Охлаждаемая турбинная лопатка (варианты) и способ охлаждения турбинной лопатки

Охлаждаемая турбинная лопатка содержит хвостовик, предназначенный для прикрепления охлаждаемой лопатки к турбинному ротору, аэродинамический профиль, концевой бандаж и один или несколько центральных охлаждающих каналов, ограниченных аэродинамическим профилем. Аэродинамический профиль проходит вдоль радиальной оси от хвостовика и ограничивает один задний охлаждающий канал, который проходит радиально через аэродинамический профиль проксимально к задней кромочной части аэродинамического профиля. Задний канал расположен в пределах расстояния от задней кромочной части, которое составляет менее 25% хордовой длины аэродинамического профиля. Концевой бандаж расположен на радиально внешнем конце аэродинамического профиля, проходит в окружном направлении от аэродинамического профиля и ограничивает внутри себя центральную полость повышенного давления и периферическую полость повышенного давления. Аэродинамический профиль ограничивает одно заднее охлаждающее впускное отверстие, предназначенное для ...

ТУРБИННЫЙ УЗЕЛ, СООТВЕТСТВУЮЩАЯ ТРУБКА СОУДАРИТЕЛЬНОГО ОХЛАЖДЕНИЯ И ГАЗОТУРБИННЫЙ ДВИГАТЕЛЬ

Турбинный узел содержит полую аэродинамическую часть, имеющую по меньшей мере одну полость с по меньшей мере одной трубкой соударительного охлаждения, предназначенную для введения внутрь полости полой аэродинамической части и используемую для соударительного охлаждения, по меньшей мере, внутренней поверхности полости, и по меньшей мере одну платформу, расположенную на радиальном конце полой аэродинамической части, и по меньшей мере одну охлаждающую камеру, используемую для охлаждения по меньшей мере одной платформы, и которая расположена на противоположной полой аэродинамической части стороне платформы. Охлаждающая камера ограничена на первом радиальном конце платформой, а на противоположном радиальном втором конце с помощью по меньшей мере одной закрывающей пластины. Трубка соударительного охлаждения выполнена из переднего элемента и заднего элемента, вставленных оба в по меньшей мере одну полость. Передний элемент расположен в направлении передней кромки полой аэродинамической части.

ЛОПАТКА ДЛЯ ТУРБОМАШИНЫ И ТУРБОМАШИНА, СОДЕРЖАЩАЯ ТАКУЮ ЛОПАТКУ.

Лопатка для турбомашины, в частности газовой турбины, расположена на турбинном роторе и содержит перо и хвостовую часть, выполненные за одно целое с лопаткой, проход для подачи охлаждающего воздуха в хвостовой части для направления охлаждающего воздуха в охладитель и отвод охлаждающего воздуха, расположенный в хвостовой части и соединенный по текучей среде с проходом для подачи охлаждающего воздуха. Перо имеет охладитель, расположенный внутри пера, а хвостовая часть имеет две узкие стороны и две широкие стороны. Отвод охлаждающего воздуха содержит сопло на одной из узких сторон хвостовой части, и сопло образовано с помощью отверстия. Хвостовая часть лопатки содержит верхнюю платформу лопатки и нижнюю платформу лопатки. Верхняя платформа лопатки и нижняя платформа лопатки выполнены в качестве частей лабиринтного уплотнения в собранном состоянии в турбомашине. Сопло расположено между верхней платформой лопатки и нижней платформой лопатки. Осевое направление отверстия наклонено вверх под углом ...

НАПРАВЛЯЮЩАЯ ЛОПАТКА ТУРБИНЫ С ОХЛАЖДАЕМОЙ ГАЛТЕЛЬЮ

Направляющая лопатка содержит полку и перо, продолжающееся от указанной полки и соединенное с полкой посредством галтели. Инжекционная трубка вставляется в перо, ограничивая охлаждающий канал между инжекционной трубкой и боковыми стенками пера. Направляющая лопатка дополнительно содержит отклоняющую структуру, расположенную смежно галтели и которая повторяет внутренний контур галтели и ограничивает первый охлаждающий проход между галтелью и отклоняющей структурой. Первое препятствие расположено внутри пера в месте соединения галтели с боковыми стенками для отделения первого охлаждающего прохода от охлаждающего канала в пере и для того, чтобы направлять охлаждающий газ из первого охлаждающего прохода в инжекционную трубку. Изобретение направлено на повышение эффективности охлаждения. 2 н. и 13 з.п. ф-лы, 8 ил.

ЭКСЦЕНТРИЧЕСКАЯ ФАСКА У ВХОДА ОТВЕТВЛЕНИЙ В ПРОТОЧНОМ КАНАЛЕ

Проточный канал содержит основной канал, ответвляющийся канал, в котором направление потока перпендикулярно направлению потока основного канала, и входное отверстие ответвляющегося канала, которое расположено в стенке основного канала и задано кромкой, содержащей верхнюю по потоку кромку и нижнюю по потоку кромку. У верхней по потоку кромки входного отверстия предусмотрена фаска. Нижняя по потоку кромка входного отверстия является острой кромкой, образованной прямым углом между стенкой ответвляющегося канала и стенкой основного канала. Фаска у верхней по потоку кромки имеет форму, которая является эксцентрической относительно продольной оси ответвляющегося канала. Изобретение направлено на уменьшение потерь давления. 2 н. и 5 з.п. ф-лы, 8 ил.

