СПОСОБ И УСТРОЙСТВО ДЛЯ МОКРОЙ ОЧИСТКИ СОПЛОВОГО КОЛЬЦА РАБОТАЮЩЕЙ НА ВЫХЛОПНОМ ГАЗЕ ТУРБИНЫ ГАЗОТУРБОНАГНЕТАТЕЛЯ (ВАРИАНТЫ)

Способ и варианты устройства для мокрой очистки соплового кольца работающей на выхлопном газе турбины газотурбонагнетателя. При реализации способа после определения необходимости очистки соплового кольца газотурбонагнетателя запускают автоматически протекающий цикл очистки, при котором воду несколько раз в течение короткого времени впрыскивают в зону перед сопловым кольцом и между впрысками выдерживают паузу для повторного нагрева соплового кольца. Устройства для реализации способа в зоне перед сопловым кольцом газотурбонагнетателя содержат радиальные отверстия с установленными в них форсунками, соединенными через соответствующие линии с подводящей линией для воды. Между измерительным звеном, регистрирующим изменения состояния выхлопных газов двигателя, и находящимся в подводящей линии исполнительным звеном расположен управляющий элемент. Изобретения позволяют улучшить очистку соплового кольца газотурбонагнетателя при использовании меньших количеств воды и при меньшем снижении мощности ...

СПОСОБ КОНВЕРТИРОВАНИЯ ТУРБОВАЛЬНОГО АВИАЦИОННОГО ДВИГАТЕЛЯ В НАЗЕМНУЮ ГАЗОТУРБИННУЮ УСТАНОВКУ

Способ конвертирования турбовального авиационного двигателя в наземную газотурбинную установку. Удаляют лопатки из проточных частей последних ступеней компрессора и первых ступеней турбины. Заменяют сопловой аппарат первой ступени (из оставшихся) конвертированной турбины на сопловой аппарат повышенной пропускной способности. В горелки камеры сгорания подают для сжигания газообразное низкокалорийное топливо типа продукта-газа или биогаза. На установившемся режиме работы конвертированного двигателя изменением расхода топлива устанавливают температуру продуктов сгорания газа в камере не выше 800 K. Уменьшают степень повышения полного давления компрессора до 3-4. Механическую энергию передают потребителю мощности через выводной вал двигателя с редуктором. Изобретение позволяет обеспечить эксплуатацию отработавших ресурс двигателей на низкокалорийных газообразных топливах из твердых бытовых отходов и биоотходов, улучшить экологию, уменьшить расходы на эксплуатацию установок и увеличить их ресурс ...

ОБЛЕГЧЕННАЯ МЕЖДУЛОПАТОЧНАЯ ПЛОЩАДКА ДЛЯ ОПОРНОГОДИСКА ЛОПАТОК ВЕНТИЛЯТОРА ТУРБОРЕАКТИВНОГО ДВИГАТЕЛЯ И ОПОРНЫЙ ДИСК ЛОПАТОК

Междулопаточная площадка для опорного диска лопаток вентилятора турбореактивного двигателя содержит отражательную часть с нижней стороной, оборудованной первой и второй крепежными лапками. Первая крепежная лапка содержит первое отверстие для прохождения первой крепежной шпильки. Вторая крепежная лапка содержит второе отверстие и третье отверстие для прохождения второй и третьей крепежных шпилек. Крепежные шпильки предназначены для жесткого закрепления крепежных лапок на опорном диске между двумя смежными лопатками. Другое изобретение группы касается опорного диска лопаток, содержащего множество указанных выше междулопаточных площадок, установленных между смежными парами лопаток. Изобретение позволяет упростить механическую обработку междулопаточной площадки и снизить ее вес, а также снизить повреждения лопаток вентилятора при попадании посторонних предметов в проточную часть двигателя. 2 н. и 7 з.п. ф-лы, 3 ил.

ГАЗОТУРБИННЫЙ ДВИГАТЕЛЬ ГТД-25СТА, КОМПРЕССОР, КАМЕРА СГОРАНИЯ, ТУРБИНА ГАЗОГЕНЕРАТОРА, СВОБОДНАЯ ТУРБИНА

... 1. Газотурбинный двигатель, состоящий из модуля газогенератора и модуля свободной турбины на собственных рамах, содержащий компрессор, камеру сгорания, турбину газогенератора, свободную турбину, опоры подшипников, вспомогательные системы, отличающийся тем, что газогенератор установлен на собственную раму при помощи опоры, содержащей цапфу для крепления двигателя и корпус опоры, который выполнен в виде крестообразного цилиндрического шарнира.2. Газотурбинный двигатель по п.1, отличающийся тем, что каждая опора газогенератора снабжена механизмом регулирования положения оси газогенератора относительно стыковочных центрирующих элементов рамы в виде винтовой пары.3. Компрессор, содержащий ротор с валом, диски с рабочими лопатками, статор с регулируемым направляющим аппаратом и клапанами перепуска воздуха, отличающийся тем, что для привода исполнительных механизмов направляющего аппарата и клапанов перепуска воздуха использован пневмопривод с отбором сжатого газа из газовой магистрали.4. Блок ...

ГАЗОТУРБИННЫЙ ДВИГАТЕЛЬ

... 1. Газотурбинный двигатель двухвальный, содержащий осецентробежный компрессор и турбину привод его между ними устройство бокового отвода отработавших газов, силовую турбину с выводным валом диски турбомашин изготовлены заодно с лопатками, противоточную камеру сгорания с сопловым аппаратом, отличающийся тем, что на наружный вал двигателя установлены пакетом диски-решетки осетангенциальноцентробежного компрессора с торцевым шлицевым соединением и осетангенциальная турбина между ними улитка отвода отработавших газов, на внутренний выводной вал осевая силовая турбина противопложного вращения с тормозным диском, лопатки турбомашин перфорированы косыми щелями, поворотный сопловой аппарат с приводом установлен перед силовой турбиной. 2. Двигатель по п.1, отличающийся тем, что радиальная лопасть центробежного компрессора на выходе из межлопастного канала загнута назад и на лобке с задней лопастью образует сопло вдува по потоку сжатого воздуха в диффузор улитки по ее выпуклой наружной поверхности ...

МЕЖДУЛОПАТОЧНАЯ ПЛОЩАДКА С КРЕПЕЖНЫМИ ШПИЛЬКАМИ С МЕНЯЮЩИМИСЯ НАПРЯЖЕНИЯМИ ИЗГИБА И СДВИГА ДЛЯ ОПОРНОГО ДИСКА ЛОПАТОК ТУРБОАКТИВНОГО ДВИГАТЕЛЯ

... 1. Междулопаточная площадка (3) для опорного диска (1) лопаток (2) вентилятора турбореактивного двигателя, содержащая отражательную часть (4), с нижней стороной (5), снабженной, по меньшей мере, одной крепежной лапкой (6, 9), содержащей, по меньшей мере, одно отверстие (7, 10) для прохождения крепежной шпильки (8, 11), предназначенной для ее соединения между двумя смежными лопатками (2) со второй крепежной лапкой (12, 15) указанного опорного диска (1), отличающаяся тем, что крепежная шпилька (8, 11) содержит стержень (17) с первой резьбовой частью (18), имеющей первый диаметр (D1), и выполненной с возможностью навинчивания на нее крепежной гайки (22) для выполнения указанного соединения, и со второй частью (19), содержащей первый участок (20), продолжающий указанную первую часть (18) и имеющий второй диаметр (D2), превышающий указанный первый диаметр (D1), и предназначенный для установки между указанными первой (6, 9) и второй (12, 15) крепежными лапками, и второй участок (21), продолжающий ...

Теплоэлектрогенератор на твердом топливе

Изобретение относится к энергетике, а именно к системам отопления и генерации электроэнергии и может быть использовано: в системах воздушного отопления и электроснабжения сельскохозяйственных объектов (фермы, теплицы, мастерские, зернохранилища, овощехранилища, сушилки фруктов, грибов), жилых домов, складских помещений, бытовок в арктических условиях эксплуатации и как автономное энергетическое средство для тепличных хозяйств. Теплоэлектрогенератор на твердом топливе состоит из воздушного компрессора, горелки топлива, электрического генератора и расширительной машины в виде турбины для привода компрессора. Он снабжен эжектором, выполненным с камерой смешения воздуха и продуктов сгорания топлива, при этом горелка выполнена в пеллетном исполнении с бункером твердого топлива и использована для получения рабочей смеси из воздуха и продуктов сгорания пеллет, которая подается в турбину для обеспечения привода компрессора и электрического генератора. В результате работы теплоэлектрогенератора ...

Способ трансформации тепловой энергии в электроэнергию свободнопоршневым энергомодулем с линейным электрогенератором и теплообменником

Изобретение относится к области энергомашиностроения. Технический результат направлен на обеспечение максимальной эффективности трансформации тепловой энергии в электроэнергию при неравномерном подводе тепла к теплообменнику. Тепловая энергия от топки, лучистая энергия солнца и т.д. подводятся к теплообменнику и нагревают воздух во внутренней полости теплообменника. Система управления отслеживает величину температуры и давления воздуха в теплообменнике. В момент времени, когда температура и давление воздуха в теплообменнике достигнут введенного в систему управления предела максимальной величины давления и температуры воздуха, система управления открывает впускной клапан цилиндра. Максимальная величина давления и температуры воздуха в теплообменнике выбирается из соображения прочностных характеристик материала теплообменника. Воздух из теплообменника через впускной клапан цилиндра поступает в рабочую полость поршня. Под действием воздуха поршень начинает движение из исходной точки движения ...

ТУРБОВАЛЬНЫЙ ГАЗОТУРБИННЫЙ ДВИГАТЕЛЬ "БЫЧОК"

Турбовальный газотурбинный двигатель, содержащий цилиндрический корпус, турбину осевого компрессора, кольцевую камеру сгорания, газовую осевую турбину, выходное устройство наработанных газов и ступень свободной газовой турбины, отличающийся тем, что короткие и широкие трапециевидной формы в плане лопатки радиально-осевой турбины компрессора и осерадиальной газовой турбины изготовлены за одно целое с дисками и образуют решетки с торцевым шлицевым соединением, которые пакетами состыкованы в рабочих колесах компрессора и газовой турбины, в межлопаточных каналах с профилем составной состыкованной лопатки, при этом на стыках внахлест носка лопатки на хвост с косыми щелями на стыке на ось двигателя установлено радиально-осевое колесо турбины осевого компрессора, в ее статор вмонтирован спрямляющий осецентростремительный диффузорный аппарат с параболически кривым профилем конических лопаток в конфузорных межлопаточных каналах с плоскими торцевыми обводами, фронтовое устройство камеры сгорания ...

УСТРОЙСТВО ДЛЯ УВЕЛИЧЕНИЯ ЭНЕРГИИ, ПРОИЗВОДИМОЙ ГАЗОВОЙ ТУРБИНОЙ

Устройство для увеличения энергии, производимой системой газовой турбины, увеличивает энергию за счет наличия предварительного компрессора, который сжимает воздух, подаваемый в главный компрессор системы, в результате температура и давление подаваемого воздуха повышаются. Воздух, подаваемый в главный компрессор, охлаждают с применением механической системы с циклом охлаждения для уменьшения температуры подаваемого воздуха на входе в главный компрессор газовой турбины. Воздух, сжимаемый в главном компрессоре, подается в камеру сгорания для нагрева сжатого воздуха и образования горячих газов, которые направляются в турбину, питающую нагрузку и приводящую главный компрессор. Горячие отходящие газы подаются в котел, производящий пар, который подается в паровую турбину, питающую нагрузку. Применение охладителя вместо испарительного охлаждения сжатого воздуха имеет преимущество, поскольку это делает устройство нечувствительным к условиям влажности окружающего воздуха.

ГАЗОТУРБИННАЯ УСТАНОВКА

Газотурбинная установка открытого цикла, содержащая компрессор, камеру сгорания, турбину для привода компрессора и установленные за ней турбину перерасширения, дожимающий компрессор и теплообменник для охлаждения газа перед дожимающим компрессором, отличающаяся тем, что, с целью упрощения конструкции и повышения экономичности, турбина перерасширения и дожимающий компрессор установлены на общем валу, механически не связанном с валом ротора турбины для привода компрессора.

Gasgenerator

Gasturbinenzyklus

Kraftwerkturbinensysteme

Geschaffen ist ein Kraftwerksturbinensystem, wobei das System einen Axialverdichter (104) aufweist, der einen Luftstrom verdichtet, der anschließend mit einem Brennstoff vermischt und in einer Brennkammer (120) verbrannt wird, so dass der sich ergebende Heißgasstrom durch eine Turbine geleitet wird; wobei die Turbine einen Niederdruckturbinenabschnitt (208) und einen Hochdruckturbinenabschnitt (204) aufweist; wobei der Hochdruckturbinenabschnitt über eine erste Welle (216) mit dem Axialverdichter (104) verbunden ist, so dass der Hochdruckturbinenabschnitt (204) im Betrieb wenigstens eine Komponente des Axialverdichters (104) antreibt; wobei der Hochdruckturbinenabschnitt (204) über die erste Welle (216) mit einem schnell laufenden Generator (802) verbunden ist, so dass der Hochdruckturbinenabschnitt (204) im Betrieb den schnell laufenden Generator (802) antreibt; und wobei der Niederdruckturbinenabschnitt (208) über eine zweite Welle (220) mit einem langsam laufenden Generator (212) verbunden ...

GASTURBINEN-TRIEBWERK MIT LEISTUNGSTURBINENSYSTEM

Pumpe oder Turbine zum Foerdern von Stroemungsmitteln

ENERGIEERZEUGER MIT GERINGEN CO2-EMISSIONEN UND ZUGEHÖRIGES VERFAHREN

Gehäuse für einen Verdichter einer Gasturbine, Verdichter und Verfahren zur Herstellung eines Gehäusesegments eines Verdichtergehäuses

Die vorliegende Erfindung betrifft ein Gehäuse für einen Verdichter einer Gasturbine, insbesondere einer Fluggasturbine, mit einem Außengehäuse (12) mit mindestens einer Luftzuführöffnung (14) und einem aus mindestens zwei Gehäusesegmenten (16) ausgebildeten Innengehäuse (18), wobei die Gehäusesegmente (16) mindestens eine Einblasungsdüse (20) zum Einblasen von über die Luftzuführöffnungen (14) angesaugter Luft in einen Strömungskanal (22) im Bereich von Schaufelspitzen (26) von Schaufeln (24) eines Rotors (28) des Verdichters aufweisen. Dabei weist das Gehäusesegment (16) mindestens einen Luftführungskanal (30) auf, wobei der Luftführungskanal (30) derart ausgebildet ist, dass eine direkte Luftführung über mindestens ein in der Luftzuführungsöffnung (14) des Außengehäuses (12) angeordnetes Luftzuführelement (32) zu der mindestens einen Einblasungsdüse (20) erfolgt. Die Erfindung betrifft weiterhin einen Verdichter einer Gasturbine sowie ein Verfahren zur Herstellung eines Gehäusesegments ...

Gas turbine powerplant

Gas turbine engine system with a condenser

In one aspect, the system comprises a compressor system 44, a combustor 6, a turbine system 8, vapour generating means 29 having a dry-run capability, for transferring heat to a fluid to be vaporised when not being run dry, a bypass damper 16 and a condenser 18. In one operational position of the damper 16 (during start up) working fluid can bypass the condenser (via line 28) while in another position of the damper 16 the working fluid passes through the condenser. Vapour (eg. steam) may be fed from the outlet 26 of the vapour generating means to the combustor 6, compression system 44 or turbine system 8. An intercooler 40, water treatment area 20, and storage tank 24 may be provided. In another aspect, heating means 50 may be provided for transferring heat from the intercooler 40 to the working fluid downstream of the condenser to suppress visible exhaust plume. The fluid for transferring heat form the intercooler to the heating means may be water, or air, in which case the heating means ...

Turbine/compressor rotor with helical blade

A rotor for a steam turbine, gas turbine, axial flow compressor or the like comprises helical blades. The blades may have a smaller radial extent at a high pressure end of the rotor than at the low pressure end, and they may be surrounded by a shroud (3, fig. 1b) to which they are connected so that the shroud rotates with the rotor. A rotor for a gas turbine engine (figs. 2a,2b,3a,3b,3c) may comprise both compressor and turbine rotor portions, each comprising helical blades. The compressor blades may be extended beyond the diameter of the engine core (fig. 3c) to serve as turbofan blades (TF). Alternatively turbofan blades may be mounted on a shroud. The blades may be cambered to provide a concave surface facing upstream. The blade leading edges may be drooped, from root to tip, in the direction of blade rotation.

Intake and exhaust arrangement in a gas turbine engine

A gas turbine engine 10 comprises at least one engine rotor including a compressor 22 driven by a turbine 44, a combustion chamber 38, a compressor inlet 12, 30 and a turbine exhaust 20, 48, the inlet 12, 30 and/or exhaust 20, 48 being located between the compressor 22 and the turbine 44, and at least one output shaft 24', 24'' extending from an axial end of the engine 10. The shaft(s) 24', 24'' may drive an engine gearbox, or one or more engine accessory. The engine 10 may be a single spool engine, or may be a multi-spool engine (figure 2) comprising a plurality of single spools cascaded together.

Gas turbine engine with secondary air flow circuit

One embodiment of the present invention is a unique gas turbine engine. Another embodiment of the present invention is a gas turbine engine having a unique secondary air flow circuit. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for gas turbine engines and secondary air flow circuits. Further embodiments, forms, features, aspects, benefits, and advantages of the present application will become apparent from the description and figures provided herewith.

Gas turbine engine compressor undercut spacer

A gas turbine engine compressor spacer (230) includes a body (229) with a hollow cylinder shape. The body (229) includes an outer surface (232) with an axially forward end. The body (229) also includes a forward face (233) extending radially inward from the axially forward end. A forward lip (237) extends axially from the body (229). The forward lip (237) includes a forward lip surface (239) located radially inward from the forward face (233). The forward lip surface (239) is the radially outer circumferential surface of the forward lip (237). The spacer (230) also includes a forward undercut (235). The forward undercut (235) is located between the forward face (233) and the forward lip surface (239).

An assembly

A shaft assembly comprises a first shaft portion 110, connected to a second shaft portion 120 by a flexible coupling 150. The first shaft portion is further connected to the second shaft portion by a frangible coupling 140 with the frangible coupling being configured to fail if a bending load through the frangible coupling exceeds a predetermined maximum load. The flexible coupling allows an angular misalignment between the first and second shaft portions not exceeding a maximum angular misalignment. The first and second shaft portions may be aligned in axial series with the frangible coupling therebetween (Figure 4) or the portions may be concentric with the frangible portion radial between (Figures 5 and 6). The flexible coupling may be a membrane coupling (150 fig. 4) or a constant velocity joint (250 fig.5).The shaft assembly may comprise a bump stop (260 fig. 5) to limit the angular misalignment between the shaft portions. The shaft assembly may be used in a gas turbine engine and ...

Device for protecting a gas turbine engine compressor against the effects of surging

... 1,076,510. Fluid-pressure servomotor systems, fluid regulation. ROLLS-ROYCE Ltd. May 13, 1966, No. 21483/66. Headings G3H and G3P. [Also in Division F1] . gas gas turbine engine is provided with means for shutting off fuel flow when surging conditions arise in the engine compressor. As shown, an engine 10 has a low pressure compressor 12 and high pressure compressor 13, some of the air from the compressor 12 being by-passed via duct 14 and mixed with the exhaust gases from the turbines (not shown). A fluid switch 21 has its power nozzle 20 supplied from compressor 13 via line 15 and filter 16, and the control nozzles 28, 31 are supplied from compressor 12 via line 26 and filter 27, a capacitor 29 and restrictor 30 being provided upstream of nozzle 31. During normal operation, the jet from nozzle 28 is more powerful than that from nozzle 31 and the power jet is deflected into exhaust line 22. If surge-promoting conditions arise, i. e. the output pressure from compressor 12 falls rapidly, ...

GAS TURBINE ENGINES

... 1520401 Gas turbine engines R S DE CHAIR 30 Dec 1975 [30 Dec 1974] 56015/74 Addition to 1435687 Headings F1G and F1L A gas turbine engine comprises a rotor 1 having blades 2, passages between each adjacent pair of blades defining compression, combustion and turbine sections 3, 4, 5 respectively. At least one compressor rotor is disposed upstream of the first rotor 1 without any intervening stator blading, the or each compressor rotor, 7, 6 being driven in contra-rotation relative to the first rotor by a turbine 9, 8 disposed downstream of the first rotor such that in operation air flows into the first rotor at a relative supersonic velocity. The engine is provided with control means comprising a row of variable inlet guide vanes 21 upstream of the compressor rotor, a valve 22 for varying fuel supply to the engine and means for varying the turbine exhaust area of the engine, which as shown comprises a variable area nozzle 23 of the pivoted flap type; alternatively if a power turbine is disposed ...

GAS TURBINE ENGINES

... 1487324 Gas turbine plant; gas generators ROLLS-ROYCE Ltd 11 Nov 1974 [15 Nov 1973] 53135/73 Heading F1G A gas turbine engine utilized as a core engine or gas generator comprises a multi-stage axial-flow compressor having at least one stage of variable pitch stator vanes, and a single-stage supersonic turbine, the turbine and the compressor being mounted on a common shaft. A supersonic turbine is shown in Fig. 2, the gas flow attaining supersonic velocity in the inlet guide vane assembly 142 as indicated by the shock wave pattern 146. The relative velocity of the gas as it enters the rotor blade assembly 143 is subsonic but the velocity becomes supersonic at the discharge throats of turbine rotor blades as indicated by the shock wave pattern 148.

Improvements in or relating to gas turbine-driven plant

... 805,633. Gas-turbine plant. POWER JETS (RESEARCH & DEVELOPMENT) Ltd. March 15, 1956 [March 17, 1955; Aug. 26, 1955], Nos. 7786/55 and 24551/55. Class 110 (3). [Also in Group XXIX] A gas-turbine plant operable to produce a variable mass flow at a pressure above the maximum gasturbine cycle pressure has a mechanism responsive to a reduction of output demand substantially to zero to over-ride the normally operative fuel supply speed governor control and to restrict the fuel input to a predetermined idling rate. The plant shown comprises a low-pressure compressor 10 the output of which is divided into two streams. One stream passes through the duct 11 to the combustion chamber 12 and the other passes through a duct 13 to a high-pressure compressor 14 which discharges to an external user through a duct 18. The compressors 10, 14 are driven by turbines 15, 16 which are supplied in series with hot gases from the combustion chamber 12. A by-pass 19 controlled by a valve 20 connects the ducts 18 ...

Rotary device with moving elements to compress, slacken or involve a fluid.

Improvements in the utilisation of methane

GAS TURBINE ENGINE WITH A SINGLE OIL CAVITY

A gas turbine engine (10) having an oil cavity architecture and bearing placement which reduce heat rejection and oil system complexity by enclosing the reduction gearbox bearings (54) and at least the shaft bearings (42, 44) supporting the high pressure shaft (38) in the same oil cavity (60).

TURBO RECUPERATOR DEVICE

A device and method for recuperating a gas turbine engine comprises a compressor [10] being configured to receive a coolant fluid stream, to compress the coolant fluid stream and to discharge the compressed coolant fluid stream to a turbine [14] in fluid communication with the compressor [10]. The compressed coolant fluid stream undergoing thermal exchange within the turbine [14], exit the turbine [14] thereafter. A source of a working fluid stream [44] is in fluid communication with the turbine [14]. The working fluid stream [44] is fluidly isolated from a portion of the coolant fluid stream and undergoing thermodynamic expansion through the turbine [14] to extract energy therefrom. Where desired, the entire coolant fluid stream is fluidly isolated from the working fluid stream [44]. At least a portion of the coolant fluid stream is channeled downstream of the turbine [14] to supply a preheated process fluid stream to another process.

TURBINE ENGINE WITH SEMI-FIXED TURBINE DRIVING A RECEIVER CONTROLLED SO AS TO PRESERVE A ROUGHLY CONSTANT ROTATION SPEED

... ²The invention relates to a turbine engine ²(1) with a semi-fixed turbine, particularly for aircraft ²driving a receiver (2) controlled so as to preserve a ²roughly constant rotation speed. The turbine engine, in ²particular via a gear system (20), drives the receiver ²and an LP compressor (6) with an LP turbine (14). ²According to the invention, the gear system has a torque ²control system (26) maintaining a constant ratio between ²the drive torque of the receiver transmitted by the gear ²system and the drive torque of the LP compressor ²transmitted by this same gear system.² ...

Brennkraftmaschine.

Verbrennungsturbinenanlage mit Gleichdruckverbrennung und Luftkühlung.

Hochleistungsschalter.

Installation à turbine à combustion interne.

Gasturbinenanlage

Schreibmaschinentisch.

Gleichdruck-Gasturbinenanlage.

Zweiteilige Abstandsschelle für Feuchtraumleitungen.

Druckfüllbleistift.

Gasturbinenanlage

Gasturbinen-Triebwerk

Turbomaschine

Gasturbinenanlage

Machine volumétrique

Zweiwellige Gasturbinenanlage

Gas turbine plant with two shafts and method for its operation

Method and system for frequency separation into a gas turbine.

