЕЁ ЭКСПЛУАТАЦИИ

Кольцевая камера сгорания газотурбинного двигателя содержит группу горелок, расположенных в одной плоскости на передней стенке камеры сгорания, по меньшей мере, двумя соосными кольцами. В пределах каждого кольца установлено одинаковое и четное число малоэмиссионных горелок. Горелки внутреннего кольца смещены в окружном направлении относительно горелок наружного кольца на их пол шага. Все горелки выполнены двухканальными. Внутренние каналы горелок служат для подачи в них только пилотного топлива, а наружные каналы горелок - для подачи в них сжатого воздуха из-за компрессора и основного топлива с образованием «бедной» топливовоздушной смеси. Наружный канал каждой горелки содержит входной направляющий аппарат, в стенках которого выполнены отверстия для подачи топлива в сносящий поток воздуха, лопаточный завихритель, установленный на выходе из канала, и проницаемый элемент с заданной пористостью, установленный между входным направляющим аппаратом и лопаточным завихрителем. Направление закрутки ...

КАМЕРА СГОРАНИЯ С ОПТИМАЛЬНЫМ РЕЖИМОМ РАБОТЫ

Камера сгорания, содержащая корпус, жаровую трубу с отверстиями для подвода воздуха в зоны горения и смешения, и фронтовое устройство с завихрителями воздуха и форсунками подачи топлива, отличающаяся тем, что жаровая труба камеры сгорания выполнена с геометрическими и газодинамическими критериями: отношение ее площади поперечного миделевого сечения к суммарной эффективной площади всех отверстий , относительная пропускная способность завихрителей , интенсивность закрутки потока воздуха завихрителями фронтового устройства и коэффициент скорости потока в отверстиях жаровой трубы , обеспечивающими оптимальный режим ее работы.

ГОРЕЛКА КАМЕРЫ СГОРАНИЯ ГАЗОТУРБИННОГО ДВИГАТЕЛЯ

Предложено устройство горелки камеры сгорания газотурбинного двигателя, в котором для уменьшения токсичности выхлопных газов двигателя в каждой горелке имеются две зоны горения: дежурная и основная. В дежурной зоне процесс горения организован с коэффициентом избытка воздуха α = 1,2 - 1,4, а в основной с a =1,8 - 2,0.

СОЕДИНИТЕЛЬНОЕ УСТРОЙСТВО, СОДЕРЖАЩЕЕ НЕСКОЛЬКО ИЗОГНУТЫХ КОНЦЕНТРИЧЕСКИХ ТРУБ

ФРОНТОВОЕ УСТРОЙСТВО ЖАРОВОЙ ТРУБЫ КОЛЬЦЕВОЙ КАМЕРЫ СГОРАНИЯ

... 1. Фронтовое устройство жаровой трубы кольцевой камеры сгорания, содержащее головку кольцевую с наружным и внутренним топливными коллекторами, с равномерно расположенными по окружности наружного топливного коллектора, штуцерами подвода газа во внутреннюю полость наружного коллектора и выполненными на головке кольцевой между коллекторами, концентрично и равномерно расположенными по окружности в один ряд воздушными фигурными окнами, подвода воздуха в первичную зону горения, с центральными отверстиями и стойками крепления горелок к головке кольцевой, при этом в стойках выполнены сквозные каналы подвода газа к горелкам, которые через горелки соединяют между собой полости наружного и внутреннего топливных коллекторов, отличающееся тем, что с целью подвода газа к горелкам равномерно с двух сторон, как со стороны наружного, так и со стороны внутреннего топливных коллекторов, в перемычках между воздушными фигурными окнами, выполнены дополнительные каналы, которые соединяют между собой полость внутреннего ...

УПЛОТНИТЕЛЬНОЕ УСТРОЙСТВО МЕЖДУ СИСТЕМОЙ ВПРЫСКА И ТОПЛИВНОЙ ФОРСУНКОЙ АВИАЦИОННОГО ГАЗОТУРБИННОГО ДАВЛЕНИЯ

ГОРЕЛКА, В ЧАСТНОСТИ, ДЛЯ ГАЗОВЫХ ТУРБИН

... 1. Горелка с центральной компоновкой (27) подачи топлива и охватывающим центральную компоновку (27) подачи топлива кольцевым воздушным каналом (17) для подачи топочного воздуха и с расположенными в кольцевом воздушном канале (17) вихревыми лопатками (19), имеющими первые газовые форсунки (21) для впрыскивания газообразного топлива в топочный воздух и вторые газовые форсунки (23) для впрыскивания газообразного топлива в топочный воздух, причем первые газовые форсунки (21) питаются от первого газораспределительного канала (29) в компоновке (27) подачи топлива, а вторые газовые форсунки (23) - от второго газораспределительного канала (31) в компоновке (27) подачи топлива, причем первый газораспределительный канал (29) и второй газораспределительный канал (31) снабжаются горючим газом от трубы (41, 141) подачи газа, расположенной со смещением к расположенной в центре пилотной горелке и рядом параллельно к ней и имеющей первый канал (35, 135) подачи газа и второй канал (37, 137) подачи газа, ...

НЕСУЩИЙ КОРПУС ТОПЛИВНОЙ ФОРСУНКИ ДЛЯ ГАЗОТУРБИННОГО ДВИГАТЕЛЯ (варианты)

... 1. Несущий корпус топливной форсунки для газотурбинного двигателя, отличающийся тем, что он имеет продольную ось и содержит удлиненный, в основном цилиндрический ствол, имеющий наружный диаметр (D0), длину (L) и выполненный симметричным относительно оси, входную торцевую часть с входными отверстиями первого и второго топливных контуров, выходную торцевую часть с выходными отверстиями первого и второго топливных контуров, концентрические продольный канал второго топливного контура и трубку первого топливного контура, причем упомянутый канал имеет внутренний диаметр (D1) и связывает входное отверстие второго топливного контура и выходное отверстие второго топливного контура, а упомянутая трубка расположена в канале с герметизацией относительно входной и выходной торцевых частей несущего корпуса и связывает входное отверстие первого топливного контура и выходное отверстие первого топливного контура, при этом обеспечиваются следующие соотношения: D0 >2D1, а L<10·(D0-D1). 2. Корпус по п.1, отличающийся ...

ТОПЛИВНАЯ ФОРСУНКА КАМЕРЫ СГОРАНИЯ ГАЗОТУРБИННОГОДВИГАТЕЛЯ

Топливная форсунка камеры сгорания газотурбинного двигателя содержит корпус с каналами и со штуцерами основного и дополнительного контуров подвода топлива на основное и дополнительное сопла, расположенные в головке форсунки, а также установочный фланец крепления форсунки к наружному корпусу камеры сгорания. Внутри канала основного контура размещена разделительная трубка, внутри которой расположен канал дополнительного контура. Разделительная трубка закреплена с одного конца в головке форсунки относительно корпуса неразъемным соединением. Разделительная трубка с другого конца установлена внутри корпуса с внешней стороны от установочного фланца, между штуцерами основного и дополнительного контуров, телескопически относительно стенок канала основного контура цилиндрическим выступом и контактирует одним торцом с упругим уплотнительным цилиндрическим кольцом, контактирующим наружным диаметром с корпусом форсунки, а другим торцом и внутренней поверхностью - с резьбовой заглушкой. Резьбовая заглушка ...

КАМЕРА СГОРАНИЯ ГАЗОВОЙ ТУРБИНЫ

В изобретении предложена камера сгорания газовой турбины, обладающая конструкционной надежностью по отношению к вибрации топливных форсунок, обусловленной действием текучей среды, и высокой экологической эффективностью за счет равномерного сгорания в секции камеры сгорания. Камера сгорания газовой турбины содержит топливную форсунку, выполненную с возможностью впрыска топлива, и пластину с топливными форсунками, включающую в себя участок с отверстием, в который вставлен вставленный участок топливной форсунки, располагающийся со стороны основания этой топливной форсунки. Топливная форсунка включает в себя наружную резьбу, по меньшей мере, на внешней окружной поверхности нижнего по потоку участка вставленного участка, если смотреть в направлении потока топлива. Пластина с топливными форсунками включает в себя внутреннюю резьбу на участке с отверстием, вкрученную в наружную резьбу вставленного участка. Топливная форсунка включает в себя вставленный участок, верхний по потоку торец которого ...

КАМЕРА СГОРАНИЯ ГАЗОВОЙ ТУРБИНЫ И ТОПЛИВНЫЙ ИНЖЕКТОР ГАЗОТУРБИННОГО ДВИГАТЕЛЯ

Камера сгорания газовой турбины имеет первичную, вторичную и третью зоны сгорания по направлению потока, канал вторичного смешивания и канал третьего смешивания. Площадь поперечного сечения каналов вторичного и третьего смешивания уменьшается от их вкускных устройств до их выпускных отверстий для создания ускоренного потока через каналы смешивания, чтобы предотвратить образование зон рециркуляции. Топливные инжекторы имеют отверстия для выпуска топлива ниже по потоку относительно любых зон рециркуляции, могущих образоваться у впускных устройств. Топливные инжекторы проходят через основной участок каналов так, чтобы эффективно подразделить каналы по, как минимум, части их длины. Участки топливных инжекторов, находящиеся внутри каналов, имеют в поперечном сечении форму гоночного трека, а участки вне каналов имеют в поперечном сечении форму профиля крыла. Топливные инжекторы имеют уменьшающиеся в поперечном направлении размеры по ширине каналов.

ФОРСУНКА КАМЕРЫ СГОРАНИЯ, ГАЗОВАЯ ТУРБИНА И СПОСОБ, ВКЛЮЧАЮЩИЙ СМЕШИВАНИЕ ВОЗДУХА И ТОПЛИВА

... 1. Форсунка (28) камеры сгорания, содержащаясекцию (30) смешивания, имеющую впуск (38) для воздуха и впуск (36) для топлива, ивыпускную секцию (32), имеющую каналы (44), выполненные с созданием определенной конфигурации на выпускной поверхности (42), причем отношение рабочей площади каналов (44) к площади выпускной поверхности превышает 0,25.2. Форсунка (28) по п.1, в которой указанная конфигурация является гексагональной конфигурацией.3. Форсунка (28) по п.1, в которой указанная конфигурация является искаженной гексагональной конфигурацией.4. Форсунка (28) по п.1, в которой указанная конфигурация является квадратичной конфигурацией.5. Форсунка (28) по п.1, в которой указанная конфигурация является круговой конфигурацией.6. Форсунка (28) по п.1, в которой отношение рабочей площади каналов (44) к площади выпускной поверхности составляет от 0,25 до 0,4.7. Форсунка (28) по п.1, в которой отношение рабочей площади каналов (44) к площади выпускной поверхности составляет от 0,4 до 0,5.8. Форсунка ...

КАМЕРА СГОРАНИЯ ГАЗОВОЙ ТУРБИНЫ

... 1. Камера сгорания (10)для газовой турбины, содержащая:переднюю панель (14);продолговатый рукав (20) с первым концом (22) и вторым концом (24); игорелку (30), установленную в рукаве (20),отличающаяся тем, что второй конец (24) рукава (20) установлен на передней панели (14) без уплотнения.2. Камера сгорания (10) по п. 1, в которой горелка (30) выполнена с возможностью перемещения в осевом направлении внутри рукава (20) и относительно рукава (20) во время работы камеры сгорания (10).3. Камера сгорания (10) по п. 1 или 2, в которой горелка (30) установлена с возможностью скольжения в рукаве (20), чтобы обеспечивать осевую регулировку горелки (30) внутри рукава (20).4. Камера сгорания (10) по любому из пп. 1 и 2, в которой рукав (20), у второго конца (24), имеет форму колокола.5. Камера сгорания (10) по любому из пп. 1 и 2, в которой горелка (30) содержит:корпус (31) горелки; иконусообразно расширяющуюся завихряющую оболочку (33), проходящую из корпуса (31), имеющую узкий первый конец (34) ...

ГОРЕЛКА И СПОСОБ ЭКСПЛУАТАЦИИ ГОРЕЛКИ

... 1. Способ эксплуатации горелки (107), которая включает в себя выходное отверстие (4) горелки с, по меньшей мере, двумя секторами (8a, 8b, 9a, 9b), причем каждому сектору (8a, 8b, 9a, 9b) соответствует, по меньшей мере, одна топливная форсунка, причем топливо подают отдельно на топливные форсунки разных секторов (8a, 8b, 9a, 9b), отличающийся тем, что ! - в режиме полной нагрузки предусмотрена равномерная подача топлива во все сектора (8a, 8b, 9a, 9b) так, что возникает однородное распределение температуры, и ! - в режиме частичной нагрузки в камере сгорания можно создавать более горячие и более холодные зоны, причем более горячие зоны создают там, где ожидаются наибольшие эффекты гашения. ! 2. Способ по п.1, отличающийся тем, что на топливные форсунки разных секторов (8a, 8b, 9a, 9b) подают топливо в соотношении от 0:100 до 100:0. ! 3. Способ по п.2, отличающийся тем, что на топливные форсунки разных секторов (8, 9) подают топливо в соотношении от 100:0 до 35:65. ! 4. Способ по любому из ...

Камера сгорания автомобильного газотурбинного двигателя

Кольцевая камера сгорания газотурбинного двигателя

Использование: в газотурбостроении. Сущность изобретения: топливные форсунки (Тф) установлены по окружности в два ряда и имеют размещенные вокруг них за- вихрители 3. Каждая из ТФ одного ряда установлена в окружном направлении между соседними ТФ другого ряда, а завихрители 3 имеют направление закрутки потока, одинаковое для всех ТФ одного ряда и противоположные для всех ТФ в другом ряду.

Brennstoffeinspritzeinrictung für mit gasförmigem oder flüssigem Brennstoff betriebene Turbine

Stabilisierung von Druckschwingen in Verbrennungsvorrichtungen sowie Verfahren hierfür

High dynamic pressure oscillations in hydrocarbon-fueled combustors typically occur when the transport time of the fuel to the flame front is at some fraction of the acoustic period. These oscillations are reduced to acceptably lower levels by restructuring or repositioning the flame front in the combustor to increase the transport time. A pilot flame front 12 located upstream of the oscillating flame 20 and pulsed at a selected frequency and duration effectively restructures and repositions the oscillating flame in the combustor to alter the oscillation-causing transport time. The pilot frame front is pulsed by intermiltently interrupting the flow of at least one of the fuel supply 18 and the oxidiser supply 15 or of the mixture of the fuel supply and the oxidiser supply.

ZWEIKRAFTSTOFFEINSPRITZDÜSE ZUM GEBRAUCH IN EINEM GASTURBINENTRIEBWERK

Verbrennungskammer mit gestufter Brennstoffeinspritzung

Brennstoffeinspritzdüse

GASTURBINENBRENNKAMMER.

INJECTION DEVICE FOR THE COMBUSTION CHAMBER OF A TURBINE ENGINE

Düse, Brenner, Brennkammer, Gasturbine und Gasturbinensystem

Eine Brennkammer (10) ist mit Folgendem vorgesehen: einer Düse (25), die darin einen Luftdüsenabschnitt (51) gebildet hat, der bewirkt, dass Luft von einem Düsenabschnitt (25s) ausgestoßen wird; einem zylindrischen Teil (26), das die Düse (25) von der Außenumfangsseite davon abdeckt und einen Luftströmungsweg zwischen dem zylindrischen Teil (26) und der Düse (25) bildet; und einem Druckverlustabschnitt (27), der am Luftströmungsweg vorgesehen ist, wobei der Druckverlustabschnitt (27) einen Druckverlust der Luft, die durch den Luftströmungsweg fließt, bewirkt. Die Düse (25) ist mit Folgendem vorgesehen: zumindest einem Lufteinlassabschnitt (52), der Luft von einer Außenumfangsfläche, die eine stromaufwärtige Seite des Druckverlustabschnitts (27) ist, einlässt; und einem Strömungswegbildungsabschnitt (50), der einen Abluft-Strömungsweg (R2) bildet, der die Luft, die von zumindest einem Lufteinlassabschnitt (52) eingelassen wird, zum Luftdüsenabschnitt (51) leitet.

GASTURBINENBRENNKAMMERVORRICHTUNG UND BRENNSTOFFDÜSENHERSTELLUNGSVERFAHREN

In der Gasturbinenbrennkammervorrichtung, die die Brennstoffdüse enthält, die durch additive 3D-Herstellung geformt ist, wird eine Gasturbinenbrennkammervorrichtung geschaffen, die eine Brennstoffdüse enthält, die eine hohe Dämpfungsleistung gegen durch instabile Verbrennung verursachte Schwingungsbelastung aufweist. In der Gasturbinenbrennkammervorrichtung, die die Brennstoffdüse enthält, die durch die additive 3D-Herstellung geformt ist, weist die Brennstoffdüse ein erstes Gebiet, in dem Metallpulver gesintert sind, und ein zweites Gebiet, das von dem ersten Gebiet umgeben ist und in dem die Metallpulver nicht gesintert sind, auf.

Steuerung einer Brennstoffversorgung für eine Gasturbine mit mehreren Magnetventilen

Reparierbare Brennstoffdüse und Reparaturverfahren

Verfahren zum Reparieren einer Brennstoffdüse (100), das die Schritte aufweist: Entfernen einer alten Pilot-Anordnung (607); Bilden einer neuen Pilot-Anordnung (808); Einbauen der neuen Pilot-Anordnung (808) in eine Reparaturbohrung (801); Vorbereiten eines primären Adapters (820) für Montage; Einsetzen des primären Adapters (820) in die Reparaturbohrung (801), so dass der primäre Adapter (820) mit einem Pilot-Strömungskanal (102) in einem Verteiler (300) und der neuen Pilot-Anordnung (808) in Strömungsverbindung steht; und Koppeln des primären Adapters (820) mit dem Verteiler (300) und der neuen Pilot-Anordnung (808).

VERWIRBELUNGSANORDNUNG FUER DIE BRENNKAMMER EINES GASTURBINENTRIEBWERKS

Gasturbinenbrennkammer mit gezielter Kraftstoffeinbringung zur Verbesserung der Homogenität des Kraftstoff-Luft-Gemisches

Die Erfindung bezieht sich auf eine Gasturbinenbrennkammer mit einem Brenner 7 und Mitteln zur Kraftstoffzufuhr sowie einem Zerstäuber 6, dadurch gekennzeichnet, dass die Mittel zur Kraftstoffzufuhr zur Einspritzung des Kraftstoffs in Bereiche mit größten Luftströmungsgeschwindigkeiten ausgebildet sind.