Es ist hierin ein Verfahren und System zur Frequenztrennung in einer Gasturbine (100) geschaffen. Die Systeme und Verfahren zur Frequenztrennung in einer Gasturbine (100) enthalten ein Bestimmen einer Eigenfrequenz einer Heissgaspfadkomponente, Bestimmen einer Verbrennungsdynamikamplitude und/oder -frequenz und Modifizieren einer Verdichterauslasstemperatur zur Trennung der Verbrennungsdynamikfrequenz von der Eigenfrequenz der Heissgaspfadkomponente.

Gas Turbine.

Eine Gasturbine, die enthält: einen durch einen Brenner und eine Turbine hindurch definierten inneren Strömungspfad; einen eine Verbindungsstelle zwischen dem Brenner (12) und der Turbine (13) bildenden hinteren Rahmen (29), wobei der hintere Rahmen (29) ein starres Strukturbauteil aufweist, das den inneren Strömungspfad umgibt, wobei der hintere Rahmen eine Innenwand enthält, die eine Aussenbegrenzung des inneren Strömungspfades definiert; einen sich längs des Umfangs erstreckenden Brennstoffsammelraum, der durch den hinteren Rahmen (29) hindurch ausgebildet ist; und durch die Innenwand des hinteren Rahmens hindurch ausgebildete Auslassöffnungen. Die Auslassöffnungen verbinden den Brennstoffsammelraum mit dem inneren Strömungspfad.

Gas turbine engine with two combustion chambers and a bypass channel.

Gasturbinenmotor bestehend aus einer Welle (20), auf welcher die Kompressoren (4) und (5), die Turbinen (6) und (7) in Reihe angeordnet sind. Die Hochdruckbrennkammer (11) ist zwischen dem Kompressor (5) und der Turbine (6) angeordnet. Der Umleitungskanal (16) für einen Teil des Arbeitsmediums verbindet den Ausgang des Kompressors (4) mit dem Eingang der Turbine (7) über die Niederdruckbrennkammer (12). Die Umverteilung des Arbeitsmediums durch den Kanal (16) erfolgt ungehindert.

System for generation of current under use of solar energy.

Es wird ein Elektrizitätserzeugungssystem (10) präsentiert. Das Elektrizitätserzeugungssystem (10) beinhaltet einen Solarvorwärmer (18) zum Vorwärmen verdichteter Auslassluft; einen Brenner (36), um die erwärmte verdichtete Luft aus dem Solarvorwärmer (18) aufzunehmen und einen Brennstoff unter Verwendung der erwärmten verdichteten Luft zu verbrennen, um heisses verbranntes Gas zu erzeugen; eine erste Turbine (28), um das heisse verbrannte Gas aus dem Brenner (36) aufzunehmen, und das heisse verbrannte Gas zu expandieren, um Abgas zu erzeugen; einen Wärmerückgewinnungs-Dampfgenerator (46), um das Abgas aus der ersten Turbine (28) aufzunehmen und Dampf zu erzeugen, indem ein kondensiertes Fluid unter Anwendung des Abgases erwärmt wird; einen Solar-Verdampfer/Überhitzer (22), um ein erwärmtes Arbeitsfluid aus dem Wärmerückgewinnungs-Dampfgenerator (46) aufzunehmen und Solar-Dampf durch Erwärmen des erwärmten Arbeitsfluids unter Nutzung eines zweiten Anteils des erwärmten Solarfluids zu erzeugen ...

Fluid-flow machine.

Eine Strömungsmaschine (2) enthält einen Verdichter (4), eine betrieblich mit dem Verdichter (4) gekoppelte Turbine (10) und eine Verbrennungsanordnung, die den Verdichter (4) und die Turbine (10) in Fluidverbindung miteinander bringt. Die Verbrennungsanordnung (5) enthält wenigstens einen Injektor mit einem Brennerrohr, das einen äusseren Wandabschnitt und einen inneren Wandabschnitt aufweist, die eine Mischungszone bilden. Ein Drallerzeuger (40, 41) ist innerhalb der Mischungszone angeordnet. Der Drallerzeuger (40, 41) weist eine Anzahl von Leitschaufeln auf, wobei wenigstens einer der mehreren Leitschaufeln einen Wandabschnitt mit einer äusseren Oberfläche und einer inneren Oberfläche aufweist, der einen hohlen inneren Bereich bildet. Ein Einsatzelement ist in dem hohlen inneren Bereich angeordnet. Das Einsatzelement weist wenigstens ein Führungselement auf, das zum Zuführen einer Fluidströmung in den hohlen inneren Bereich zum Strömen über den Wandabschnitt der wenigstens einen der ...

ГАЗОТУРБИННАЯ УСТАНОВКА

Изобретение относится к области газотурбинных установок, предназначенных для использования на газотурбовозах, передвижных и стационарных электрических станциях, и отличается использованием криогенного газового топлива. Масляные системы газотурбинного двигателя и исполнительных агрегатов выполнены по отдельным регулируемым циркуляционным контурам со своими топливомасляными теплообменниками, охлаждающей средой которых является криогенное газовое топливо, нагнетающим насосом и баком для масла. Охлаждающие полости топливомасляных теплообменников соединены топливными трубопроводами на входе с устройством подачи и регулирования топлива, а на выходе они соединены с подогревателем топлива, установленным в выхлопном патрубке газотурбинного двигателя. Техническим результатом изобретения является создание экономичной газотурбинной установки с компактными системой охлаждения масла и подогревателем топлива. Предлагаемое изобретение позволяет производить возврат в термодинамический цикл газотурбинного ...

SYSTEMS AND METHODS TRAPPING CARBON DIOXIDE AND GENERATION OF ENERGY IN TYRBINY SYSTEMS WITH LOW EMISSION OF

ГАЗОТУРБИННАЯ УСТАНОВКА

Изобретение относится к области газотурбинных установок, предназначенных для использования на газотурбовозах, передвижных и стационарных электрических станциях, и отличается использованием криогенного газового топлива. Масляные системы газотурбинного двигателя и исполнительных агрегатов выполнены по отдельным регулируемым циркуляционным контурам со своими топливомасляными теплообменниками, охлаждающей средой которых является криогенное газовое топливо, нагнетающим насосом и баком для масла. Охлаждающие полости топливомасляных теплообменников соединены топливными трубопроводами на входе с устройством подачи и регулирования топлива, а на выходе они соединены с подогревателем топлива, установленным в выхлопном патрубке газотурбинного двигателя. Техническим результатом изобретения является создание экономичной газотурбинной установки с компактными системой охлаждения масла и подогревателем топлива. Предлагаемое изобретение позволяет производить возврат в термодинамический цикл газотурбинного ...

СПОСОБ ИСПОЛЬЗОВАНИЯ ПОЛУЧАЕМОГО В ГАЗОГЕНЕРАТОРЕ СИНТЕЗ-ГАЗА

Синтез-газ (Н2+СО), который образовывается в газогенераторе, должен быть экономическим и оптимальным способом использован, прежде всего для выработки электроэнергии, при этом полученный одновременно как побочный продукт СО2 должен отводиться на хранение. Это достигнуто за счет того, что синтез-газ (Н2+СО) и кислород (О2) из установки для распределения воздуха сжигают в горелке и с помощью газовой турбины (с приведением в действие генератора) подвергают снижению давления, СО2 в потоке выходящих газов отделяют и направляют к компрессору, который приводится в действие газовой турбиной, и в виде сжатого СО2 направляют в хранилище СО2.

ГАЗОТУРБІННИЙ ДВИГУН

Газотурбінний двигун із двоопорним ротором містить компресор 1, камеру згоряння 2, турбіну 3. Ротор складається з консольної барабанної секції 4, зварних барабанно-дискових секцій 5 і 6 компресора 1, барабанно-дискової секції 7 турбіни 3. Ротор виконаний з конічними, що звужуються до опор, цапфами з вмонтованими дисками. Зварна барабанно-дискова секція 6 компресора 1 з’єднана безпосередньо з барабанно-дисковою секцією 7 турбіни 3.

Турбина (варианты)

Изобретение относится к гидравлическим машинам и двигателям, в частности к турбинам. Турбина состоит из вала, пары боковых пластин, смонтированных на валу на расстоянии одна от другой, и лопастей, выполненных в виде прямоугольного равнобедренного треугольника. Вершина прямого угла лопастей ориентирована к наружной части турбины. Вершины острых углов лопастей прикреплены к боковым пластинам. Лопасти изогнуты в форме U, V или W. На валу турбины смонтированы дополнительно по меньшей мере еще одна или более боковых пластин, или пара боковых пластин к которым присоединены другие лопасти той же формы U, V или W, высота и радиус которых равны или отличаются от высоты и радиуса других лопастей той же формы U, V или W. Некоторые боковые пластины могут быть установлены на расстоянии одна от другой внутри или снаружи других боковых пластин, получая таким образом турбину составленную из одной исходной и одной или более дополнительных. Один из вариантов предусматривает прикрепление вершины прямого угла ...

GAS-TURBINE ENGINE

INTERBLADE PLATFORM FOR A DISC SUPPORTING BLADES ON A TURBOJET BLOWER AND BLADE SUPPORT DISC

SYSTEMS AND METHODS TRAPPING CARBON DIOXIDE AND GENERATION OF ENERGY IN TYRBINY SYSTEMS WITH LOW EMISSION OF

СИСТЕМА И СПОСОБ ГЕНЕРАЦИИ ЭНЕРГИИ

Предлагаются системы, способы и устройства для генерации энергии в турбинных системах с низкими выбросами и разделения выхлопа на обогащенный CO2 поток и обедненный CO2 поток. В одном или нескольких вариантах осуществления выхлоп разделяют при повышенном давлении, например, между ступенью расширения высокого давления и ступенью расширения низкого давления.

СИСТЕМА И СПОСОБ ГЕНЕРАЦИИ ЭНЕРГИИ

Предлагаются системы, способы и устройства для генерации энергии в турбинных системах с низкими выбросами и разделения выхлопа на обогащенный СО2 поток и обедненный СО2 поток. В одном или нескольких вариантах осуществления выхлоп разделяют при повышенном давлении, например, между ступенью расширения высокого давления и ступенью расширения низкого давления.

ИНТЕГРИРОВАННАЯ ГАЗОТУРБИННАЯ СИСТЕМА И СПОСОБ ВЫРАБОТКИ ЭНЕРГИИ

Предложены системы, способы и устройства для управления подачей окислителя в турбинных системах с низким уровнем выбросов для поддержания стехиометрических или по существу стехиометрических условий горения. В одном или более вариантах осуществления подобное управление осуществляется посредством способов или систем, которые обеспечивают подачу окислителя с постоянным массовым расходом в камеру сгорания.

TWIN SPOOL INDUSTRIAL GAS TURBINE ENGINE WITH VARIABLE INLET GUIDE VANES

Micro gas turbine distributed power generation system based on turbocharger

Inner culvert turbofan and ram-type double-mode engine

The invention discloses a turbofan and a ram-type double-mode engine. In a propulsion working mode of the turbofan, an inner culvert turbofan engine is adopted, and then a fan blade and a main shaft form a certain angle to make the rotating fan blade propel the air backwards. When the external air velocity reaches high subsonic velocity or supersonic velocity, the inner culvert turbofan stops active rotation; the angle between the fan blade and the main shaft is adjusted to be small so as to reduce the windward sides of the inner culvert turbofan and the fan blade and adjust the airflow; the airflow enters a rear spray pipe through the inner culvert turbofan and the fan blade, and is mixed, burned and expanded with the fuel oil in the rear spray pipe to generate a backward thrust; and then the rear spray pipe is equivalent to the ram-type engine. By adjusting the angle between the fan blade and the main shaft, the engine has two different working modes and reaches an optimal working state ...

GAS TURBINE POWER PLANT

LINEAR MOTOR HAS COMPRESSION DEMULTIPLIEE

ASSEMBLY OF BLADE OF TURBINE WHEEL AND PROCESS OF CONFIGURATION

Un assemblage de rotor pour moteur à turbine, l'assemblage de rotor incluant : une pale de turbine qui inclut un pied située entre des moyens de fixation et un profil aérodynamique, le pied ayant une partie avant et une partie arrière ; et une plate-forme comprenant un côté de pression de plate-forme et un côté d'aspiration de plate-forme, chacun comprenant des composants qui ne sont pas d'un seul bloc avec la pale de turbine. La plate-forme peut comprendre une interface entre le côté de pression de plate-forme et le côté d'aspiration de plateforme . Et, la plate-forme peut être configurée de telle manière que l'interface soit alignée avec au moins l'une de la partie avant et de la partie arrière du pied.

Rotary engine explosion and applications of detonating explosives and gas turbines in general

RECUPERATOR TYPE GAS TURBINE CYCLE DEVICE CAPABLE OF IMPROVING A HEAT TRANSMISSION BY INCREASING A HEAT EXCHANGE AREA

PURPOSE: A recuperator type gas turbine cycle device is provided to improve the heat exchange and heat transmission efficiency by increasing a heat exchange area due to the formation of a louver unit or a dimple unit on both sides of an uneven unit. CONSTITUTION: A recuperator type gas turbine cycle device comprises a compressor(10), a combustor(20), a turbine(30), a generator(40), and a recuperator type heat exchanger(50). The compressor compresses and transmits air which is drawn from the outside. The combustor combusts and heats the air and gas by mixing the air transmitted from the compressor with the gas drawn from the outside. The turbine is connected to the combustor and is operated by high temperature and high pressure combustion gas. The generator is connected to the turbine and generates electricity according to the operation of the turbine. The recuperator type heat exchanger is connected to the turbine and transfers the high temperature combustion gas used in the turbine. The ...

VACUUM COMPRESSOR AND A TURBINE GENERATOR STRUCTURE IN WHICH STRONG VACUUM AND HIGH PRESSURE ARE GENERATED WHILE AN IMPELLER IS ROTATED AT A LOW SPEED

PURPOSE: A vacuum compressor and a turbine generator structure are provided to increase the cross section of an impeller and to reduce the size of the impeller as the impeller is formed into multiple stages. CONSTITUTION: A vacuum compressor and a turbine generator structure comprise a motor, a generator, a vacuum compressor, a turbine, an impeller, a guide vane, a body, and a lid. The vacuum compressor is mounted to both sides of the motor. The vacuum compressor vacuumizes or compresses the pressure of fluids with the rotation of the impeller. The turbine absorbs the pressure of the fluids, and generates electricity by operating the generator. The impeller is connected to a rotary shaft or the generator or the motor in the turbine. The guide vane is fixed to the body, which makes the fluid flow smoothly. The body fixes the guide vane and the motor or the generator. The guide vane and the impeller are mounted to the body. The lid is connected to the body with bolts to prevent the fluids ...

MOTOR MOUNT TURBOFAN, AIRCRAFT AND METHOD FOR MOUNTING A MOTOR MOUNT

SYSTEM AND METHOD FOR GENERATING ELECTRIC ENERGY

An object of the present invention is to provide a method and a system for implementing the method so as to alleviate the disadvantages of a reciprocating combustion engine and gas turbine in electric energy production. The invention is based on the idea of arranging a combustion chamber (10) outside a gas turbine (22) and providing compressed air to the combustion chamber (10) in order to carry out a combustion process in controlled and optimal conditions.

CUSTOMER BLEED AIR PRESSURE LOSS REDUCTION

A bleed air supply system for a gas turbine engine comprising a duct having an inlet end and extending to an outlet end. The inlet end of the duct is provided with a central insert. In another feature, there may be a plurality of ducts, and inlet ends of the plurality of ducts being spaced by at least 90. In another feature, a compressor may have a diffuser with a shroud ending upstream of the downstream end of an inner shroud, having an outer shroud ending at a location upstream of a downstream end of an inner shroud at locations circumferentially aligned with an inlet end of the duct.

GAS TURBINE ENGINE MID TURBINE FRAME WITH FLOW TURNING FEATURES

A gas turbine engine includes first and second stages having a rotational axis. A mid turbine frame is arranged axially between the first and second stages. The mid turbine frame includes a circumferential array of airfoils, and the airfoils each have a curvature provided equidistantly between pressure and suction sides. The airfoils extend from a leading edge to a trailing edge at a midspan plane along the airfoil. An angle is defined between first and second lines respectively tangent to the intersection of the plane and the curvature at airfoil leading and trailing edges. The angle is equal to or greater than about 10°, for example. In one example, an airfoil aspect ratio is less than 1.5.

MULTI-COVER GAS TURBINE ENGINE COMPONENT

An airfoil for a gas turbine engine includes an airfoil body extending between leading and trailing edges in a chordwise direction and extending from a root section in a spanwise direction, and the airfoil body defining pressure and suction sides separated in a thickness direction. The airfoil body defines a recessed region extending inwardly from at least one of the pressure and suction sides, and the airfoil body includes one or more ribs that define a plurality of pockets within a perimeter of the recessed region. A plurality of cover skins is welded to the airfoil body along the one or more ribs to enclose respective ones of the plurality of pockets.

HEAT PIPE ENGINE

A heat pipe engine, comprising a housing (1), a shaft sleeve (4) disposed in the middle of the housing and fixedly connected with the housing, a spindle (5) disposed in the shaft sleeve through a bearing, and a combustion chamber (6) disposed outside the shaft sleeve and fixedly connected with the housing. A vortex generator (8) is connected at the end of the spindle, the vortex generator being in a truncated cone shape, and a plurality of conical helix shaped working medium passages (9) being provided on the vortex generator. According to the heat pipe engine, as the cone of the working medium passages is helix-shaped, a working medium moves along the helix; and as the working medium passages are 0.2-4 mm in diameter, the working medium can be divided into a plurality of small working medium flows. Thus a fluid boundary layer effect is made full use of to convert the energy of the working medium into the rotating kinetic energy for the spindle to be output. The heat pipe engine therefore ...

FLUID FLOW DEVICES WITH VERTICALLY SIMPLE GEOMETRY AND METHODS OF MAKING THE SAME

A micro-turbine engine, consisting of at least a compressor, combustor, and turbine, is a complicated fluid flow device that controls the flow rate and thermodynamic properties of a working fluid in order to generate shaft power. Existing micro-turbines are costly to manufacture because they are designed with sophisticated contours and exotic materials. The present invention discloses a method for designing a micro-turbine with stacked layers of structure, each of which is designed with vertically simple geometry such that it can be manufactured using conventional machining technology. The resulting micro-turbine is low cost compared to existing alternatives in the target range of power outputs and applications. The present invention also describes a method for connecting the micro-turbine to an electrical generator to generate power. Lastly, the method for designing the micro-turbine is applied to heat exchangers, Rankine engines, fluid mixers, and other fluid flow devices.

GAS TURBINE POWER PLANT WITH SUPERSONIC GAS COMPRESSOR

A gas turbine engine. The engine is based on the use of a gas turbine driven rotor having a compression ramp traveling at a local supersonic inlet velocity (based on the combination of inlet gas velocity and tangential speed of the ramp) which compresses inlet gas against a stationary sidewall. The supersonic compressor efficiently achieves high compression ratios while utilizing a compact, stabilized gasdynamic flow path. Operated at supersonic speeds, the inlet stabilizes an oblique/normal shock system in the gasdyanamic flow path formed between the rim of the rotor, the strakes, and a stationary external housing. Part load efficiency is enhanced by use of a lean pre-mix system, a pre-swirl compressor, and a bypass stream to bleed a portion of the gas after passing through the pre-swirl compressor to the combustion gas outlet. Use of a stationary low NOx combustor provides excellent emissions results.

THERMAL ENERGY STORAGE SYSTEMS AND METHODS

Thermal energy storage systems and methods are provided including a bed, a blend of aggregates packed in the bed, and a high-density heat transfer fluid flowing through the blend of aggregates. The blend of aggregates includes rock materials and may also include non-rock materials. The heat transfer fluid flows through the blend of aggregates such that heat is transferred between the heat transfer fluid and the blend of aggregates. The porosity of the aggregates increases heat transfer and the high density of the heat transfer fluid reduces the pressure gradient of the heat transfer fluid. In exemplary embodiments, the heat transfer fluid is a liquid comprised of carbon-based molecules. Methods of safely storing and releasing energy are provided in which axial thermal conductivity of the bed is minimized and inadvertent pressure release failures are mitigated.

TURBINE ENGINE INCLUDING BALANCED LOW PRESSURE STAGE COUNT

A turbine engine includes at least a compressor section and a turbine section, each having at least a first and second portion. A ratio of turbine section second portion stages to compressor section second portion stages is less than or equal to 1.

FLOWPATH COMPONENT FOR A GAS TURBINE ENGINE INCLUDING A CHORDAL SEAL

A flow path component includes a platform having at least one radially aligned face. A chordal seal extends axially from the radially aligned face. The chordal seal includes a first curved face configured to prevent edge line contact under deflection conditions while the flow path component is installed in an engine.

MULTI-ENGINE COORDINATION DURING GAS TURBINE ENGINE MOTORING

A system is provided for multi-engine coordination of gas turbine engine motoring in an aircraft. The system includes a controller operable to determine a motoring mode as a selection between a single engine dry motoring mode and a multi-engine dry motoring mode based on at least one temperature of a plurality of gas turbine engines and initiate dry motoring based on the motoring mode.

Control device, system, control method, power control device, gas turbine, and power control method

Provided is a control device of a gas turbine including a compressor, a combustor, and a turbine. The control device executes load control of allowing an operation control point for operation control of a gas turbine to vary in response to a load of the gas turbine. The operation of the gas turbine is controlled on the basis of a rated temperature adjustment line for temperature adjustment control of a flue gas temperature at a predetermined load to a rated flue gas temperature at which performance of the gas turbine becomes rated performance, a preceding setting line for setting of the flue gas temperature at the predetermined load to a preceding flue gas temperature that becomes lower in precedence to the rated flue gas temperature, and a limit temperature adjustment line for temperature adjustment control.

Enhanced protection for aluminum fan blade via sacrificial layer

A component is described which may comprise a structure formed from a material selected from the group consisting of aluminum and an aluminum alloy. The component may further comprise a sacrificial layer in electrical contact with at least a portion of a surface of the structure. The sacrificial layer may protect the surface from localized corrosion and may comprise an alloy that is more anodic than the material forming the structure. The alloy may be selected from the group consisting of an aluminum alloy and a zinc alloy.

Gas turbine engine airfoil with variable pitch cooling holes

An airfoil includes a ceramic airfoil section that defines leading and trailing edges, pressure and suction sides, and radially inner and outer ends. The span of the airfoil section has first, second, and third radial span zones. There is a row of cooling holes in an aft 25% of the axial span. The row of cooling holes extends though the first, second, and third radial span zones. The cooling holes in the first radial span zone define a first pitch P1, the cooling holes in the second radial span zone define a second pitch P2, and the cooling holes in the third radial span zone define a third pitch P3, wherein P2 Подробнее

AERO BOOST - GAS TURBINE ENERGY SUPPLEMENTING SYSTEMS AND EFFICIENT INLET COOLING AND HEATING, AND METHODS OF MAKING AND USING THE SAME

The invention relates generally to electrical power systems, including generating capacity of a gas turbine, and more specifically to pressurized air injection that is useful for providing additional electrical power during periods of peak electrical power demand from a gas turbine system power plant, as well as to inlet heating to allow increased engine turn down during periods of reduced electrical demand. 1. A power augmentation system for a gas turbine engine comprising:one or more compressors coupled to one or more turbines by a shaft, the one or more compressors having an inlet region; a fueled engine;', 'one or more auxiliary compressors; and,', 'a recuperator;, 'an auxiliary source of compressed air comprisingwherein the auxiliary source of compressed air is in fluid communication with the inlet of the one or more compressors.2. The power augmentation system of claim 1 , wherein the one or more compressors comprises a low pressure compressor and a high pressure compressor.3. The power augmentation system of claim 2 , wherein the one or more turbines comprises a low pressure turbine and a high pressure turbine.4. The power augmentation system of claim 1 , wherein the inlet region to the one or more compressors is located at an engine inlet.5. The power augmentation system of claim 1 , wherein the auxiliary source of compressed air discharges a flow of hot compressed air.6. The power augmentation system of claim 5 , wherein the flow of hot compressed air is formed by a heat exchange process between air from the one or more auxiliary compressors and exhaust from the fueled engine.7. The power augmentation system of claim 1 , wherein the auxiliary source of compressed air provides inlet heating to the gas turbine engine.8. The power augmentation system of claim 1 , wherein the auxiliary source of compressed air is generated separate from the gas turbine engine.9. A method of augmenting power in a gas turbine engine comprising:placing an auxiliary source of ...

Endwall countouring trench

A gas turbine engine component comprises a casing, and a component fixed to the casing to extend from a first edge to a second edge. The component has a first side and a second side. A trench is formed within the casing adjacent the second side of the component. The trench has a maximum depth that is positioned at or aft of the first edge.

GAS TURBINE

A gas turbine is provided and includes a compressor, which via an air intake inducts and compresses air; a combustion chamber, in which a fuel is combusted using the compressed air, producing a hot gas; and a turbine, equipped with turbine blades, in which the hot gas is expanded, performing work. A first device is provided in order to cool turbine blades with compressed cooling air. The first device includes at least one separate compressor stage which produces compressed cooling air independently of the compressor. 1203012111314153232222512. A gas turbine ( , ) , comprising a compressor () , which via an air intake () inducts and compresses air; a combustion chamber () in which a fuel () is combusted , using the compressed air , producing a hot gas; and a turbine () , equipped with turbine blades () , in which the hot gas is expanded , performing work , wherein a first device is provided in order to cool turbine blades () with compressed cooling air , the first device comprises at least one separate compressor stage ( , ) which produces compressed cooling air independently of the compressor ().2222515. The gas turbine as claimed in claim 1 , wherein the at least one compressor stage ( claim 1 , ) is arranged aft of the turbine () in the flow direction.315162422251624. The gas turbine as claimed in claim 1 , wherein the turbine () has a rotor or a shaft ( claim 1 , ) claim 1 , and the at least one compressor stage ( claim 1 , ) is integrated into the rotor or into the shaft ( claim 1 , ).4222525. The gas turbine as claimed in claim 1 , wherein the at least one compressor stage ( claim 1 , ) is configured as a radial compressor ().52225322225282722253132. The gas turbine as claimed in claim 3 , wherein the at least one compressor stage ( claim 3 , ) is arranged directly aft of the last turbine rotor blades () in the flow direction claim 3 , the at least one compressor stage ( claim 3 , ) inducts a cooling air mass flow () from a rotor-bearing plenum () claim 3 , and ...

Rotor, a steam turbine and a method for producing a rotor

A rotor, a steam turbine having a rotor, and a method of producing a rotor are disclosed. The rotor disclosed includes a shaft high pressure section. The high pressure section includes a first high pressure section, a second high pressure section, the second high pressure section being joined to the first pressure section, and a third high pressure section, the third high pressure section being joined to the second high pressure section. At least a portion of the second high pressure section is formed of a high-chromium alloy steel comprising 0.1-1.2 wt % of Mn, up to 1.5 wt % of Ni, 8.0-15.0 wt % of Cr, up to 4.0 wt % of Co, 0.5-3.0 wt % of Mo, 0.05-1.0 wt % of V, 0.02-0.5 wt % of Cb, 0.005-0.15 wt % of N, up to 0.04 wt % of B, up to 3.0 wt % of W, and balance Fe and incidental impurities.

SYSTEM AND METHOD FOR GENERATING POWER USING A SUPERCRITICAL FLUID

A dual cycle system for generating shaft power using a supercritical fluid and a fossil fuel. The first cycle is an open, air breathing Brayton cycle. The second cycle is a closed, supercritical fluid Brayton cycle. After compression of air in the first cycle, the compressed air flows through a first cross cycle heat exchanger through which the supercritical fluid from the second cycle flows after it has been compressed and then expanded in a turbine. In the first cross cycle heat exchanger, the compressed air is heated and the expanded supercritical fluid is cooled. Prior to expansion in a turbine, the compressed supercritical fluid flows through a second cross cycle heat exchanger through which also flows combustion gas, produced by burning a fossil fuel in the compressed air in the first cycle. In the second cross cycle heat exchanger, the combustion gas is cooled and the compressed supercritical fluid is heated. 1. A method of generating shaft power in a system comprising an air cycle and supercritical fluid cycle , comprising the steps of:a) burning a fossil fuel in air so as to produce a combustion gas;b) expanding said combustion gas in at least a first turbine so as to produce an expanded combustion gas, said expansion of said combustion gas generating shaft power;c) compressing a supercritical fluid in a first compressor;d) flowing at least a portion of said compressed supercritical fluid through a first cross cycle heat exchanger, and flowing said combustion gas through said first cross cycle heat exchanger so as to transfer heat from said combustion gas to said compressed supercritical fluid so as to produce a heated compressed supercritical fluid;e) expanding at least a portion of said heated compressed supercritical fluid in a second turbine so as to produce an expanded supercritical fluid, said expansion of said supercritical fluid generating additional shaft power; andf) flowing at least a portion of said expanded supercritical fluid through a second ...

BLADE GROUP ARRANGEMENT AS WELL AS TURBOMACHINE

A blade group arrangement for a turbomachine in order to form a blade-row group, whereby a front blade and a rear blade each form an overlapping area that has a contraction ratio of at least 1.2, and it also relates to a turbomachine having such a contraction ratio between a front blade and a rear blade. 1. A blade group arrangement for a turbomachine in order to form a blade-row group , comprising:{'sub': min,1', 'min,2, 'a front blade and a rear blade arranged offset from each other in the axial and circumferential directions and forming an overlapping area running between a pressure side of the front blade and a suction side of the rear blade, the front and rear blades in the overlapping area converging with a contraction ratio KV of KV≧1.2 between an inlet surface Dand an outlet surface D.'}2. The blade group arrangement as recited in wherein the contraction ratio KV is ≦2.8.3. The blade group arrangement as recited in wherein the contraction ratio KV=1.7.4. The blade group arrangement as recited in wherein the suction side of the rear blade has a greater curvature downstream from the outlet surface Dthan upstream from the outlet surface D.5. The blade group arrangement as recited in wherein the curvature has a maximum ranging from 1.6 to 1.7 times a mean curvature of the suction side of the rear blade.6. The blade group arrangement as recited in wherein the curvature maximum amounts to approximately 5% to 25% of the relative skeleton line length behind the outlet surface D.7. A turbomachine comprising at least one blade-row group having a plurality of blade group arrangements as recited in . This claims the benefit of European Patent Application EP 12154944.8, filed Feb. 10, 2012 and hereby incorporated by reference herein.The invention relates to a blade group arrangement as well as to a turbomachine.The maximum deflection of a row of blades of a turbomachine and thus its aerodynamic load capacity are limited, for one thing, by a separation of the flow along ...

GAS TURBINE APPARATUS WITH IMPROVED EXERGY RECOVERY

A gas turbine apparatus comprising a compressor adapted to compress air; a combustion chamber, disposed downstream of the compressor and adapted to receive and combust a first portion of the compressed air is disclosed. The gas turbine apparatus further comprises a turbine, disposed downstream of the combustion chamber and adapted to receive and perform work from the combusted and compressed air, and a heat exchanger comprising a cooler, disposed parallel to the combustion chamber, downstream of the compressor and upstream of the gas turbine, which is adapted to receive and cool at least a second portion of the compressed air and provide the cooled compressed air to at least one thermally loaded part of the gas turbine. In addition, the gas turbine apparatus comprises at least one super-heater assembly, disposed downstream of the compressor and upstream of the cooler, adapted to extract and transfer energy from the compressed air to the combustion chamber and/or the turbine and/or an auxiliary energy recovering system, utilizing at least one auxiliary working medium adapted to transfer the energy. 1. A gas turbine apparatus comprising:a compressor adapted to compress air;a combustion chamber, disposed downstream of said compressor and adapted to receive and combust a first portion of said compressed air;a turbine, disposed downstream of said combustion chamber and adapted to receive and perform work from said combusted and compressed air;a heat exchanger system comprising a cooler, disposed parallel to said combustion chamber, downstream of said compressor and upstream of said gas turbine, adapted to receive and cool at least a second portion of said compressed air and provide said cooled compressed air to at least one thermally loaded part of said gas turbine; andfurther comprising at least one super-heater assembly, disposed downstream of said compressor and upstream of said cooler, adapted to extract and transfer energy from said compressed air to said combustion ...

AIRCRAFT ENGINE WITH TURBINE HEAT EXCHANGER BYPASS

An aircraft engine, in particular a helicopter engine, having a one or multi-stage compressor system (V), a combustion chamber (BK) connected downstream therefrom, and a one- or multi-stage turbine system (HT, NT) connected downstream therefrom, a compressor heat exchanger system (WV), a turbine heat exchanger system (WT), and a bypass means for optionally guiding the working medium through or past at least one turbine heat exchanger (WT) of the turbine heat exchanger system. 1. An aircraft engine comprising:a one- or multi-stage compressor system;a combustion chamber connected downstream from the compressor system;a one- or multi-stage turbine system connected downstream from the combustion chamber;a compressor heat exchanger system;a turbine heat exchanger system having at least one turbine heat exchanger; anda bypass for optionally guiding the working medium through or past the at least one turbine heat exchanger.2. The aircraft engine as recited in wherein the compressor and/or the turbine heat exchanger system is/are designed for partial load operation.3. The aircraft engine as recited in wherein one or multiple stages of all stages of the compressor system and/or of the turbine system are designed to be adjustable.4. The aircraft engine as recited in wherein one or multiple stages of all stages of the compressor system and/or of the turbine system have adjustably designed stationary and/or moving blades.5. The aircraft engine as recited in wherein at least 30% of all of the stages of the compressor system and/or of the turbine system are designed to be adjustable.6. The aircraft engine as recited in wherein at least 50% of all of the stages of the compressor system and/or of the turbine system are designed to be adjustable.7. The aircraft engine as recited in wherein the turbine system includes a high-pressure turbine or turbine stage having a variable capacity.8. The aircraft engine as recited in further comprising having a control unit configured to guide a ...

PROPFAN ENGINE

The present disclosure relates to a propfan engine comprising: one or more rotor stages comprising a plurality of rotors; and an outer wall comprising an outer profile, at least a portion of the outer profile defining a substantially circular cross-section, wherein the diameter of the substantially circular cross-section increases in the direction of flow over the outer wall and downstream of a leading edge of the rotors, and the diameter increases at substantially all points defining the circumference of the substantially circular cross-section. 1. A propfan engine comprising:one or more rotor stages comprising a plurality of rotors; andan outer wall comprising an outer profile, at least a portion of the outer profile comprising a substantially circular cross-section, whereinthe diameter of the substantially circular cross-section increases in the direction of flow over the outer wall and downstream of a leading edge of the rotors, and the diameter increases at substantially all points defining the circumference of the substantially circular cross-section.2. The propfan engine of claim 1 , wherein the diameter increases at an increasing rate in a first part of the portion of the outer profile claim 1 , and the diameter increases at a decreasing rate in a second part of the portion of the outer profile.3. The propfan engine of claim 2 , wherein the second part is downstream of the first part of the portion of the outer profile and there is a point of inflection in the outer profile between the first and second parts of the portion of the outer profile.4. The propfan engine of wherein the radius of the curvature in the first part of the portion of the outer profile is substantially 20% of a maximum diameter of the propfan outer wall.5. The propfan engine of claim 1 , wherein in a third part of the portion of the outer profile the diameter of the substantially circular cross-section reduces in the direction of flow over the outer wall.6. The propfan engine of claim 2 ...

STAND-BY OPERATION OF A GAS TURBINE

The invention relates a method for operating a power plant with a single shaft gas turbine, in which the gas turbine is operated at a constant speed which is below the speed at which the gas turbine is turning when the first generator is synchronized to an electric grid. The proposed method ensures good stable combustion with low emissions, a high turbine exhaust temperature, and minimized fuel consumption. 2. Method according to claim 1 , wherein the gas turbine is operating at a speed between 20% and 85% of the speed when the first generator is synchronized to an electric grid.3. Method according to claim 1 , wherein the gas turbine speed is controlled by at least one of the following: applying a torque on the shaft by the first generator which is controlled by a static frequency converter claim 1 ,adjusting a variable inlet guide vane,adjusting the temperature of the compressor inlet air,adjusting a blow off control valve position.4. Method according to claim 1 , wherein the first generator is operated to deliver power to the electric grid via the static frequency converter.5. Method according to claim 1 , wherein the first generator is operated to apply a positive torque on the shaft to drive the gas turbine via the frequency converter.6. Method according to claim 1 , wherein saidgas turbine plus steam turbine are arranged on a single shaft and steam turbine power is used to drive the gas turbine.7. Method according to claim 1 , wherein the temperature of the compressor inlet gas is increased relative to the temperature of ambient air.8. Method according to claim 7 , wherein the temperature of the compressor inlet gas is increased by at least one of following:admixing compressor bleed air to the compressor inlet gas,heating the inlet gas by heat exchange in an air pre-heater,admixing recirculated flue gas to the compressor inlet gas .9. Method according to claim 3 , wherein the gas turbine emissions are controlled by at least one of the following:adjusting the ...

REGENERATIVE TURBINE FOR POWER GENERATION SYSTEM

Disclosed is a power generation system and method. The system includes a combustor, and a turbine driven by products of the combustor. The turbine includes at least one disk supporting a plurality of airfoils, and the airfoils each have an internal passage formed therein. The system further includes a passage for routing a coolant within the system. A portion of the passage is provided by the internal passages of the airfoils, and another portion of the passage is provided between the airfoils and the combustor. The system also includes a generator driven by the turbine to generate electric power. 1. A closed power generation system comprising:a turbine having an inlet, an outlet, and a cooling passage, the cooling passage having an inlet and an outlet;a heat exchanger having an inlet and an outlet, the inlet of the heat exchanger in fluid communication with the outlet of the cooling passage, the outlet of the heat exchanger in fluid communication with the inlet of the turbine; anda compressor having an inlet and an outlet, the outlet of the compressor in fluid communication with both the inlet of the heat exchanger and the inlet of the cooling passage.2. The system as recited in claim 1 , wherein the heat exchanger is configured to heat a working fluid.3. The system as recited in claim 1 , wherein the heat exchanger is a first of two heat exchangers within the system claim 1 , the second of the two heat exchangers having an inlet and an outlet claim 1 , the inlet of the second heat exchanger in fluid communication with the outlet of the turbine claim 1 , and the outlet of the second heat exchanger in fluid communication with the inlet of the compressor.4. The system as recited in claim 3 , wherein the second heat exchanger is configured to provide heat rejection relative to a working fluid.5. The system as recited in claim 3 , including a main system loop for directing a working fluid between the compressor claim 3 , the first heat exchanger claim 3 , the turbine ...

COMBUSTOR SHELL AIR RECIRCULATION SYSTEM IN A GAS TURBINE ENGINE

A shell air recirculation system for use in a gas turbine engine includes one or more outlet ports located at a bottom wall section of an engine casing wall and one or more inlet ports located at a top wall section of the engine casing wall. The system further includes a piping system that provides fluid communication between the outlet port(s) and the inlet port(s), a blower for extracting air from a combustor shell through the outlet port(s) and for conveying the extracted air to the inlet port(s), and a valve system for selectively allowing and preventing air from passing through the piping system. The system operates during less than full load operation of the engine to circulate air within the combustor shell but is not operational during full load operation of the engine. 1. A gas turbine engine including a longitudinal axis defining an axial direction of the engine , the engine comprising:a compressor section where air pulled into the engine is compressed;a combustion section comprising a plurality of combustor apparatuses where fuel is mixed with at least a portion of the compressed air from the compressor section and burned to create hot combustion gases;a turbine section where the hot combustion gases from the combustion section are expanded to extract energy from the combustion gases;a casing having a portion disposed about the combustion section, the casing portion comprising a casing wall having a top wall section defining a top dead center, left and right side wall sections, and a bottom wall section defining a bottom dead center, the casing portion further defining an interior volume in which the combustor apparatuses and air compressed by the compressor section are located; and at least one outlet port located at the bottom wall section of the casing wall;', 'first and second inlet ports located at the top wall section of the casing wall, the inlet ports being circumferentially spaced apart from one another and located at generally the same axial ...

GAS TURBINE ENGINE AND METHOD OF OPERATING THEREOF

A turbine system and method of operating is provided. The system includes a compressor configured to generate a compressed low-oxygen air stream and a combustor configured to receive the compressed low-oxygen air stream and to combust a fuel stream to generate a post combustion gas stream. The turbine system also includes a turbine for receiving the post combustion gas stream to generate a low-NOexhaust gas stream, a heat recovery system configured to receive the low-NOexhaust gas stream and generate a cooled air stream and an auxiliary compressor configured to generate an oxygen and water vapor deficient cooled and compressed air stream. A portion of the oxygen and water vapor deficient cooled and compressed air stream is directed to the combustor to generate an Oxygen and HO deficient film on exposed portions of the combustor, and another portion is directed to the turbine to provide a cooling flow. 1. A turbine system , comprising:a compressor configured to generate a compressed low-oxygen air stream;a combustor configured to receive the compressed low-oxygen air stream from the compressor and to combust a fuel stream to generate a post combustion gas stream;{'sub': 'x', 'a turbine for receiving the post combustion gas stream from the combustor to generate power and a low-NOexhaust gas stream;'}{'sub': 'x', 'a heat recovery system configured to receive the low-NOexhaust gas stream from the turbine for generation of a cooled gas stream; and'}an auxiliary compressor configured to receive at least a portion of the cooled gas stream from the heat recovery system for generation of an oxygen and water vapor deficient cooled and compressed gas stream,{'sub': '2', 'wherein a portion of the oxygen and water vapor deficient cooled and compressed gas stream is directed to the combustor to generate an oxygen and HO deficient film on exposed portions of the combustor, and'}wherein a portion of the oxygen and water vapor deficient cooled and compressed gas stream is directed ...

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 ...

AIRFOIL WITH VARIABLE PROFILE RESPONSIVE TO THERMAL CONDITIONS

An airfoil includes an airfoil body having a first section and a second section that differ in coefficient of thermal expansion by at least 10% and that each have a through-thickness that is 20% or greater than a through-thickness of the airfoil body. The second section is arranged in thermomechanical juxtaposition with the first section such that the first section and the second section cooperatively thermomechanically control a profile of the airfoil body responsive to varying thermal conditions. 1. An airfoil comprising:an airfoil body including a first section and a second section that differ in coefficient of thermal expansion by at least 10% and that each have a through-thickness that is 20% or greater than a through-thickness of the airfoil body, the second section being arranged in thermomechanical juxtaposition with the first section such that the first section and the second section cooperatively thermomechanically control a profile of the airfoil body responsive to varying thermal conditions.2. The airfoil as recited in claim 1 , wherein the first section and the second section are different composition metallic materials.3. The airfoil as recited in claim 2 , including a compositional gradient boundary between the first section and the second section.4. The airfoil as recited in claim 1 , wherein the first section and the second section are different compositions selected from aluminum claim 1 , aluminum alloys claim 1 , titanium claim 1 , titanium alloys claim 1 , iron claim 1 , iron alloys claim 1 , nickel claim 1 , nickel alloys claim 1 , cobalt claim 1 , cobalt alloys and combinations thereof.5. The airfoil as recited in claim 1 , including a distinct boundary between the first section and the second section.6. The airfoil as recited in claim 1 , wherein the first section is a core and the second section is a shell circumscribing the core.7. The airfoil as recited in claim 6 , wherein the core terminates at an intermediate span between ends of the ...

Airfoil with Thickened Root and Fan and Engine Incorporating Same

In accordance with one aspect of the disclosure, an airfoil is disclosed. The airfoil may include a platform and a blade extending from the platform. The blade may have a root proximate the platform and a tip radially outward from the platform. The root may have a greater thickness than a cross-section at about a quarter-span of the blade or greater. 1. An airfoil , comprising:a platform; anda blade extending from the platform, where the blade has a root proximate the platform and a tip radially outward from the platform, the root having a greater thickness than a cross-section at about a quarter-span of the blade or greater.2. The airfoil of claim 1 , wherein the root of the blade has a thickness about twenty percent greater than a cross-section at about a quarter-span of the blade or greater.3. The airfoil of claim 1 , wherein the root of the blade includes about twenty-five percent of a radial height of the blade.4. The airfoil of claim 1 , further including a transition zone between the tip and the root of the blade claim 1 , the transition zone being aerodynamically smooth.5. The airfoil of claim 1 , further including a fillet joining the blade with the platform of the airfoil.6. The airfoil of claim 5 , wherein the fillet has a width that varies along an axial length of the blade.7. The airfoil of claim 6 , wherein the blade includes a leading edge claim 6 , a central portion claim 6 , and a trailing edge; the leading edge interacting with incoming airflow before other surfaces of the blade claim 6 , the trailing edge interacting with outgoing airflow claim 6 , and the central portion extending between the leading and trailing edges; and the leading and trailing edges of the blade have a steeper fillet than that of the central edge of the blade.8. A fan of a gas turbine engine claim 6 , comprising:a hub; anda plurality of airfoils radially extending from the hub, each airfoil having a platform and a blade radially extending from the platform, the blade having ...

GAS TURBINE ENGINE STATOR VANE PLATFORM REINFORCEMENT

A stator vane for a gas turbine engine includes a first platform and a second platform radially spaced apart from one another. The first and second airfoils are circumferentially spaced from one another and interconnect the first and second platforms. The first platform has a gas path side facing the airfoils and a non-gas path side opposite the gas path side. A circumferentially extending rail provided on the first platform extends radially outward from the gas path side to the non-gas path side to form a pocket on the non-gas path side between the first platform and the rail. A reinforcement is arranged in the pocket and joins the first platform and the rail. The reinforcement includes a variable thickness in the circumferential direction and is arranged generally centrally between the first and second airfoils. 1. A stator vane for a gas turbine engine , comprising:a first platform and a second platform radially spaced apart from one another;first and second airfoils circumferentially spaced from one another and interconnecting the first and second platforms;the first platform having a gas path side facing the airfoils and a non-gas path side opposite the gas path side, and a circumferentially extending rail provided on the first platform and extending radially outward from the gas path side to the non-gas path side to form a pocket on the non-gas path side between the first platform and the rail; anda reinforcement arranged in the pocket and joining the first platform and the rail, the reinforcement includes a variable thickness in the circumferential direction and is arranged generally centrally between the first and second airfoils.2. The stator vane according to claim 1 , comprising spaced apart hooks supported by the first platform.3. The stator vane according to claim 1 , comprising circumferentially spaced apart side walls supported on the first platform and forming the pocket.4. The stator vane according to claim 1 , wherein the hooks extend radially ...

GAS TURBINE ENGINE HAVING CONFIGURABLE BYPASS PASSAGE

A gas turbine engine is disclosed which includes a bypass passage that in some embodiments are capable of being configured to act as a resonance space. The resonance space can be used to attenuate/accentuate/etc a noise produced elsewhere. The bypass passage can be configured in a number of ways to form the resonance space. For example, the space can have any variety of geometries, configurations, etc. In one non-limiting form the resonance space can attenuate a noise forward of the bypass duct. In another non-limiting form the resonance space can attenuate a noise aft of the bypass duct. Any number of variations is possible. 1. An apparatus comprising:a gas turbine engine having a core flow path and a bypass duct structured to bypass a working flow around a combustor of the gas turbine engine; anda moveable component associated with the bypass duct and structured to change a flow area by moving between a first position and a second position, the moveable component actuated to the second position such that a geometry of the duct forms a resonance space tuned to attenuate a noise.2. The apparatus of claim 1 , which further includes an actuator coupled to the moveable component and energized by a controller.3. The apparatus of claim 2 , which further includes a sensor structured to detect a noise information claim 2 , the sensor in communication with the controller claim 2 , wherein the controller is structured to operate upon the noise information when energizing the controller.4. The apparatus of claim 1 , wherein the bypass duct includes a plurality of bypass ducts claim 1 , and wherein various of the plurality of bypass ducts can form resonance spaces.5. The apparatus of claim 1 , wherein the resonance space includes a first portion forward of the moveable component and a second portion aft of the moveable component claim 1 , each of the first portion and second portion capable of generating a noise.6. The apparatus of claim 1 , wherein the resonance space ...

COMBUSTOR AND GAS TURBINE

A combustor includes a first tube which supplies fuel and air from an opening formed in a distal end thereof, and a second tube which is configured such that flames are formed on an inner circumferential side thereof by the fuel and the air and the distal end of the first tube is inserted into an inner circumferential side of a proximal end thereof. The first tube includes a first tube main body and a ring part forming the distal end of the first tube. The ring part has a main body of a tubular shape and a plurality of protrusions that are integrally formed with the main body on an outer circumferential surface of the main body and protrude radially outward. When viewed in an axial direction, the outer circumferential surface of the main body has a polygonal cross-sectional shape within a range within which the protrusions are formed. 1. A combustor comprising:a first tube which supplies fuel and air from an opening formed in a distal end thereof; anda second tube which is configured such that the distal end of the first tube is inserted into an inner circumferential side of a proximal end thereof,wherein the first tube includes a first tube main body and a ring part forming the distal end of the first tube,the ring part has a main body of a tubular shape and a plurality of protrusions that are integrally formed with the main body on an outer circumferential surface of the main body and protrude radially outward, andwhen viewed in an axial direction along an axis of the first tube, the outer circumferential surface of the main body has a polygonal cross-sectional shape within a range within which the protrusions are formed.2. The combustor according to claim 1 , wherein the main body includes a polygonal ring part on which the protrusions are formed claim 1 , and a restrictor part which is provided at a distal end side relative to the polygonal ring part and an outer circumferential surface of which is formed in a cylindrical shape at a radially outer ...

SPINNER FOR A GAS TURBINE ENGINE

A spinner for use in a gas turbine engine includes a body and an aft ring coupled to the body. The body is symmetrically formed by filament winding around a central axis. The aft ring includes a band coupled to the body and a plurality of fairings that extend outward in a radial direction from the band away from the central axis. 1. A spinner for use in a gas turbine engine comprisinga filament-wound body formed substantially symmetrically around a central axis, the body including a side wall having a diameter that increases along the central axis from a forward side of the body to an aft side of the body, andan aft ring including a band that overlaps the side wall along the aft side of the body and a plurality of fairings that extend outward in a radial direction from the band away from the central axis.2. The spinner of claim 1 , wherein the aft ring is formed from a composite material and is bonded to the body by a resin.3. The spinner of claim 1 , wherein the aft ring includes a continuous braided sock of woven fabric that extends around the aft side of the body without an axial seam that extends along the central axis.4. The spinner of claim 3 , wherein the fairings included in aft ring are formed by prearranged portions of the continuous braided sock filled with a plug.5. The spinner of claim 1 , wherein the aft ring includes a plurality of segments each arranged to extend around a portion of the diameter of the aft ring.6. The spinner of claim 5 , wherein each segment is formed to include a first fairing and a second fairing.7. The spinner of claim 5 , wherein each segment is formed to include only one fairing.8. The spinner of claim 5 , wherein each segment includes a plurality of layers in which at least one radially-outer layer extends further in the axial direction along the body toward the forward side of the body than at least one radially-inner layer.9. The spinner of claim 5 , wherein each segment is spaced circumferentially apart from adjacent ...