Verfahren und Vorrichtung zur Beeinflussung thermoakustischer Schwingungen in Verbrennungssystemen

Die vorliegende Erfindung betrifft ein Verfahren und eine Vorrichtung (1) zur Beeinflussung thermoakustischer Schwingungen in einem Verbrennungssystem (6), umfassend wenigstens einen Brenner (7) und wenigstens eine Brennkammer (8). DOLLAR A Um die Beeinflussung der thermoakustischen Schwingungen zu verbessern, DOLLAR A - wird eine sich im Bereich des Brenners (7) ausbildende Gasströmung akustisch angeregt, DOLLAR A - erfolgt eine Eindüsung von Brennstoff moduliert, DOLLAR A - sind die akustische Anregung der Gasströmung und die modulierte Eindüsung des Brennstoffs zur Beeinflussung derselben Störfrequenz abgestimmt.

Fuel preparation device for gas turbine combustion chamber

The device has two mixing shells (1), spaced in a conical shape. Fuel in applied to their outer walls and is entrained by an air stream impinging onto the shells and forming swirl apices downstream of the shells. To the latter is coupled a mixing, a pre-evaporation tube (5) of variable cross-section over its length. Pref. the cross-sectional variation is formed by a surface variation. Typically the tube inlet cross-section is elliptical, while the outlet cross-section is circular, or also elliptical, but with inverse ratio of the ellipse axes w.r.t. inlet ellipse.

Aerodynamisches Einspritzsystem zur Einspritzung eines Brennstoff-Luft-Gemisches

KRAFTSTOFFDUESE

System mit einer Vielrohr-Brennstoffdüse

Ein System enthält mehrere Vielrohr-Brennstoffdüsen, die jeweils mehrere Rohre aufweisen, die sich in einer axialen Richtung erstrecken, wobei jedes Rohr von den mehreren Rohren einen Lufteinlass, einen Brennstoffeinlass und einen Brennstoff-Luft-Mischungsauslass aufweist, ein Brennstoffdüsengehäuse mit einer ersten Außenwand, die sich in Umfangsrichtung um eine Mittelachse erstreckt, wobei die mehreren Vielrohr-Brennstoffdüsen in dem Brennstoffdüsengehäuse angeordnet sind, einen Einlassströmungskonditionierer, der mit einem ersten Endabschnitt der ersten Außenwand lösbar verbunden ist, wobei der Einlassströmungskonditionierer mehrere Luftöffnungen enthält, und eine rückseitige Plattenanordnung, die mit einem zweiten Endabschnitt der ersten Außenwand lösbar verbunden ist, wobei die rückseitige Plattenanordnung eine rückseitige Platte mit mehreren Rohröffnungen aufweist und die mehreren Rohre sich zu den mehreren Rohröffnungen erstrecken.

BRENNER FUER EIN GASTURBINENTRIEBWERK

Verfahren zum Betrieb einer Brennkammer und Brennkammer

Düsenbaugruppe für eine Brennkammer eines Triebwerks

Die vorliegende Erfindung betrifft eine Düsenbaugruppe für eine Brennkammer (3) eine Triebwerks (T), mit mehreren nebeneinander angeordneten Düsen (2A, 2B, 2C) zur Einbringung von Treibstoff in die Brennkammer (3), wobei jede Düse (2A, 2B, 2C) eine Düsenaustrittsöffnung (210) und einen Treibstoffkanal (220) zur Förderung von Treibstoff in Richtung der Düsenaustrittsöffnung (210) aufweist. Erfindungsgemäß ist vorgesehen, dass mindestens zwei unterschiedliche Typen von Düsen (2A, 2B, 2C) vorgesehen sind, wobei- die Düsen (2A, 2B, 2C) unterschiedlicher Typen Düsenaustrittöffnungen (210) mit identischem Querschnitt aufweisen, und- zur Vorgabe unterschiedlicher Durchflussmengen an Treibstoff durch die Treibstoffkanäle der Düsen (2A, 2B, 2C) unterschiedlicher Typen ein Querschnitt eines Treibstoffkanals (220) eines Typs von Düse (2A, 2B, 2C) zu einem Querschnitt eines Treibstoffkanals (220) eines anderen Typs von Düse (2B, 2C, 2B) verschieden ist.

Brenner mit einer zentralen Brennstoffzufuhranordnung

Die Erfindung betrifft einen Brenner (1) mit einer zentralen Brennstoffzufuhranordnung (2), einem die zentrale Brennstoffzufuhranordnung (2) umgebenden Ringluftkanal (3) zur Zuführung von Verbrennungsluft, ersten Brennstoffdüsen (4) zum Eindüsen eines im Wesentlichen gasförmigen Brennstoffs in den Ringluftkanal (3), wobei die ersten Brennstoffdüsen (4) von einem ersten Brennstoffverteiler (5) in der zentralen Brennstoffzufuhranordnung (2) gespeist werden, und zweiten Brennstoffdüsen (6) zum Eindüsen eines im Wesentlichen flüssigen Brennstoffs in den Ringluftkanal (3), wobei die zweiten Brennstoffdüsen (6) von einem zweiten Brennstoffverteiler (7) in der zentralen Brennstoffzufuhranordnung (2) gespeist werden, wobei der zweite Brennstoffverteiler (7) in der zentralen Brennstoffzufuhranordnung (2) im Wesentlichen vom Rest der zentralen Brennstoffzufuhranordnung (2) thermisch entkoppelt ist.

Method for charging burner for gas turbines with pilot gas involves supplying pilot gas at end of burner cone in two different flow directions through pilot gas pipes set outside of burner wall

The burner is operated with a premix gas and/or pilot gas wherein the premix gas is prepared in a burner cone (6) and the pilot gas is supplied at the end of the burner cone facing the combustion zone. The pilot gas can be supplied in at least two different flow directions. Independent claim describes burner where pilot gas supply system (3) is connected to two pilot gas pipes (9) set outside of the burner wall and bent inwards at their free ends.

Verfahren und Vorrichtung zur Kühlung von Gasturbinenbrennkammern

Brenneranordnung mit primärem und sekundärem Pilotbrenner

System und Verfahren zur Steuerung der Verbrennungsdynamik in einem Verbrennungssystem

Die vorliegende Offenbarung betrifft allgemein ein System mit einer Gasturbine, die eine erste Brennkammer und eine zweite Brennkammer enthält. Die erste Brennkammer enthält eine erste Endabdeckung mit einer ersten Geometrie, und die zweite Brennkammer enthält eine zweite Endabdeckung mit einer zweiten Geometrie. Die erste Geometrie hat relativ zu der zweiten Geometrie einen oder mehrere geometrische Unterschiede.

Emulsifying residual oil with water - for use as fuel in gas turbine, giving clean, efficient combustion

In the processing of residual oil for use in gas turbines, esp. in jet transmission appts. in aircraft or ships, 3-8 (5) wt.% of water is added to the residual oil, an oil-water emulsion is formed in which the water droplets are dispersed in a continuous oil phase and the emulsion is led to a gas turbine for combustion. Cheap residual oils can be utilised. Overheating of the jet appts. is avoided. The emulsion is stable and can be stored for a relatively long time without sepn. In the gas turbine chamber the water droplets undergo 'micro-explosion' into steam, leading to formation of fuel droplets of 25 mu, compared with 70150 mu in known processes. Complete, clean combustion is achieved in the short combustion chamber of the gas turbine. Addn. of water leads to a water gas reaction during combustion, increasing the combustion efficiency, and also lowers the peak temp. in the combustion chamber.

Mehrstufiges, mehrebenes Vebrennungssystem für Gasturbine

Düsencluster

Die Erfindung betrifft einen Düsencluster umfassend eine Vielzahl von Paaren aus jeweils einer ersten (1) und einer zweiten Gasauslassöffnung (3). Bei jedem Paar ist die zweite Gasauslassöffnung (3) um die erste Gasauslassöffnung (1) zumindest teilweise umlaufend angeordnet. Die Erfindung betrifft weiterhin einen mindestens einen erfindungsgemäßen Düsencluster umfassenden Brenner, eine einen oder mehrere erfindungsgemäße Brenner umfassende Gasturbine, einen mindestens einen erfindungsgemäßen Düsencluster umfassenden Mischer sowie ein Verfahren zur Herstellung eines erfindungsgemäßen Düsenclusters durch additive Fertigung mittels 3D Druck.

Gasturbinenbrennkammer

Brennkammer

Vorrichtung zur Führung einer Einrichtung in eine Öffnung einer Gasturbinenbrennkammerwand

Doppelbrennstoffmischer für eine Gasturbinenbrennkammer

Fuel/air mixing arrangement for combustion apparatus

Improvements in or relating to burner assemblies for internal combustion turbines

... 614,553. Burners. OULIANOFF, G. March 26, 1946, No. 9344. [Class 75 (i)] [Also in Group XII] In a combustion chamber and burner assembly for a gas turbine engine of the kind in which the combustion chamber comprises an outer casing within which is supported a flame tube and wherein the fuel burner jet introduces fuel into the flame tube at its inlet end, the burner jet support is in the form of a spider comprising a rim portion supported by the outer casing of the combustion chamber, a hub portion supporting the burner. jet parts and the flame tube at its inlet end, and a plurality of spokes extending between the rim and hub one at least of these spokes being formed as a conduit to convey liquid fuel to the burner jet. A burner assembly 10, comprising a circular rim 21, three spider arms or spokes 22, 22a, and a hub 23 which supports the burner and the inlet end of the flame tube, mounted in the cylindrical recess 25 of the inlet casting 24 of a combustion chamber. The rim is provided with ...

CHAMBER ASSEMBLY FOR USE IN A CONTINUOUS COMBUSTION PROCESS

... 1414412 Gas turbine combustion chambers FORD MOTOR CO Ltd 7 May 1973 [25 May 1972] 21556/73 Heading F1L Cumbustion equipment comprises a mixing chamber having an inlet and an outlet there being means adjacent the inlet for producing a toroidal swirl in air under pressure flowing through the mixing chamber also means for supplying fuel adjacent the inlet, the mixing chamber opening into a combustion chamber, there being ignition means to cause ignition of the fuel/air mixture. A further chamber is provided in which the gases from the combustion chamber are cooled. In Fig. 1 the combustion equipment comprises a mixing chamber A, a combustion chamber B, and a dilution chamber C, the mixing chamber extending around the combustion chamber. The mixing chamber A has an air inlet 16 comprising swirl vanes 16a and a fuel injector 15 disposed at the axis 14. Air under pressure is supplied from a comcompressor to the space between the outer casing 28 and the mixing chamber and passes to the inlet ...

Variable spread fluid dispersal systems

In a modification of the fluid dispersal system claimed in Specification No. 995,244 in which a first fluid and another fluid meet at a common boundary in a conduit immersed in a flow of a second fluid, means are provided to increase the ratio of the perimeter of the confined path within the conduit to the cross-sectional area of the confined path at points along the length of the conduit between the fluid entries so as to reduce the risk of the first and the other fluid mixing at the common boundary. The entry 4 of conduit 1 mounted in a duct wall 2, such as the casing of a ram jet or gas turbine jet propulsion engine reheat chamber, is supplied with liquid fuel through a metering device 15 comprising a movable obturator 18 and diaphragms 16, 19 with orifices 17, 20 respectively. A second entry 6 faces in the upstream direction and is subjected to the dynamic pressure of the air, combustion gases, or other fluid 3 flowing through duct 2. Alternatively, entry 6 may be connected to an engine ...

Improvements in or relating to gas-turbine engines

... 694,448. Generating combustion products under pressure. ROLLS-ROYCE, Ltd. Jan. 7, 1952 [Oct. 9, 1950], No. 24656/50. Class 51(i) In an air jacketed type of combustion chamber for a gas turbine engine, the air flow is divided into two parts, one of which enters the flame tube at its inlet end and the other of which flows around the outside of the flame tube, by means of a duct part which is supported from the air casing by one or more supports, the duct part affording at its outlet end a support for the inlet end of the flame tube. A fuel burner or nozzle is supported from the air casing by a strut which extends through an aperture or cut-away in the duct part, and has a shaped flanged portion in the cut-away so as to afford a continuation of the wall of the duct part, the fuel nozzle and support being removable from the air casing without removal of the duct part. The duct part 11 in Fig. 1 has struts 13 formed integrally therewith which are secured in recesses 14 in the air casing wall ...

GAS TURBINE ENGINE GASEOUS FUEL INJECTOR

A gaseous fuel injector for an industrial gas turbine plant is arranged to operate on fuel produced from a coal gasifier for normal running or natural gas for starting purposes. The injector is self purging to prevent the ingestion of natural gas or combustion products into the passage of the fuel injector for the lower calorific value fuel. The fuel injector has fuel ducts and gas flow passages in a duct assembly attached to the head of each flame tube of a gas generator. For starting, natural gas flows through the duct and the central passage, while air flows through the outer passage preventing ingestion of natural gas and combustion products into the outer fuel duct. When running on fuel from a coal gasifier both ducts run full of fuel, as do the passages. The air for the coal gasifier may be provided by the gas generator.

APPARATUS FOR A FUEL SYSTEM

Fuel injector

A fuel injector for a gas turbine conbustor, operates on gas and liquid fuels selectively. The injector has a central liquid fuel duct 23 and jets 25 an annular gas duct 1 and gas jets 3, and a outer annular combustion air duct 9. Operation on the liquid fuel tends to cause combustion products to be ingested into the gas orifices so impeding efficient gas combustion. The invention provides apertures 17 between the outer air duct and the intermediate gas duct whereby during operation on liquid fuel air bleeds into the gas duct and purges the gas orifices 3. As an added advantage in a transition from liquid fuel to gas, the gas pressure is increased to a point where the air bleed is reversed and gas bleeds into the air duct. The pre-mixed gas/air is then emitted from swirlers surrounding the ring of gas jets. More efficient gas combustion results. The arrangement prevents spillage, flashback and fouling of the gas orifices during secondary fuel operation. ...

Fuel staging system

The system comprises a main fuel manifold (58) being coupled downstream of a sequence valve (28), and a plurality of main fuel nozzles (66) each coupled to the main fuel manifold (58) through a respective main nozzle shut-off valve (96). A first set of pilot nozzles (82) is coupled to the main fuel manifold (58) through the sequence valve (28), and a second set of pilot nozzles (86) is coupled to the main fuel manifold (58) through the sequence valve (28). At low engine speeds, in the first and/or second pilot open positions, fuel flows to either pilot nozzle (82, 86) through the main fuel manifold (58), and the main fuel nozzles (66) are isolated from the main fuel manifold (58) by the main nozzle shut-off valves (96). Then, at higher engine speeds, the main nozzle shut-off valves (96) are opened, and the sequence valve (28) splits the fuel flow from a fuel inlet line (14) between the main fuel manifold (58) and the first and second pilot nozzles (82, 86). The fuel in the main fuel manifold ...

Fuel/air mixing arrangement for combustion apparatus

An apparatus/method for modifying a gaseous fuel

An apparatus/method for modifying a gaseous fuel comprises supplying a first portion of the fuel and an oxidant to a combustion device 1, 9, 10 to produce products of partial combustion, mixing the products of partial combustion with the remaining portion of the fuel which has not been combusted to provide a modified fuel. The fuel may be methane supplied via a pipe 2 and the oxidant may be air supplied via air inlet feeds 3. The combustion device may comprise of a burner 1 supported by the inlet feeds 3, ignitor 9 and flame tube 10. Burner 1 may comprise of a plate 6, a radial or axial swirler 5 and a pre-chamber 7 within which the air and fuel mix. Flame tube 10 may have effusion holes 11 and quench holes 13 for the passage of the remaining un-combusted fuel. Effusion holes 11 permit the fuel to be modified into a fuel rich mixture so that only partial combustion occurs because of an insufficient supply of air that gives rise to intermediate combustion products such as carbon monoxide ...

Liner Component for a Combustor

A sacrificial liner 54 is provided for an access aperture in a combustor casing (16 fig 3), the aperture permitting location of a fuel injector 28. The liner has an outer annular surface 62 and an inner annular surface 59, and both surfaces are arranged in an eccentric relationship. The combustor casing and the liner are dimensioned such that the liner is an interference fit within the combustor aperture, and the liner may also include a flange 55 to locate the liner at a predetermined axial location. The liner serves to protect the combustor outer casing from wear by a series of piston rings 50 mounted in a seal carrier 52. The eccentricity of the surfaces prevents excess contact pressure between the piston rings and the seal carrier, and therefore improves component life.

Fuel nozzle with central body cooling system

A fuel nozzle for turbine engine includes a cooling shroud located at the downstream end of the fuel nozzle to help cool the downstream end of the fuel nozzle. The cooling shroud surrounds the exterior circumference of the downstream end of the fuel nozzle. A flow of air is admitted into the cooling shroud and the flow of air travels in the downstream direction through a first passageway which covers the exterior of the fuel nozzle. The cooling air flow then turns 180° and travels in the upstream direction through a second passageway which is located concentrically outside the first passageway. The airflow then leaves the upstream end of the cooling shroud and enters the interior of the fuel nozzle.

Passive air-fuel mixing prechamber

A gas turbine combustion system passive air-fuel mixing prechamber includes one or more fuel passages. Each fuel passage includes at least one downstream fuel injection orifice. One or more fluid conduits connect an upstream portion of at least one fuel passage with one or more air passages such that pressure drops across each fuel injection orifice substantially self-equalize in a passive manner with corresponding air passage pressure drops over a broad range of fuel lower heating value (LHV) from about 150 Btu/scf to about 900 Btu/scf of fuel passing through the fuel passage while mixing with air passing through one or more connected fluid conduits. The effective area of each fluid conduit relative to the corresponding fuel and air passages is dependent upon the desired fuel LHV operating range.