GAS TURBINE SILENCER, AND GAS TURBINE PROVIDED WITH SAME

Provided are a gas turbine silencer for avoiding the formation of gaps between an upstream silencer panel and a downstream silencer panel and suppressing the occurrence of secondary noise, and a gas turbine provided with this silencer. A silencer panel () has a structure that can be divided into an upstream silencer panel () and a downstream silencer panel () in an airflow direction, a stepped part () shorter in length along the alignment direction of the silencer panel () is formed in an opening-side portion of the downstream silencer panel (), and the upstream silencer panel () and downstream silencer panel () are linked by the stepped part () fitting into substantially the entire opening () in the upstream silencer panel (). 1. A gas turbine silencer installed between an air intake port and a compressor of a gas turbine , the gas turbine silencer comprising:a plurality of plate-shaped divided silencer panels aligned at predetermined intervals in a direction orthogonal to a flow direction of a fluid from the air intake port toward the compressor;the divided silencer panels wherein a surface thereof with a greatest plate area is arranged in orientation along a flow of the fluid, the divided silencer panels comprising an upstream silencer panel arranged on an upstream side in the flow direction of the fluid, and a downstream silencer panel arranged on an downstream side of the upstream silencer panel and linked with the upstream silencer panel;one silencer panel out of the upstream silencer panel and the downstream silencer panel being formed with an opening opened on a side facing the other silencer panel;the other silencer panel being formed with a fitting section fitting into the opening; andthe upstream silencer panel and the downstream silencer panel being linked by the fitting section of the other silencer panel fitting into the opening of the one silencer panel.2. The gas turbine silencer according to claim 1 , wherein the divided silencer panels are ...

Turbofan jet engine

A turbofan engine includes a core engine, having a high-pressure compressor, a combustion chamber and a high-pressure turbine which are coupled to one another via a high-pressure shaft, at least one fan from which gas is supplied into both a primary flow duct and a secondary flow duct of the turbofan engine, at least one low-pressure turbine arranged behind the core engine, and at least one low-pressure shaft, with each low-pressure shaft coupling a fan to a low-pressure turbine. It has been provided that no low-pressure shaft of the turbofan engine passes through the core engine.

ACOUSTIC LINER HEAT EXCHANGER

The present disclosure provides an acoustic liner for a gas turbine engine including an acoustic structure for the absorption of acoustic excitation and a heat exchanger. The heat exchanger is able to exchange heat across the acoustic structure.

FAN UNIT FOR A TURBOFAN ENGINE

A fan unit for a turbofan engine includes a plurality of fan blades mounted to a fan disc unit and a plurality of vortex-generating elements each of which is arranged to generate a respective vortex extending axially between a respective pair of adjacent fan blades during rotation of the fan unit about its rotation axis. During operation of a turbofan engine comprising the fan unit, fan-wake-induced mechanical degradation of compressor blades in a compressor downstream of the fan unit is eliminated, or the rate at which such degradation occurs is reduced. 1. A fan unit for a turbofan engine , the fan unit comprising a plurality of fan blades mounted to a fan disc unit and a plurality of vortex-generating elements each of which is arranged to generate a respective vortex extending axially between a respective pair of adjacent fan blades during rotation of the fan unit about its rotation axis.2. A fan unit according to wherein an outer surface of the fan disc unit defines an inner annulus profile surface of the fan unit and each of the vortex-generating elements is arranged to generate a vortex extending axially between a respective pair of adjacent fan blades at or adjacent the inner annulus profile surface during rotation of the fan unit about its rotation axis.3. A fan unit according to wherein each vortex-generating element is located axially upstream of the leading edges of a respective pair fan blades.4. A fan unit according to wherein each vortex-generating element is located in azimuth between azimuthal positions of a respective pair of adjacent fan blades.5. A fan unit according to wherein the number of vortex-generating elements is equal to the number of fan blades.6. A fan unit according to any of wherein the fan unit has n fan blades claim 1 , where n is even claim 1 , and(i) there are n/2 vortex-generating elements, with one vortex-generating element being located in azimuth between the azimuthal positions of the roots of every other pair of fan blades; ...

TRANSIENT LIQUID PAHSE BONDED TURBINE ROTOR ASSEMBLY

A rotor assembly for a turbine engine includes a rotor disk constructed of a first material. Multiple rotor blades constructed of a second material are connected to the rotor disk via a diffusion material. 1. A turbine engine comprising:a compressor section;a combustor in fluid communication with the compressor section;a turbine section in fluid communication with the combustor; a rotor disk constructed of a first material;', 'a plurality of rotor blades constructed of a second material; and', 'a transient liquid phase bond connecting a bond surface of said rotor disk and a bond surface of each of said rotor blades., 'a gas path defined by said compressor section, said combustor and said turbine section, wherein said gas path includes at least one rotor assembly, wherein said rotor assembly includes2. The turbine engine of claim 1 , wherein said transient liquid phase bond is a partial transient liquid phase bond.3. The turbine engine of claim 2 , wherein said transient liquid phase bond is a combined transient liquid phase bond and partial transient liquid phase bond.4. The turbine engine of claim 1 , wherein said transient liquid phase bond is a diffusion layer formed of material diffused from a thin foil interlayer material.5. The turbine engine of claim 1 , further comprising at least one cover plate connected to a cover plate mounting feature of said rotor disk claim 1 , wherein said cover plate is spaced from a root portion of the rotor blade.6. The turbine engine of claim 1 , further comprising a compressor blade cooling flow passage disposed entirely within said rotor blade.7. The turbine engine of claim 1 , further comprising a blade cooling flow passage having a blade cooling flow passage inlet disposed on an inner diameter surface of said rotor blade claim 1 , and defined entirely by said rotor blade.8. The turbine engine of claim 1 , wherein each of said rotor blades is constructed of a high temperature claim 1 , low ductility first material claim 1 , ...

TURBINE ENGINE FACE SEAL ARRANGEMENT INCLUDING ANTI-ROTATION FEATURES

A turbine engine includes a main shaft bearing compartment seal. The seal includes at least an approximately circular seal portion and a seal carrier disposed about the approximately circular seal portion. A plurality of anti-rotation pins maintain the seal carrier in position relative to a housing and are received in an anti-rotation slot of the seal carrier. 1. A turbine engine comprising:a compressor section;a combustor in fluid communication with the compressor section;a turbine section in fluid communication with the combustor;a main shaft interconnecting each of the compressor section, the combustor and the turbine section; and an at least approximately circular seal;', 'a seal carrier disposed about said seal, wherein said seal carrier maintains said seal in a position;', 'a housing surrounding said seal carrier, wherein said seal carrier is maintained in position relative to said housing via a plurality of springs;', 'a plurality of anti-rotation pin assemblies rigidly connected to said housing on a first end and received in an anti-rotation slot of said seal carrier on a second end, wherein said anti-rotation pins are aligned axially with said seal carrier and said seal; and', 'wherein each of said anti-rotation slots comprises an elongated seal carrier contact surface contacting said anti-rotation pin assembly., 'a main shaft bearing compartment including a main shaft bearing compartment seal, wherein said main shaft bearing compartment seal further comprises2. The turbine engine of claim 1 , wherein each of said anti-rotation pin assemblies comprises a pin body claim 1 , a sleeve disposed about said pin body and an end cap connected to said pin body claim 1 , wherein said pin body and said end cap are arranged such that said sleeve cannot be removed while said end cap is attached to said pin body.3. The turbine engine of claim 2 , wherein said sleeve contacts said elongated seal carrier contact surface.4. The turbine engine of claim 3 , wherein said ...

COMPACT MULTI-PIECE SPRING-LOADED CROSSFIRE TUBE

A crossfire tube assembly is positioned between adjacent combustors, the crossfire tube assembly having a primary body made up of a first telescoping sleeve slidably engaged with a second telescoping sleeve. An interlocking raceway is configured to limit axial travel length of the telescoping sleeves and lock the telescoping sleeves to each other. A bias is positioned between the first telescoping sleeve and the second telescoping sleeve. First and second floating collars are removably disposed to the first and second telescoping sleeves at a first and second floating collar annulus. First and second liner collars are disposed between the first and second floating collars on the first and second combustors. The crossfire tube assembly is adapted to provide fluid communication from the first combustor to the second combustor serving a gas turbine. 1. A crossfire tube assembly between adjacent combustors , the crossfire tube assembly comprising; a first telescoping sleeve slidably engaged with a second telescoping sleeve defining axial and circumferential relative movement of the telescoping sleeves,', 'an interlocking raceway configured to limit axial travel length of the telescoping sleeves and configured to lock the telescoping sleeves to each other, and', 'a bias between the first telescoping sleeve and the second telescoping sleeve,, 'a primary body, comprising;'}a first floating collar removably disposed to the first telescoping sleeve at a first floating collar annulus,a second floating collar removably disposed to the second telescoping sleeve at a second floating collar annulus,a first liner collar disposed between the first floating collar and a first liner of a first combustor,a second liner collar disposed between the second floating collar and a second liner of a second combustor, andwherein the crossfire tube assembly is adapted to provide fluid communication from the first combustor to the second combustor.2. The crossfire tube assembly as in claim 1 , ...

APPARATUS FOR A GAS TURBINE ENGINE

Apparatus for a gas turbine engine, the apparatus comprising: a low pressure compressor; a high pressure compressor; a first electrical machine configured to provide torque to the low pressure compressor; a second electrical machine configured to receive torque from the high pressure compressor; and wherein the high pressure compressor does not comprise a sub-idle bleed valve. 1. Apparatus for a gas turbine engine , the apparatus comprising:a high pressure compressor;an electrical machine configured to receive torque from the high pressure compressor; and wherein the high pressure compressor comprises only bleed valves having three or more operational positions.2. Apparatus as claimed in claim 1 , wherein the high pressure compressor does not comprise a sub-idle bleed valve.3. Apparatus as claimed in claim 1 , wherein the high pressure compressor has a pressure ratio of at least eight to one.4. Apparatus as claimed in claim 1 , wherein the high pressure compressor comprises a plurality of compressor stages including a first compressor stage and a final compressor stage claim 1 , the high pressure compressor comprising no bleed valves between a first position halfway between the first compressor stage and the final compressor stage claim 1 , and a second position at the final compressor stage.5. Apparatus as claimed in claim 1 , further comprising: a low pressure compressor; and another electrical machine configured to provide torque to the low pressure compressor.6. Apparatus for a gas turbine engine claim 1 , the apparatus comprising:a high pressure compressor comprising a plurality of compressor stages including a first compressor stage and a final compressor stage;an electrical machine configured to receive torque from the high pressure compressor; andwherein the high pressure compressor comprises no bleed valves between a first position halfway between the first compressor stage and the final compressor stage, and a second position at the final compressor stage. ...

VARIABLE-PITCH ROTOR WITH REMOTE COUNTERWEIGHTS

A pitch control mechanism includes: a rotor structure configured for rotation about a longitudinal axis; a row of blades carried by the rotor structure, each blade having an airfoil and a trunnion mounted for pivoting movement relative to the rotor structure, about a trunnion axis which is perpendicular to the longitudinal axis; a unison ring interconnecting the blades; an actuator connected to the unison ring and the rotor structure, operable to move the unison ring relative to the rotor structure; at least one moveable counterweight carried by the rotor structure, remote from the blades; and an interconnection between the blades and the counterweight, such that movement of the counterweight causes a change in the pitch angle of the blades. 1. A pitch control mechanism , comprising:a rotor structure configured for rotation about a longitudinal axis;a row of blades carried by the rotor structure, each blade having an airfoil and a trunnion mounted for pivoting movement relative to the rotor structure, about a trunnion axis which is perpendicular to the longitudinal axis;a unison ring interconnecting the blades;an actuator connected to the unison ring and the rotor structure, operable to move the unison ring relative to the rotor structure;at least one moveable counterweight carried by the rotor structure, remote from the blades; andan interconnection between the blades and the counterweight, such that movement of the counterweight causes a change in the pitch angle of the blades.2. The pitch control mechanism of wherein the actuator is configured to produce rotary movement between the rotor structure and the unison ring.3. The pitch control mechanism of wherein the unison ring and counterweights are interconnected by gears.4. The pitch control mechanism of wherein the rotor structure carries an array of counterweight assemblies each including: a pinion shaft claim 1 , a pinion gear claim 1 , and a counterweight with an offset mass.5. The pitch control mechanism of ...

TURBINE AND GAS TURBINE

A flow path width at a hub endwall of a blade main body decreases toward a minimum width from a leading edge, and increases toward a trailing edge from the minimum width, the flow path width at a reference blade height aparted toward a tip side from the hub endwall of the blade main body gradually decreases toward the trailing edge from the leading edge, and an axial chord length position of the minimum flow path widths at respective blade height shift toward a transition to the trailing edge side from the hub endwall toward the tip side of the blade main body and coincides with the trailing edge at the reference blade height. 1. A turbine comprising:a plurality of blades including blade main bodies extending radially outwards from an axis, a flow path being formed between the neighboring blade main bodies by the blades being arranged in a circumferential direction of the axis,wherein a width of the flow path at a hub endwall of the blade main body decreases toward a minimum width from a leading edge, and increases toward a trailing edge from the minimum width, and the minimum width is located between the leading edge and the trailing edge,wherein the flow path width at a reference blade height aparted toward a tip side from the hub endwall gradually decreases toward the trailing edge from the leading edge, andwherein an axial chord length position of the minimum flow path widths at respective blade height shift toward the trailing edge side from the hub endwall toward the tip side of the blade main body and coincides with the trailing edge at the reference blade height.2. The turbine according to claim 1 ,wherein the reference blade height is located between 5% and 25% of a blade height toward the tip side from the hub endwall.3. The turbine according to claim 1 ,wherein the blade height at a transition position is positioned in a region within 10% of the blade height toward the tip side from the hub endwall when the flow path width at each of the blade height on ...

METHOD OF MAKING AN AERO-DERIVATIVE GAS TURBINE ENGINE

A method of making an aero-derivative gas turbine engine () is provided. A combustor outer casing () is removed from an existing aero gas turbine engine (). An annular combustor () is removed from the existing aero gas turbine engine. A first row of turbine vanes () is removed from the existing aero gas turbine engine. A can annular combustor assembly () is installed within the existing aero gas turbine engine. The can annular combustor assembly is configured to accelerate and orient combustion gasses directly onto a first row of turbine blades of the existing aero gas turbine engine. A can annular combustor assembly outer casing () is installed to produce the aero-derivative gas turbine engine (). The can annular combustor assembly is installed within an axial span () of the existing aero gas turbine engine vacated by the annular combustor and the first row of turbine vanes. 1. A method of making an aero-derivative gas turbine engine , the method comprising:removing a combustor outer casing of an existing aero gas turbine engine;removing an annular combustor of the existing aero gas turbine engine;removing a first row of turbine vanes of the existing aero gas turbine engine;installing a can annular combustor assembly within the existing aero gas turbine engine, the can annular combustor assembly configured to accelerate and orient combustion gasses directly onto a first row of turbine blades of the existing aero gas turbine engine, the can annular combustor assembly comprising a plurality of combustors and associated transition ducts; andinstalling a can annular combustor assembly outer casing to produce the aero-derivative gas turbine engine, wherein the can annular combustor assembly is installed within an axial span of the existing aero gas turbine engine vacated by the annular combustor and the first row of turbine vanes.2. The method of claim 1 , wherein the step of installing the can annular combustor assembly does not involve changing at least one component ...

METHOD FOR REDUCING THE CO EMISSIONS OF A GAS TURBINE, AND GAS TURBINE

A method for reducing the CO emissions of a gas turbine having a compressor, a turbine and an air preheater positioned upstream of the compressor, that permits technically simpler regulation without losses in terms of the quality of the reduction of the CO emissions. The heat transfer power of the air preheater is regulated on the basis of a minimum value for the inlet temperature of the compressor, wherein the minimum value is predefined as a function of the absolute power of the gas turbine. 1. A method for reducing the CO emissions of a gas turbine having a compressor , a turbine and an air preheater connected upstream from the compressor , the method comprising:regulating a heat transfer capacity of the air preheater with the aid of a minimum value for the inlet temperature of the compressor, andpredetermining the minimum value as a function of an absolute capacity of the gas turbine.2. The method as claimed in claim 1 ,wherein the function is determined with the aid of a model calculation for the gas turbine.3. The method as claimed in claim 1 ,wherein the function is monotonously decreasing.4. The method as claimed in claim 1 ,wherein the function is constant below a lower limit value and/or above an upper limit value for the capacity of the gas turbine.5. The method as claimed in claim 4 ,wherein the function is linear between the limit values.6. The method as claimed in claim 4 ,wherein the function has a value of approximately 20° C. below the lower limit value and has a value of approximately 20° C. above the upper limit value.7. A gas turbine comprisinga compressor,a turbine,an air preheater connected upstream from the compressor,a temperature measurement device, arranged between the compressor and the air preheater, which is connected on a data output side to a control device of the gas turbine,a capacity measurement device which is connected on the data output side to the control device, anda flow control valve for regulating the heat transfer capacity ...

COMPRESSOR BLADE ASSEMBLY STRUCTURE, GAS TURBINE HAVING SAME, AND COMPRESSOR BLADE ASSEMBLY METHOD

A compressor blade assembly structure, a gas turbine having the same, and a method of assembling compressor blade are provided. The compressor blade assembly structure includes a compressor blade having an airfoil, a platform part, and a dovetail part, a compressor rotor disk having a dovetail slot into which the dovetail part is inserted, and a locking key mounted in a key slot formed in the dovetail slot to support the compressor blade in an axial direction. 1. A compressor blade assembly structure comprising:a compressor blade having an airfoil, a platform part, and a dovetail part;a compressor rotor disk having a dovetail slot into which the dovetail part is inserted; anda locking key mounted in a key slot formed in the dovetail slot to support the compressor blade in an axial direction.2. The compressor blade assembly structure according to claim 1 , wherein the locking key comprises:a main body formed elongated in the axial direction;a protrusion formed in a center of a lower surface of the main body; anda pair of ribs formed on both sides of an upper surface of the main body.3. The compressor blade assembly structure according to claim 2 , wherein the main body is formed of an elastically deformable rectangular plate.4. The compressor blade assembly structure according to claim 3 , wherein the key slot is formed on a bottom of the dovetail slot to accommodate the main body claim 3 , wherein the key slot includes an insertion groove into which the protrusion is inserted and an accommodation groove for accommodating one end of the elastically deformed locking key.5. The compressor blade assembly structure according to claim 4 , wherein the dovetail part of the compressor blade includes a pair of key grooves on both sides in the axial direction to accommodate the pair of ribs.6. The compressor blade assembly structure according to claim 5 , wherein the locking key is mounted in the key slot and elastically deformed on one side so that the dovetail part of the ...

METHOD OF ENGINE SPLIT AND REASSEMBLY

A method of axially separating an annular system, such as a gas turbine engine (), comprising first () and second () annular components. An annular array of fastenings () couples the first () and second () components together axially. The first () and second () components are supported by support tooling. The fastenings () are removed to leave one fastening () located on each side of the system. One of the remaining fastenings () is removed. The relative height of the first () and second () components is adjusted so that the apertures () for the fastening () removed at the previous step are aligned. Then the final fastening () is removed. 1101034364234364648. A method of axially separating an annular system (); the system () comprising a first annular component () , a second annular component () and an annular array of fastenings () that couple the first and second components ( , ) together axially through aligned apertures ( , ); the method comprising steps to:{'b': 34', '36, 'a) support the first and second components (, ) with support tooling;'}{'b': 42', '42', '10, 'b) remove fastenings () to leave one fastening () located on each side of the system ();'}{'b': '42', 'c) remove one of the remaining fastenings ();'}{'b': 34', '36', '46', '48', '42, 'd) adjust the relative height of the first and second components (, ) so that the apertures (, ) for the fastening () removed at step c) are aligned; and'}{'b': '42', 'e) remove the final fastening ().'}215242. A method as claimed in wherein the locations in step .b) are each less than or equal to 15° from a horizontal plane () through the diameter of the array of fastenings ().3142541042561011. A method as claimed in in which step .b) is modified to comprise leaving one fastening () located on one side () of the system () and two fastenings () located on the second side () of the system () claim 1 , the method having a further step between steps .b) and .c) to:{'b': 42', '56', '34', '36', '46', '48', '42, 'a) remove ...

GAS TURBINE ENGINE COOLING FLUID METERING SYSTEM

A cooling fluid system for a gas turbine engine includes a fluid source. A turbine section includes first and second components. A fluid supply system has a primary pipe that is configured to provide a cooling supply fluid from the fluid source to a fluid fitting having a fluid junction. The fluid junction is in fluid communication with and is configured to supply a first cooling fluid to the first component. The fluid junction is in fluid communication with and is configured to supply a second cooling fluid to the second component. A flow meter is upstream from the fluid junction and is configured to receive the cooling supply fluid. 1. A cooling fluid system for a gas turbine engine comprising:a fluid source;a turbine section includes first and second components; anda fluid supply system has a primary pipe configured to provide a cooling supply fluid from the fluid source to a fluid fitting having a fluid junction, the fluid junction in fluid communication with and configured to supply a first cooling fluid to the first component, and the fluid junction in fluid communication with and configured to supply a second cooling fluid to the second component, and a flow meter upstream from the fluid junction and configured to receive the cooling supply fluid.2. The cooling fluid system according to claim 1 , wherein the fluid source is bleed air from a compressor section.3. The cooling fluid system according to claim 2 , wherein the fluid source is a high pressure compressor.4. The cooling fluid system according to claim 1 , wherein the first and second components are provided by a mid-turbine frame.5. The cooling fluid system according to claim 4 , wherein the first component is a mid-turbine frame vane.6. The cooling fluid system according to claim 5 , wherein the first component is configured to supply the first cooling fluid to a turbine rotor.7. The cooling fluid system according to claim 5 , wherein the second component is an I-rod that extends radially through the ...

NOZZLE HAVING AN ORIFICE PLUG FOR A GAS TURBOMACHINE

A gas turbomachine nozzle includes a base portion having a first fluid inlet and a second fluid inlet, and an outlet portion having one or more outlets. A connection section fluidically connects the base portion and the outlet portion. An orifice plug is arranged in the base portion at the second fluid inlet. The orifice plug includes one or more openings fluidically connecting the second fluid inlet and the connection section. 1. A gas turbomachine nozzle comprising:a base portion including a first fluid inlet and a second fluid inlet;an outlet portion having one or more outlets;a connection section fluidically connecting the base portion and the outlet portion; andan orifice plug arranged in the base portion at the second fluid inlet, the orifice plug including one or more openings fluidically connecting the second fluid inlet and the connection section.2. The gas turbomachine nozzle according to claim 1 , wherein the orifice plug includes a base section claim 1 , the one or more openings being formed in the base section.3. The gas turbomachine nozzle according to claim 1 , wherein the base portion includes a base surface defined by an annular claim 1 , outer edge claim 1 , the first and second fluid inlets being formed in the base surface.4. The gas turbomachine nozzle according to claim 3 , further comprising: an orifice plug receiving zone having an opening formed in the annular claim 3 , outer edge.5. The gas turbomachine nozzle according to claim 4 , wherein the orifice plug includes a base section and a plug section joined by an intermediate section claim 4 , the plug section extending into claim 4 , and closing claim 4 , the opening.6. The gas turbomachine nozzle according to claim 5 , wherein the base section and the plug section include a first diameter and the intermediate section includes a second diameter that is smaller than the first diameter.7. The gas turbomachine nozzle according to claim 1 , further comprising: a first passage extending from the ...

EXPANDABLE EXHAUST CONE

An expandable exhaust cone assembly is described which is able to move from a collapsed position to an expanded position. 1. An expandable exhaust cone assembly adapted for use in an associated gas turbine engine , the assembly comprisinga fixed cone segment arranged around a central axis and adapted to be mounted to a structural component of the associated gas turbine engine,a plurality of movable cone segments configured to move from a collapsed position, associated with a first overall length of the expandable exhaust cone along the central axis, to an expanded position, associated with a second overall length of the expandable exhaust cone along the central axis that is longer than the first length, each of the plurality of movable cone segments nested within the fixed cone segment when the expandable exhaust cone assembly is in the collapsed position and arranged to extend outward of the fixed cone segment along the central axis when the expandable exhaust cone assembly is in the expanded position, anda cone mover configured to drive the plurality of movable cone segments from the collapsed position to the expanded position in response to startup of the associated gas turbine engine.2. The expandable exhaust cone assembly of claim 1 , wherein the cone mover includes a threaded plate fixed relative to the fixed cone segment and a mover rod adapted to be coupled to a turbine rotor for rotation therewith claim 1 , the mover rod including a threaded portion engaging threads of the threaded plate and the mover rod coupled to the plurality of movable cone segments so that the mover rod is configured to drive the plurality of movable cone segments to the expanded position in response to rotation of the turbine rotor.3. The expandable exhaust cone assembly of claim 2 , wherein the cone mover includes an end cap claim 2 , the end cap defining the end of the expandable exhaust cone assembly and the end cap not engaging the threaded portion of the mover rod so that the ...

COMBUSTOR SHAPE COOLING SYSTEM

A gas turbine engine includes a combustor which has at least one annular wall defining a combustion chamber therein. The annular wall is formed by a circumferential array of panels overlapping one with another to define a plurality of radial gaps between respective adjacent two panels. The radial gaps are configured in a spiral pattern and are in fluid communication with the combustion chamber and a space outside the combustor to allow air surrounding the annular wall to enter the combustion chamber via the radial gaps for film cooling. 1. A gas turbine engine having a gas generation section including a combustor , the combustor comprising at least one annular wall defining a combustion chamber therein , the at least one annular wall formed by a circumferential array of panels in a loadbearing configuration for bearing loads generated by a combustion reaction taking place in the combustion chamber , the panels extending from an upstream end to a downstream end of the at least one annular wall and overlapping one with another to define a plurality of radial gaps between respective adjacent two of the panels , the radial gaps being configured in a spiral pattern and being in fluid communication with the combustion chamber and a space outside the combustor to allow air surrounding the at least one annular wall to enter the combustion chamber via the radial gaps for film cooling a hot side of the at least one annular wall.2. The gas turbine engine as defined in wherein the panels each comprise opposite first and second surfaces claim 1 , a portion of the first surface which is free of said overlapping being exposed to the air surrounding the at least one annular wall and forming a part of a cold side of the at least one annular wall claim 1 , and a portion of the second surface which is free of said overlapping being exposed to combustion gases in the combustion chamber and forming a part of said hot side of the at least one annular wall.3. The gas turbine engine as ...

GAS TURBINE BLADE

A gas turbine blade according to an embodiment of the present invention includes a turbine blade provided in a gas turbine; and a plurality of film coolers formed in a section from a leading edge to a trailing edge of the turbine blade, in which the film cooler includes a cooling channel through which cooling air is introduced and formed in an oval shape; and an outlet through which the cooling air passing through the cooling channel is discharged and extending from an extended end of the cooling channel toward an outer side thereof and formed in an oval shape. 1. A gas turbine blade , comprising:a turbine blade provided in a gas turbine; anda plurality of film coolers formed in a section from a leading edge to a trailing edge of the turbine blade,wherein each film cooler includes a cooling channel through which cooling air is introduced and formed in an oval shape; and an outlet extending from one end of the cooling channel to an outer surface of the turbine blade and formed in an oval shape that becomes longer from the one end of the cooling channel to the outer surface of the turbine blade.2. The gas turbine blade of claim 1 , wherein the cooling channel and the outlet extend for the same length.3. The gas turbine blade of claim 1 , wherein the cooling channel extends for a length longer than that of the outlet.4. The gas turbine blade of claim 1 , wherein the cooling channel extends for a length shorter than that of the outlet.5. The gas turbine blade of claim 1 , wherein when an opening height of the cooling channel is H claim 1 , and a width of the cooling channel is W claim 1 , a ratio of the width W to the height H is maintained to be 2.5 to 3 times.6. The gas turbine blade of claim 1 , wherein the cooling channel and the outlet are processed to have a smooth inner surface.7. The gas turbine blade of claim 1 , wherein a diffusion angle of the outlet extended from a left side and a right side of the cooling channel is maintained within a range of 10° to 13°.8 ...

HIGH EFFICIENCY SELF-CONTAINED MODULAR TURBINE ENGINE POWER GENERATOR

A high efficiency self-contained modular turbine engine unit for generating power includes a housing defining an air intake and an exhaust port. A turbine engine is positioned and operable within the housing. The turbine engine includes a drive shaft a compressor rotor assembly, a compressor vane assembly, a combustor and diffuser assembly, a turbine vane assembly and a turbine rotor assembly. The combustor and diffuser assembly is a one-piece unit defining a shroud extending forwardly therefrom and a plurality of combustion flow channels extending rearwardly and radially inwardly thereby forming a flowpath angle in the range from about 15° to about 35° with the drive shaft. An igniter is positioned in each flow channel to ignite a fuel/oxygen mixture introduced into the compressor rotor assembly. External components required for operation of turbine engine are mounted within the housing. 1. A high efficiency self-contained modular turbine engine unit for generating power , the modular turbine engine unit comprising:a housing having a housing frame, a top panel, a bottom panel, a first side panel, a second side panel, a third side panel and a fourth side panel, each of the panels being removeably secured to the housing frame;an air intake defined in the housing; andan exhaust port defined in the housing; a drive shaft defining a drive shaft centerline,', 'at least one compressor rotor assembly mounted on the drive shaft,', 'at least one compressor vane assembly mounted on the drive shaft proximate to and downstream from the at least one compressor rotor assembly,', 'a combustor and diffuser assembly mounted on the drive shaft proximate to and downstream from the at least one compressor vane assembly, the combustor and diffuser assembly comprising a one-piece unit defining a shroud extending forwardly therefrom to define a flowpath for compressed air exiting the at least one compressor vane assembly, and a plurality of combustion flow channels extending rearwardly ...

APPARATUS AND METHOD FOR AIR PARTICLE SEPARATOR IN GAS TURBINE ENGINE

An engine and particle decelerator is provided herein. The engine having: an inlet opening for directing air towards a compressor of the engine; and a particle decelerator located between the inlet opening and the compressor such that air travelling towards the compressor from the inlet opening must travel through the particle decelerator and wherein an area of the particle decelerator is greater than an inlet and an outlet of the particle decelerator. 1. An engine , comprising:an inlet opening for directing air towards a compressor of the engine; anda particle decelerator located between the inlet opening and the compressor such that air travelling towards the compressor from the inlet opening must travel through the particle decelerator and wherein an area of the particle decelerator is greater than an inlet and an outlet of the particle decelerator.2. The engine of claim 1 , wherein the engine is a gas turbine engine and the gas turbine engine further comprises a fan for directing the air to the compressor and wherein the engine has a primary air flow path from the fan to the particle decelerator and a secondary air flow path from the fan and away from the particle decelerator and wherein the gas turbine engine is configured for use in an aircraft.3. The engine of claim 1 , wherein the area of the particle decelerator further comprises a transition area located between the inlet and the outlet of the particle decelerator and wherein the area of the particle decelerator is an annular area located about an axis of the engine.4. The engine of claim 1 , wherein the particle decelerator further comprises an opening located in an outer shell of the particle decelerator claim 1 , and a door for sealing the opening in the outer shell of the particle decelerator.5. The engine of claim 1 , wherein the particle decelerator is configured such that the inlet is offset from the outlet such that an air flow path from the inlet is offset from the outlet.6. The engine of claim 5 ...

Low Pressure Shaft

A low pressure shaft is provided for interconnecting a fan or low pressure compressor of a gas turbine engine with a low pressure turbine of the engine. The shaft has an aft section whose diameter expands progressively with distance along the rearward direction of the engine. The shaft further has an annular joint portion at the rear end of the aft section for joining the shaft to a drive arm of the low pressure turbine. The rate of expansion of the diameter of the aft section decreases with distance along the rearward direction of the engine. 1. A low pressure shaft for interconnecting a fan or low pressure compressor of a gas turbine engine with a low pressure turbine of the engine , the shaft having an aft section whose diameter expands progressively with distance along the rearward direction of the engine , and an annular joint portion at the rear end of the aft section for joining the shaft to a drive arm of the low pressure turbine;wherein the rate of expansion of the diameter of the aft section decreases with distance along the rearward direction of the engine.2. A low pressure shaft according to claim 1 , wherein on a longitudinal cross-section through the shaft the aft section has an elliptical section profile.3. A low pressure shaft according to claim 1 , wherein the aft section is formed as an axial row of two or more frustoconical sub-sections claim 1 , for each pair of adjacent frustoconical sub-sections the cone angle of the rearward sub-section being less than the cone angle of the forward sub-section.4. A low pressure shaft according to claim 1 , wherein the annular joint portion is configured to form a curvic coupling with the drive arm.5. A low pressure shaft according to claim 1 , wherein the wall thickness of the aft section decreases with distance along the rearward direction of the engine.6. A low pressure shaft according to further comprising an annular stiffening formation which bridges the rear end of the aft section and the joint portion ...

RIM SEAL FOR GAS TURBINE ENGINE

A rim seal for a rotor of a gas turbine engine includes a seal portion extending circumferentially across a rim cavity of a rotor, the sealing portion configured to seal the rim cavity and a first foot portion extending radially inwardly from a first end of the sealing portion. A rotor assembly for a gas turbine engine includes a rotor disc and a plurality of rotor blades secured to the rotor disc defining a rim cavity between the rotor disc and a rim portion of the plurality of rotor blades. A rim seal is located in the rim cavity and includes a seal portion extending circumferentially across the rim cavity, the sealing portion configured to seal the rim cavity. The seal portion has an increasing radial thickness with increasing distance from a first end of the rim seal and from a second end opposite the first end. 1. A rim seal for a rotor of a gas turbine engine , comprising:a seal portion extending circumferentially across a rim cavity of a rotor, the sealing portion configured to seal the rim cavity; anda first foot portion extending radially inwardly from a first end of the sealing portion.2. The rim seal of claim 1 , further comprising a second foot portion extending radially inwardly from a second end of the sealing portion opposite the first end.3. The rim seal of claim 1 , wherein the seal portion has an increasing radial thickness with increasing distance from the first end and from a second end opposite the first end.4. The rim seal of claim 1 , wherein the seal portion includes a seal surface configured to abut a rim portion of the plurality of rotor blades.5. The rim seal of claim 4 , wherein the seal surface has a cross-sectional shape matching a cross-sectional shape of the rim portion.6. A rotor assembly for a gas turbine engine comprising:a rotor disc;a plurality of rotor blades secured to the rotor disc defining a rim cavity between the rotor disc and a rim portion of the plurality of rotor blades; anda rim seal disposed in the rim cavity ...

SYSTEM AND METHOD FOR REDUCED TURBINE DEGRADATION BY CHEMICAL INJECTION

A gas turbine injection system having a gas turbine with an inlet section, a compressor section, at least one combustor in a combustion section, and a turbine section is disclosed. Air supply piping, water supply piping, and chemical reactant supply piping is in fluid communication with the injection system. A mixing chamber is in fluid communication with at least one of the water supply piping, air supply piping, and the chemical reactant supply piping to produce a chemical mixture. Chemical mixture supply piping is in fluid communication with the mixing chamber and at least one spray nozzle configured to selectively combine the chemical mixture with the air and inject an atomized chemical mixture into at least one section of the turbine. 1. A gas turbine injection system , comprising:a gas turbine having an inlet section, a compressor section, at least one combustor in a combustion section, and a turbine section;air supply piping in fluid communication with a supply of air and at least one spray nozzle;water supply piping in fluid communication with a supply of water;chemical reactant supply piping in fluid communication with the supply of a chemical reactant;a mixing chamber in fluid communication with the water supply piping and the chemical reactant supply piping, the mixing chamber configured to receive water from the water supply piping and the chemical reactant from the chemical reactant supply piping to produce a chemical mixture; andchemical mixture supply piping in fluid communication with the mixing chamber and the at least one spray nozzle, the at least one spray nozzle configured to selectively combine the chemical mixture with the air and inject an atomized chemical mixture into at least one section of the turbine.2. The injection system of claim 1 , further comprising a retractable manifold in fluid communication with at least one retractable spray nozzle claim 1 , the chemical mixture supply piping claim 1 , and the air supply piping.3. The ...

System and Method for Creating A Fluidic Barrier With Vortices From the Upstream Splitter

A turbofan engine has a fan portion in fluid communication with a core stream and a bypass stream of air separated by splitters disposed both upstream and downstream of the fan portion. A fluid passage is defined between the splitters. The turbofan engine has a plurality of surface interruptions on a core side surface of the upstream splitter and a plurality of vortices originating from the surface interruptions, thereby restricting the migration of the core stream into the bypass stream through the fluid passage. 1. A turbofan engine having a fan portion in fluid communication with a core stream and a bypass stream of air; the core stream being:compressed by the fan portion and a core compressor portion, heated and expanded through a core turbine portion;the core turbine portion driving the fan and the compressor portion; the core turbine portion connected to a shaft;the bypass stream being compressed by the fan portion;the core and the bypass streams separated by an upstream splitter and a downstream splitter with the fan portion disposed axially between the upstream and downstream splitters wherein a fluid passage between the core and bypass streams is defined between the splitters; the fan portion having a plurality of blades; each of the blades of the fan portion having a high pressure side and a low pressure side; anda plurality of surface interruptions on a core side surface of the upstream splitter and a plurality of vortices originating from the surface interruptions thereby restricting the migration of the core stream into the bypass stream through the fluid passage.2. The turbofan engine of claim 1 , wherein the plurality of surface interruptions are selected from the group consisting of a plurality of ridges extending into the core stream; a plurality of blades or flaps extending into the core stream; a plurality of grooves recessed into the upstream splitter claim 1 , a plurality of ramps extending into the core stream and a plurality of wedges ...

CMC VANE WITH SUPPORT SPAR AND BAFFLE

A vane includes a ceramic airfoil section that has an airfoil wall defining a leading edge, a trailing edge, a pressure side, and a suction side. The ceramic airfoil section has an internal cavity. A support spar extends through the internal cavity for supporting the ceramic airfoil section. The support spar is spaced from the airfoil wall such that there is a gap there between. The support spar has an internal through-passage that is fluidly isolated from the gap in the ceramic airfoil section. A baffle is disposed in the gap and is spaced apart from the airfoil wall and the support spar so as to divide the gap into a plenum space between the support spar and the baffle and an impingement space between the baffle and the airfoil wall. The baffle has impingement holes directed toward the airfoil wall that connect the plenum space and the impingement space. 1. A vane for a gas turbine engine , comprising:a ceramic airfoil section having an airfoil wall defining a leading edge, a trailing edge, a pressure side, and a suction side, the ceramic airfoil section having an internal cavity;a support spar extending through the internal cavity for supporting the ceramic airfoil section, the support spar being spaced from the airfoil wall such that there is a gap there between, the support spar having an internal through-passage that is fluidly isolated from the gap in the ceramic airfoil section; anda baffle disposed in the gap, the baffle being spaced apart from the airfoil wall and the support spar so as to divide the gap into a plenum space between the support spar and the baffle and an impingement space between the baffle and the airfoil wall, the baffle having impingement holes directed toward the airfoil wall and connecting the plenum space and the impingement space.2. The vane as recited in claim 1 , wherein the baffle circumscribes the structural spar.3. The vane as recited in claim 1 , wherein the baffle defines first and second ends claim 1 , and the second end is ...

HIGH YIELD STRENGTH NICKEL ALLOY WITH AUGMENTED PRECIPITATION HARDENING

An embodiment of a superalloy composition includes 0 to 3.5 wt % Al; 0.005 to 0.05 wt % B; 0.005 to 0.25 wt % C; 10 to 27 wt % Co; 12 to 20 wt % Cr; 6 to 10 wt % Ti; 0.01 to 0.1 wt % Zr; and balance Ni and incidental impurities. 1. A superalloy composition consisting essentially of:0 to 3.5 wt % Al;0.005 to 0.05 wt % B;0.005 to 0.25 wt % C;10 to 27 wt % Co;12 to 20 wt % Cr;6.00 to 10.00 wt % Ti;0.01 to 0.1 wt % Zr; andbalance Ni and incidental impurities.2. (canceled)3. The composition of claim 1 , wherein the composition includes 1.00 wt % Al.4. The composition of claim 1 , wherein the composition includes 0.01 wt % B.5. The composition of claim 1 , wherein the composition includes 0.03 wt % C.6. The composition of claim 1 , wherein the composition includes 26.24 wt % Co.7. The composition of claim 1 , wherein the composition includes 13.43 wt % Cr.8. The composition of claim 1 , wherein the composition includes 7.93 wt % Ti.9. The composition of claim 1 , wherein the composition includes 0.08 wt % Zr.10. A gas turbine engine component formed from an alloy having a composition consisting essentially of:0 to 3.5 wt % Al;0.005 to 0.05 wt % B;0.005 to 0.25 wt % C;10 to 27 wt % Co;12 to 20 wt % Cr;6.00 to 10.00 wt % Ti;0.01 to 0.1 wt % Zr; andbalance Ni and incidental impurities.11. The component of claim 10 , wherein the component is a rotor disk for a compressor section or a turbine section of the gas turbine engine.12. The component of claim 11 , wherein the rotor disk is adapted to be installed in a high pressure compressor section or a high pressure turbine section of the gas turbine engine claim 11 , immediately upstream or immediately downstream of a combustor section.13. (canceled)14. The component of claim 10 , wherein the composition includes 1.00 wt % Al.15. The component of claim 10 , wherein the composition includes 0.01 wt % B.16. The component of claim 10 , wherein the composition includes 0.03 wt % C.17. The component of claim 10 , wherein the ...

Disc Turbine Engine

A disc turbine engine with a multi disc engine where each disc engine includes a turbine blade, a low-pressure compressor blade, a high-pressure compressor blade, and a bearing. Each disc engine runs freely and individually, counter-rotating directions from each other and around a fixed shaft. Each disc engine has its own cooling system. The compressor's blades act as cooling fins for the turbine blade. A coolant, such as liquid hydrogen is used to fill the hollow body of the high-pressure compressor, the hollow body of the lower pressure compressor, and the hollow body of the turbine blade with a connection chamber in between the hollow blades as a sealed system. The nozzle is filled with a coolant inside the hollow bodies of the nozzle and guide fan as a sealed system.

TURBINE MINIDISK BUMPER FOR GAS TURBINE ENGINE

An assembly for a gas turbine engine includes a minidisk that includes an axial extension extending from a disc. The axial extension includes an inner diameter surface and a recess arranged radially opposite the inner diameter surface. The recess provides a radially outwardly extending flange and a bumper extending radially inward from and proud of the inner diameter surface. A method of working on a gas turbine engine section includes inserting a tool into a cavity beneath a seal assembly, and engaging a flange of a minidisk with the tool to manipulate first and second rotors with respect to one another. 1. An assembly for a gas turbine engine comprising:a minidisk including an axial extension extending from a disc, the axial extension including an inner diameter surface and a recess arranged radially opposite the inner diameter surface, the recess providing a radially outwardly extending flange, and a bumper extending radially inward from and proud of the inner diameter surface.2. The assembly according to claim 1 , wherein the recess is axially elongated compared to a radial depth of the recess.3. The assembly according to claim 2 , wherein the recess axially overlaps the bumper.4. The assembly according to claim 1 , comprising a turbine rotor having a hub claim 1 , the minidisk mounted on the hub claim 1 , and the hub operatively supported relative to an engine static structure by a bearing.5. The assembly according to claim 4 , comprising a seal assembly arranged between the minidisk and the bearing to create a bearing compartment.6. The assembly according to claim 5 , comprising a sleeve supported by the hub claim 5 , the bearing mounted to the sleeve.7. The assembly according to claim 6 , wherein the hub includes circumferentially spaced radially outwardly extending first tabs claim 6 , and the axial extension includes circumferentially spaced radially inwardly extending second tabs claim 6 , the first and second tabs axially aligned with one another claim 6 ...

ADVANCED SOLAR THERMALLY DRIVEN POWER SYSTEM AND METHOD

A proposed thermally driven solar power generating system includes a solar air heater receiving an airflow, a gas turbine driving a generator and at least one exhaust compressor, a humidifying air recuperator communicating with the heater. The recuperator includes a water pipeline, product channels, dry working channels, wet working channels connected with the dry channels, having pair-wise heat transfer relations. The airflow passes via the dry and wet channels, a portion thereof moves and expands through the turbine converting into a product stream, drawn into the product channels, cooled and condensed there. Condensate water moves from the product channels and pipeline into the wet channels. The product stream is sucked by the compressor and converted into a product air stream discharged therefrom. In embodiments, it incorporates an air cooler, auxiliary burner chamber, heat accumulator, solar desorber, additional cooler or M-Cooler, heat exchangers communicating with a number of the exhaust compressors. 4. The thermally driven solar power generating system according to claim 3 , wherein: said system is so controlled that is capable of operating in one of the following three modes:{'b': 1', '15, '(a) using only said solar air heater () for heating up said airflow (); or'}{'b': 20', '15, '(b) using only said auxiliary burner chamber () for heating up said airflow (); or'}{'b': 1', '20', '15, '(c) using a combination of said solar air heater () and said auxiliary burner chamber () for heating up said airflow ().'}7. The thermally driven solar power generating system according to claim 1 , further comprising:{'b': 37', '16', '37', '40', '16', '5', '37, 'claim-text': [{'b': 38', '16', '7', '37', '12, 'a condensing channel () passing the product stream () therethrough before further passing thereof into said exhaust compressor (); said additional cooler () is capable of collecting the condensate water (); and'}, {'b': 39', '40', '38, 'a cooling channel () passing the ...

PANELS OF A FAN OF A GAS TURBINE

An aircraft gas turbine with a fan disk on which are fastened fan blades spread over the circumference and forming an intermediate space with one another, with a sealing disk arranged at the rear of the fan disk and with an inlet cone mounted at the front of the fan disk, as well as with filler elements arranged in the intermediate spaces, where the sealing disk has an annular groove and where the filler elements are at the rear inserted into the annular groove and at the front held underneath a rim area of the inlet cone, characterized in that the filler elements are designed as bending beams and that on the radially inner side of the filler element at least one rib-like reinforcing area is provided, extending in the axial direction and longitudinally to the filler element. 1. An aircraft gas turbine with a fan disk on which are fastened fan blades spread over the circumference and forming an intermediate space with one another , with a sealing disk arranged at the rear of the fan disk and with an inlet cone mounted at the front of the fan disk , as well as with filler elements arranged in the intermediate spaces , where the sealing disk has an annular groove and where the filler elements are at the rear inserted into the annular groove and at the front held underneath a rim area of the inlet cone , characterized in that the filler elements are designed as bending beams and that on the radially inner side of the filler element at least one rib-like reinforcing area is provided , extending in the axial direction and longitudinally to the filler element.2. The aircraft gas turbine in accordance with claim 1 , wherein the reinforcing area is designed in one piece with the filler element.3. The aircraft gas turbine in accordance with claim 1 , wherein several reinforcing areas are provided.4. The aircraft gas turbine in accordance with claim 1 , wherein the reinforcing area is designed as a hollow section.5. The aircraft gas turbine in accordance with claim 1 , wherein ...

COOLING APPARATUS FOR A FUEL INJECTOR

An annular air swirler configured to receive a fuel injector in a central bore. The swirler has one or more annular channels, defined by radially facing channel walls and having an inlet for receiving compressed air and an axially distal outlet. The channel walls converge inwardly towards the outlet and swirl vanes extend between opposing faces of the walls. The swirler turns incoming air to create a vortex at the channel outlet. An annular cooling apparatus associated with the air swirler is arranged axially adjacently downstream of the channel outlet(s), and includes a skirt portion radially spaced from a converging portion of the outermost channel wall defining a converging portion of a bowed coolant channel. A radially outwardly extending wall connects with the outermost channel wall and, with a face of the skirt portion, defines a radially outwardly extending portion of the bowed coolant channel adapted for increased heat exchange. 1. An annular air swirler configured to receive a fuel injector in a central bore thereof , the air swirler having one or more annular channels , with each annular channel defined by radially facing channel walls and having an inlet for receiving compressed air and an axially distal outlet , the channel walls converging radially inwardly towards the outlet , swirl vanes extending between opposing faces of the radially facing walls and configured for turning incoming air to create a vortex at the channel outlet , andan annular cooling apparatus associated with the air swirler and arranged axially adjacently downstream of the channel outlet(s), the cooling apparatus comprising a skirt portion radially spaced from a converging portion of the radially outermost channel wall defining a radially converging portion of a bowed coolant channel, and a radially outwardly extending wall connected with the radially outermost channel wall adjacent the outlet and defining with a radially extending face of the skirt portion, a radially outwardly ...

MINICORE COOLING PASSAGE NETWORK HAVING TRIP STRIPS

A gas turbine engine article includes an article wall that defines leading and trailing ends and first and second sides that join the leading and trailing ends. The article wall defines a cavity. A cooling passage network is embedded in the article wall between inner and outer portions of the article wall. The cooling passage network has an inlet orifice through the inner portion of the article wall to receive cooling air from the cavity, a plurality of sub-passages that extend axially from the at least one inlet orifice, at least one outlet orifice through the outer portion of the airfoil wall, and trip strips for mixing cooling air in the cooling passage network. 1. A gas turbine engine article comprising:an article wall defining a cavity; anda cooling passage network embedded in the article wall between inner and outer portions of the article wall, the cooling passage network having an inlet orifice through the inner portion of the airfoil outer wall to receive cooling air from the cavity, a plurality of sub-passages that extend axially from the at least one inlet orifice, at least one outlet orifice through the outer portion of the article wall, and trip strips for mixing cooling air in the cooling passage network.2. The article as recited in claim 1 , wherein at least one of the trip strips is elongated.3. The article as recited in claim 1 , wherein a first one of the trip strips is a partial ring that extends around an interior region greater than 180°.4. The article as recited in claim 3 , wherein one or more of the trip strips are within the interior region.5. The article as recited in claim 4 , wherein one of the trip strips that is within the interior region is a ring.6. The article as recited in claim 1 , wherein the trips strips include a first set of trip strips that have a first common orientation and a second set of trip strips that have a second common orientation that is different than the first common orientation.7. The airfoil as recited in claim ...