Pegless secondary fuel nozzle

A unitary fuel injection manifold for a secondary fuel nozzle improves fuel-air mixing and offers flexibility to alter the mixing profile through adaptability to a variety of number, types, and orientation of discharge outlets to the combustion air mixing space around the secondary fuel nozzle. An aerodynamic surface with reduced extension into the mixing space reduces pressures drop and interference with design airflow. Manifold integrity is enhanced by elimination of fillet welds to mount external pegs.

Cooling flowpath dirt deflector in fuel nozzle

A fuel nozzle assembly includes a chamfered leading edge of an annular wall section disposed between an outer pilot swirler and an inlet to an injector cooling flowpath surrounding the second pilot swirler. A radially inwardly facing conical chamfered surface of the chamfered leading edge deflects dirt from cooling flowpath.

Fuel atomization dual orifice fuel nozzle

A pilot fuel injector tip includes concentric primary and secondary pilot fuel nozzles having a circular primary exit axially aft and downstream of an annular secondary exit respectively. A fuel nozzle assembly includes a pilot swirler flowpath section having an annular inwardly tapering conical flowpath section surrounding primary and secondary exits. An inwardly tapering conical wall section radially inwardly bounding flowpath section defines a conical surface. Exits are located at or axially forward or upstream of the conical surface. An annular secondary fuel supply passage in secondary pilot fuel nozzle includes a secondary fuel swirler with an array of helical spin slots that may have rectangular cross sections. A chamfered leading edge of an annular wall section disposed between an outer pilot swirler and an inlet to an injector cooling flowpath surrounding the second pilot swirler includes a radially inwardly facing conical chamfered surface for deflecting dirt from cooling flowpath.

Fuel Nozzle Passive Purge Cap Flow

A cooling circuit for a fuel nozzle in a gas turbine includes an end cap cavity receiving passive purge flow from a compressor of the turbine, and fuel nozzle swozzles disposed in a swozzle shroud that impart swirl to incoming fuel and air. Purge slots are formed in the swozzle shroud and through the fuel nozzle swozzles in fluid communication with the end cap cavity. The purge slots are positioned upstream of a quat fuel injection passage, and the passive purge flow enters fuel nozzle tip cavities of the fuel nozzle to provide tip cooling and tip purge volume without mixing the passive purge flow with quat fuel.

Fuel injector

A fuel injector is provided and includes a first tube, having first and second opposing ends, which is supplied with fuel, and one or more second tubes disposed within the first tube, each of the one or more second tubes being supplied with air and having sidewalls defining injection holes through which the fuel enters the one or more second tubes to mix with the air, and an outlet end of the sidewalls corresponding to the second end of the first tube.

Fuel injector assembly

A fuel injector assembly for a combustor is provided, including a fuel nozzle having an axial inflow swirler and one or more radial inflow swirlers spaced radially outward of the downstream end of the fuel nozzle and mounted to the combustor, wherein the airstreams produced by the swirlers airblast atomize fuel films produced by the fuel nozzle.

Fuel injection assembly for use in turbine engines and method of assembling same

A method for assembling a fuel injection assembly for use in a turbine engine is provided. The method includes providing a cap assembly that has at least one first opening extending at least partially through it and a plurality of second openings extending at least partially through it. Moreover, a plurality of tube assemblies are coupled within the cap assembly. Each tube assembly includes a plurality of tubes. Further, at least one injection system is coupled to the cap assembly to enable a fluid from a fluid source to be discharged through at least one of the plurality of second openings. The fluid flows between at least two adjacent tube assemblies to facilitate at least one of reducing a temperature within the cap assembly and reducing dynamic pressure oscillations within a combustor during operation of the turbine engine.

Gas Turbine Combustor

A combustor of the prior art that defines the outlet position and direction of an air hole and suppresses adhesion of flame to an air hole outlet can reduce a discharge amount of NOx by increasing a distance over which fuel and air are mixed with each other. However, such a combustor is not sufficiently discussed for measures to suppress the occurrence of combustion oscillation resulting from the variation of a flame surface. A combustor 2 according to the present invention includes a combustion chamber 5 to which fuel and air are supplied; air holes 32 adapted to supply air to the combustion chamber 5; fuel nozzles 25 adapted to supply gaseous fuel to the air holes 32; and orifices 24 adapted to allow the gaseous fuel supplied to the air holes 32 to cause a pressure drop.

Apparatus for mixing fuel in a gas turbine

A combustor nozzle includes an inlet surface and an outlet surface downstream from the inlet surface, wherein the outlet surface has an indented central portion. A plurality of fuel channels are arranged radially outward of the indented central portion, wherein the plurality of fuel channels extend through the outlet surface.

Apparatus for injecting fluid into a combustion chamber of a combustor

A combustor is disclosed having a combustion liner defining a combustion chamber. The combustor may also include a liner cap disposed upstream of the combustion chamber. The liner cap may include a first plate and a second plate. Additionally, the combustor may include a fluid conduit extending between the first and second plates. The fluid conduit may be configured to receive fluid flowing adjacent to the first plate and inject the fluid into the combustion chamber.

Gas turbine combustor having a fuel nozzle for flame anchoring

A combustor includes an end cover having a nozzle. The nozzle has a front end face and a central axis. The nozzle includes a plurality of fuel passages and a plurality of oxidizer passages. The fuel passages are configured for fuel exiting the fuel passage. The fuel passages are positioned to direct fuel in a first direction, where the first direction is angled inwardly towards the center axis. The oxidizer passages are configured for having oxidizer exit the oxidizer passages. The oxidizer passages are positioned to direct oxidizer in a second direction, where the second direction is angled outwardly away from the center axis. The plurality of fuel passages and the plurality of oxidizer passages are positioned in relation to one another such that fuel is in a cross-flow arrangement with oxidizer to create a burning zone in the combustor.

Flat fan air assist injectors

An injector for injecting a flat fan of liquid includes an injector body defining a pair of air channels, with each air channel fluidly connected to a respective air inlet. The air channels join one another at a common throat defined in the injector body and are separated by a land defined in the injector body extending from the air inlets to a point proximate the throat. The air channels and a liquid inlet are in proximity to draw liquid out of the liquid inlet into the throat with air flowing through the air channels. A diverging diffuser is provided in fluid communication with the throat. The diffuser includes an impingement surface defined in the injector body opposed to the liquid inlet. The liquid inlet is configured to inject liquid against the impingement surface to form a fan of liquid diverging outward through the diffuser.

Gas turbine combustion burner

In a gas-turbine combustion burner which is provided, at a distal end thereof, with a fuel spraying hole that sprays fuel into a combustion region formed inside a combustion cylinder of a gas-turbine combustor and in which a fuel flow path that guides the fuel, which is supplied from a fuel source, to the fuel spraying hole is formed in the interior thereof, the impedance of a fuel supply system that guides the fuel from the fuel source to the fuel spraying hole is set so that propagation of pressure fluctuations from the combustion region to the fuel supply system becomes an allowable level or less.

Sector nozzle mounting systems

Systems are provided for mounting sector nozzles within gas turbine combustors. In one embodiment, a sector nozzle includes a nozzle portion configured to mix fuel and air to produce a fuel-air mixture and a shell coupled to the nozzle portion. The sector nozzle also includes a first longitudinal strut and a second longitudinal strut coupled to a first surface of the shell on opposite sides of a window within the first surface. A third longitudinal strut is coupled to a second surface of the shell, and the second surface is disposed opposite of the first surface.

Multi-fuel injection nozzle

A multi-fuel nozzle ( 90 ) for a gas turbine engine. The nozzle includes: an annular main body ( 68 ) having a plurality of fuel gas channels ( 22 ), all disposed circumferentially about a main body longitudinal axis ( 14 ); an annular fuel oil body ( 30 ) disposed within the annular main body ( 68 ) and having a central oil channel ( 36 ) coaxial with the main body longitudinal axis ( 14 ); an annular cooling air channel ( 42 ) between the annular main body ( 68 ) and the fuel oil body ( 30 ); a discrete cooling air body ( 70, 100 ) having a guide ( 74, 104 ), the guide ( 74, 104 ) supported independent of a downstream end ( 84 ) of the main body ( 68 ) and configured to direct cooling air traveling downstream in the annular cooling air channel ( 42 ) radially inward at a location immediately downstream of a central oil channel downstream end ( 34 ).

Systems and methods involving improved fuel atomization in air blast fuel nozzles of gas turbine engines

Systems and methods involving improved fuel atomization in air-blast fuel nozzles of gas turbine engines are provided. In this regard, a representative method includes: providing fuel to a chamber defined by an inner surface; and continuously atomizing a portion of the fuel via interaction of the fuel with the inner surface.

COMBUSTOR WITH A PRE-NOZZLE MIXING CAP ASSEMBLY

The present application and the resulting patent provide a pre-nozzle mixing cap assembly positioned about a cap member of a combustor for mixing a flow of air and a flow of fuel. The pre-nozzle mixing cap assembly may include a fuel plenum in communication with the flow of fuel and a number of tubes in communication with the flow of air and extending through the fuel plenum. Each of the tubes may include a number of fuel holes therein such that the flow of fuel in the fuel plenum passes through the fuel holes and mixes with the flow of air. 1. A pre-nozzle mixing cap assembly positioned about a cap member of a combustor for mixing a flow of air and a flow of fuel , comprising:a fuel plenum in communication with the flow of fuel; anda plurality of tubes in communication with the flow of air and extending through the fuel plenum;each of the plurality of tubes comprising a plurality of fuel holes therein such that the flow of fuel in the fuel plenum passes through the plurality of fuel holes and mixes with the flow of air.2. The pre-nozzle mixing cap assembly of claim 1 , further comprising a plurality of plates with the plurality of tubes extending therethrough.3. The pre-nozzle mixing cap assembly of claim 2 , wherein the plurality of plates comprises a first plate and a second plate and wherein the first plate and the second plate define the fuel plenum therebetween.4. The pre-nozzle mixing cap assembly of claim 3 , wherein the plurality of plates comprises a third plate and wherein the second plate and the third plate define an air plenum therebetween.5. The pre-nozzle mixing cap assembly of claim 1 , wherein each of the plurality of tubes comprises a chamfered edge.6. The pre-nozzle mixing cap assembly of claim 1 , further comprising a fuel inlet in communication with the flow of fuel and the fuel plenum.7. The pre-nozzle mixing cap assembly of claim 1 , further comprising a cooling circuit therein.8. The pre-nozzle mixing cap assembly of claim 7 , wherein the ...

Turbomachine combustor assembly including a combustion dynamics mitigation system

A turbomachine combustor includes a combustor cap having a cap surface and a wall that extends about the cap surface to define a cap volume, and a plurality of nozzle members that extend from the cap surface. The plurality of nozzle members include a center nozzle member and one ore more outer nozzle members. A combustor dynamics mitigation system is arranged in the combustor cap and includes plurality of divider members that extend from the wall toward the center nozzle member. The plurality of divider members define a plurality of parallel resonator volumes. The combustor dynamics mitigation system also includes a plurality of tubes that extend into corresponding ones of the plurality of parallel resonator volumes.

Burner assembly

A burner assembly for a gas turbine is provided. The burner assembly has a combustor, a centrally arranged pilot burner and plurality of main burners surrounding the pilot burner. Each main burner has a cylindrical housing having a lance which is centrally arranged therein and has a fuel channel for liquid fuel. The lance is supported on the housing by swirl blades and an attachment is arranged on the lance in the direction of the combustor. The liquid fuel nozzle is arranged in the attachment downstream of the swirl blades and connected to the fuel channel. For the improved mixing of the fuel with the air, the liquid fuel nozzle is designed as a full jet nozzle and the full jet nozzle has a length and a diameter, the ratio of the length to the diameter is at least 1.5.

Combustor and method for supplying fuel to a combustor

A combustor includes an end cap having an upstream surface axially separated from a downstream surface. A cap shield circumferentially surrounds the upstream and downstream surfaces, tubes extend from the upstream surface through the downstream, and a plenum is inside the end cap. A first baffle extends radially across the plenum toward the cap shield, and a plate extends radially inside the plenum between the first baffle and the upstream surface. A method for supplying fuel to a combustor includes flowing a working fluid through tubes, flowing a fuel into a plenum between upstream and downstream surfaces, radially distributing the fuel along a first baffle, and axially flowing the fuel across a plate that extends radially inside the plenum.

GAS TURBINE COMBUSTION CHAMBER

A gas turbine combustion chamber is provided including a pilot fuel nozzle arranged in the central section of a cylinder that opens at one end towards a combustion chamber. The pilot fuel nozzle includes a fuel nozzle and a cylindrical outer casing around the outer circumference of the fuel nozzle. A pilot swirl element is arranged between fuel nozzle and outer casing, including a plurality of main burners which are arranged around the pilot fuel nozzle, and including a pilot cone having an inner side and an outer side. The pilot cone is arranged on the pilot fuel nozzle and an opening, such that a pilot flame is formed in the pilot cone by mixing air and pilot fuel in order to ignite a fuel injected by the main burners, wherein the pilot cone has turbulence generators on the inner side and/or outer side thereof. 15-. (canceled)6. A gas turbine combustion chamber comprising: a fuel nozzle and a cylindrical outer casing around the outer circumference of the fuel nozzle and at a radial distance therefrom,', 'a pilot swirl element arranged between fuel nozzle and outer casing; and, 'a pilot fuel nozzle arranged in a central section of a cylinder which opens at one end towards a combustion chamber, the pilot fuel nozzle comprisesa pilot cone having an inner side and an outer side,wherein a plurality of main burners are arranged around the pilot fuel nozzle with respect to the radial direction,wherein the pilot cone is arranged on the pilot fuel nozzle at a combustion chamber end and having an opening at the combustion chamber end, such that a pilot flame is formed in the pilot cone by mixing air and pilot fuel in order to ignite a fuel injected by the plurality of main burners,wherein the pilot cone includes a plurality of turbulence generators on its inner side and/or its outer side,wherein the plurality of turbulence generators are trapezoidal and/or triangular strips which are arranged at an opening of the pilot cone over the entire circumference of the opening of ...

AIR FUEL PREMIXER HAVING ARRAYED MIXING VANES FOR GAS TURBINE COMBUSTOR

A fuel-air premixer for use in a combustor of a gas turbine includes an air inlet, a fuel inlet, a shroud, a central body and a cascade of vanes. The premixer mixes fuel and air in the annular mixing passage into a uniform mixture for injecting into a combustor reaction zone. The air from a compressor is injected into the mixer through an air inlet. The fuel is introduced into air stream via fuel injection holes that pass through the walls of the vanes which contain internal fuel flow passages. The flow field inside the premixer is broken up by the arrayed vanes into a series of small regions each containing a well designed small size mixing eddy which is steadily attached to the surface of the vanes. 114122113. A premixer for use in a combustor of a gas turbine , the premixer comprises: a circular shroud () and a cylinder central body () which contains internal fuel flow passages () , wherein the shroud and the central body confine an annular mixing passage () between them , wherein{'b': '11', 'an air inlet is located at an upstream end of the mixing passage and a fuel inlet () is located at an upstream end of a fuel flow passage inside the central body, wherein fuel and air are premixed in the annular mixing passage into a mixture for injecting into a combustor reaction zone, and'}the mixing passage comprises a cascade of arrayed vanes including, from upstream to downstream:{'b': 25', '15, 'a plurality of fuel nozzle vanes (), each comprising multiple fuel injection holes () on the wall thereof and an internal fuel flow passage communicated with the fuel passage inside the central body receiving the fuel from the fuel inlet,'}{'b': '24', 'a plurality of mixing vanes () breaking up the flow field in the mixing passage and forming eddies, and'}{'b': '23', 'a plurality of turning vanes () imparting swirl to the incoming mixture.'}23040. The premixer according to claim 1 , wherein each of the fuel nozzle vanes comprises a bluff forehead () and a suddenly constringent ...

SYSTEM AND METHOD FOR REDUCING COMBUSTION DYNAMICS IN A TURBOMACHINE

A turbomachine includes a combustion chamber, and at least one pre-mixer mounted to the combustion chamber. The at least one pre-mixer includes a main body having a first end portion that extends to a second end portion. The first end portion is configured to receive an amount of fuel and an amount of air and the second end portion defines an exit plane from which a fuel-air mixture discharges into the combustion chamber. The turbomachine also includes a combustion dynamics reduction system operatively coupled to the at least one pre-mixer. The combustion dynamics reduction system includes at least one of a boundary layer perturbation mechanism and an acoustic wave introduction system which disrupt a flow pattern of the fuel-air mixture within the at least one pre-mixer. 1. A turbomachine comprising:a combustion chamber;at least one pre-mixer mounted to the combustion chamber, the at least one pre-mixer including a main body including a first end portion that extends to a second end portion, the first end portion being configured to receive an amount of fuel and an amount of air and the second end portion defining an exit plane from which a fuel-air mixture discharges into the combustion chamber; anda combustion dynamics reduction system operatively coupled to the at least one pre-mixer, the combustion dynamics reduction system including at least one of a boundary layer perturbation mechanism and an acoustic wave introduction system that disrupt a flow pattern of the fuel-air mixture within the at least one pre-mixer.2. The turbomachine according to claim 1 , wherein the combustion dynamics reduction system includes a boundary layer perturbation mechanism claim 1 , the boundary layer perturbation mechanism including one of an air/inert injection system operatively coupled to the pre-mixer and mechanical member mounted in the at least one pre-mixer.3. The turbomachine according to claim 2 , wherein the boundary layer perturbation mechanism includes an air/inert ...

Gas turbine combustor and gas turbine

A gas turbine combustor includes: an external cylinder ( 31 ); an inner cylinder ( 32 ) provided inside the external cylinder ( 31 ) to form an air passage ( 30 ) between the external cylinder ( 31 ) and the inner cylinder ( 32 ); a pilot nozzle ( 35 ) provided in a center part of the inner cylinder ( 32 ) along a direction of a combustor axis (S); a plurality of main nozzles ( 36 ) provided on an inner peripheral surface of the inner cylinder ( 32 ) along a circumferential direction thereof so as to surround the pilot nozzle ( 35 ), the plurality of main nozzles ( 36 ) premixing fuel with combustion air introduced to the air passage ( 30 ) and ejecting the fuel into the inner cylinder ( 32 ); and a top hat nozzle ( 41 ) provided inside the air passage ( 30 ) across a circumferential direction to mix fuel with the combustion air prior to reaching the plurality of main nozzles ( 36 ).