SEALING ARRANGEMENT AT THE INTERFACE BETWEEN A COMBUSTOR AND A TURBINE OF A GAS TURBINE AND GAS TURBINE WITH SUCH A SEALING ARRANGEMENT

A sealing arrangement is provided at the interface between a combustor and a turbine of a gas turbine. The turbine includes guiding vanes at its inlet. The guiding vanes are each mounted within the turbine at their outer diameter by means of rear outer diameter vane hook and are each at their inner diameter in sealing engagement by means of a front inner diameter vane tooth with a honeycomb seal arranged at the corresponding inner diameter part of the outlet of said combustor. The rear outer diameter vane hook allows a relative movement of the guiding vane in form of a rotation around the rear outer diameter vane hook. The sealing adapts the front inner diameter vane tooth and the corresponding honeycomb seal in their configuration to the rotating relative movement of the guiding vane, such that the compression of said honeycomb seal through transients of the gas turbine is reduced. 1. A sealing arrangement at the interface between a combustor and a turbine of a gas turbine , said turbine comprising guiding vanes at its inlet , which guiding vanes are each mounted within said turbine at their outer diameter by means of an outer diameter vane hook and are each at their inner diameter in sealing engagement by means of a front inner diameter vane tooth with a honeycomb seal arranged at the corresponding inner diameter part of the outlet of said combustor , whereby said rear outer diameter vane hook allows a relative movement of said guiding vane in form of a rotation around said outer diameter vane hook , wherein said front inner diameter vane tooth and the corresponding honeycomb seal are adapted in their configuration to said rotating relative movement of said guiding vane , such said the compression of said honeycomb seal through transients of the gas turbine is reduced.216. The sealing arrangement as claimed in claim 1 , wherein said front inner diameter vane tooth and the corresponding honeycomb seal are adapted in their shape to said rotating relative movement () ...

FAN BLADE TIP WITH FRANGIBLE STRIP

Disclosed is a fan blade for a gas turbine engine, the fan blade having: an airfoil having a leading edge, a trailing edge, a tip, and a frangible strip connected to the blade tip and extending outwardly therefrom, the frangible strip being less resistant to plastic deformation than the fan blade tip. 1. A fan blade for a gas turbine engine , the fan blade comprising:an airfoil having a leading edge, a trailing edge, a tip, anda frangible strip connected to the blade tip and extending outwardly therefrom, the frangible strip being less resistant to plastic deformation than the fan blade tip.2. The fan blade of wherein the fan blade tip and the frangible strip are connected by a concave mating portion in one of the fan blade tip and the frangible strip and a convex mating portion in another of the fan blade tip and the frangible strip.3. The fan blade of wherein an abradable portion of the frangible strip extends over a pressure side edge and a suction side edge of the fan blade tip and the frangible strip extends in a fan blade chord direction over the fan blade tip.4. The fan blade of wherein one or more dowels fix the frangible strip to the fan blade tip.5. The fan blade of wherein the frangible strip is anisotropic and the abradable portion is less resistant to plastic deformation than the mating portion.6. The fan blade of wherein the frangible strip includes a plurality of honeycomb segments that comprise internal stiffening tubes extending and tapering in one or both of a fan blade span direction and the fan blade chord direction.7. A gas turbine engine comprisinga fan module including a fan case, a fan rotor, a plurality of fan blades connected to the fan rotor, each fan blade comprising:an airfoil having a leading edge, a trailing edge, a tip, anda frangible strip connected to the blade tip and extending outwardly therefrom, the frangible strip being less resistant to plastic deformation than the fan blade tip.8. The engine of wherein the fan blade tip and ...

ADVANCED STATIONARY SEALING COOLED CROSS-SECTION FOR AXIAL RETENTION OF CERAMIC MATRIX COMPOSITE SHROUDS

In one aspect, the present subject matter is directed to a gas turbine sealing assembly that includes a first static gas turbine wall and a second static gas turbine wall. A seal is disposed between the first static gas turbine wall and the second static gas turbine wall. The seal includes a first seal layer defining a first seal layer aperture extending therethrough. A second seal layer defines an elongated slot extending therethrough. The elongated slot includes a first end and a second end. A third seal layer defines a third seal layer aperture extending therethrough. The second seal layer is positioned between the first seal layer and the third seal layer such that the first seal layer aperture is in fluid communication with the first end and the third seal layer aperture is in fluid communication with the second end. 1. A gas turbine sealing assembly , comprising:a first static gas turbine wall;a second static gas turbine wall; and a first seal layer defining a first seal layer aperture extending through the first seal layer;', 'a second seal layer defining an elongated slot extending through the second seal layer; and', 'a third seal layer defining a third seal layer aperture extending through the third seal layer, wherein the second seal layer is positioned between the first seal layer and the third seal layer such that the first seal layer aperture is in fluid communication with a first position in the elongated slot and the third seal layer aperture is in fluid communication with a second position in the elongated slot., 'a seal disposed between the first static gas turbine wall and the second static gas turbine wall, the seal comprising2. The gas turbine sealing assembly of claim 1 , wherein the first static gas turbine wall comprises a turbine shroud assembly mount and the second gas static gas turbine wall comprises a stator vane assembly mount.3. The gas turbine sealing assembly of claim 1 , wherein the first seal layer claim 1 , the second seal layer ...

COUNTER-ROTATING COMPRESSOR

A compressor is described which includes a non-axial first rotor, and a second rotor disposed immediately downstream from the first rotor and being co-axial therewith about a longitudinal axis of rotation. The second rotor rotates in a direction opposite the non-axial first rotor to discharge fluid flow into an uninterrupted passage between an outlet of the second rotor and one of a downstream combustor or a further compression stage. The uninterrupted passage is free of a diffusing passage. A gas turbine engine including such a compressor and a method of compressing fluid flow is also described. 1. A compressor comprising: a non-axial first rotor; and a second rotor disposed immediately downstream from the first rotor and being co-axial therewith about a longitudinal axis of rotation , the second rotor rotating in a direction opposite the non-axial first rotor to discharge fluid flow into an uninterrupted passage between an outlet of the second rotor and one of a downstream combustor or a further compression stage , the uninterrupted passage being free of a diffusing passage.2. The compressor as defined in claim 1 , comprising an annular gap between rotor blades of the non-axial first rotor and rotor blades of the second rotor.3. The compressor as defined in claim 2 , comprising a curved fluid conduit located in the annular gap claim 2 , the curved fluid conduit turning fluid flow discharged from the rotor blades of the non-axial first rotor to reduce a radial direction flow component therefrom.4. The compressor as defined in claim 3 , wherein the rotor blades of the second rotor extend into the fluid conduit.5. The compressor as defined in claim 1 , wherein rotation of the second rotor turns fluid flow discharged from the non-axial first rotor in a substantially axial direction.6. The compressor as defined in claim 1 , wherein the non-axial first rotor rotates about the longitudinal axis at a first rotational speed and the second rotor rotates about the ...

CARRIER RING

The invention relates to a carrier ring for a high-pressure gas turbine of a gas turbine plant, which has a high-pressure combustion chamber upstream of the high-pressure gas turbine, a compressor upstream of the high-pressure combustion chamber, a low-pressure combustion chamber downstream of the high-pressure gas turbine, a low-pressure gas turbine downstream of the low-pressure combustion chamber, and a rotor that carries rotor blades for the compressor, for the high-pressure gas turbine, and for the low-pressure gas turbine. The carrier ring carries guide blades and/or heat shields of the high-pressure gas turbine and can be fastened to the high-pressure combustion chamber. The installation of the carrier ring can be simplified by segmenting the carrier ring at least in the area of the guide blades thereof and/or of the heat shields thereof in the circumferential direction, wherein the segmented carrier ring has at least two ring segments that carry the guide blades and/or the heat shields. 1. A carrier ring for a gas turbine installation , which gas turbine installation consists of at least one compressor , at least one combustion chamber operated downstream of the compressor and at least one turbine located downstream of the combustion chamber and acted upon by hot gases from the combustion chamber , wherein the rotor blades of the compressor and of the turbine are arranged on a common rotor shaft , said carrier ring comprising: the guide vanes and/or the heat shields of the turbine and is attached to the combustion chamber , and in that the carrier ring is segmented circumferentially at least in the region of its guide vanes and/or of its heat shields and has at least two ring segments which carry the guide vanes and/or the heat shields.2. The carrier ring as claimed in claim 1 , wherein the gas turbine installation consists essentially of a compressor unit consisting of at least one compressor claim 1 , a first combustion chamber operating downstream of the ...

STRUCTURE FOR COOLING TURBINE BLADES AND TURBINE AND GAS TURBINE INCLUDING THE SAME

A structure for cooling turbine blades, and a turbine and gas turbine including the same, enhance efficiency in cooling turbine blades by improving the structures of a turbine disk and a retainer for securing a turbine blade. The structure includes a turbine blade connected to a turbine blade root; a turbine disk including a slot for receiving the turbine blade root, a cooling passage through which cooling air flows to the turbine blade root, and a branch passage communicating at one end with the cooling passage; and a retainer fixed to the turbine disk on each of opposite sides of the turbine blade to prevent separation of the turbine blade from the turbine disk, the fixed retainer having a chamber communicating with the other end of the branch passage so that the cooling air of the cooling passage is introduced into the chamber through the branch passage. 1. A structure for cooling turbine blades , comprising:a turbine blade connected to a turbine blade root;a turbine disk including a slot for receiving the turbine blade root, a cooling passage through which cooling air flows to the turbine blade root, and a branch passage communicating at one end with the cooling passage; anda retainer fixed to the turbine disk on each of opposite sides of the turbine blade to prevent separation of the turbine blade from the turbine disk, the fixed retainer having a chamber communicating with the other end of the branch passage so that the cooling air of the cooling passage is introduced into the chamber through the branch passage.2. The structure according to claim 1 , wherein the cooling air introduced into the chamber flows between the retainer and the turbine blade.3. The structure according to claim 1 , wherein the cooling air introduced into the chamber flows between the retainer and the turbine blade root.4. The structure according to claim 1 , wherein the chamber is divided into a plurality of chamber partitions communicating with each other claim 1 , at least one chamber ...

FUEL DRIVEN NEAR ISOTHERMAL COMPRESSOR

A gas compressor system includes a compression liquid holding tank in fluid communication with a combustion tank. A combustible fluid is directed to the combustion tank. An ignition system is provided for igniting the combustible fluid. A compression liquid flows between the liquid holding tank, the combustion tank, and a compression tank. A compressible gas is provided in the compression tank. The ignition of the combustible fluid drives the compression liquid from the combustion tank to the compression tank, compressing the compressible liquid. An HVAC&R system and a method of compressing gas are also disclosed. 1. A gas compressor system , comprising:a compression liquid holding tank;a combustion tank having a compression liquid inlet in fluid communication with the compression liquid holding tank, a combustible fluid inlet for fluid communication with a combustible fluid source, an ignition system for igniting the combustible fluid, and a compression liquid flow opening;a compression liquid;a pump for pumping the compression liquid from the compression liquid holding tank to the combustion tank;a compression tank having a compression liquid flow opening, the compression liquid flow opening of the compression tank being in liquid communication with the compression liquid flow opening of the combustion tank;a valve for controlling flow of the compression liquid between the compression liquid holding tank and the combustion tank;a valve for controlling fluid communication between the combustion tank and the compression tank;a valve for controlling the flow of combustible fluid with the combustion tank;a compressible gas;wherein, the combustible fluid is flowed into the combustion tank and compressible gas is provided in the compression tank, the compression liquid is pumped by the pump from the compression liquid holding tank to the combustion tank, compressing the combustible fluid, the combustible fluid is ignited by the ignition system causing the compression ...

TURBINE BLADE, TURBINE INCLUDING SAME TURBINE BLADE, AND GAS TURBINE INCLUDING SAME TURBINE

A turbine blade is provided with a rotatable leading edge to reduce pressure loss as a flow direction of combustion gas varies. The turbine blade includes an airfoil including as separate bodies a trailing-edge portion and a leading-edge portion being linked to the trailing-edge portion and disposed upstream of the trailing-edge portion, the leading-edge portion including a front surface arranged on an upstream side of the leading-edge portion; and a rotary unit connected to the leading-edge portion and configured to rotate the leading-edge portion according to an inflow angle of the combustion gas such that the front surface faces a flow of the combustion gas. A barrier wall extends from a side surface of the leading-edge portion toward an open end of the trailing edge portion, and when the leading-edge portion is rotated, the bather wall portion prevents formation of a gap between the side surface and the open end. 1. A turbine blade mounted on a turbine disk in a turbine casing of a turbine , the turbine blade rotating the turbine when combustion gas flows in the turbine casing , the turbine blade comprising:an airfoil including as separate bodies a trailing-edge portion and a leading-edge portion being linked to the trailing-edge portion and disposed upstream of the trailing-edge portion, the leading-edge portion including a front surface arranged on an upstream side of the leading-edge portion; anda rotary unit connected to the leading-edge portion and configured to rotate the leading-edge portion according to an inflow angle of the combustion gas such that the front surface faces a flow of the combustion gas.2. The turbine blade according to claim 1 , wherein the trailing-edge portion includes an open end linked to the leading-edge portion and a closed end having a streamlined shape.3. The turbine blade according to claim 2 , further comprising:a barrier wall portion extending from a side surface of the leading-edge portion toward the open end of the trailing ...

COMPRESSOR ENDWALL TREATMENT TO DELAY COMPRESSOR STALL

A compressor is provided including a casing, a hub, a flowpath, a plurality of blades defining a plurality of axially extending compressor stages and an endwall treatment formed in the casing on at least two downstream most stages of the plurality of compressor stages. The remaining stages of the plurality of compressor stages located upstream of the at least two downstream most stage are devoid of any endwall treatment. Each of the endwall treatments faces a tip of each blade in the at least two downstream most stages. The tip of each blade and the endwall treatment are configured to move relative to each other. The endwall treatment formed in the casing on at least two downstream most stages of the plurality of compressor stages is configured to extend a stall margin to delay stall due to ice ingestion. A method and engine application are disclosed. 1. A compressor comprising:a casing;a hub;a flow path formed between the casing and the hub;a plurality of blades positioned in the flow path and defining a plurality of axially extending compressor stages; andan endwall treatment formed in the casing on at least two downstream most stages of the plurality of compressor stages, each of the endwall treatments facing a tip of each blade in the at least two downstream most stages, wherein the tip of each blade and the endwall treatment are configured to move relative to each other,wherein the remaining stages of the plurality of compressor stages located upstream of the at least two downstream most stages are devoid of any endwall treatment, andwherein the endwall treatments are configured to extend a stall margin to delay stall due to ice ingestion.2. The compressor of claim 1 , wherein each of the endwall treatments include a geometrical modification of the casing.3. The compressor of claim 2 , wherein each of the endwall treatments is formed into an interior surface of the casing and disposed circumferentially thereabout.4. The compressor of claim 2 , wherein each of ...

FUEL INJECTOR FOR A GAS TURBINE ENGINE COMBUSTION CHAMBER

A fuel injector includes a central member arranged on axis of fuel injector and a first member arranged around central member. A first swirler is arranged between central member and first member. A pilot fuel injector is arranged to supply fuel onto inner surface of first member. A second member is arranged between downstream end of first member and downstream end of shroud. A second swirler is arranged between upstream end of first member and upstream end of shroud and between downstream end of first member and upstream end of second member. A third swirler is arranged between downstream end of shroud and second member. A main fuel injector is arranged to supply fuel into an annular passage between first member and second member and a fourth swirler extends through first member. 1. A fuel injector comprising a pilot fuel injector , a main fuel injector and a plurality of air swirlers , wherein a shroud is arranged around the pilot fuel injector , the main fuel injector and the plurality of air swirlers , the fuel injector has an axis , the shroud has a radially inner surface , a central member arranged on the axis of the fuel injector , a first member arranged coaxially around the central member , a first air swirler arranged radially between the central member and the first member , the pilot fuel injector arranged within the first member to supply fuel into a passage at least defined by the first member , a second member arranged coaxially between a downstream portion of the first member and a downstream portion of the shroud , a second air swirler arranged radially between an upstream portion of the first member and an upstream portion of the shroud and radially between a downstream portion of the first member and the second member , the second air swirler comprising a plurality of circumferentially spaced vanes , each of the vanes of the second air swirler extending radially from the upstream portion of the first member to the upstream portion of the shroud and ...

GAS TURBINE PLANT AND EXHAUST CARBON DIOXIDE RECOVERY METHOD THEREFOR

A gas turbine plant includes an exhaust line, a carbon dioxide recovery device configured to recover carbon dioxide contained in an exhaust gas, a circulation line connected to a gas turbine, a first valve device, a bypass line bypassing the carbon dioxide recovery device, a second valve device provided on the bypass line, a third valve device provided at a position between the bypass line and the carbon dioxide recovery device, a densitometer configured to detect a carbon dioxide concentration in the exhaust gas, and a control device configured to adjust opening degrees of the first valve device, the second valve device, and the third valve device based on an operation state of the gas turbine and the carbon dioxide concentration.

SYSTEM AND METHOD FOR PURGING FUEL FROM TURBOMACHINE

Systems and methods include one or more fluid lines configured to flow a fluid in a first direction in the gas turbine system. The systems and methods also include an eductor configured to reverse flow of the fluid in the one or more fluid lines during a reverse purge of the gas turbine system. 1. A gas turbine system , comprising:one or more fluid lines configured to flow a fluid in a first direction in the gas turbine system; andan eductor configured to reverse flow of the fluid in the one or more fluid lines during a reverse purge of the gas turbine system.2. The gas turbine system of claim 1 , wherein the fluid comprises a coolant.3. The gas turbine system of claim 2 , wherein the coolant comprises water.4. The gas turbine system of claim 1 , comprising:a shaft;a compressor coupled to the shaft and configured to compress air for an air and fuel mixture; anda combustor coupled to the shaft and configured to consume fuel in generating rotational energy on the shaft.5. The gas turbine system of claim 4 , wherein the fluid comprises coolant that is injected into air prior to compression in the compressor.6. The gas turbine system of comprising a cooling system configured to cool the liquid fuel when reverse purged from the gas turbine system.7. The gas turbine system of claim 4 , wherein the fluid comprises liquid fuel to be consumed in the combustor.8. The gas turbine system of claim 1 , wherein the one or more fluid lines comprises:one or more relatively low pressure lines that use a relatively high amount of urging from the eductor; andone or more relatively high pressure lines that use at least a portion of pressure in the one or more relatively high pressure lines to urge the fluid back through the one or more relatively high pressure lines.9. A gas turbine system claim 1 , comprising:a shaft;a compressor coupled to the shaft and configured to compress ambient air for a mixture of air and liquid fuel;a combustor configured to receive the mixture from the ...

ELECTRONIC FUEL CONTROL FOR GAS TURBINE ENGINES

A fuel injector for a gas turbine engine includes a feed arm defining a conduit extending between an inlet end and an outlet end and a plunger. The plunger is disposed within the conduit and is movable between a plunger first position and a plunger second position. A flow area defined between the plunger and the feed arm is smaller in the plunger first position than in the plunger second position to bias fuel flow through the fuel injector. Fuel systems, gas turbine engines, and methods of controlling fuel flow in gas turbine engine fuel systems are also described. 1. A fuel injector for a gas turbine engine , comprising:a feed arm defining a conduit extending an inlet end and an outlet end; anda plunger disposed within the conduit and movable between a plunger first position and a plunger second position,wherein the plunger and conduit define between one another a flow area that is smaller in the plunger second position than in the plunger first position.2. The fuel injector as recited in claim 1 , further comprising an electrical actuator operably connected to the plunger and arranged to move the plunger relative to the feed arm.3. The fuel injector as recited in claim 2 , wherein the electrical actuator is appurtenant to the fuel injector.4. The fuel injector as recited in claim 2 , wherein the electrical actuator is severable from the fuel injector.5. The fuel injector as recited in claim 2 , further comprising a rotary drive coupling the electrical actuator to the plunger and arranged to rotate the plunger relative to the feed arm.6. The fuel injector as recited in claim 2 , wherein the electrical actuator includes a solenoid claim 2 , a rotary motor claim 2 , or a stepper motor.7. The fuel injector as recited in claim 1 , wherein the feed arm has female threads engagable with the plunger.8. The fuel injector as recited in claim 7 , wherein the plunger has male threads engagable with the female threads of the feed arm.9. The fuel injector as recited in claim 7 ...

HIGH OVERALL PRESSURE RATIO GAS TURBINE ENGINE

A gas turbine engine includes a compressor section and turbine section. The gas turbine engine of the present disclosure is a relatively small gas turbine engine, configured to generate less than 2,000 horsepower during peak operations. The compressor section defines an overall compressor ratio of at least 15:1 in order to increase in overall efficiency of the gas turbine engine. 1. A gas turbine engine defining an axial direction and a radial direction , the gas turbine engine comprising:a stage of variable inlet guide vanes;a compressor section comprising a compressor located downstream of the stage of variable inlet guide vanes and defining an overall compressor ratio between 15:1 and 16:1 during operation of the gas turbine engine; anda turbine section located downstream of the compressor section, the turbine section comprising a high pressure turbine and a low pressure turbine, the high pressure turbine comprising two stages of turbine rotor blades;wherein the gas turbine engine generates less than 2,000 horsepower during peak operation; andwherein the compressor of the compressor section comprises between three stages of radially oriented compressor rotor blades and four stages of radially oriented compressor rotor blades, and one stage of centrifugal compressor rotor blades.2. The gas turbine engine of claim 1 , further comprising:a first spool coupling the compressor to the high pressure turbine; anda second spool coupled to the low pressure turbine.3. The gas turbine engine of claim 2 , wherein the second spool is mechanically coupled to a drive shaft.4. The gas turbine engine of claim 3 , further comprising:a gearbox, wherein the second spool is mechanically coupled to the drive shaft through the gearbox, andwherein the gas turbine engine has an airflow of 10.5 pounds per second or less during peak operation.5. The gas turbine engine of claim 1 , wherein the compressor section further comprises a stage of variable stator vanes.6. The gas turbine engine of ...

Systems and Methods for Wheel Space Temperature Management

The present application provides a gas turbine engine intended to be used in part in hot ambient conditions. The gas turbine engine may include a compressor, a turbine with a wheel space adjacent to a number of rotor wheels, and a wheel space water cooling system in communication with the wheel space to provide a flow of water thereto. 1. A gas turbine engine , comprising:a compressor;a turbine;the turbine comprising a wheel space adjacent to a plurality of rotor wheels; anda wheel space water cooling system in communication with the wheel space to provide a flow of water thereto.2. The gas turbine engine of claim 1 , wherein the compressor comprises a compressor discharge case with one or more bore holes therethrough.3. The gas turbine engine of claim 2 , wherein the wheel space water cooling system comprises a water line extending through one of the one or more bore holes in whole or in part.4. The gas turbine engine of claim 3 , wherein the water line comprises a concentric line.5. The gas turbine engine of claim 2 , wherein the one or more bore holes in the compressor discharge case are plugged to prevent a flow of air to the wheel space.6. The gas turbine engine of claim 3 , wherein the water line comprises a thermal barrier coating.7. The gas turbine engine of claim 1 , wherein the wheel space water cooling system comprises a source of water.8. The gas turbine engine of claim 1 , wherein the wheel space water cooling system comprises a pump.9. The gas turbine engine of claim 1 , wherein the wheel space cooling system comprises a discharge nozzle.10. The gas turbine engine of claim 9 , wherein the discharge nozzle comprises a counter-flow position claim 9 , a cross-flow position claim 9 , or an angled position.11. The gas turbine engine of claim 1 , wherein the flow of water comprises a flow of water in a liquid state and/or a flow of water in a gaseous state.12. The gas turbine engine of claim 11 , wherein the flow of water comprises the gaseous state during ...

POWER GENERATION SYSTEM AND METHOD OF STOPPING POWER GENERATION SYSTEM

A gas turbine including a compressor and a combustor, an SOFC including an air electrode (cathode) and a fuel electrode (anode), a first compressed air supply line adapted to supply a compressed air compressed by the compressor to the combustor, a second compressed air gas supply line adapted to supply a part of a compressed air compressed by the compressor to the air electrode (cathode), a first fuel gas supply line adapted to supply a fuel gas to the combustor, a second fuel gas supply line adapted to supply a fuel gas to the fuel electrode (anode), a fuel gas recirculation line adapted to return an exhausted fuel gas discharged from the fuel electrode (anode) to the fuel electrode (anode), a cooler provided in the fuel gas recirculation line are provided. 1. A power generation system comprising:a gas turbine including a compressor and a combustor;a fuel cell including an cathode and a anode;a first compressed oxidant supply line adapted to supply a compressed oxidant compressed by the compressor to the combustor;a second compressed oxidant supply line adapted to supply a part of the compressed oxidant compressed by the compressor to the cathode;a first fuel gas supply line adapted to supply a fuel gas to the combustor;a second fuel gas supply line adapted to supply a fuel gas to the anode;a fuel gas recirculation line adapted to return an exhausted fuel gas discharged from the anode to the anode;a cooler provided in the fuel gas recirculation line; anda control unit adapted to operate the cooler when the control unit has stopped the fuel cell and has cut the compressor and the second compressed oxidant supply line.2. The power generation system according to claim 1 , wherein a recirculation blower is provided in the fuel gas recirculation line claim 1 , and the control unit operates the cooler at the same time as the control unit causes the fuel cell stop.3. The power generation unit according to claim 1 , wherein a reductant supply line adapted to supply a ...