COMBUSTION CHAMBER FOR AN AIRCRAFT ENGINE, AND METHOD FOR ATTACHING AN INJECTION SYSTEM IN A COMBUSTION CHAMBER OF AN AIRCRAFT ENGINE

A combustion chamber in which the injection system is attached so as to prevent any axial motion of the injection system, the combustion chamber includes: a baffle including a tubular portion having a first upstream end surrounded by a first radially projecting collar; and an injection system including a bowl which includes a cylindrical portion inserted into the tubular portion, the cylindrical portion including a second upstream end surrounded by a second radially projecting collar, the second collar axially bearing against the first collar, wherein the combustion chamber includes a clamp which axially clamps the first and the second collars against one another, and a fastener which retains the clamp clamped on either side of the first and second collars. 1. A combustion chamber of an aircraft engine comprising:a baffle comprising at least one tubular portion, the tubular portion extending in the direction of a reference axis, the tubular portion having a first upstream end surrounded by a first radially projecting collar;an injection system comprising at least one bowl which comprises a cylindrical portion inserted into the tubular portion, the cylindrical portion comprising a second upstream end that exits the first upstream end of the tubular portion, the second upstream end of the cylindrical portion being surrounded by a second radially projecting collar, the second collar being braced axially against the first collar;a clamp configured to clamp the first and the second collars against one another axially, anda fastener configured to retain the clamp clamped on either side of the first and second collars.2. The combustion chamber according to claim 1 , wherein the clamp comprises: a retaining ring bearing axially against the first collar, and', 'a peripheral retaining skirt projecting axially from the retaining ring, the peripheral retaining skirt surrounding the first and second collars radially,, 'a retaining bushing split into at least two bushing sections ...

Combustor assembly in a gas turbine engine

A combustor assembly in a gas turbine engine includes a combustor device, a fuel injection system, a transition duct, and an intermediate duct. The combustor device includes a flow sleeve for receiving pressurized air and a liner surrounded by the flow sleeve. The fuel injection system provides fuel to be mixed with the pressurized air and ignited in the liner to create combustion products. The intermediate duct is disposed between the liner and the transition duct so as to define a path for the combustion products to flow from the liner to the transition duct. The intermediate duct is associated with the liner such that movement may occur therebetween, and the intermediate duct is associated with the transition duct such that movement may occur therebetween. The flow sleeve includes structure that defines an axial stop for limiting axial movement of the intermediate duct.

Combustor fuel nozzle and method for supplying fuel to a combustor

A combustor fuel nozzle includes a center body and an inner shroud that circumferentially surrounds at least a portion of the center body. The inner shroud has a downstream surface. The fuel nozzle includes an inner passage between the center body and the inner shroud, an outer passage that circumferentially surrounds at least a portion of the inner shroud and a first plurality of fuel ports extending substantially radially outward through the center body. The first plurality of fuel ports is upstream from the downstream surface of the inner shroud. A method for supplying fuel to a combustor fuel nozzle includes flowing a working fluid through an inner passage between a center body and an inner shroud, injecting a fuel from the center body against the inner shroud, and flowing a portion of the working fluid through an outer passage that surrounds at least a portion of the inner shroud.

Fuel nozzel

The present application provides a fuel nozzle for mixing a flow of fuel and a flow of air. The fuel nozzle may include a downstream face, a number of fuel passages positioned about the downstream face for the flow of fuel, and a nozzle collar position about the downstream face. The nozzle collar may include a number of air vanes for the flow of air and one or more purge holes therethrough.

Annular combustion chamber of a gas turbine

The present invention relates to an annular combustion chamber of a gas turbine with—relative to the engine axis—a radially outer combustion chamber wall and a radially inner combustion chamber wall, with the combustion chamber walls forming an annular combustion space, with a combustion chamber head having a plurality of fuel nozzles and air inlet openings, with the respective central axes of the fuel nozzles forming an envelope rotationally symmetrical to the engine axis, the envelope dividing the combustion chamber into an annular and radially outer area and an annular and radially inner area, with the radially outer area and the radially inner area having the same volumes. 1. Annular combustion chamber of a gas turbine with—relative to the engine axis—a radially outer combustion chamber wall and a radially inner combustion chamber wall , with the combustion chamber walls forming an annular combustion space , with a combustion chamber head having a plurality of fuel nozzles and air inlet openings , with the respective central axes of the fuel nozzles forming an envelope rotationally symmetrical to the engine axis , the envelope dividing the combustion chamber into an annular and radially outer area and an annular and radially inner area , with the radially outer area and the radially inner area having the same volumes.2. Annular combustion chamber of a gas turbine in accordance with claim 1 , characterized in that the respective central axes of the fuel nozzles are inclined relative to the engine axis by an angle.3. Annular combustion chamber of a gas turbine in accordance with claim 2 , characterized in that the radially outer area and the radially inner area in the axial direction along the envelope have identical areas in a respectively conical plane which is vertical to the conical envelope. This invention relates to an annular combustion chamber of a gas turbine. An annular combustion chamber of this type has an upper/radially outer combustion chamber wall ...

Gas Turbine Combustor and Operating Method Thereof

A gas turbine combustor has a chamber supplied with fuel and air and a multi-burner having a plurality of burners provided with an air hole plate having a plurality of air holes, and fuel nozzles for supplying fuel to the air holes in the air hole plate; the multi-burner is made up of a center burner disposed in the center and a plurality of outer burners around the center burner, the outer burners are divided into inner fuel nozzles and outer fuel nozzles to separately supply fuel through fuel systems, and the fuel is supplied to the fuel nozzles in the center burner or to the fuel nozzles in the center burner and the inner fuel nozzles in the outer burners disposed around the center burner in a partial load condition in which the load is lower than that of when all the fuel systems are used to supply fuel. 1. A gas turbine combustor comprising: a chamber for mixing and burning supplied fuel and supplied air to generate combustion gas;an air hole plate located in an upstream side of the chamber, forming a plurality of air holes for supplying the air;a plurality of fuel nozzles for supplying the fuel to the plurality of air holes formed in the air hole plate,wherein the air holes are disposed in the downstream side of the fuel nozzles that one of the fuel nozzles is paired with one of the air holes; anda plurality of burners made up of the plurality of air holes and the plurality of fuel nozzles in pairs,characterized in that,the plurality of burners are comprised a center burner disposed on an axis of the gas turbine combustor and a plurality of outer burners installed around the center burner,the center burner is fixed to a first fuel supply system for supplying the fuel to the fuel nozzles in the center burner,the plurality of outer burners is fixed to a second fuel supply system for supplying the fuel to the fuel nozzles in outer burner's inner portions which are inner regions of specific outer burners among the outer burners,the plurality of outer burners is ...

Jet micro-induced flow reversals combustor

A jet micro-induced flow reversals combustor is used to reduce NO x emissions. The combustor has a nozzle disposed at the head end of the combustion chamber. The nozzle includes a plurality of jets for injecting a fuel and oxidant mixture stream into the combustion chamber. A combustion liner is disposed within the casing on one side of the nozzle and a plenum chamber is disposed on another side of the nozzle and configured to provide an input of a fuel and oxidant. The nozzle and the combustion liner are sized and shaped to input the fuel and oxidant mixture stream into the combustion liner at a high velocity ratio wherein a jet velocity is greater than a combustion mean velocity within the combustion liner, to increase turbulence within the combustion liner and reduce combustion emissions.

Combustor Assembly with Trapped Vortex Cavity

Embodiments of the present application include a combustor assembly. The combustor assembly may include an annular trapped vortex cavity located adjacent to a downstream end of a bundle of air/fuel premixing injection tubes. The annular trapped vortex cavity may include an opening at a radially inner portion of the annular trapped vortex cavity adjacent to the head end of the bundle of premixing tubes. The annular trapped vortex cavity may also include one or more air injection holes and one or more fuel sources disposed about the annular trapped vortex cavity such that the one or more air injection holes and the one or more fuel sources are configured to drive a vortex within the annular trapped vortex cavity.

UNKNOWN

The present invention relates to a premix burner of a combustion chamber of a gas turbine with at least one annular duct for supplying air and fuel, including a radially outer and a radially inner combustion chamber wall relative to a burner central axis and with at least one swirler arranged in the duct, said swirler including several flow-guiding elements distributed around the circumference of the duct cross-section, characterized in that at least one radially inner duct wall is provided in the area of the flow-guiding elements with a concave recess of the annular groove type. 1. Premix burner of a combustion chamber of a gas turbine with at least one annular duct for supplying air and fuel , including a radially outer and a radially inner combustion chamber wall relative to a burner central axis and with at least one swirler arranged in the duct , said swirler including several flow-guiding elements distributed around the circumference of the duct cross-section , characterized in that at least one radially inner duct wall is provided in the area of the flow-guiding elements with a concave recess of the annular groove type.2. Premix burner in accordance with claim 1 , characterized in that a radially outer duct wall is provided in the area of the flow-guiding elements with a concave recess of the annular groove type.3. Premix burner in accordance with claim 1 , characterized in that the cross-sectional area of the radially inner and/or the radially outer recess is equal to or less than a total of the thicknesses of the flow-guiding elements established in the respective cross-sectional plane perpendicular to the burner axis.4. Premix burner in accordance with claim 1 , characterized in that each flow-guiding element has a flow profile that is provided between an inflow-side leading edge and an outflow-side trailing edge claim 1 , with the leading edge and the trailing edge each being arranged on a cross-sectional plane perpendicular to the burner axis and that ...

Gas Turbine Combustor and Method for Operating Same

Disclosed is a gas turbine combustor for stably burning low-BTU gases, such as blast furnace gases, that are heavily laden with CO. The gas turbine combustor includes a double-swirling burner with an inner swirler and an outer swirler, the burner having a configuration with gas fuel injection holes and air injection holes arranged at alternate positions in the inner swirler and with gas injection holes arranged in the outer swirler. In addition, fuel injection holes for enhancing flame stability are provided at positions radially inward of the inner swirler. An inner flame by the inner swirler and an outer flame by the outer swirler interact with each other to stably burn the low-BTU gas. In the inner swirler, the gas injection holes and air injection holes arranged at alternate positions contributes to raising a temperature of the inner flame to a level required for flame stabilization. 1. A gas turbine combustor comprising:a combustion chamber for burning a fuel and air in a mixed condition; anda burner provided upstream in a gas flow direction of the combustor, for supplying the fuel and the air to inside of the combustion chamber and thus stabilizing a flame; a first swirler in which both of a plurality of gas injection holes for injecting the fuel, and a plurality of air injection holes for injecting air are arranged at alternate positions in a circumferential direction of the swirler, and', 'a second swirler provided at an outer periphery of the first swirler, the second swirler being fitted only with a plurality of gas injection holes to inject the fuel., 'wherein the burner includes'}2. The gas turbine combustor according to claim 1 , wherein:a plurality of fuel injection holes for enhancing flame stability are arranged more radially inward of the burner than the gas injection holes and air injection holes arranged in the first swirler.3. The gas turbine combustor according to claim 1 , further comprising:a first fuel flow control valve provided in a first ...

Exhaust temperature based threshold for control method and turbine

A gas turbine, computer software and a method for controlling an operating point of the gas turbine that includes a compressor, a combustor and at least a turbine is provided. The method comprises: determining an exhaust pressure at an exhaust of the turbine; measuring a compressor pressure discharge at the compressor; determining a turbine pressure ratio based on the exhaust pressure and the compressor pressure discharge; calculating a primary to lean-lean mode transfer threshold reference curve as a function of the turbine pressure ratio, where the primary to lean-lean mode transfer threshold curve includes points at which an operation of the gas turbine is changed between a primary mode to a lean-lean mode; and controlling the gas turbine to change between the primary mode and the lean-lean mode.

System and method for reducing combustion dynamics in a combustor

A system for reducing combustion dynamics in a combustor includes an end cap having an upstream surface axially separated from a downstream surface, and tube bundles extend from the upstream surface through the downstream surface. A divider inside a tube bundle defines a diluent passage that extends axially through the downstream surface, and a diluent supply in fluid communication with the divider provides diluent flow to the diluent passage. A method for reducing combustion dynamics in a combustor includes flowing a fuel through tube bundles, flowing a diluent through a diluent passage inside a tube bundle, wherein the diluent passage extends axially through at least a portion of the end cap into a combustion chamber, and forming a diluent barrier in the combustion chamber between the tube bundle and at least one other adjacent tube bundle.

GAS TURBINE FUEL INJECTOR WITH INSULATING AIR SHROUD

A fuel injector for a gas turbine engine is disclosed. The fuel injector includes an injector housing extending from a first end to a second end along a longitudinal axis. The second end of the housing is fluidly coupled to a combustor of the turbine engine and the housing includes a liquid fuel gallery annularly disposed about the longitudinal axis. The fuel injector also includes a stem extending longitudinally from the first end of the housing to a third end. The stem includes a liquid tube configured to deliver liquid fuel to the fuel injector. The fuel injector also includes an annular shell extending along the longitudinal axis from the first end to the third end and circumferentially disposed about the stem. The fuel injector further includes an insulating air shroud formed inside the shell. The air shroud includes a layer of air between the shell and the stem. 19-. (canceled)10. A method of operating a gas turbine engine , comprising:delivering liquid fuel to a combustor of the turbine engine through one or more liquid fuel carrying components of a fuel injector coupled to the combustor;combusting the liquid fuel in the combustor;providing an insulating air shroud around one or more of the liquid fuel carrying components, the insulating air shroud including a layer of atmospheric air;generating eddy air currents in the insulating air shroud in response to the combustion, the eddy air currents assisting in expelling heated atmospheric air from the insulating air shroud and drawing cooler atmospheric air into the insulating air shroud; andmaintaining a temperature of the one or more liquid fuel carrying components below a threshold temperature as a result of the generation of the eddy air currents.11. The method of claim 10 , wherein providing an insulating air shroud includes providing an insulating air shroud between the one or more liquid fuel carrying components and a shell.12. The method of claim 11 , wherein the delivering of liquid fuel includes ...

BURNER FOR A GAS COMBUSTOR AND A METHOD OF OPERATING THE BURNER THEREOF

A burner for a gas combustor and a method of operating the burner are disclosed. The burner includes a front surface area divided into a plurality of subareas and inlets arranged on the front surface area such that each subarea is encircled by at least four inlets and such that during operation of the burner, a gas recirculation in the combustor is facilitated corresponding to each subarea.

Combustor cap mounting structure for a turbine engine

A mounting structure for mounting a combustor cap in a combustor of a turbine engine includes support struts that are connected between the combustor cap and a concentric combustor cap barrel flange. The support struts may have an airfoil shape to minimize wakes created in a flow of compressed air that is passing over the support struts. Also, the support struts may have an interior passageway that allows a portion of the compressed air to flow though the support strut. The flow of air passing through the support struts may also pass through corresponding vent apertures in the combustor cap barrel flange so that the flow of air passing through the support struts is delivered into a space between the exterior of the combustor cap barrel flange and a forward casing of the combustor.

DEVICE FOR INJECTING A MIXTURE OF AIR AND FUEL INTO A TURBINE ENGINE COMBUSTION CHAMBER

A device for injecting a mixture of air and fuel into a turbine engine combustion chamber, the device including a mechanism for centering a fuel injector, which mechanism is movable radially in a support mechanism fastened to a wall of the chamber, the support mechanism carrying a retaining mechanism for axially retaining the centering mechanism on a side opposite from the chamber wall, the retaining mechanism being fastened in releasable manner to the support mechanism. 1. A device for injecting a mixture of air and fuel into a turbine engine combustion chamber , the device comprising centering means for centering a fuel injector , said centering means being movable radially in support means for fastening to a wall of the chamber , the support means carrying retaining means for axially retaining the centering means on a side opposite from the chamber wall , the retaining means are fastened in releasable manner to the support means , wherein the retaining means comprise at least one peg carried by the support means , a latch engaged on the peg and holding the centering means against the support means , and a pin having one end portion engaged and prevented from moving in a housing of the peg and having an opposite end portion bearing against the latch to prevent the pin from moving relative to the peg.2. A device according to claim 1 , wherein the support means comprise at least one radially outer wall or tab for guiding or bearing against the centering means claim 1 , and wherein the retaining means are releasably fastened to said wall or tab.3. A device according to claim 2 , wherein the support means comprise two diametrically opposite tabs extending radially outwards.4. A device according to claim 3 , wherein the centering means comprise two radially outer tabs that are diametrically opposite and that are held against the tabs of the support means by the retaining means that pass through orifices or notches in the tabs of the centering means with circumferential ...

Fuel nozzle

A fuel nozzle includes a center body and a shroud circumferentially surrounding at least a portion of the center body to define an annular passage between the center body and the shroud. A plurality of vanes extend radially between the center body and the shroud in the annular passage. A first ceramic extension extends downstream from the shroud to define at least a portion of the annular passage downstream from the shroud.

INJECTOR FOR THE COMBUSTION CHAMBER OF A GAS TURBINE HAVING A DUAL FUEL CIRCUIT, AND COMBUSTION CHAMBER PROVIDED WITH AT LEAST ONE SUCH INJECTOR

A starting injector usable in all flight modes without additional cost or weight. The starting injector includes a dual fuel circuit and an air circuit. An injector for a combustion chamber of a gas turbine includes a dual fuel injection circuit including a starting fuel circuit for ignition and then for all the flight modes, and a main fuel circuit for all the flight modes after starting. The circuits include parallel pipes in a common tube having an axis. The pipe of the starting circuit is substantially in communication with a center of a spherical injector body. At the end, the pipe accommodates an injection manifold coupled to a central channel passing through a central wall of a swirler. The pipe of the main circuit is in communication with an annular channel opposite jet channels. An air circuit is guided between two portions shaped as concentric spheres. 111-. (canceled)12. An injector for a gas turbine combustion chamber , comprising:a dual fuel supply circuit and an air circuit;fuel injection circuits including a starting fuel supply circuit configured to trigger chamber ignition, and then to operate in any flight mode, and a main fuel supply circuit configured to operate in any flight mode further to starting,wherein the fuel supply circuits include parallel conducts made in a common tube with a longitudinal axis,wherein the conduct of the starting circuit opens, on one end, substantially into a center of a spherical injector body extending the common tube, and, on the one end, the conduct houses an injection ring configured to drive fuel into rotation before projecting the fuel inside the chamber through a central channel passing through a central wall of a swirling device,wherein the conduct of the main circuit opens into an annular channel made in the body facing jet channels radially configured in a main wall around the central channel, andwherein the air circuit is guided between two concentric sphere portions including an injector body and a sheath ...

Fuel Plenum Premixing Tube with Surface Treatment

The present application provides a micro-mixer fuel plenum for mixing a flow of fuel and a flow of air in a combustor. The micro-mixing fuel plenum may include an outer barrel and a number of mixing tubes positioned within the outer barrel. The mixing tubes may include one or more heat transfer features thereon.

Turbomachine combustor nozzle and method of forming the same

A turbomachine combustor nozzle includes a first plate member having a first plurality of openings, and a second plate member having a second plurality of openings that are configured and disposed to be in alignment with the first plurality of openings. A plurality of nozzle members extends through corresponding ones of the first plurality of openings and the second plurality of openings. Each of the plurality of nozzle members includes a solid core.

Combustor assembly having a fuel pre-mixer

A combustor assembly having a fuel pre-mixer including a duct for mixing an airflow and a fuel therein. Also included is a center body coaxially aligned within the duct for receiving the fuel from a fuel source and configured to distribute the fuel to at least one axial location within the duct. Further included is a planar vane section in communication with the airflow and the fuel to provide a first injection of fuel and a flow conditioning effect on the airflow. Yet further included is a swirler vane section disposed downstream of the planar vane section, wherein the swirler vane section is configured to provide a second injection of fuel and a mixing of the fuel and the airflow.

FUEL INJECTOR BEARING PLATE ASSEMBLY AND SWIRLER ASSEMBLY

A bearing plate assembly for a turbine engine fuel injector includes a bearing plate with an opening bordered by a race A swivel ball nests inside the race and is rotatable relative thereto. A lock, which may be a tip bushing resists disengagement of the swivel ball from the race. A fuel injector nozzle extends through an opening in the swivel ball. During engine operation, the ball can swivel inside the race to accommodate rotational movement of the nozzle about lateral and radial axes. 1. A bearing plate assembly , comprising:a bearing plate having an opening penetrating therethrough, the opening being bordered by a race having an inner surface with a curved profile;a swivel ball nested inside the race, the swivel ball having a curved profile of the same shape as the inner surface of the race; anda lock for resisting disengagement of the swivel ball from the race, wherein the lock defines an inner peripheral surface that is configured to engage an outer peripheral surface of a cylindrical portion of a fuel injector nozzle for a gas turbine engine.2. The assembly of wherein each of the curved profiles is spherical.3. The assembly of wherein the race includes loading slots for receiving the swivel ball during assembly.4. The assembly of wherein the lock comprises a bushing circumscribed by the swivel ball.5. A bearing plate assembly claim 1 , comprising:a bearing plate having an opening penetrating therethrough, the opening being bordered by a race having an inner surface with a curved profile;a swivel ball nested inside the race, the swivel ball having a curved profile of the same shape as the inner surface of the race; anda lock for resisting disengagement of the swivel ball from the race wherein the lock comprises a bushing circumscribed by the swivel ball, and wherein the race includes loading slots, and one end of the bushing has at least one ear residing in one of the loading slots, another end of the bushing being deformed to grasp the swivel ball.6. The ...

Non-symmetric arrangement of fuel nozzles in a combustor

A fuel nozzle arrangement includes an annular combustor having four quadrants. Fuel injectors are located in the four quadrants and each fuel injector has either a first or second type nozzle attached to it. The first and second type fuel nozzles are located in an array so that at least two of one nozzle type are located together to allow for non-uniform injector spacing with uniform downstream temperatures.

Gas Turbine Combustor and Operating Method for Gas Turbine Combustor

Disclosed is a gas turbine combustor equipped with a burner constructed to fire a plurality of combustors at the same time at a fuel/air ratio suitable for gas turbine ignition. 1. A gas turbine combustor comprising:a combustion chamber that burns a fuel and air to generate combustion gases;a fuel header with a plurality of fueling nozzles disposed thereupon to inject the fuel;an air injection hole plate with a plurality of air injection holes formed therein to deliver to the combustion chamber the air along with the fuel injected from the fueling nozzles;cross tire tubes that each transport the combustion gases to an adjacent combustor and ignite the adjacent combustor during gas turbine ignition; andsupports for fixing the air injection hole plate to the fuel header,wherein the supports are provided so as to foe of the same phase as that of the cross fire tubes.2. The gas turbine combustor according to claim 1 , further comprising:a firing burner that supplies and burns the fuel during gas turbine ignition; anda non-firing burner inactivated during gas turbine ignition;wherein:part of the non-firing burner is formed by a row of the air injection holes annularly arranged around an outer circumference of the firing burner; andat the same phase position as that in which the cross fire tubes are disposed, the non-firing burner has a region particularly large in hole pitch between adjacent air-injection holes arranged in the row of the air injection holes that forms part of the non-firing burner.3. The gas turbine combustor according to claim 1 , further comprising porous plates placed downstream of the supports so as to extend in a direction parallel to a flow of air moving from an outer circumferential side of the air injection hole plate claim 1 , towards an inner circumferential side thereof.4. The gas turbine combustor according to claim 1 , wherein a plurality of air injection holes arranged at the same phase position as that of the cross fire tubes have a ...

COMBUSTION CHAMBER AND A METHOD OF MIXING FUEL AND AIR IN A COMBUSTION CHAMBER

A combustion chamber including a first fuel injector and a second fuel injector, the first and second fuel injectors being arranged to inject fuel into a mainstream flow of air with the second fuel injector arranged downstream of the first fuel injector. A method of mixing fuel and air in a combustion chamber, including injecting fuel into a mainstream flow of air with a first fuel injector; injecting fuel into the mainstream flow of air with a second fuel injector, which is arranged downstream of the first fuel injector; injecting fuel into the mainstream flow with the first fuel injector such that the resulting mixture between the first and second fuel injectors has an equivalence ratio less than the lean flame stability limit; and injecting fuel into the mainstream flow with the second fuel injector such that a combustion zone is provided downstream of the second fuel injector. 1. A combustion chamber comprising a first fuel injector and a second fuel injector , the first and second fuel injectors being arranged to inject fuel into a mainstream flow of air with the second fuel injector arranged downstream of the first fuel injector ,wherein the first fuel injector is configured to inject fuel into the mainstream flow such that the resulting mixture between the first and second fuel injectors has an equivalence ratio less than the lean flame stability limit and the second fuel injector is configured to inject fuel into the mainstream flow such that a combustion zone is provided downstream of the second fuel injector.2. The combustion chamber of claim 1 , wherein the combustion chamber comprises a longitudinal axis and the second fuel injector is arranged downstream of the first fuel injector in a substantially longitudinal direction.3. The combustion chamber of claim 1 , wherein the resulting mixture between the first and second fuel injectors has an equivalence ratio less than 0.5.4. The combustion chamber of claim 1 , wherein the combustion chamber further ...

PREMIX BURNER OF THE MULTI-CONE TYPE FOR A GAS TURBINE

The invention relates to a premix burner of the multi-cone type for a gas turbine that includes a plurality of shells which are arranged around a central burner axis and are parts of a virtual, axially extending common cone, which opens in a downstream direction, whereby said parts are displaced perpendicular to said burner axis such that a tangential slot is defined between each pair of adjacent shells. The flame front of such a burner is stabilized by providing a virtual common cone with a cone angle, which varies in axial direction. 1. A premix burner of the multi-cone type for a gas turbine , said premix burner comprising:a plurality of shells, which are arranged around a central burner axis and are parts of a virtual, axially extending common cone, which opens in a downstream direction,wherein said parts are displaced perpendicular to said burner axis such that a tangential slot is defined between each pair of adjacent shells, andwherein said virtual common cone has a cone angle which varies in axial direction.2. The premix burner according to claim 1 , wherein the cone angle of the virtual common cone increases in the downstream direction.3. The premix burner according to claim 1 , wherein the variation of the cone angle of the virtual common cone is generated by twisting said common cone around the central burner axis.4. The premix burner according to claim 3 , wherein the surface area of the twisted common cone is generated by rotating a meridian around the central burner axis claim 3 , one end of which is rotated around the central burner axis relative to the other end by a predetermined twist angle claim 3 , and wherein the shells are generated by cutting said virtual common cone along respective meridians.5. The premix burner according to claim 4 , wherein the twist angle is equal to or larger than 30°.6. The premix burner according to claim 4 , wherein the twist angle is equal to or larger than 60°.7. The premix burner according to wherein the common ...

PREMIX BURNER OF THE MULTI-CONE TYPE FOR A GAS TURBINE

The invention relates to a premix burner of the multi-cone type for a gas turbine, that includes a plurality of shells, which are arranged around a central burner axis and are parts of a virtual, axially extending common cone, which opens in a downstream direction, whereby said parts are displaced perpendicular to said burner axis such that a tangential slot is defined between each pair of adjacent shells. A disadvantageous transition piece between the shells and a downstream mixing tube is avoided by bordering the downstream ends of the shells by intersecting planes, which are defined by intersecting said shells with a virtual coaxial cylinder of a predetermined radius. 1. A premix burner of the multi-cone type for a gas turbine , said premix burner comprising:a plurality of shells which are arranged around a central burner axis and are parts of a virtual, axially extending common cone, which opens in a downstream direction, andwherein said parts are displaced perpendicular to said burner axis such that a tangential slot is defined between each pair of adjacent shells, andwherein the downstream ends of the shells are bordered by intersecting planes, which are defined by intersecting said shells with a virtual coaxial cylinder of a predetermined radius.2. The premix burner according to claim 1 , wherein each of the shells is equipped with a premix gas channel extending along an axially oriented edge of the respective shell such that a gas can be injected from said premix gas channel through gas injection holes into a stream of air entering the interior of the arrangement of shells through the adjacent slot claim 1 , and that the downstream ends of the premix gas channels are bordered by intersecting planes claim 1 , which are defined by intersecting said premix gas channels with said virtual coaxial cylinder of said predetermined radius.3. The premix burner according to claim 1 , further comprising a cylindrical burner ring with an inner radius similar to said ...

COMBUSTOR ARRANGEMENT, ESPECIALLY FOR A GAS TURBINE

The invention relates to a combustor arrangement that includes a combustion chamber with a front panel, and a premix burner of the multi-cone type, which is connected to said front panel though an elongated mixing zone in an axially moveable fashion by means of a sealed sliding joint. A wide range of axial variation of the burner with a minimized influence of the leakage air flow on the oxidation process within the flame is achieved by positioning said sealed sliding joint upstream of said mixing zone. 1. A combustor arrangement , especially for a gas turbine , comprising:a combustion chamber with a front panel, anda premix burner of the multi-cone type, which is connected to said front panel though an elongated mixing zone in an axially moveable fashion by means of a sealed sliding joint,wherein said sealed sliding joint is positioned upstream of said mixing zone.2. The combustor arrangement according to claim 1 ,wherein the sealed sliding joint is made up by a coaxial sliding arrangement of a cylindrical burner ring and an essentially cylindrical burner sleeve, andwherein said burner ring is fixed to said burner and said burner sleeve is fixed to and part of said front panel, and a seal is provided between said burner ring and burner sleeve.3. The combustor arrangement according to claim 2 , wherein said burner ring is surrounded by said burner sleeve.4. The combustor arrangement according to claim 3 , wherein said burner ring extends upstream of the downstream end of the burner claim 3 , and the seal is positioned at the upstream end of the burner ring.5. The combustor arrangement according to claim 2 , wherein the burner sleeve has a conically widening burner outlet at the transition to said combustion chamber.6. The combustor arrangement according to claim 4 , wherein the burner sleeve includes purge air holes upstream of the seal to purge the gap between burner ring and burner sleeve with air.7. The combustor arrangement according to wherein said premix burner ...

COMBUSTOR AND GAS TURBINE PROVIDED WITH SAME

A combustor () according to the invention includes: a combustor basket () in which compressed air and fuel are mixed with each other and the mixture is combusted; a transition piece () in which a tip portion of the combustor basket () is inserted with a gap (C) therebetween; a spring clip () that seals the gap between the combustor basket () and the transition piece (); a throttle section () that is provided in an opening portion (Ck) of the gap (C) that is opened to the transition piece () on the tip side of the combustor basket (), and narrows an opening area of the opening portion (Ck), compared to the base end side; and cooling device () for cooling the throttle section (). 1. A combustor comprising:a combustor basket in which compressed air and fuel are mixed with each other and the mixture is combusted;a transition piece in which a tip portion of the combustor basket is inserted with a gap therebetween;a spring clip that seals the gap between the combustor basket and the transition piece;a throttle section that is provided in an opening portion of the gap that is opened to the transition piece on the tip side of the combustor basket, and narrows an opening area of the opening portion, compared to the base end side of the combustor basket; andcooling device for cooling the throttle section.2. The combustor according to claim 1 , wherein the throttle section is provided by projecting an inner surface of the transition piece to the combustor basket.3. The combustor according to claim 1 , wherein the throttle section is provided by projecting an outer surface of the combustor basket to the transition piece.4. The combustor according to claim 1 , wherein the cooling device cools the throttle section by injecting a cooling fluid to the throttle section.5. The combustor according to claim 4 , wherein the cooling device has an injection section that injects the cooling fluid toward the opening portion claim 4 , and a guide section that guides the injected cooling ...

COMBUSTION CHAMBER WITH A WALL SECTION AND A BRIM ELEMENT

A combustion chamber for a gas turbine is proposed. The combustion chamber has a wall section and a brim element. The wall section has an inlet aperture for injecting a cooling medium into the combustion chamber. The brim element is mounted to an inner face of the wall section. The brim element is formed in such a way that a projected area of the brim element onto the inner face along a direction of a normal of the inner face at least partially covers the inlet channel. 113.-. (canceled)14. A combustion chamber for a gas turbine , comprising:a wall section comprising an inlet channel for injecting a cooling medium into the combustion chamber, anda brim element mounted to an inner face of the wall section,wherein the brim element is formed in such a way that a projected area of the brim element onto the inner face along a direction of a normal of the inner face at least partially covers the inlet channel,wherein the wall section further comprises a fuel injection aperture through which fuel is injectable into the combustion chamber,wherein the wall section further comprises a groove formed within the inner face,wherein the groove forms a part of an end section of the inlet channel,wherein the groove is formed such that the groove runs at least partially along a predefined direction for guiding the cooling medium streaming through the inlet channel,wherein the groove has a curved shape along a circumferential direction around a centre axis of the fuel injection aperture,wherein the inlet channel forms together with an edge of the brim element and the inner face an inlet aperture through which the cooling medium streams inside the combustion chamber, andwherein the brim element is formed in such a way that the cooling medium streaming through the inlet aperture is guided partially along the circumferential direction around the centre axis of the fuel injection aperture.15. The combustion chamber according to claim 14 ,wherein the wall section further comprises at least ...

Combustor nozzle assembly, combustor equipped with the same, and gas turbine

A combustor nozzle assembly includes: a nozzle mounting base which blocks a combustor insertion opening formed in a turbine casing; a nozzle rod which passes through the nozzle mounting base and has a rod tip portion and a rod base end portion; an oil fuel pipe which is as a whole inserted into the nozzle rod, which has a pipe tip portion and a pipe base end portion, in which fuel is supplied to the inside through the rod base end portion, and which injects the fuel from the pipe tip portion through the rod tip portion; and an O-ring which is disposed in the rod base end portion and suppresses leakage of fuel to the pipe tip portion side between the inner periphery side of the nozzle rod and the outer periphery side of the oil fuel pipe.

BURNER FOR A GAS TURBINE

A burner for a gas turbine includes a burner housing, and a pilot combustor having a supply module providing pilot fuel and pilot air into a pilot combustion room being enclosed by a pilot combustor housing having a tapered exit with a throat of a defined length into the resulting main flow direction. The throat discharges a concentration of radicals are and heat generated in the pilot combustion room into a main combustion room enclosed by the burner housing. The interior cross section area of the throat deviates from a circle by means of flow guiding elements provided as protrusions with a defined radial height or as recesses with a defined depth extending longitudinally along the direction of the burner axis to give the discharging flow a defined velocity distribution with regard to a circumferential direction. 110-. (canceled)11. A burner for a gas turbine , comprising:a burner housing, anda pilot combustor, comprising a supply module providing pilot fuel and pilot air into a pilot combustion room being enclosed by a pilot combustor housing comprising a tapered exit with a throat of a defined length into the resulting main flow direction, said throat discharging a concentration of radicals are and heat generated in said pilot combustion room into a main combustion room enclosed by said burner housing,wherein a burner axis is defined by a centre line of said throat extending in the direction of the resulting main flow through said throat,wherein the interior cross section area of said throat deviates from a circle by means of flow guiding elements provided as protrusions with a defined radial height or as recesses with a defined depth extending longitudinally along the direction of the burner axis to give the discharging flow a defined velocity distribution with regard to a circumferential direction.12. The burner according to claim 11 , wherein said cross section protrusions or recesses form a cross or a star shape of said radial cross-section of the throat.13. ...

REHEAT BURNER AND METHOD OF MIXING FUEL/CARRIER AIR FLOW WITHIN A REHEAT BURNER

The invention refers to a reheat burner that includes a flow channel for a hot gas flow with a lance arranged along said flow channel, protruding into the flow channel for injecting a fuel over an injection plane perpendicular to a channel longitudinal axis, wherein the channel and lance define a vortex generation zone upstream of the injection plane and a mixing zone downstream of the injection plane in the hot gas flow direction. The mixing zone provides at least one axially region having different cross sectional areas along its longitudinal axis with continuously changing shape, or having non circular cross section areas which change location along its longitudinal axis by continuously rotation around the longitudinal axis. 2. The reheat burner of claim 1 , wherein the at least one axially region extends in one related piece over the entire mixing zone.4. The reheat burner of claim 3 , wherein the first and second axially regions are related axially directly or indirectly.5. The reheat burner of wherein the axially region having the different cross sectional areas along its longitudinal axis with continuously changing shape provides different cross sectional areas which cannot be brought into line by scaling only.6. The reheat burner of wherein the non-circular cross section areas which changes location along the longitudinal axis by continuously rotation around the longitudinal axis are constant in shape.7. The reheat burner of wherein at least two of the non-circular cross section areas which changes location along its longitudinal axis by continuously rotation around the longitudinal axis differ in size.8. The reheat burner of wherein the mixing zone provides at least one axially region having changing cross sectional areas along its longitudinal axis which change shape and/or location in flow direction starting at a first cross section area shape CSASand ending at a last cross section area shape CSASin one of the following manner:{'sub': first', 'last, 'a) ...