METHOD OF DETECTING FLAMEOUT IN A COMBUSTOR AND TURBINE SYSTEM

The method allows to detect flameout in a combustor of a turbine system; it includes the steps of: A) measuring angular acceleration of a shaft of the or each turbine of the turbine system, B) calculating a flameout indicator as a function of the angular acceleration, and C) carrying out a comparison between the flameout indicator and at least one threshold. 1. A method of detecting flameout in a combustor of a turbine system , wherein the turbine system comprises a compressor upstream of said combustor and a first turbine downstream of said combustor , the method comprising the steps of:A) measuring a first angular acceleration of a first shaft of said first turbine, and measuring a second angular acceleration of a second shaft of a second turbine downstream of the first turbine, the first shaft being mechanically disconnected from the second shaft;B) calculating a flameout indicator as a function of said first angular acceleration and said second angular acceleration,C) carrying out a comparison between said flameout indicator and a threshold, andD) tripping said turbine system, in response to said flameout indicator exceeding the threshold in the comparison.2. The method of claim 1 , wherein steps A claim 1 , B claim 1 , and C are cyclically repeated claim 1 , and wherein a set of consecutive incidents of said flameout indicator exceeding the threshold in of said comparison indicates flameout.3. The method of claim 1 , further comprising calculating said threshold during operation of the turbine system.4. The method of claim 1 , further comprising a step E of signaling an alarm claim 1 , to be performed after step C or step D.5. The method of claim 1 ,wherein the calculating at step B said flameout indicator further includes calculating as a function of a pressure measured at an outlet of said compressor.6. The method of claim 1 , wherein the calculating at step B of said flameout indicator further includes calculating as a function of a thermal power generated ...

MULTI-PIECE TURBINE AIRFOIL

A gas turbine engine includes a compressor, a combustor fluidly connected to the compressor via a flow path, and a turbine fluidly connected to the combustor via the flow path. At least one multi-piece structure is disposed within the flow path such that the multi-piece structure at least partially radially spans the flow path. The at least one multi-piece structure includes a fore portion defining a leading edge, a consumable aft portion defining a trailing edge, a pressure surface at least partially defined by a first surface of the fore portion and a first surface of the consumable aft portion, and a suction surface at least partially defined by a second surface of the fore portion and a second surface of the consumable aft portion. 1. A gas turbine engine comprising:a compressor;a combustor fluidly connected to the compressor via a flow path;a turbine fluidly connected to the combustor via the flow path; a fore portion defining a leading edge;', 'a consumable aft portion defining a trailing edge;', 'a pressure surface at least partially defined by a first surface of said fore portion and a first surface of said consumable aft portion;', 'a suction surface at least partially defined by a second surface of said fore portion and a second surface of said consumable aft portion., 'at least one multi-piece structure disposed within said flow path such that the multi-piece structure at least partially radially spans the flow path, and wherein the at least one multi-piece structure includes2. The gas turbine engine of claim 1 , further comprising an axial gap defined between said fore portion and said consumable aft portion.3. The gas turbine engine of claim 1 , wherein the fore portion has at least one internal cooling cavity.4. The gas turbine engine of claim 1 , wherein the consumable aft portion is a solid component.5. The gas turbine engine of claim 1 , wherein the fore portion has a first replacement frequency claim 1 , the consumable aft portion has a second ...

FIRE MITIGATION SYSTEM FOR GAS TURBINE ENGINE

The described gas turbine engine includes a fire mitigation system in fluid system thereof, and includes an air-introduction component located in a fluid conveying conduit upstream of a liquid pump of the fluid system. The air-introduction component has a sacrificial element which remains in place during normal operation of the gas turbine engine but has a heat-induced failure point lower than tat of a remainder of the fluid conveying conduit such that it fails when exposed to a threshold temperature greater than the heat-induced failure point. Failure of the sacrificial element allows air introduction into the fluid conveying conduit upstream of the liquid pump, thereby starving the liquid pump in the event of a fire condition generating the threshold temperature. 1. A gas turbine engine comprising:a fluid system including at least one liquid pump providing motive flow of a liquid through the fluid system; anda fire mitigation system including an air-introduction component located in a fluid conveying conduit upstream of the at least one liquid pump, the air-introduction component having a sacrificial element which remains in place during normal operation of the gas turbine engine, the sacrificial element having a heat-induced failure point lower than that of a remainder of the fluid conveying conduit, the sacrificial element configured to fail when exposed to a threshold temperature greater than the heat-induced failure point to allow air entry into the fluid conveying conduit upstream of the liquid pump, thereby starving the liquid pump in the event of a fire condition generating said threshold temperature.2. The gas turbine engine as defined in claim 1 , wherein the sacrificial element and a remainder of the fluid conveying conduit are formed of different materials.3. The gas turbine engine as defined in claim 2 , wherein the sacrificial element has a lower melting point than the remainder of fluid conveying conduit.4. The gas turbine engine as defined in claim ...

SYSTEM AND METHOD FOR PROTECTING COMPONENTS IN A GAS TURBINE ENGINE WITH EXHAUST GAS RECIRCULATION

A system includes a gas turbine engine that includes a turbine section having one or more turbine stages between an upstream end and a downstream end, an exhaust section disposed downstream from the downstream end of the turbine section, and a fluid supply system coupled to the exhaust section. The fluid supply system is configured to route an inert gas to the exhaust section. 1. A system , comprising: a turbine section having one or more turbine stages between an upstream end and a downstream end;', 'an exhaust section disposed downstream from the downstream end of the turbine section; and', 'a fluid supply system coupled to the exhaust section, wherein the fluid supply system is configured to route an inert gas to the exhaust section., 'a gas turbine engine, comprising2. The system of claim 1 , wherein the exhaust section comprises:an exhaust passage in fluid communication with the turbine section; andan inert gas passage coupled to the fluid supply system.3. The system of claim 2 , wherein the inert gas passage is fluidly coupled to the exhaust passage.4. The system of claim 3 , wherein the exhaust section comprises a wall disposed along the exhaust passage claim 3 , and the inert gas passage is fluidly coupled to the exhaust passage through a plurality of openings in the wall.5. The system of claim 2 , wherein the inert gas passage extends through at least one of an outer shroud cavity surrounding the exhaust passage claim 2 , an inner shroud cavity surrounded by the exhaust passage claim 2 , a vane protruding into the exhaust passage claim 2 , a bearing cavity having a bearing assembly claim 2 , or a combination thereof.6. The system of claim 2 , wherein the exhaust section comprises:an outer shroud extending circumferentially about the exhaust passage; anda casing extending circumferentially about the outer shroud to define an outer shroud cavity, wherein the inert gas passage extends through the outer shroud cavity.7. The system of claim 1 , wherein the gas ...

AFT FRAME ASSEMBLY FOR GAS TURBINE TRANSITION PIECE

An aft frame assembly has a main body with an upstream facing surface, a downstream facing surface, a radially outer facing surface and a radially inner facing surface. Feed hole inlets are located on the upstream facing surface and radially outward of the outer sleeve so that the feed hole inlets are located to receive input from a high pressure plenum. The feed hole inlets are coupled to cooling channels that pass through the main body. Microchannels are formed in or near the radially inner facing surface and the downstream facing surface. The cooling channels are connected to and terminate in the microchannels. Exit holes are connected to the plurality of microchannels, and the exit holes are located radially outward of the transition piece and radially inward of the outer sleeve. The exit holes are located to exhaust into the cooling annulus. 1. An aft frame assembly for a transition piece of a gas turbine , the transition piece is located within an outer sleeve having a plurality of cooling holes , a cooling annulus is formed in a space between the transition piece and the outer sleeve , a high pressure plenum surrounds an exterior of the outer sleeve , the aft frame assembly comprising:a main body comprising an upstream facing surface, a downstream facing surface, a radially outer facing surface and a radially inner facing surface;a plurality of feed hole inlets are located on the upstream facing surface and radially outward of the outer sleeve so that the feed hole inlets are located to receive input from the high pressure plenum, the feed hole inlets coupled to a plurality of cooling channels passing through the main body;a plurality of microchannels are formed in or near the radially inner facing surface and the downstream facing surface, the cooling channels are connected to and terminate in the microchannels;a plurality of exit holes are connected to the plurality of microchannels, the exit holes are located radially outward of the transition piece and ...

BRACKET FOR A TUBE OF AN ENGINE IN A SPACE-LIMITED COMPARTMENT

Aspects are directed to a bracket comprising: a first portion corresponding to a J-blade, a second portion integral with the first portion, the second portion corresponding to an L-bracket, and a third portion integral with the second portion, the third portion configured to rotate about a fold-line that coincides with an interface between the second portion and the third portion. Aspects are directed to a system comprising: a tube, a first housing that includes a first flange, a second housing that includes a second flange, the second flange abutting the first flange, and a bracket that includes a first portion corresponding to a J-blade that at least partially seats the tube, and a second portion integral with the first portion, the second portion corresponding to an L-bracket, and the second portion abutting the second flange. 1. A bracket comprising:a first portion corresponding to a J-blade;a second portion integral with the first portion, the second portion corresponding to an L-bracket; anda third portion integral with the second portion, the third portion configured to rotate about a that coincides with an interface between the second portion and the third portion.2. The bracket of claim 1 , wherein the bracket is made of at least one of steel or nickel.3. The bracket of claim 1 , wherein the first portion includes a sub-portion that has a radius of curvature within a range of 0.1 to 10 millimeters.4. The bracket of claim 1 , wherein the second portion comprises:a first sub-portion that is integral with the first portion;a second sub-portion that is integral with the third portion; anda curved third sub-portion that spans the first sub-portion and the second sub-portion.5. The bracket of claim 1 , wherein the first portion seats a tube claim 1 , and wherein the second portion is coupled to a flange of a housing claim 1 , and wherein the third portion is a washer tab that provides for anti-rotation with respect to a nut.6. A system comprising:a tube;a first ...

FUEL NOZZLE WALL SPACER FOR GAS TURBINE ENGINE

A fuel nozzle configured to channel fluid towards a combustion chamber defined within a gas turbine engine is provided. The fuel nozzle includes a first hollow tube and a second hollow tube concentrically aligned with the first hollow tube and defining a gap therebetween. The first hollow tube has a central passageway configured to channel fuel therethrough. The second hollow tube is typically in contact with compressor discharge gases and is therefore at a higher temperature than the first hollow tube. Thus, the fuel nozzle includes at least one detached or free spacer retained within the gap so as to minimize heat transfer between the first and second hollow tubes. Accordingly, the detached spacer(s) is un-joined or free within the gap where thermal energy transfer is disadvantageous. 1. A fuel nozzle for channeling fluid towards a combustion chamber defined within a gas turbine engine , the fuel nozzle comprising:a first hollow tube comprising a central passageway configured to channel fuel therethrough;a second hollow tube configured with the first hollow tube and defining a gap therebetween, the second hollow tube at a higher temperature than the first hollow tube; andat least one detached spacer retained within the gap so as to minimize heat transfer between the first and second hollow tubes.2. The fuel nozzle of claim 1 , wherein the first and second hollow tubes are concentrically aligned.3. The fuel nozzle of claim 2 , wherein the first and second hollow tubes are oriented substantially linearly claim 2 , the detached spacer configured to maintain linear separation between the first and second hollow tubes.4. The fuel nozzle of claim 1 , wherein the at least one spacer is free within the gap such that the spacer is not joined to the first and second hollow tubes.5. The fuel nozzle of claim 1 , further comprising a plurality of spacers configured within the gap between the first and second hollow tubes.6. The fuel nozzle of claim 5 , wherein the plurality of ...

Compact Compressor

Methods, apparatus, systems and articles of manufacture for compact compressors are disclosed including a gas turbine engine defining an axial direction and a radial direction, the gas turbine engine including an axial flow compressor and a radial flow compressor, wherein the axial flow compressor is located axially forward of the radial flow compressor, a blade assembly including a splitter shroud to divide incoming air into axial air flow for the axial flow compressor and radial air flow for the radial flow compressor, the blade assembly rotating relative to the axial flow compressor and counter-rotating relative to the radial flow compressor, and wherein the blade assembly is located axially aft of the radial flow compressor. 1. A gas turbine engine defining an axial direction and a radial direction , an axial flow compressor;', 'a radial flow compressor, wherein the axial flow compressor is located axially forward of the radial flow compressor; and', 'a blade assembly including a splitter shroud to divide incoming air into axial air flow for the axial flow compressor and radial air flow for the radial flow compressor, the blade assembly rotating relative to the axial flow compressor and counter-rotating relative to the radial flow compressor, wherein the blade assembly is located axially aft of the radial flow compressor., 'the gas turbine engine comprising2. The gas turbine engine of claim 1 , wherein outer airfoils of the blade assembly define a fan disposed in a bypass flow passage of the gas turbine engine claim 1 , and wherein inner airfoils of the blade assembly are disposed within a primary flow passage of the gas turbine engine common to the inner airfoils claim 1 , the axial flow compressor claim 1 , and the radial flow compressor.3. The gas turbine engine of claim 1 , wherein the radial flow compressor is a centrifugal compressor including an impeller and at least one diffuser passage.4. The gas turbine engine of claim 3 , wherein the impeller includes ...

AFTERBURNER SYSTEM

A turbofan engine has: an engine core producing a core exhaust flow; an upstream fan; a bypass duct carrying a bypass air flow produced by the fan; an exhaust assembly. The turbofan engine afterburner system has: a core plug; second core flow duct, and entrance(s) from the first core flow duct for diversion into the second of a portion of the core exhaust flow, and exit for re-joining the diverted portion of the core exhaust flow to the undiverted portion of the core exhaust flow; door(s) open during normal operation of the engine but closable, during a reheat operation, to block the entrances to the second core flow duct; and fuel injector(s) and igniter(s) operable, during the reheat operation, to inject and ignite fuel within the second core flow duct, providing a reheat flow into the core exhaust flow from the exit of the second core flow duct. 1. A turbofan engine having:an engine core comprising a compressor, a combustor, a turbine, and a core shaft connecting the turbine to the compressor, and producing a core exhaust flow;a fan upstream of the engine core;a bypass duct surrounding the engine core and carrying a bypass air flow produced by the fan;an exhaust assembly comprising a first exhaust nozzle defining a downstream end of a first core flow duct which receives the core exhaust flow produced by the engine core; and a core plug defining a radially inner surface of the first core flow duct;', 'a second core flow duct located in the core plug, and having one or more entrances from the first core flow duct for diversion into the second core flow duct of a portion of the core exhaust flow, and further having an exit therefrom for re-joining the diverted portion of the core exhaust flow to the undiverted portion of the core exhaust flow;', 'one or more respective doors for the entrances to the second core flow duct, the doors being open during normal operation of the engine but being closable, during a reheat operation, to at least partially block the ...

PRESSURE SENSOR SYSTEM FOR CALCULATING COMPRESSOR MASS FLOW RATE USING SENSORS AT PLENUM AND COMPRESSOR ENTRANCE PLANE

A pressure sensor system for a compressor including an inlet bellmouth is disclosed. The system includes a first static pressure sensor positioned within a plane of a plenum that is upstream of the inlet bellmouth; and a second static pressure sensor positioned at an entrance plane of the compressor. A mass flow rate calculator may calculate a mass flow rate based on a pressure differential between the plane of the plenum and the entrance plane of the compressor. 1. A pressure sensor system for a compressor including an inlet bellmouth , the system comprising:a first static pressure sensor positioned within a plane of a plenum that is upstream of the inlet bellmouth; anda second static pressure sensor positioned at an entrance plane of the compressor.2. The pressure sensor system of claim 1 , wherein the first static pressure sensor includes a plurality of sensors that are substantially non-uniformly arranged about the plane of the plenum.3. The pressure sensor system of claim 1 , wherein the first static pressure sensor includes a plurality of sensors that are substantially uniformly arranged about the plane of the plenum.4. The pressure sensor system of claim 1 , wherein the first static pressure sensor includes a plurality of sensors and a position of each of the plurality of sensors about the plane of the plenum is based on a computational fluid dynamic simulation.5. The pressure sensor system of claim 1 , wherein the first static pressure sensor includes a plurality of sensors and each sensor includes a differential type sensor.6. The pressure sensor system of claim 1 , further comprising a calculator for calculating a mass flow rate of a flow through the compressor inlet based on a pressure differential between the plane of the plenum and the entrance plane of the compressor.8. The pressure sensor of claim 1 , wherein the compressor further includes an inlet guide vane claim 1 , and the entrance plane of the compressor is substantially aligned with an entrance ...

Compressor and gas turbine

This compressor is provided with: a rotor casing that encircles a rotor, which rotates around a rotational axis; an air bleed chamber casing that is provided to the outer peripheral side of the rotor casing and demarcates an air bleed chamber interconnecting to a primary duct via a slot; and an air bleed tube that is connected to the air bleed chamber casing from the outer peripheral side and is provided with an air bleed pathway. In the slot, at which an opening to the primary duct is formed, a large opening, at which the opening area of the opening is locally larger than that of the other positions in the peripheral direction of the opening, is formed at a position in the peripheral direction corresponding to the position at which the air bleed tube is provided.

Transition Manifolds for Cooling Channel Connections in Cooled Structures

A cooled structure of a gas turbine engine having a main body with a leading edge, a trailing edge, a first side portion, a second side portion, and a cavity. A first set of cooling air micro-channels extends from the cavity and are arranged along the first side portion. A second set of cooling air micro-channels extends from the cavity and are arranged along the second side portion. Each set of cooling air micro-channels has at least one transition manifold in fluid communication with an adjacent micro-channel and also in fluid communication with at least one of an intake end, an exhaust end, and mixtures thereof. The cooled structure described above is also embodied in a gas turbine. 1. A cooled structure of a gas turbine engine , comprising;a main body having a leading edge, a trailing edge, a first side portion, a second side portion, and a cavity;a first set of cooling air micro-channels extending from said cavity and arranged along the first side portion;a second set of cooling air micro-channels extending from said cavity and arranged along the second side portion; andwherein each set of cooling air micro-channels further comprises at least one transition manifold in fluid communication with an adjacent micro-channel and in fluid communication with at least one of an intake end, an exhaust end, and mixtures thereof.2. The cooled structure of comprising a third set of cooling air micro-channels extending from said cavity and arranged along the leading edge and trailing edge.3. The cooled structure of claim 1 , wherein the transition manifold cross sectional area is greater than or equal to the adjacent micro-channel cross sectional area.4. The cooled structure of claim 1 , wherein the transition manifold extends in a generally radially outward direction.5. The cooled structure of claim 1 , wherein the intake end is in fluid communication with the cavity.6. The cooled structure of claim 1 , wherein the exhaust end is in fluid communication with a gas turbine ...

DIFFUSER REMOVAL TOOL

A diffuser removal tool for displacing an exhaust diffuser from a gas turbine engine is provided, which includes a pair of support arms, a clamp and wedge assembly coupled to an exhaust diffuser, and a lead screw assembly coupled between the support arms and to the clamp and wedge assembly via a wedge of the clamp and wedge assembly. The tool also includes a pair of linear bearing assemblies coupled to the support arms and to the clamp and wedge assembly via wedges of the clamp and wedge assembly. The lead screw assembly is configured to displace the diffuser in an axial direction relative to a longitudinal axis of the gas turbine engine while the linear bearing assemblies are configured to both enable the displacement of the diffuser in the axial direction while vertically supporting the diffuser to maintain alignment between the diffuser and the gas turbine engine. 1. A system , comprising:a gas turbine engine comprising an exhaust diffuser;an exhaust collector coupled to the exhaust diffuser, wherein the exhaust diffuser is partially disposed within the exhaust collector; anda diffuser removal tool configured to couple to both the exhaust collector and the exhaust diffuser and to separate the exhaust diffuser from the gas turbine engine by displacing the exhaust diffuser further into the exhaust collector in an axial direction relative to a longitudinal axis of the gas turbine engine while vertically supporting the exhaust diffuser to maintain axial alignment between the exhaust diffuser and the gas turbine engine.2. The system of claim 1 , comprising: a compressor claim 1 , a combustor section claim 1 , and a turbine.3. The system of claim 1 , wherein the diffuser removal tool comprises a clamp and wedge assembly coupled to the exhaust diffuser claim 1 , wherein the clamp and wedge assembly comprises a plurality of wedges and a clamp claim 1 , the clamp is circumferentially disposed relative to the longitudinal axis about the exhaust diffuser claim 1 , and the ...

GAS TURBINE POWER GENERATION DEVICE

A gas turbine power generation device includes: an inlet pipe (), a compressor (), an air storage compartment (), a compressor rotor (), a compressor gear shift (), a compressor exhaust pipeline (), a combustion chamber intake pipeline (), a combustion chamber intake cone (), a combustion chamber pneumatic valve (), a spark plug (), a combustion chamber (), a Tesla turbine (), a gas collection compartment (), an outlet pipe (), a turbine rotor (), a generator gear shift () and a generator (). 1. A gas turbine power generation device , comprising:{'b': '1', 'an inlet pipe () for allowing air to flow in the gas turbine power generation device;'}{'b': 2', '1', '4, 'a compressor (), connected to the inlet pipe () and having a compressor rotor ();'}{'b': 3', '2, 'an air storage compartment (), located downstream of the compressor ();'}{'b': 6', '4', '30, 'a compressor gear shift (), having a left end connected to the compressor rotor () and a right end connected to a turbine rotor ();'}{'b': 8', '3, 'a compressor exhaust pipeline (), connected to the air storage compartment ();'}{'b': 10', '8, 'a combustion chamber intake pipeline (), connected to the compressor exhaust pipeline ();'}{'b': 11', '10, 'a combustion chamber intake cone (), located in the combustion chamber intake pipeline ();'}{'b': 14', '11, 'a combustion chamber pneumatic valve (), located on the combustion chamber intake cone ();'}{'b': 21', '10', '19, 'a combustion chamber (), connected to the combustion chamber intake pipeline (), and having a head provided with a spark plug ();'}{'b': 22', '21, 'a Tesla turbine (), located at an outlet of the combustion chamber ();'}{'b': 27', '22, 'a gas collection compartment (), located on a side of the Tesla turbine ();'}{'b': 28', '27, 'an outlet pipe (), connected to the gas collection compartment ();'}{'b': '33', 'a generator (); and'}{'b': 32', '30', '33, 'a generator gear shift (), having a left end connected to the turbine rotor () and a right end connected ...

FUEL NOZZLE ASSEMBLY AND GAS TURBINE INCLUDING THE SAME

Disclosed herein are a fuel nozzle assembly and a gas turbine including the same. The fuel nozzle assembly includes a fuel nozzle, a head end plate to which the fuel nozzle is fixedly fastened, and a nozzle casing to which the head end plate is fixed. The fuel nozzle includes an injection cylinder having a nozzle flange fixed to the head end plate, and a nozzle shroud surrounding the injection cylinder and forming a passage between its inner wall and the injection cylinder. The injection cylinder is fastened to the head end plate by a plurality of fixing bolts arranged in a circumferential direction of the nozzle flange. Each of the fixing bolts is fitted to a pressure plate, which extends in the circumferential direction of the nozzle flange and is pressed against a flat surface thereof, to fasten the injection cylinder to the head end plate. 1. A fuel nozzle assembly comprising:a fuel nozzle;a head end plate to which the fuel nozzle is fixedly fastened; anda nozzle casing to which the head end plate is fixed, wherein the fuel nozzle comprises:an injection cylinder that supplies fuel to a combustion chamber and having a nozzle flange fixed to the head end plate; anda nozzle shroud separated from the injection cylinder to surround the injection cylinder and forming a passage between its inner wall and the injection cylinder,the injection cylinder is fixedly fastened to the head end plate by a plurality of fixing bolts arranged in a circumferential direction of the nozzle flange, andeach of the fixing bolts is fitted to a pressure plate, which extends in the circumferential direction of the nozzle flange and is pressed against a flat surface of the nozzle flange, to fixedly fasten the injection cylinder to the head end plate.2. The fuel nozzle assembly according to claim 1 , wherein the nozzle flange has a recess portion formed on the flat surface thereof such that the pressure plate is accommodated in the recess portion.3. The fuel nozzle assembly according to claim ...