NOZZLE, GAS TURBINE COMBUSTOR AND GAS TURBINE

A pilot nozzle, a gas turbine combustor and a gas turbine are provided with a nozzle main body having a fuel passage, a cover ring arranged at an outside of a front end-outer peripheral portion of the nozzle main body at a predetermined interval to form an inner air passage and capable of injecting air toward a front side of the nozzle main body, a plurality of nozzle tips that includes a fuel injection nozzle attached to a front end portion of the cover ring at a predetermined interval in a circumferential direction to communicate with the fuel passage and is able to inject fuel toward an outside of injection air from the inner air passage, and a swirling force application unit that applies a swirling force to air injected through the inner air passage. 1. A nozzle , comprising:a nozzle main body having a fuel passage;a cover ring arranged at an outside of a front end-outer peripheral portion of the nozzle main body at a predetermined interval to form an inner air passage, for injecting air toward a front side of the nozzle main body;fuel injection nozzles attached to a front end portion of the cover ring at a predetermined interval in a circumferential direction to communicate with the fuel passage; anda swirling force application unit for applying a swirling force to air injected through the inner air passage.2. The nozzle according to claim 1 , wherein the swirling force application unit has guide portions provided at an outlet of the inner air passage.3. The nozzle according to claim 2 , wherein the fuel injection nozzles are provided at a plurality of nozzle tips for injecting fuel to an outside of injection air from the inner air passage claim 2 , and the guide portions are provided at the plurality of nozzle tips.4. The nozzle according to claim 1 , wherein guide portions are provided so as not to be positioned at the same row as the plurality of nozzle tips of the cover ring in the circumferential direction.5. The nozzle according to claim 1 , wherein a ...

LEAN BURN INJECTORS HAVING MULTIPLE PILOT CIRCUITS

A fuel injector for a gas turbine engine includes a nozzle body defining a central axis and having a main fuel circuit. An air circuit is formed within the nozzle body inboard of the main fuel circuit. A primary pilot fuel circuit is formed within the nozzle body inboard of the air circuit. A secondary pilot fuel circuit is formed within the nozzle body inboard of the air circuit and outboard of the primary pilot fuel circuit. 1. A fuel injector for a gas turbine engine comprising:a) a nozzle body defining a central axis and having a main fuel circuit;b) an air circuit formed within the nozzle body inboard of the main fuel circuit;c) a primary pilot fuel circuit formed within the nozzle body inboard of the air circuit; andd) a secondary pilot fuel circuit formed within the nozzle body inboard of the air circuit and outboard of the primary pilot fuel circuit.2. An injector as recited in claim 1 , further comprising a pilot air circuit inboard of the secondary pilot fuel circuit claim 1 , wherein the primary pilot fuel circuit is inboard of the pilot air circuit.3. An injector as recited in claim 2 , wherein at least one of the main fuel circuit and secondary pilot fuel circuit includes a diverging prefilming air-blast atomizer.4. An injector as recited in claim 2 , wherein the primary pilot fuel circuit includes a pressure swirl atomizer.5. An injector as recited in claim 4 , wherein the pressure swirl atomizer is defined in an inner air swirler along the axis of the nozzle body.6. An injector as recited in claim 1 , wherein the primary pilot fuel circuit includes a primary pressure swirl atomizer on the axis of the nozzle body.7. An injector as recited in claim 6 , wherein the secondary pilot fuel circuit includes a secondary pressure swirl atomizer outboard of the primary pressure swirl atomizer.8. An injector as recited in claim 7 , wherein the primary and secondary pressure swirl atomizers are combined as a dual orifice atomizer.9. An injector as recited in claim ...

METHOD FOR MIXING A DILUTION AIR IN A SEQUENTIAL COMBUSTION SYSTEM OF A GAS TURBINE

The invention concerns a method for mixing a dilution air with a hot main flow in sequential combustion system of a gas turbine, wherein the gas turbine essentially comprises at least one compressor, a first combustor which is connected downstream to the compressor, and the hot gases of the first combustor are admitted to at least one intermediate turbine or directly or indirectly to at least one second combustor. The hot gases of the second combustor are admitted to a further turbine or directly or indirectly to an energy recovery, wherein at least one combustor runs under a caloric combustion path having a can-architecture. At least one dilution air injection is introduced into the first combustor, and wherein the direction of the dilution air injection is directed against or in the direction of the original swirl flow inside of the first combustor. 1. A method for mixing a dilution air with a hot main flow in a sequential combustion system of a gas turbine , wherein the gas turbine essentially comprises at least one compressor , a first combustor which is connected downstream to the compressor , and the hot gases of the first combustor are admitted to at least one intermediate turbine or directly or indirectly to at least one second combustor , wherein the hot gases of the second combustor are admitted to a further turbine or directly or indirectly to an energy recovery , wherein at least one combustor runs under a caloric combustion path having a can architecture , and wherein at least one dilution air injection is introduced into the first combustor , and wherein the resulting swirl flow through the dilution air injection is directed against or in the direction of the original swirl flow inside of the first combustor.2. The method as claimed in claim 1 , wherein the first and second combustor run under a caloric combustion path having a can-architecture.3. The method as claimed in claim 1 , wherein the first combustor runs under a caloric combustion path having ...

Seal for a perforated plate

A cooling circuit of a gas turbine passes an airflow through a combustor section that includes a plurality of mixing tubes for transporting a fuel/air mixture and a perforated plate including a plurality of impingement holes and a plurality of tube holes for accommodating the mixing tubes. The tube holes and the mixing tubes form a plurality of annulus areas between the perforated plate and the mixing tubes. The impingement holes and the annulus areas are configured to pass the airflow through the perforated plate. A flow management device modifies an effective size of the annulus areas to control a distribution of the airflow through the impingement holes and the annulus areas of the perforated plate to enhance cooling efficiency.

BURNER ARRANGEMENT

The invention refers to burner arrangement for producing hot gases to be expanded in a gas turbine, including a burner inside a plenum, where the burner has means for fuel injection, means for air supply and means for generating an ignitable fuel/air mixture inside the burner, and a combustion chamber following downstream said burner having an outlet being fluidly connected to the gas turbine. The invention is characterized in that the means for air supply includes at least two separate flow passages, and that the one of the two flow passages is fed by a first supply pressure and the other flow passage is fed by a second supply pressure. 1. A burner arrangement for producing hot gases to be expanded in a gas turbine , comprising:a burner inside a plenum, said burner has means for fuel injection, means for air supply and means for generating an ignitable fuel/air mixture inside the burner, anda combustion chamber following downstream said burner having an outlet being fluidly connected to the gas turbine, andwherein the means for air supply comprise at least two separate flow passages, wherein one of the two flow passages is fed by a first supply pressure and the other flow passage is fed by a second supply pressure.2. A burner arrangement according to claim 1 , wherein one of the two flow passages is fluidly connected to the plenum in which the first pressure prevails which is fluidly connected to a compressor and the other flow passage is fluidly connected to an interspace in which the second pressure prevails and which is bordered by a combustor liner having at least one fluidly access to the plenum.3. A burner arrangement according to claim 2 , wherein the at least one fluidly access of the combustor liner to the plenum is in a downstream region of the combustion chamber.4. A burner arrangement according to wherein one of the two flow passages is an outer flow passage which surrounds the other flow passage claim 1 , which is a so called inner flow passage.5. A ...

COOLED PILOT FUEL LANCE

A device for injecting fuel into a combustion chamber of a gas turbine is provided, having a distribution section to which a first fuel channel, a second fuel channel and an injection channel are coupled. The first fuel channel and the second fuel channel are arranged such that a) fuel is transportable by one of the first fuel channel and the second fuel channel to the distribution section, and b) a first quantity of fuel is transportable by the other one of the first fuel channel and the second fuel channel out of the distribution section. The injection channel is arranged such that a second quantity of fuel is injectable from the distribution section into the combustion chamber. The device further comprises an end cap with a protrusion having the injection channel inside, and extending inside the inner tube. 1. A device for injecting fuel into a combustion chamber of a turbine , comprisinga distribution section,a first fuel channel which is coupled to the distribution section,a second fuel channel which is coupled to the distribution section,an injection channel which is coupled to the distribution section, a) fuel is transportable by one of the first fuel channel and the second fuel channel to the distribution section, and', 'b) a first quantity of fuel is transportable by the other one of the first fuel channel and the second fuel channel out of the distribution section, and, 'wherein the first fuel channel and the second fuel channel are arranged in such a way that'}wherein the injection channel is arranged in such a way that a second quantity of fuel is injectable from the distribution section into the combustion chamber,a pressure control arrangement which is arranged for controlling a pressure of the fuel in at least one of the first fuel channel and the second channel such that the first quantity and the second quantity is controllable,an inner tube,an outer tube which surrounds the inner tube,wherein the inner tube and the outer tube are arranged coaxial ...

MULTIPOINT FUEL INJECTION ARRANGEMENTS

A multipoint fuel injection system includes a plurality of fuel manifolds. Each manifold is in fluid communication with a plurality of injectors arranged circumferentially about a longitudinal axis for multipoint fuel injection. The injectors of separate respective manifolds are spaced radially apart from one another for separate radial staging of fuel flow to each respective manifold. 1. A multipoint fuel injection system comprising:a plurality of fuel manifolds, wherein each manifold is in fluid communication with a plurality of injectors arranged circumferentially about a longitudinal axis for multipoint fuel injection in an annular combustor, wherein the injectors of separate respective manifolds are spaced radially apart from one another for separate radial staging of fuel flow to each respective manifold.2. A multipoint fuel injection system as recited in claim 1 , wherein each manifold includes a conduit extending circumferentially around a combustor claim 1 , and wherein the injectors in fluid communication with each manifold are spaced apart circumferentially from one another.3. A multipoint fuel injection system as recited in claim 1 , further comprising a plurality of feed arms for fluid communication between the manifolds and the injectors claim 1 , wherein each feed arm forms an injector tree with at least one injector mounted thereto corresponding to each manifold claim 1 , and wherein each feed arm includes fuel conduits connected to provide fluid communication between respective manifolds and injectors.4. A multipoint fuel injection system as recited in claim 1 , where the injectors of each of the manifolds are circumferentially staggered from radially adjacent injectors of other manifolds.5. A multipoint fuel injection system as recited in claim 1 , further comprising a combustor having an annular upstream wall claim 1 , an outboard wall extending downstream from the upstream wall claim 1 , and an inboard wall extending downstream from the upstream ...

GAS TURBINE COMBUSTOR

An object of the present invention is to provide a combustor having a premixing burner, wherein a conical flame can be formed and the metal temperature at a liner and a burner end face can be reduced. 1. A gas turbine combustor comprising:at least one premixing burner for premixing gaseous fuel with air and jetting the mixed gas into a chamber;a cylinder disposed on an outer circumference of the premixing burner so as to surround the premixing burner and connected to a burner outlet end face which is an end face of the premixing burner on the chamber side; anda plurality of air supply holes formed in the cylinder;wherein an interval defined between the adjacent air supply holes is smaller than a quenching distance in the premixed gas jetted from the premixing burner, andwherein an interval defined between each air supply hole and the burner outlet end face is smaller than the quenching distance in the premixed gas jetted from the premixing burner.2. The gas turbine combustor according to claim 1 , whereinthe premixing burner includes an air hole plate with a plurality of air holes and fuel nozzles adapted to jet gaseous fuel into the air hole of the air hole plate, andthe gas turbine combustor has at least one burner configured by arranging, as a set, a plurality of the fuel nozzles and of the air holes such that each of the fuel nozzles are paired with each of the air holes.3. The gas turbine combustor according to claim 1 ,wherein the interval between the adjacent air supply holes and the interval between the air supply hole and the outlet end face of the burner are each narrower than 1 cm.4. The gas turbine combustor according to claim 1 ,wherein the gas turbine combustor has a multi-burner composed of a plurality of the premixing burners.5. The gas turbine combustor according to claim 1 , further comprising:a liner for surrounding the chamber; anda seal member disposed between the liner and the cylinder inserted into the liner to secure the cylinder and the ...

GAS TURBINE COMBUSTOR

An annular type gas turbine combustor having a plurality of fuel nozzle assemblies () on a circumference includes a pilot nozzle unit () for spraying a fuel for diffusive combustion from a pilot outer peripheral nozzle () into a combustion chamber (), a main nozzle unit () provided so as to surround the pilot nozzle unit () for spraying a fuel for premix combustion, and a flow guide () disposed on a downstream side of each of the fuel nozzle assemblies () and having a sectional area of a passage for air and air-fuel mixture from each of the fuel nozzle assemblies (), which gradually increase in a downstream direction. 1. An annular gas turbine combustor comprising:a plurality of fuel nozzle assemblies disposed on a circumference; anda flow guide mounted on a downstream side of the fuel nozzle assembly and having a sectional area of a passage for an air and an air-fuel mixture from the fuel nozzle assembly, which sectional area is gradually increased towards the downstream side; whereineach of the fuel nozzle assemblies includesa first fuel injection unit to spray a fuel from a spraying nozzle into a combustion chamber, anda second fuel injection unit provided so as to surround the first fuel injection unit and operable to spray a fuel, anda main outlet flare which forms an outlet of the fuel nozzle assembly, the main outlet flare flaring outwardly toward the downstream side of the fuel nozzle assembly,the flow guide being disposed radially outwardly of the main outlet flare.2. The annular gas turbine combustor as claimed in claim 1 , whereinthe flow guide has a transverse sectional shape that is round and has an upstream end of an inner diameter which is equal to or somewhat greater than an air outlet diameter of the fuel nozzle assembly.3. The annular gas turbine combustor as claimed in claim 1 , wherein the flow guide has a conical portion of a shape flared in a conical shape from the upstream side towards the downstream side.4. The annular gas turbine combustor ...

SEAL FOR FUEL DISTRIBUTION PLATE

A fuel flow passes through a micromixer section of a gas turbine that includes a plurality of mixing tubes for transporting a fuel/air mixture and a distribution plate including a plurality of distribution holes and a plurality of tube holes for accommodating the mixing tubes. Each of the mixing tubes includes a plurality of fuel holes through which fuel enters the mixing tubes. The tube holes and the mixing tubes form a plurality of annulus areas between the distribution plate and the mixing tubes. The distribution holes and the annulus areas are configured to pass the fuel flow through the distribution plate toward the fuel holes. A flow management device modifies an effective size of the annulus areas to control a distribution of the fuel flow through the distribution holes and the annulus areas of the distribution plate to provide a uniform fuel/air composition in each of the mixing tubes. 1. A gas turbine combustor , comprising:a plurality of mixing tubes arranged to transport a fuel/air mixture to a reaction zone for ignition, each mixing tube including a plurality of fuel holes through which fuel enters the respective mixing tube;a plate having a plurality of tube holes formed therein, the plurality of tube holes being configured to accommodate the plurality of mixing tubes thereby forming a plurality of annulus areas between the plate and the plurality of mixing tubes, the plurality of annulus areas being configured such that the fuel flows through the plurality of annulus areas, the plurality of fuel holes being arranged on a downstream side of the plate with respect to the fuel flow; anda flow management device engaging at least one of the plate and the plurality of mixing tubes and including a portion situated within the plurality of annulus areas to control a distribution of the fuel to the plurality of fuel holes.2. The gas turbine combustor of claim 1 , wherein the flow management device includes a plurality of metering elements for controlling a flow ...

ANNULAR COMBUSTION CHAMBER FOR A TURBINE ENGINE

An annular combustion chamber for a turbine engine, the chamber including an annular row of fuel injectors including heads engaged in fuel injection systems mounted in openings in the chamber end wall, each injector head including at least one fuel-passing helical channel for causing fuel to rotate about the longitudinal axis of the head, and each injection system including at least one swirler including air-passing channels of sections with axes that are inclined relative to the plurality axis of the swirler at an angle that is substantially equal to a helix angle of the helical channel, to within ±10°, and are oriented in a same direction as the channel about the longitudinal axis of the swirler. 113-. (canceled)14. An annular combustion chamber for a turbine engine , the chamber comprising:inner and outer coaxial annular walls connected together at their upstream ends by an annular wall forming a chamber end wall; andan annular row of fuel injectors including heads engaged in fuel injection systems mounted in openings in the chamber end wall, each injector head including at least one fuel-passing helical channel for causing fuel to rotate about the longitudinal axis of the head, and each injection system including at least one swirler on a same axis as the injector head and including substantially radial air-passing channels of elongate section having respective longitudinal axes,wherein the longitudinal axes of the sections of the channels are inclined relative to the longitudinal axis of the swirler at an angle that is substantially equal to a helix angle of the helical channel of the injector head, to within ±10°, and are oriented in a same direction as the channel about the longitudinal axis of the swirler.15. A chamber according to claim 14 , wherein the axes of the sections of the channels of the swirler are inclined at an angle lying in a range of 20° to 40° approximately relative to the longitudinal axis of the swirler.16. A chamber according to claim 14 ...

Variable length combustor dome extension for improved operability

The present invention discloses a novel apparatus and method for operating a gas turbine combustor having a structural configuration proximate a pilot region of the combustor which seeks to minimize the onset of thermo acoustic dynamics. The pilot region of the combustor includes a generally cylindrical extension having an outlet end with an irregular profile which incorporates asymmetries into the system so as to destroy any coherent structures.

Flamesheet combustor dome

The present invention discloses a novel apparatus and way for controlling a velocity of a fuel-air mixture entering a gas turbine combustion system. The apparatus comprises a hemispherical dome assembly which directs a fuel-air mixture along a portion of the outer wall of a combustion liner and turns the fuel-air mixture to enter the combustion liner in a manner coaxial to the combustor axis and radially outward of a pilot fuel nozzle so as to regulate the velocity of the fuel-air mixture.

COMBUSTER WITH RADIAL FUEL INJECTION

A combustor for a gas turbine engine includes an forward fuel injection system in communication with a combustion chamber and a downstream fuel injection system that communicates with the combustion chamber downstream of the forward fuel injection system. 1. A combustor for a gas turbine engine comprising:an forward fuel injection system in communication with a combustion chamber; anda downstream fuel injection system that communicates with said combustion chamber downstream of said forward fuel injection system.2. The combustor as recited in claim 1 , wherein said downstream fuel injection system at least partially surrounds said combustion chamber.3. The combustor as recited in claim 1 , wherein said downstream fuel injection system is radially inboard of said combustion chamber.4. The combustor as recited in claim 1 , wherein said downstream fuel injection system is radially outboard of said combustion chamber.5. The combustor as recited in claim 1 , wherein said downstream fuel injection system is radially outboard and radially inboard of said combustion chamber.6. The combustor as recited in claim 1 , wherein said downstream fuel injection system includes a multiple of fuel nozzle assemblies axially upstream of a necked region of said combustor.7. The combustor as recited in claim 1 , wherein said downstream fuel injection system includes a multiple of fuel nozzle assemblies within a first two-thirds of said combustor.8. The combustor as recited in claim 1 , wherein said downstream fuel injection system is radially inboard of said combustion chamber claim 1 , a main supply line of a radially inner fuel injection manifold extends through a forward assembly.9. The combustor as recited in claim 1 , wherein said downstream fuel injection system is radially inboard of said combustion chamber claim 1 , a main supply line of a radially inner fuel injection manifold extends through a vane in a turbine section downstream of said combustion chamber.10. A gas turbine ...

Combustor with radially staged premixed pilot for improved

The present invention discloses a novel apparatus and method for a mixing fuel and air in a gas turbine combustion system. The mixer helps to mix fuel and air while being able to selectively increase the fuel flow to a shear to a shear layer of a pilot flame in order to reduce polluting emissions. The mixer directs a flow of air radially inward into the combustion system and includes two sets of fuel injectors within each radially-oriented vane. A first plurality of fuel injectors operate independent of a second plurality of fuel injectors and the second plurality of fuel injectors are positioned to selectively modulate the fuel flow to the shear layer of the resulting pilot flame.

Variable flow divider mechanism for a multi-stage combustor

The present invention discloses a novel apparatus and way for altering the airflow to a gas turbine combustion system. The apparatus comprises a flow divider mechanism which splits the airflow surrounding a combustion liner into two distinct portions, one directed towards a pilot and one directed towards a main stage combustion. The flow divider mechanism is interchangeable so as to provide a way of altering airflow splits between stages of the combustion system.

Burner for a gas turbine

A burner for a gas turbine is provided. The burner has a pilot combustor, a supply module providing pilot fuel and air into a pilot combustion room enclosed by a pilot burner housing having a tapered exit throat discharging radicals and heat generated in a pilot combustion zone into a main combustion room. An equalizer has holes for main flow air entering a cavity in a radial direction with regard to a burner axis defined by centers of pilot combustion zone and main combustion zone. A fuel injector is downstream the equalizer to supply flow fuel into flow air. A swirler is downstream the injector to give flow distribution to the flow fuel and air entering the main combustion room. A channel leads from the equalizer to the swirler arranged circumferentially around the pilot burner housing to direct the main air flow from the equalizer in axial and circumferential direction.

COMBUSTOR APPARATUS IN A GAS TURBINE ENGINE

A combustor apparatus defining a combustion zone where air and fuel are burned to create high temperature combustion products. The combustor apparatus comprises an outer wall including a fuel inlet opening for receiving a fuel feed pipe. A coupling assembly is engaged with the fuel feed pipe at the fuel inlet opening to attach the fuel feed pipe to the outer wall. A fuel injection system is located in the interior volume of the outer wall and comprises fuel supply structure including a fuel feed block having a fuel intake passage aligned with the outlet portion of the fuel feed pipe. A coupling fastener is engaged against an exterior outer face of the fuel feed block to create a sealed coupling for containing fuel passing from the fuel feed pipe into the fuel feed block, and to secure the fuel feed block relative to the coupling assembly. 1. A combustor apparatus in a gas turbine engine , the combustor apparatus defining a combustion zone where air and fuel are burned to create high temperature combustion products , the combustor apparatus comprising:an outer wall including an inner face and an outer face, and defining an interior volume and comprising a fuel inlet opening;a fuel feed pipe that extends through the fuel inlet opening in the outer wall, the fuel feed pipe including an inlet portion and an outlet portion; coupling structure on the outer face of the outer wall adjacent to the fuel inlet opening, the coupling structure comprising a threaded inner coupling portion;', create a first sealed coupling with the coupling structure; and', 'secure the fuel feed pipe relative to the outer wall;, 'a fitting member disposed about and engaged with the inlet portion of the fuel feed pipe and comprising a threaded outer coupling portion, the threaded outer coupling portion being threadedly engaged with the inner coupling portion of the coupling structure to], 'a coupling assembly comprising a fuel intake passage aligned with the outlet portion of the fuel feed pipe; ...

Cap assembly for a bundled tube fuel injector

A cap assembly for a bundled tube fuel injector includes an impingement plate and an aft plate that is disposed downstream from the impingement plate. The aft plate includes a forward side that is axially separated from an aft side. A tube passage extends through the impingement plate and the aft plate. A tube sleeve extends through the impingement plate within the tube passage towards the aft plate. The tube sleeve includes a flange at a forward end and an aft end that is axially separated from the forward end. A retention plate is positioned upstream from the impingement plate. A spring is disposed between the retention plate and the flange. The spring provides a force so as to maintain contact between at least a portion of the aft end of the tube sleeve and the forward side of the aft plate.

SYSTEM AND METHOD FOR A FUEL NOZZLE

A system includes an oxidant compressor and a gas turbine engine turbine, which includes a turbine combustor, a turbine, and an exhaust gas compressor. The turbine combustor includes a plurality of diffusion fuel nozzles, each including a first oxidant conduit configured to inject a first oxidant through a plurality of first oxidant openings configured to impart swirling motion to the first oxidant in a first rotational direction, a first fuel conduit configured to inject a first fuel through a plurality of first fuel openings configured to impart swirling motion to the first fuel in a second rotational direction, and a second oxidant conduit configured to inject a second oxidant through a plurality of second oxidant openings configured to impart swirling motion to the second oxidant in a third rotational direction. The first fuel conduit surrounds the first oxidant conduit and the second oxidant conduit surrounds the first fuel conduit. 1. A system , comprising:an oxidant compressor; and [ a first oxidant conduit configured to inject a first oxidant through a plurality of first oxidant openings, wherein the plurality of first oxidant openings are configured to impart swirling motion to the first oxidant in a first rotational direction;', 'a first fuel conduit configured to inject a first fuel through a plurality of first fuel openings, wherein the first fuel conduit surrounds the first oxidant conduit, and the plurality of first fuel openings are configured to impart swirling motion to the first fuel in a second rotational direction; and', 'a second oxidant conduit configured to inject a second oxidant through a plurality of second oxidant openings, wherein the second oxidant conduit surrounds the first fuel conduit, and the plurality of second oxidant openings are configured to impart swirling motion to the second oxidant in a third rotational direction;, 'a plurality of diffusion fuel nozzles, wherein each of the plurality of diffusion fuel nozzles comprises, 'a ...

Combustor nozzle, and combustor and gas turbine including the same

A combustor nozzle capable of injecting fuel uniformly, and a combustor and gas turbine including the same are provided. The combustor nozzle includes a main cylinder having a fuel passage through which fuel flows, a nozzle shroud surrounding the main cylinder, and a fuel injection module disposed between the main cylinder and the nozzle shroud to inject fuel, wherein the fuel injection module includes a plurality of first struts protruding from the main cylinder and having strut injection holes to inject fuel, a first support tube coupled to outer ends of the first struts, and a plurality of second struts protruding from the first support tube and having strut injection holes to inject fuel, and each of the first and second struts includes a swirl guide inclined with respect to a longitudinal direction of the main cylinder.

GAS TURBINE FUEL COMPONENTS

Gas turbine combustion systems and fuel cartridge assemblies are provided. An exemplary combustion system may comprise a combustor including one or more components, such as a cylindrical combustion liner, a flow sleeve, a main mixer, a radial inflow swirler, a combustor dome, and a fuel cartridge assembly. An exemplary fuel cartridge assembly may comprise first and second fuel manifolds which are connected to respective fuel circuits which supply fuel, such as liquid fuel, through a plurality of fuel passages within the fuel cartridge assembly or to other locations within an associated combustor. The fuel cartridge assembly may further include a plurality of fuel injector tips located at a tip plate of the fuel cartridge assembly through which fuel may be supplied to an associated combustor. 1. A fuel cartridge assembly , comprising: an aft portion comprising at least a first fuel manifold and a second fuel manifold;', 'a main body extending from the aft portion and having a first passageway contained therein;', 'a plurality of fuel passages extending axially from the aft portion to a tip plate, each fuel passage of the plurality of fuel passages in communication with the first fuel manifold or the second fuel manifold; and', 'the tip plate coupled to an end of the main body opposite the aft portion, the tip plate having at least one ignition opening and a plurality of openings corresponding to the plurality of axially extending fuel passages., 'a centerbody comprising2. The assembly of claim 1 , wherein the centerbody further comprises a plurality of fuel injector tips coupled to the respective plurality of fuel passages claim 1 , each of the plurality of fuel injector tips circumscribed at least partially by one of the plurality of openings in the tip plate.3. The assembly of claim 2 , wherein the fuel injector tips are positioned in the respective plurality of openings such that the fuel injector tips are movable relative to the tip plate.4. The assembly of claim ...

FUEL DISPENSING APPARATUS AND METHOD OF OPERATION

A fluid dispensing apparatus that may be additive manufactured as one unitary piece and may be a fuel injector for a gas turbine engine includes a radial displacement bellows having an outer surface that faces and may be exposed to a surrounding environment and an interior surface that faces and may define at least in-part a flowpath extending along a centerline. The radial displacement bellows is constructed and arranged to move between an expanded state when a pressure differential between the environment and the flowpath is low to a restricted state when the pressure differential is high. 1. An additive manufactured fluid dispensing apparatus comprising:a radial displacement bellows having an outer surface exposed to a surrounding environment and an interior surface defining at least in-part a flowpath extending along a centerline, and wherein the radial displacement bellows is constructed and arranged to move between an expanded state when a pressure differential between the environment and the flowpath is low to a restricted state when the pressure differential is high.2. The additive manufactured fluid dispensing apparatus set forth in claim 1 , wherein the radial displacement bellows is made of a metal.3. The additive manufactured fluid dispensing apparatus set forth in claim 2 , wherein the radial displacement bellows generally has a wall thickness of about 0.004 inches to 0.008 inches.4. The additive manufactured fluid dispensing apparatus set forth in claim 1 , wherein the radial displacement bellows has a plurality of axially displaced convolutions.5. The additive manufactured fluid dispensing apparatus set forth in further comprising:a fluid dispensing spray nozzle defining in-part the flowpath.6. The additive manufactured fluid dispensing apparatus set forth in further comprising:an axial displacement device defining in-part the flowpath, and wherein the device is axially extended when the radial displacement device is in the restricted state and ...

ANNULAR COMBUSTION CHAMBER IN A TURBINE ENGINE

The invention relates to a device for supporting and centring a fuel injector in a turbine engine combustion chamber, which includes means for centring a fuel injector along an axis, which are movable in a plane that is radial to the centring axis () in supporting means intended for being attached to the bottom of an annular chamber (). According to the invention, the centring means include at least two radially external tabs () each inserted respectively in a circumferential recess () of the supporting means, the device including circumferential abutment means () of the radial tabs () of the centring means in the circumferential recesses (), the circumferential abutment means being configured such as to enable a greater angular displacement of a first () one of the radial tabs in a first circumferential recess () relative to a second () one of the radial tabs in a second circumferential recess (). 1. Device for supporting and centring a fuel injector in a turbine engine combustion chamber , which includes means for centring a fuel injector along an axis , which are movable in a plane that is radial to the centring axis in supporting means intended for being attached to the bottom of an annular chamber , wherein the means of centring include at least two radially external tabs , each inserted respectively in a circumferential recess of the supporting means , wherein the device includes circumferential abutment means of the radial tabs of the means of centring in the circumferential recesses , wherein the circumferential abutment means are configured such as to enable a greater angular displacement of a first one of the radial tabs in a first circumferential recess relative to a second one of the radial tabs in a second circumferential recess.2. Device according to claim 1 , wherein the supporting means include an annular edge extending along the centring axis claim 1 , in which the first and second circumferential recesses accommodating the first and second radial ...

FUEL NOZZLE WITH FLEXIBLE SUPPORT STRUCTURES

A fuel nozzle apparatus for a gas turbine engine includes: a fuel discharge element having a discharge orifice communicating with a fuel supply connection; a static supporting structure; and a cantilevered flexible support structure interconnecting the supporting structure and the fuel discharge element, the flexible support structure having a first end connected to the static supporting structure, and a second end connected to the fuel discharge element. 1. A fuel nozzle apparatus , comprising:a fuel discharge element having a discharge orifice communicating with a fuel supply connection;a static supporting structure; anda cantilevered flexible support structure interconnecting the supporting structure and the fuel discharge element, the flexible support structure having a first end connected to the static supporting structure, and a second end connected to the fuel discharge element.2. The fuel nozzle apparatus of wherein the fuel discharge element claim 1 , the static supporting structure claim 1 , and the flexible support structure all form part of a single monolithic construction.3. A fuel nozzle apparatus claim 1 , comprising:an annular inner wall disposed coaxially along a centerline axis and having a fuel discharge orifice at a first end thereof;an annular outer wall surrounding the inner wall; anda support arm interconnecting the inner wall and the outer wall, wherein the support arm extends at an acute angle to the centerline axis.4. The apparatus of wherein:a forward end of the support arm joins the outer wall at a forward junction;an aft end of the support arm joins the inner wall at an aft junction; andeach of the forward and aft junctions has a smoothly-curved, arcuate shape.5. The apparatus of wherein the support arm is a single claim 3 , fully-annular structure.6. The apparatus of wherein the inner wall has a metering plug disposed therein claim 3 , the metering plug including at least one spray hole communication with the fuel discharge orifice.7. A ...

FUEL NOZZLE STRUCTURE FOR AIR ASSIST INJECTION

A fuel nozzle includes an outer body extending parallel to a centerline axis, having a generally cylindrical exterior surface, forward and aft ends, and a plurality of openings through the exterior surface. The fuel nozzle further includes an inner body inside the outer body, cooperating with the outer body to define an annular space, and a main injection ring inside the annular space, the main injection ring including fuel posts extending therefrom. Each fuel post is aligned with one of the openings and separated from the opening by a perimeter gap which communicates with the annular space. There is a circumferential main fuel gallery in the main injection ring, and a plurality of main fuel orifices, wherein each orifice communicates with the main fuel gallery and extends through one of the fuel posts. 1. A fuel nozzle apparatus , comprising:an annular outer body, the outer body extending parallel to a centerline axis, the outer body having a generally cylindrical exterior surface extending between forward and aft ends, and having a plurality of openings passing through the exterior surface;an annular inner body disposed inside the outer body, cooperating with the outer body to define an annular space;an annular main injection ring disposed inside the annular space, the main injection ring including an annular array of fuel posts extending radially outward therefrom;each fuel post being aligned with one of the openings in the outer body and separated from the opening by a perimeter gap which communicates with the annular space;a main fuel gallery extending within the main injection ring in a circumferential direction; anda plurality of main fuel orifices, each main fuel orifice communicating with the main fuel gallery and extending through one of the fuel posts.2. The apparatus of claim wherein:each opening communicates with a conical well inlet formed on an inner surface of the outer body; andeach fuel post is frustoconical in shape and includes a conical lateral ...

FUEL NOZZLE ASSEMBLY

A system includes a fuel nozzle including a first wall disposed about a central axis, a second wall disposed about the first wall, and a plurality of lobes disposed between the first and second walls, wherein the plurality of lobes is spaced about the central axis to define a plurality of flow passages, and the plurality of flow passages is configured to output a plurality of flows into a flame region. 1. A system , comprising: a first wall disposed about a central axis;', 'a second wall disposed about the first wall; and', 'a plurality of lobes disposed between the first and second walls, wherein the plurality of lobes is spaced about the central axis to define a plurality of flow passages, and the plurality of flow passages is configured to output a plurality of flows into a flame region., 'a fuel nozzle, comprising2. The system of claim 1 , wherein the plurality of lobes extends to a downstream end portion of the fuel nozzle.3. The system of claim 1 , wherein the plurality of lobes extends from the second wall toward the first wall claim 1 , or the plurality of lobes extends from the first wall toward the second wall claim 1 , or a combination thereof.4. The system of claim 1 , wherein each lobe of the plurality of lobes has a radial gap adjacent a tip of the respective lobe.5. The system of claim 1 , wherein at least one lobe of the plurality of lobes has a constant radial lobe dimension along an axial lobe dimension of the at least one lobe claim 1 , or at least one passage of the plurality of flow passages has a constant radial passage dimension along an axial passage dimension of the at least one passage claim 1 , or a combination thereof.6. The system of claim 1 , wherein at least one lobe of the plurality of lobes has a variable radial lobe dimension along an axial lobe dimension of the at least one lobe claim 1 , or at least one passage of the plurality of flow passages has a variable radial passage dimension along an axial passage dimension of the at ...

Gas turbine control system

Gas turbine combustion systems and fuel cartridge assemblies are provided. An exemplary combustion system may comprise a combustor including a cylindrical combustion liner, a flow sleeve, a main mixer, a radial inflow swirler, a combustor dome, and a fuel cartridge assembly. An exemplary combustor and/or fuel cartridge assembly may comprise first and second fuel circuits or manifolds. Methods and systems are also provided for staging and controlling a flow of fuel and/or water through different fuel circuits and pilot injectors, to allow purging and ignition using different fuel circuits, pilot injectors, and fuel sources.