DEGRADED GAS TURBINE TUNING AND CONTROL SYSTEMS, COMPUTER PROGRAM PRODUCTS AND RELATED METHODS

Various embodiments include a system having: at least one computing device configured to tune a set of gas turbines (GTs) by performing actions including: commanding each GT in the set of GTs to a base load level, based upon a measured ambient condition for each GT; commanding each GT in the set of GTs to adjust a respective output to match a nominal mega-watt power output value, and subsequently measuring an actual emissions value for each GT; adjusting an operating condition of each GT in the set of GTs based upon a difference between the respective measured actual emissions value and a nominal emissions value at the ambient condition; and calculating a degradation for each GT in the set of GTs over a period. 1. A system comprising: commanding each GT in the set of GTs to a base load level, based upon a measured ambient condition for each GT;', 'commanding each GT in the set of GTs to adjust a respective output to match a nominal mega-watt power output value, and subsequently measuring an actual emissions value for each GT;', 'adjusting an operating condition of each GT in the set of GTs based upon a difference between the respective measured actual emissions value and a nominal emissions value at the ambient condition; and', 'calculating a degradation for each GT in the set of GTs over a period., 'at least one computing device configured to tune a set of gas turbines (GTs) by performing actions including2. The system of claim 1 , wherein the base load is associated with a mega-watt power output value and an emissions value for the measured ambient condition.3. The system of claim 1 , wherein in response to commanding each GT in the set of GTs to the base load level claim 1 , each GT does not attain at least one of the nominal MW output value or the nominal emissions value.4. The system of claim 1 , wherein the at least one computing device is further configured to convert the difference between the respective measured actual emissions value and the nominal ...

AIRFOIL LEADING EDGE IMPINGEMENT COOLING

An airfoil including a spar and a cover sheet. Standoffs, a leading edge wall, and a separator wall extend away from an outer surface of the spar. The standoffs are arranged to define leading grooves disposed at the pressure side of the leading edge. The cover sheet is coupled to the leading edge wall and the standoffs over the leading grooves to define cooling passageways. The cooling passageways are in communication with one or more inlet ports formed in the spar, which are in communication with a plenum disposed within the spar. The cover sheet is arranged to define outlet ports or a slot in communication with the cooling passageway. Cooling air is delivered from the cooling air plenum through the inlet port for impingement cooling at the cover sheet, and traverses downstream through the cooling passageway to the outlet ports or slot for film cooling of the leading edge. 1. An airfoil for use in a gas turbine engine , the airfoil having a pressure side , a suction side , a leading edge , and a trailing edge , the airfoil comprising:a spar including a cooling air plenum disposed along an airfoil axis extending radially through the airfoil, a plurality of standoffs disposed along the pressure side of the airfoil, a leading edge wall disposed at the leading edge of the airfoil and extending away from the outer surface of the spar, and a leading separator wall extending away from the outer surface of the spar and spaced apart from the leading edge wall, the leading edge wall and the leading separator wall extending in a direction along the airfoil axis, the standoffs disposed spaced apart from another to define a plurality of grooves, the standoffs and the grooves extending in a direction transversing the leading separator wall, wherein the leading separator wall divides the grooves into a plurality of body section leading grooves and a plurality of body section trailing grooves, the body section leading grooves being closer in proximity to the leading edge than the ...

AIRFOIL PIECES SECURED WITH ENDWALL SECTION

An airfoil includes an airfoil section that defines an airfoil profile, and a first endwall section with which the airfoil section is attached. First and second airfoil pieces form different portions of the airfoil profile. The first and second airfoil pieces include respective first ends. The first ends interlock with the first endwall section such that the first and second airfoil pieces are retained with the first endwall section. 1. An airfoil comprising:an airfoil section defining an airfoil profile;a first endwall section with which the airfoil section is attached; andfirst and second airfoil pieces, the first airfoil piece forming a portion of the airfoil profile and the second airfoil piece forming another, different portion of the airfoil profile, the first and second airfoil pieces including respective first ends, the first ends interlocking with the first endwall section such that the first and second airfoil pieces are retained with the first endwall section.2. The airfoil as recited in claim 1 , further comprising a second endwall section claim 1 , the first and second airfoil pieces include respective second ends opposed from the respective first ends claim 1 , and the second ends interlock with the second endwall section.3. The airfoil as recited in claim 1 , wherein the first endwall section includes a cavity claim 1 , and the first ends are disposed in the cavity.4. The airfoil as recited in claim 3 , wherein the cavity is divided into a plurality of sub-cavities claim 3 , and the respective first ends are disposed in different sub-cavities.5. The airfoil as recited in claim 1 , wherein the first airfoil piece forms a leading end of the airfoil profile claim 1 , and the second airfoil piece forms at least one of a trailing end claim 1 , a pressure side claim 1 , or a suction side of the airfoil profile.6. The airfoil as recited in claim 1 , further comprising a bias member between the first airfoil piece and the first endwall section.7. The airfoil ...

AIRFOIL WITH CERAMIC AIRFOIL PIECE HAVING INTERNAL COOLING CIRCUIT

An airfoil includes an airfoil section that has radially inner and outer ends and defines an airfoil profile. The airfoil profile has a leading end, a trailing end, a suction side, and a pressure side. The airfoil section includes a ceramic airfoil piece that defines a portion of the airfoil profile. The ceramic airfoil piece includes an exterior wall that has an internal cooling circuit. 1. An airfoil comprising:an airfoil section having radially inner and outer ends and defining an airfoil profile having a leading end, a trailing end, a suction side, and a pressure side, the airfoil section including a ceramic airfoil piece that defines a portion of the airfoil profile, the ceramic airfoil piece including an exterior wall having an internal cooling circuit.2. The airfoil as recited in claim 1 , wherein the ceramic airfoil piece includes a core cavity claim 1 , and a plurality of inlet holes that open at one end thereof to the core cavity and open at another end thereof to the internal cooling circuit.3. The airfoil as recited in claim 1 , wherein the exterior wall includes a plurality of outlet holes that open on one end thereof to the internal cooling circuit and open on another end thereof to an exterior surface of the exterior wall.4. The airfoil as recited in claim 1 , wherein the exterior wall includes inner and outer wall portions that define a passage there between of the internal cooling circuit claim 1 , at least one of the inner or outer wall portions including a plurality of flow guides that project into the passage toward the other of the inner or outer wall portions.5. The airfoil as recited in claim 4 , wherein the flow guides are radially elongated ridges.6. The airfoil as recited in claim 5 , wherein the flow guides are staggered in a radial direction.7. The airfoil as recited in claim 1 , wherein the exterior wall includes inner and outer wall portions that define a passage there between of the internal cooling circuit claim 1 , and a plurality of ...

AIRFOIL WITH DUAL PROFILE LEADING END

An airfoil includes an airfoil section with a dual airfoil profile. The airfoil section includes a double wall. The double wall has an outer wall that defines a primary leading end of the dual airfoil profile and an inner wall that defines a secondary leading end of the dual airfoil profile. The inner wall is spaced from the outer wall. 1. An airfoil comprising: an outer wall that defines a primary leading end of the dual airfoil profile, and', 'an inner wall spaced from the outer wall, and the inner wall defines a secondary leading end of the dual airfoil profile., 'an airfoil section having a dual airfoil profile, the airfoil section including a double wall having'}2. The airfoil as recited in claim 1 , wherein the outer wall is formed of a first material composition claim 1 , and the inner wall is formed of a second claim 1 , different material composition.3. The airfoil as recited in claim 2 , wherein the first material composition is ceramic and the second material composition is metal.4. The airfoil as recited in claim 1 , wherein the outer wall has an exterior side and an interior side claim 1 , and the inner wall has a plurality of cooling holes that open to the interior side of the outer wall.5. The airfoil as recited in claim 4 , wherein the airfoil section includes an internal passage and at least one baffle disposed in the internal passage.6. The airfoil as recited in claim 5 , wherein the internal passage is adjacent the cooling holes such that the cooling holes open to the internal passage claim 5 , and the outer wall has a plurality of cooling holes that open on one end to the exterior side of the outer wall and on another end to the interior side of the outer wall.7. The airfoil as recited in claim 1 , further comprising first and second endwall sections claim 1 , the first and second endwall sections trapping the outer wall there between claim 1 , and at least one of the first or second endwall sections engages the outer wall in a joint claim 1 , ...

COOLING SYSTEM FOR A TURBINE ENGINE

A gas turbine engine includes a compressor section having a high pressure compressor and a core casing surrounding the compressor section and defining an inlet. The gas turbine engine also includes a cooling system for cooling air in or to the compressor section. The cooling system includes a fluid tank for storing a volume of cooling fluid and a fluid line assembly in fluid communication with the fluid tank. The fluid line assembly includes an outlet positioned upstream of the high pressure compressor and downstream of the inlet defined by the core casing for injecting cooling fluid into an airflow upstream of the high pressure compressor. 1. A gas turbine engine defining an axial direction and a radial direction , the gas turbine engine comprising:a compressor section for progressively compressing air, the compressor section including a high pressure compressor;a core casing surrounding the compressor section and defining an inlet; anda cooling system for cooling air in or to the compressor section, the cooling system comprising a fluid line assembly, the fluid line assembly including an outlet positioned upstream of the high pressure compressor and downstream of the inlet defined by the core casing for injecting cooling fluid into an airflow upstream of the high pressure compressor.2. The gas turbine engine of claim 1 , wherein the compressor section further includes a low pressure compressor claim 1 , and wherein the outlet of the fluid line assembly is positioned downstream of the low pressure compressor.3. The gas turbine engine of claim 2 , further comprising:an inner flowpath liner extending between the low pressure compressor and the high pressure compressor; andan outer flowpath liner also extending between the low pressure compressor and the high pressure compressor at a location outward from the inner flowpath liner along the radial direction, wherein the outlet of the fluid line assembly is positioned at the inner flowpath liner.4. The gas turbine ...

VANE ARC SEGMENT WITH RADIALLY PROJECTING FLANGES

A vane arc segment includes an airfoil fairing that has first and second platforms and an airfoil section. The platforms have first and second openings that open into the airfoil section. The platforms each define first and second circumferential mate faces, forward and aft sides, a gaspath side, a non-gaspath side. The first platform has a first flange that projects radially from the non-gaspath side aft of the first opening and a second flange that projects radially from the non-gaspath side forward of the first opening. The first and second flanges are exclusive flanges on the first platform. The second platform has a third flange that projects radially from the non-gaspath side aft of the second opening. 1. A vane arc segment comprising:an airfoil fairing having first and second platforms and an airfoil section extending there between, the first and second platforms having, respectively, first and second openings that open into the airfoil section,the first and second platforms each defining first and second circumferential mate faces, forward and aft sides, a gaspath side, a non-gaspath side, a first flange projecting radially from the non-gaspath side of the first platform aft of the first opening, the first flange extending in a first flange length-wise direction between the first and second circumferential mate faces of the first platform, and', 'a second flange projecting radially from the non-gaspath side of the first platform forward of the first flange, the first and second flanges being exclusive flanges on the first platform, and, 'the first platform having 'a third flange projecting radially from the non-gaspath side of the second platform aft of the second opening, the third flange extending in a third flange length-wise direction between the first and second circumferential mate faces of the second platform.', 'the second platform having2. The vane as recited in claim 1 , wherein the second flange extends in a second flange length-wise direction ...

COMPRESSOR AND GAS TURBINE

A compressor includes a rotor including a plurality of disks, a shaft portion connected on a downstream side of the disks; and rotor blade rows fixed to the plurality of disks; a stator including a compressor casing; and a plurality of stator vane rows each provided between corresponding adjacent ones of the rotor blade rows; an outlet guide vane including blade main bodies disposed at an interval in a circumferential direction on the downstream side of the disk located most downstream, and inner shrouds connecting the blade main bodies in the circumferential direction, on an inner side in a radial direction; and an inner casing disposed on the downstream side of the disk located most downstream with a gap between the disk and the inner casing. The inner casing includes an outer peripheral wall surface having recesses accommodating the inner shrouds and forming, together with an inside surface of the compressor casing, a diffuser on the downstream side of the recesses, and an inner peripheral wall surface forming an air extraction cavity. An air extraction hole is formed in a portion, in the recesses, on the downstream side. 2. The compressor according to claim 1 , wherein a plurality of the outlet guide vanes are arranged at an interval in the axial direction claim 1 , and the recesses are provided for each of the outlet guide vanes.3. The compressor according to claim 2 , wherein the air extraction hole is formed in one of the plurality of recesses located most downstream in the axial direction.4. The compressor according to claim 1 , wherein a portion of the inner casing that is more downstream in the axial direction than the outlet guide vane extends inward in the radial direction toward the downstream side in the axial direction.5. A compressor comprising:a rotor including a plurality of disks stacked in an axial direction, a shaft portion connected on a downstream side in the axial direction of the disks, and a plurality of rotor blade rows fixed to the ...

AIRFOIL SHAPE FOR TURBINE ROTOR BLADES

A turbine rotor blade having an airfoil that includes a pressure side portion of a nominal airfoil profile substantially in accordance with Cartesian coordinate values of X, Y, and Z of a pressure side as set forth in Table I. The Cartesian coordinate values of X, Y, and Z are non-dimensional values from 0% to 100% convertible to dimensional distances by multiplying the Cartesian coordinate values of X, Y and Z by a height of the airfoil defined along the Z axis. The X and Y values of the pressure side are coordinate values that, when connected by smooth continuing arcs, define pressure side sections of the pressure side portion of the nominal airfoil profile at each Z coordinate value. The pressure side sections may be joined smoothly with one another to form the pressure side portion. 1. A turbine rotor blade including an airfoil that comprises a pressure side portion of a nominal airfoil profile substantially in accordance with Cartesian coordinate values of X , Y , and Z of a pressure side as set forth in Table I , wherein:the Cartesian coordinate values of X, Y, and Z are non-dimensional values from 0% to 100% convertible to dimensional distances by multiplying the Cartesian coordinate values of X, Y and Z by a height of the airfoil defined along a Z axis;the X and Y values of the pressure side are coordinate values that, when connected by smooth continuing arcs, define pressure side sections of the pressure side portion of the nominal airfoil profile at each Z coordinate value; andthe pressure side sections are joined smoothly with one another to form the pressure side portion.2. The turbine rotor blade of claim 1 , wherein the nominal airfoil profile lies in an envelope within +/− 5% of a chord length in a direction normal to any of the pressure side sections.3. The turbine rotor blade of claim 1 , wherein a height of the pressure side portion is defined along the Z axis claim 1 , and wherein the height of the pressure side portion is less than or equal to ...

GAS TURBINE WITH PLURALITY OF TIE RODS AND METHOD OF ASSEMBLING THE SAME

Disclosed herein is a gas turbine with a plurality of tie rods. The gas turbine includes a compressor section having a plurality of compressor-side rotor disks, a turbine section having a plurality of turbine-side rotor disks arranged downstream of the compressor-side rotor disks, a first tie rod penetrating central portions of the rotor disks provided in the compressor section to press the compressor-side rotor disks against each other, a plurality of second tie rods penetrating vicinities of edges of the rotor disks provided in the turbine section to press the turbine-side rotor disks against each other, and a torque transfer member coupled to the first and second tie rods so as to transfer rotational torque generated by the turbine section to the compressor section. 1. A gas turbine comprising:a compressor section having a plurality of compressor-side rotor disks;a turbine section having a plurality of turbine-side rotor disks arranged downstream of the compressor-side rotor disks;a first tie rod penetrating central portions of the rotor disks provided in the compressor section to press the compressor-side rotor disks against each other;a plurality of second tie rods penetrating vicinities of edges of the rotor disks provided in the turbine section to press the turbine-side rotor disks against each other; anda torque transfer member coupled to the first and second tie rods so as to transfer rotational torque generated by the turbine section to the compressor section.2. The gas turbine according to claim 1 , wherein the torque transfer member comprises a first end facing the compressor section claim 1 , and the first end has a first through-hole through which the first tie rod passes.3. The gas turbine according to claim 2 , further comprising a first pressure nut fastened to one side of the first tie rod claim 2 ,wherein the first pressure nut presses an inside surface of the first end toward the compressor section.4. The gas turbine according to claim 3 , ...

Gas Turbine Combustor

The invention provides a gas turbine combustor having structural reliability against vibration of fuel nozzles caused by fluid force and high environmental performance by uniform combustion in a combustion chamber. The gas turbine combustor comprises a fuel nozzle configured to inject fuel, and a fuel nozzle plate including a hole section into which an insertion section located in a root part of the fuel nozzle is inserted. The fuel nozzle includes a male screw section at least on an outer circumferential surface of a downstream portion viewed from a flowing direction of the fuel in the insertion section. The fuel nozzle plate includes a female screw section in the hole section. The female screw section screws with the male screw section. The fuel nozzle includes the insertion section. An upstream end portion of the insertion section is metallurgically joined to an upstream end portion of the fuel nozzle plate. 1. A gas turbine combustor comprising:a fuel nozzle configured to inject fuel; anda fuel nozzle plate including a hole section into which an insertion section located in a root part of the fuel nozzle is inserted, whereinthe fuel nozzle includes a male screw section at least on an outer circumferential surface of a downstream portion viewed from a flowing direction of the fuel in the insertion section,the fuel nozzle plate includes a female screw section in the hole section, the female screw section screwing with the male screw section, andthe fuel nozzle includes the insertion section, an upstream end portion of the insertion section being metallurgically joined to an upstream end portion of the fuel nozzle plate.2. The gas turbine combustor according to claim 1 , whereinthe fuel nozzle includes the insertion section including the downstream portion and an upstream portion,the fuel nozzle includes the male screw section only on the outer circumferential surface of the downstream portion of the insertion section, anda root diameter of the male screw section ...

TURBINE CLEARANCE CONTROL UTILIZING LOW ALPHA MATERIAL

A turbine module comprises a stator assembly disposed annularly about a rotor assembly. The rotor assembly includes a plurality of turbine blades circumferentially distributed about a turbine disk. The stator assembly includes at least one case segment and an abradable surface disposed radially adjacent to a tip of each of the plurality of rotor blades. The turbine blades each include an airfoil section with a first gamma-phase titanium aluminide (gamma-TiAl) substrate, and the at least one case segment has a second gamma-TiAl substrate. 1. A turbine module comprising:a rotor assembly including a plurality of turbine blades circumferentially distributed about a turbine disk, the plurality of turbine blades each including an airfoil section with a first gamma-phase titanium aluminide (gamma-TiAl) substrate; anda stator assembly disposed annularly about the rotor assembly, the stator assembly including an abradable surface disposed radially adjacent to a tip of each of the plurality of rotor blades, the stator assembly including at least one case segment with a second gamma-TiAl substrate.2. The turbine module of claim 1 , wherein the first gamma-TiAl substrate includes a first composition and the second gamma-TiAl substrate includes a second composition.3. The turbine module of claim 2 , wherein at least one of the first composition and the second composition includes less than about 15 vol % alpha-TiAl.4. The turbine module of claim 2 , wherein at least one of the first composition and the second composition includes less than about 5 vol % alpha-TiAl.5. The turbine module of claim 2 , wherein the second composition is substantially different from the first composition.6. The turbine module of claim 2 , wherein the second composition is substantially identical to the first composition.7. The turbine module of claim 1 , wherein the turbine disk includes a gamma-TiAl substrate with a third composition.8. The turbine module of claim 1 , wherein the third composition is ...

COMPRESSOR END-WALL TREATMENT WITH MULTIPLE FLOW AXES

A compressor is provided including a casing, a hub, a flow path, a plurality of blades, and an end-wall treatment formed in at least one of the casing and the hub, and facing a tip of each blade. The flow path is formed between the casing and the hub, and the plurality of blades is positioned in the flow path. The tip of each blade and the end-wall treatment are configured to move relative to each other. Such end-wall treatment includes a first recess portion extending along a first axis to maintain a fluid flow substantially straight through the first recess portion. The end-wall treatment further includes a plurality of second recess portions spaced apart from each other and extending from the first recess portion along a second axis different than the first axis to maintain the fluid flow substantially straight through the plurality of second recess portions. 1. A compressor comprising:a casing;a hub;a flow path formed between the casing and the hub;a plurality of blades positioned in the flow path; andan end-wall treatment formed in at least one of the casing and the hub, and facing a tip of each blade among the plurality of blades, wherein the tip of each blade and the end-wall treatment are configured to move relative to each other,wherein the end-wall treatment comprises a first recess portion extending along a first axis to maintain a fluid flow substantially straight through the first recess portion, and a plurality of second recess portions spaced apart from each other and extending from the first recess portion along a second axis different than the first axis to maintain the fluid flow substantially straight through the plurality of second recess portions.2. The compressor of claim 1 , wherein the plurality of second recess portions extends along a first direction of the second axis.3. The compressor of claim 2 , wherein the plurality of second recess portions extends along a second direction opposite to the first direction claim 2 , of the second axis.4 ...

SUPPORT STRUCTURE FOR RADIAL INLET OF GAS TURBINE ENGINE

The compressor inlet can have two walls forming an annular fluid path with a radial inlet end, and a support structure extending axially between the two walls, the support structure having a plurality of circumferentially-interspaced supports, each one of the plurality of supports extending freely between the two walls across the radial inlet end of the annular fluid path, each support having at least one node at an intermediary location between the two walls, at least one branch extending from the node to a first one of the walls, and at least two branches branching off from the node and leading to the second one of the walls. 1. A compressor inlet for a gas turbine engine , the compressor inlet having two walls forming an annular fluid path with a radial inlet end , and a support structure extending axially between the two opposite walls , the support structure having a plurality of circumferentially-interspaced supports , the supports extending freely between the two walls across the radial inlet end of the annular fluid path , the supports having at least one node at an intermediary location between the two walls and a plurality of branches extending therefrom , at least one of said branch extending from the node to a first one of the walls , and at least two of said branches branching off from the node and leading to the second one of the walls.2. The compressor inlet of wherein at least one support has said branches arranged in a Y shape claim 1 , with a single branch leading from the node to the first wall and two branches extending from the node to the second wall.3. The compressor inlet of wherein the single branch is closer to the compressor stage than the two branches extending from the node to the second wall.4. The compressor inlet of wherein at least one support has said branches arranged in an X-shape claim 1 , with two branches extending from the node to the first wall and two branches extending from the node to the second wall.5. The compressor ...

GAS TURBINE ENGINE SYSTEM AND SUPERSONIC EXHAUST NOZZLE

One embodiment of the present invention is a unique gas turbine engine system. Another embodiment is a unique exhaust nozzle system for a gas turbine engine. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for gas turbine engine systems and exhaust nozzle systems for gas turbine engines. Further embodiments, forms, features, aspects, benefits, and advantages of the present application will become apparent from the description and figures provided herewith. 1. A gas turbine engine system , comprising:a fan system;a compressor system in fluid communication with the fan system;a combustion system in fluid communication with the compressor system;a turbine system in fluid communication with the combustion system, wherein the turbine system is configured to discharge a first stream flow in the form of an engine core flow;and wherein the fan system is configured to discharge a second stream flow in the form of a bypass flow and to discharge a third stream flow in the form of another bypass flow; and a first nozzle configured to discharge the first stream flow and the second stream flow; and', 'an ejector configured to entrain ambient free stream air into the third stream flow., 'an exhaust nozzle system in fluid communication with the fan system and the turbine system, including2. (canceled)3. The gas turbine engine system of claim 1 , wherein the exhaust nozzle system includes a mixer configured to mix the first stream flow and the second stream flow.4. The gas turbine engine system of claim 1 , wherein the ejector includes a plurality of ejector doors that are mounted for movement among various positions.5. (canceled)6. The gas turbine engine system of claim 4 , wherein the ejector doors are configured to vary an amount of the ambient free stream air entrained into the third stream flow.7. The gas turbine engine system of claim 6 , wherein the ejector doors are configured to selectively close and prevent entrainment of the ...

Combined high pressure turbine case and turbine intermediate case

A disclosed gas turbine engine includes a compressor section, a combustor section, a first turbine section and a second turbine section. An outer case structure for the gas turbine engine includes a single-piece case structure with a turbine case portion and a transition case portion. The transition case portion is integrally formed with the turbine case portion as a single part module. A combustor case houses the combustor and an aft turbine case supports the low pressure turbine. The outer case includes a forward end attachable to the combustor case and an aft end attachable to the aft turbine.

SYSTEMS AND METHODS FOR COMPRESSOR ANTICORROSION TREATMENT

The present application provides a gas turbine engine. The gas turbine engine may include a compressor, a compressor wash system in communication with the compressor, a condensate or boiler feed water system in communication with the compressor, and a dosing system in communication with the condensate or boiler feed water system. 1. A gas turbine engine , comprising:a compressor;a compressor wash system in communication with the compressor;a condensate or boiler feed water system in communication with the compressor; anda dosing system in communication with the condensate or boiler feed water system.2. The gas turbine engine of claim 1 , wherein the compressor wash system comprises a deionized wash fluid.3. The gas turbine engine of claim 1 , wherein the compressor wash system comprises a plurality of nozzles positioned about a bellmouth of the compressor.4. The gas turbine engine of claim 1 , wherein the condensate or boiler feed water system comprises a flow of condensate or boiler feed water with a concentration of an anticorrosion agent therein.5. The gas turbine engine of claim 1 , wherein the condensate or boiler feed water system is in communication with the compressor wash system upstream of the compressor.6. The gas turbine engine of claim 1 , wherein the dosing system comprises a flow of amine therein.7. The gas turbine engine of claim 1 , wherein the dosing system is in communication with the compressor wash system upstream of the compressor.8. The gas turbine engine of claim 1 , further comprising a temperature and concentration management system in communication with the compressor wash system claim 1 , the condensate or boiler feed water system claim 1 , and the dosing system.9. The gas turbine engine of claim 8 , wherein the temperature and concentration management system comprises a flow sensor and a modulating valve in communication with the compressor wash system.10. The gas turbine engine of claim 8 , wherein the temperature and concentration ...