PROCESS OF ASSEMBLING FUEL NOZZLE END COVER

A process of assembling a fuel nozzle end cover includes machining a base material of the fuel nozzle end cover, then positioning one or more ring inserts in contact with the base material, then welding the ring insert(s) to the base material to define one or more ledge features within the fuel nozzle end cover. The ring insert(s) have a net shape or near-net shape. 1. A process of assembling a fuel nozzle end cover , the process comprising:machining a base material of the fuel nozzle end cover to define a cylindrical region, the cylindrical region including ports for fluid transport within the fuel nozzle end cover; thenpositioning a ring insert in contact with the base material within the cylindrical region in a position that permits the fluid transport through the ports; thenwelding the ring insert to the base material to define one or more ledge features within the fuel nozzle end cover in the position that permits the fluid transport through the ports;wherein the ring insert has a net shape or near-net shape.2. The process of claim 1 , further comprising masking the ports.3. The process of claim 2 , wherein the masking includes a technique selected from the group consisting of positioning one or more copper chill blocks claim 2 , positioning sheet metal claim 2 , ceramic masking claim 2 , and combinations thereof.4. The process of claim 1 , wherein the welding is selected from the group consisting of gas tungsten arc welding claim 1 , gas metal arc welding claim 1 , cold metal transfer claim 1 , and combinations thereof.5. The process of claim 1 , wherein the welding is selected from the group consisting of beam welding claim 1 , friction welding claim 1 , and combinations thereof.6. The process of claim 1 , further comprising machining a body to form the ring insert prior to the welding of the ring insert.7. The process of claim 1 , wherein the process is devoid of generating weld spatter within the fuel nozzle end cover.8. The process of claim 1 , further ...

COMBUSTOR INLET MIXING SYSTEM WITH SWIRLER VANES HAVING SLOTS

A combustor inlet mixing system () formed from a plurality of circumferentially spaced swirler vanes () extending radially outward from a nozzle hub. Each of the swirler vanes () may have a length () that extends downstream along at least a portion of the combustor inlet mixing system (), and may further have a thickness () that extends along a circumference of the nozzle hub. At least one of the swirler vanes () may further have at least one slot () cut entirely through the thickness () of a portion of the swirler vane (). The slot () may separate the swirler vane () from the nozzle hub along a portion of the length () of the swirler vane (). 117-. (canceled)18. A turbine engine , comprising:at least one combustor positioned upstream from a rotor assembly, wherein the rotor assembly includes at least one row of turbine blades extending radially outward from a rotor;a compressor positioned upstream from the at least one combustor;at least one compressor exhaust plenum extending between the compressor and the at least one combustor; andat least one combustor inlet mixing system formed from a plurality of circumferentially spaced swirler vanes extending radially outward from a nozzle hub, each of the plurality of swirler vanes having a length that extends downstream along at least a portion of the at least one combustor inlet mixing system and further having a thickness that extends along a circumference of the nozzle hub, wherein at least one swirler vane of the plurality of swirler vanes further has at least one slot cut entirely through the thickness of a portion of the at least one swirler vane, the at least one slot separating the at least one swirler vane from the nozzle hub along a portion of the length of the at least one swirler vane.19. The turbine engine of claim 18 , wherein the at least one slot is configured to add a layer of at least partially non-swirling air around the nozzle hub.20. The turbine engine of claim 18 , wherein the at least one slot is ...

SEALING DEVICE BETWEEN AN INJECTION SYSTEM AND A FUEL INJECTION NOZZLE OF AN AIRCRAFT TURBINE ENGINE

An arrangement for an aircraft turbine engine combustion chamber including an injection system and a fuel injector is provided. The injection system includes an injector nozzle guide, the inner surface of which delimits an opening for centering the nozzle, which includes an outer casing. The arrangement further includes a sealing device between the inner surface of the guide and the outer casing. The sealing device includes a first part accommodated in a groove of the outer casing, the groove being delimited, in part, by a downstream delimiting surface, the first part having a first sealing surface and bearing axially against the downstream delimiting surface; and a second part having a second sealing surface bearing radially against the inner surface of the guide. 110-. (canceled)11. An arrangement for a combustion chamber for an aircraft turbine engine , the arrangement comprising:a system for injection of an air-fuel mix into the combustion chamber; anda fuel injector comprising an injector nozzle,the injection system comprising an injector nozzle guide, an inner surface of the injector nozzle guide delimits a centering opening in which there is the injector nozzle that is composed of an outer casing centered on a longitudinal axis of the injector nozzle, a first part accommodated in a groove in the outer casing, said groove extending around said longitudinal axis and being delimited partly by a downstream delimiting surface, the first part having a first sealing surface and bearing axially against said downstream delimiting surface of the groove; and', 'a second part having a second sealing surface bearing radially against said inner surface of the injector nozzle guide., 'wherein the arrangement further comprises a sealing device between the inner surface of the injector nozzle guide and the outer casing of the injector nozzle, the sealing device comprising12. The arrangement according to claim 11 , wherein said first and second parts of the sealing device are ...

Fuel Nozzle of Gas Turbine Combustor and Manufacturing Method Thereof, and Gas Turbine Combustor

To provide a fuel nozzle for a gas turbine combustor, offering favorable durability and strength reliability. In a fuel nozzle for a gas turbine combustor, jetting fuel into a combustion chamber of the gas turbine combustor, the fuel nozzle is metallurgically and integrally bonded with a base plate that supports the fuel nozzle, and an interface between the fuel nozzle and the base plate includes a surface in which bonding is performed by a fusion joint or a brazing joint and an inside part in which bonding is performed by pressure bonding.

BLUFF BODY FUEL MIXER

A fuel injection system may comprise a mixer and a fuel injector disposed within the mixer. The mixer may comprise an outer housing with an exit port and a bluff body extending across the exit port of the outer housing. A flared surface of the mixer may match a contour of the bluff body. 1. A fuel and air mixer , comprising:an outer housing with an exit port; anda bluff body extending across the exit port of the outer housing.2. The fuel and air mixer of claim 1 , comprising a turbulator disposed in the outer housing.3. The fuel and air mixer of claim 1 , wherein the bluff body comprises a v-shaped gutter.4. The fuel and air mixer of claim 3 , wherein the outer housing has a cylindrical geometry.51212. The fuel and air mixer of claim 4 , wherein the bluff body comprises an outlet radius (R) at the exit port of the outer housing and an inlet radius (R) at an apex of the v-shaped gutter claim 4 , and wherein R/R ranges from 1.1 to 1.6.6111. The fuel and air mixer of claim 5 , wherein the v-shaped gutter comprises a width (W) claim 5 , and wherein W/R ranges from 0.2 to 0.45.7. The fuel and air mixer of claim 1 , wherein the bluff body comprises at least one of a circular claim 1 , multi-radial claim 1 , squared claim 1 , or irregularly shaped gutter.8. A fuel injection system claim 1 , comprising:a mixer having a bluff body at an exit port of the mixer; anda fuel injector disposed within the mixer.9. The fuel injection system of claim 8 , further comprising a turbulator disposed in the mixer.10. The fuel injection system of claim 8 , wherein the bluff body comprises a gutter.11. The fuel injection system of claim 8 , wherein the mixer comprises a flared surface matching a contour of the bluff body. This application is a divisional of and claims priority to U.S. application Ser. No. 14/601,389, filed Jan. 21, 2015 and titled “BLUFF BODY FUEL MIXER,” which is hereby incorporated by reference in its entirety.This disclosure was made with government support under contract ...

COMBUSTOR FLOATING COLLAR MOUNTING ARRANGEMENT

A floating collar assembly is configured to receive a fuel nozzle or an igniter projecting through an opening defined in a combustor shell lined with heat shields having studs projecting through the combustor shell for engagement with corresponding fasteners outside the combustor shell. A floating collar is mounted outside the combustor shell with an opening in alignment with the opening in the combustor shell for receiving the fuel nozzle or the igniter. An external retaining bracket is mounted to the heat shield studs or other studs projecting outwardly from the combustor shell so as to trap the floating collar between the combustor shell and the bracket.

CMC Combustor Deflector

Combustor dome assemblies having combustor deflectors are provided. For example, a combustor dome assembly comprises a combustor dome defining an opening; a ceramic matrix composite (CMC) deflector positioned adjacent the combustor dome on an aft side of the assembly; a fuel-air mixer defining a groove about an outer perimeter thereof; and a seal plate including a key. The CMC deflector includes a cup extending forward through the opening in the combustor dome that defines one or more bayonets and a slot. The bayonets are received in the fuel-air mixer groove, and the seal plate key is received in the CMC deflector slot. In another embodiment, where the seal plate may be omitted, a spring is positioned between the fuel-air mixer and the CMC deflector to hold the CMC deflector in place with respect to the combustor dome. Methods of assembling combustor dome assemblies having CMC deflectors also are provided. 1. A combustor dome assembly having a forward side and an aft side , the combustor dome assembly comprising:a combustor dome defining an opening;a ceramic matrix composite (CMC) deflector positioned adjacent the combustor dome on the aft side of the combustor dome assembly, the CMC deflector including a cup extending forward through the opening in the combustor dome, the cup defining one or more bayonets and a slot;a fuel-air mixer defining a groove about an outer perimeter of the fuel-air mixer, the bayonets received in the groove; anda seal plate including a key, the key received in the slot of the CMC deflector.2. The combustor dome assembly of claim 1 , wherein the seal plate is attached to the combustor dome.3. The combustor dome assembly of claim 1 , wherein the fuel-air mixer is attached to the seal plate.4. The combustor dome assembly of claim 1 , wherein the fuel-air mixer defines one or more slots configured for the passage of the one or more bayonets therethrough.5. The combustor dome assembly of claim 4 , wherein the CMC deflector includes a flare ...

COMBUSTION CHAMBER OF A GAS TURBINE, GAS TURBINE AND METHOD FOR OPERATING THE SAME

A combustion chamber assembly of a gas turbine, for combusting a fuel in the presence of combustion air, includes: a combustion chamber, in which combustion of fuel occurs; a precombustion chamber upstream of the combustion chamber; an atomization device that feeds a liquid fuel to the precombustion chamber; and a swirl body that feeds combustion air and gaseous fuel to the precombustion chamber. The combustion chamber assembly is configured as a dual-fuel combustion chamber assembly, which, in a gas fuel operating mode, feeds a mixture of a gaseous fuel and combustion air to the combustion chamber via the swirl body, and which, in a liquid fuel operating mode, feeds liquid fuel to the combustion chamber via the atomization device and combustion air to the combustion chamber via the swirl body. The atomization device includes an atomization lance with a central atomization nozzle, and plural decentralized atomization nozzles. 1. A combustion chamber assembly of a gas turbine , for combusting a fuel in the presence of combustion air , the combustion chamber assembly comprising:{'b': 1', '1', '2, 'a combustion chamber (), in which combustion of fuel occurs, the combustion chamber () being delimited by a wall ();'}{'b': 9', '1, 'a precombustion chamber (), arranged upstream, in a fuel feeding direction, of the combustion chamber ();'}{'b': 4', '9, 'an atomization device () configured to feed a liquid fuel to the precombustion chamber (); and'}{'b': 3', '9, 'a swirl body () configured to feed combustion air and gaseous fuel to the precombustion chamber (),'}wherein:{'b': 1', '3', '1', '4', '1', '3, 'the combustion chamber assembly is configured as a dual-fuel combustion chamber assembly, which, in a gas fuel operating mode, feeds a mixture of a gaseous fuel and combustion air to the combustion chamber () via the swirl body (), and which, in a liquid fuel operating mode, feeds liquid fuel to the combustion chamber () via the atomization device () and combustion air to ...

AIR-SHIELDED FUEL INJECTION ASSEMBLY TO FACILITATE REDUCED NOX EMISSIONS IN A COMBUSTOR SYSTEM

An air-shielded fuel injection assembly for use in a combustion chamber of a turbine assembly. The air-shielded fuel injection assembly generally includes a fuel manifold including a plurality of fuel injection ports and an air manifold including a plurality of air injection ports. Each of the plurality of fuel injection ports is configured to introduce a fuel column into an annular cavity of a mixer assembly. Each of the plurality of air injection ports is configured to introduce an air curtain about an associated fuel injection column to minimize recirculation upstream of the fuel injection column and increase penetration of the fuel injection column into the cavity. Also disclosed are a mixer assembly and a turbine assembly including the air-shielded fuel injection assembly. 1. An air-shielded fuel injection assembly for use in a combustion chamber of a turbine , the assembly comprising:a fuel manifold including a plurality of fuel injection ports, each fuel injection port configured to inject a fuel radially outward into an annular cavity of a mixer assembly to introduce a fuel injection column extending radially outward into the annular cavity; andan air manifold including a plurality of air injection ports, each air injection port associated with a fuel injection port of the plurality of fuel injection ports and wherein each air injection port is separated from the associated fuel injection port by a distance, each air injection port configured to inject a flow of air radially outward into the annular cavity of the mixer assembly and downstream of the plurality of fuel injection ports, to introduce a downstream air curtain extending radially outward into the annular cavity and downstream of an associated fuel injection column, to minimize recirculation of the fuel injection column and an upstream swirl crossflow and increase penetration of the fuel injection column into the annular cavity.2. The air-shielded fuel injection assembly as claimed in claim 1 , ...

GAS TURBINE COMBUSTOR

A combustor for a gas turbine that includes a front panel, an elongated sleeve with first end and second ends and a burner mounted in the sleeve. The second end of the sleeve seallessly mounted on the front panel. The sleeve and burner are configured to enable slidable mounted of the burner in the sleeve.

COOLING OF A MAIN LINE IN A MULTIPOINT FUEL INJECTION SYSTEM

The invention relates to a fuel system () for a turbine engine, adapted for injecting fuel in a combustion chamber () of the turbine engine, comprising: 2. The fuel system according to claim 1 , wherein the insulating material can have a thickness comprised between around 1 mm and around 40 mm claim 1 , preferably between 1 mm and 10 mm claim 1 , for example of the order of 3 mm.3. The fuel system according to claim 1 , wherein the thermal conductor comprises heat-transfer fluid.4. The fuel system according to claim 3 , wherein the heat-transfer fluid is air.5. The fuel system according to claim 1 , wherein the thermal conductor comprises thermally conductive material in the solid state.6. The fuel system according to claim 5 , wherein the thermally conductive material has thermal conductivity greater than that of air.7. The fuel system according to claim 5 , wherein the thermally conductive material comprises rubber or an elastomer claim 5 , such as silicone.8. The fuel system according to claim 1 , wherein the thermal conductor extends over all or part of the length of the main pipe upstream of the combustion chamber.9. A turbine engine comprising a combustion chamber and a fuel system according to .11. The fuel system according to claim 10 , wherein the thermally conductive material comprises silicone. The present invention relates to the field of turbine engine fuel systems, aircraft in particular, and more particularly relates to fuel injection systems in these combustion chambers.The invention relates more precisely to injection systems with dual circuit fuel injection, which comprise a central nozzle, currently called pilot nozzle, delivering a permanent fuel flow rate optimised for low speeds, as well as a peripheral nozzle, sometimes called main nozzle, which delivers an intermittent fuel flow rate optimised for high speeds. These injection systems have been developed for improved adaptation of the injection of air and fuel at different operating speeds of ...

FUEL INJECTOR

A fuel injection system for a gas turbine engine () comprises; a pilot fuel injector section () and a main airblast fuel injector section (), the main airblast fuel injector section having an aft end () facing a combustion chamber (). A surface of the injection system exposed to air flow through an injection system is non-axisymmetric or non-planar in a reference circumferential plane and/or is configured to generate controlled and varying acoustic impedance at or adjacent the aft end where, in use, the air flow collides with an oncoming acoustic wave. 1. A fuel injection system for a gas turbine engine comprising;a main airblast fuel injector section,the main airblast fuel injector section having an aft end facinga combustion chamber and whereina surface exposed to air flow through the injection system is non-axisymmetric, or, non-planar in a reference circumferential plane, andconfigured to generate acoustic impedance at or adjacent the aft end where, in use, the air flow collides with an oncoming acoustic wave.2. A fuel injection system as claimed in further comprising a pilot fuel injector section and wherein the surface is a surface of the pilot fuel injector section.3. A fuel injection system as claimed in wherein the exposed surface is a surface of an air swirler or prefilmer of the main airblast fuel injector section.4. A fuel injection system as claimed in wherein the exposed surface is an annular wall of an air swirler and the wall is extended axially along part of its circumference.5. A fuel injection system as claimed in wherein the exposed surface is an annular wall of an air swirler and comprises an annular array of swirl vanes which includes an axial step.6. A fuel injection system as claimed in wherein the exposed surface is an annular wall of an air swirler and an annular array of vanes on the annular wall includes a first plurality of vanes with a first length claim 1 , pitch and thickness and a second plurality of vanes having a second length ...

CAVITY STAGING IN A COMBUSTOR

A combustor assembly including a combustor liner defining therein a combustion chamber for the downstream flow of a main fluid. At least two annular trapped vortex cavities are located on the combustor liner and staged axially spaced apart. A cavity opening is located at a radially inner end of each of the at least two annular trapped vortex cavities spaced apart from a radially outer wall and extending between an aft wall and a forward wall of each cavity. A plurality of injectors are configured tangentially relative to the circular radially outer wall to provide for an injection of air and fuel to form an annular rotating trapped vortex of a fuel and air mixture within a respective annular trapped vortex cavity. The annular rotating trapped vortex of the fuel and air mixture at the cavity openings is substantially perpendicular to the downstream flow of the main fluid. A gas turbine engine including the combustor assembly is disclosed. 1. A combustor assembly comprising:a combustor disposed axially about a central axis and including a combustor liner having defined therein a combustion chamber for the downstream flow of a main fluid;at least two annular trapped vortex cavities located on the combustor liner and staged axially spaced apart, each of the at least two annular trapped vortex cavities defined between an annular aft wall, an annular forward wall, and a circular radially outer wall formed therebetween;a cavity opening at a radially inner end of each of the at least two annular trapped vortex cavities spaced apart from the circular radially outer wall and extending between the annular aft wall and the annular forward wall;a plurality of fuel injectors and a plurality of air injectors disposed in the circular radially outer wall of one or more of the at least two annular trapped vortex cavities, the plurality of fuel injectors and a plurality of air injectors configured tangentially relative to the circular radially outer wall to provide for an injection of ...