СИСТЕМА РЕГУЛИРОВАНИЯ ТЕМПЕРАТУРЫ ОТКЛЮЧЕНИЯ ТУРБИННОГО ДВИГАТЕЛЯ С ВСПРЫСКИВАЮЩИМ СОПЛОМ ДЛЯ ГАЗОТУРБИННОГО ДВИГАТЕЛЯ

ЛОПАТКИ ЛОПАТОЧНОГО КОЛЕСА ГАЗОТУРБИННОГО ДВИГАТЕЛЯ, ОСНАЩЕННЫЕ КАНАВКАМИ ДЛЯ ОХЛАЖДЕНИЯ

... 1. Лопатка (10) лопаточного колеса (100) газотурбинного двигателя, содержащая аэродинамический профилированный элемент (50), имеющий нижнюю поверхность (56) и платформу (60), проходящую от одного из концов этого аэродинамического профилированного элемента (50) в направлении, в целом перпендикулярном к продольному направлению аэродинамического профилированного элемента, причем платформа содержит, по меньшей мере, один канал (16) впрыскивания воздуха, отличающаяся тем, что платформа (60) содержит канавку (40, 140, 240), проходящую вдоль нижней поверхности (56) вблизи от, по меньшей мере, задней по потоку части (57) этой поверхности, причем в этой канавке выполнен, по меньшей мере, один канал (16) впрыскивания воздуха. ! 2. Лопатка по п.1, выполненная с возможностью размещения совместно с множеством других, по существу, идентичных лопаток, для формирования венца вокруг оси (А) этого венца, причем в этом венце профилированные аэродинамические элементы (50) располагаются по существу в радиальном ...

УСТРОЙСТВО СЕКЦИОННОГО ОХЛАЖДЕНИЯ И СПОСОБ ОХЛАЖДЕНИЯ СОПЛОВОЙ ЛОПАТКИ ТУРБИНЫ

... 1. Устройство (100) секционного охлаждения для подачи охлаждающего потока (170) в турбине (40) с потоком (35) газообразных продуктов сгорания, содержащее:турбинную сопловую лопатку (105), имеющую вставку (150), расположенную в ее аэродинамической части, и наружную боковую поверхность (140), идефлектор (200) для охлаждающей среды, имеющий проход (180) для охлаждающей среды высокого давления, сообщающийся с указанной вставкой (150), в первом контуре (300), и инжекционную пластину (230), расположенную над указанной наружной боковой поверхностью (140), во втором контуре (330).2. Устройство (100) по п.1, в котором дефлектор (200) для охлаждающей среды содержит сегменты (210).3. Устройство (100) по п.2, в котором дефлектор (200) для охлаждающей среды имеет шлицевое уплотнение (220), расположенное между каждой парой его указанных сегментов (210).4. Устройство (100) по п.1, в котором дефлектор (200) для охлаждающей среды содержит камеру (265) послеударного давления, расположенную между инжекционной ...

КАРТЕР ТУРБИНЫ, СОДЕРЖАЩИЙ СРЕДСТВА КРЕПЛЕНИЯ СЕКЦИЙ КОЛЬЦА

... 1. Картер (30) турбины летательного аппарата, предназначенный для установки на нем блока секций кольца (28), которое частично ограничивает канал прохождения потока газа через турбину, при этом картер (30) содержит средства динамической регулирования радиального положения секций кольца (28) посредством контролируемого нагнетания потока воздуха на участки (36) кольцевой стенки (34) картера (30), при этом картер (30) содержит радиальную входную лапку (38), которая соединяет входной концевой участок каждой секции кольца (28) с картером (30) в направлении течения потока газа, и выходную радиальную лапку (40), которая связывает выходной концевой участок каждой секции кольца (28) с картером (30),отличающийся тем, что, по меньшей мере, одна входная радиальная лапка (38) выполнена как одно целое с картером (30) и связана непосредственно с входным концевым участком каждой секции кольца (28).2. Картер (30) по п. 1, отличающийся тем, что две радиальные лапки (38, 40) выполнены как единое целое с картером ...

ОХЛАЖДАЕМАЯ ТУРБИНА

... 1. Охлаждаемая турбина, содержащая рабочее колесо с установленными на нем рабочими лопатками с двумя контурами охлаждения, последовательно соединенные с воздушными каналами в рабочем колесе, с независимыми кольцевыми диффузорными каналами, образованными на поверхности рабочего колеса, соединенными с сопловыми аппаратами закрутки и транзитными воздуховодами на их входе, сопловые лопатки, каждая из которых выполнена в виде конструктивного элемента, ограниченного верхней и нижней полками, и пространства между ними, ограниченного вогнутой и выпуклой стенками пера сопловой лопатки, в виде расположенных вдоль ее оси раздаточного коллектора входной кромки и раздаточной полости, причем раздаточный коллектор входной кромки соединен на входе с воздушной полостью камеры сгорания, а на выходе через перфорационные отверстия во входной кромке сопловой лопатки с проточной частью турбины, теплообменник, соединенный на входе с воздушной полостью камеры сгорания, а на выходе последовательно сообщенный с ...

ВНУТРЕННЯЯ ПЛАТФОРМА СОПЛА И СОПЛОВАЯ ЛОПАТКА (ВАРИАНТЫ)

... 1. Внутренняя платформа сопла, включающая:полость платформы,инжекционную камеру, расположенную в упомянутой полости платформы;удерживающую пластину, расположенную на первой стороне упомянутой инжекционной камеры; иэластичное уплотнение, расположенное на второй стороне упомянутой инжекционной камеры.2. Внутренняя платформа по п.1, в которой упомянутая удерживающая пластина включает держатель уплотнения.3. Внутренняя платформа по п.1, в которой упомянутая полость платформы включает один или более зацепов платформы, а удерживающая пластина включает один или более зацепов пластины, чтобы удерживать удерживающую пластину в полости платформы.4. Внутренняя платформа по п.1, в которой упомянутая удерживающая пластина включает цилиндрический контур, чтобы удерживать удерживающую пластину в полости платформы.5. Внутренняя платформа по п.1, также включающая один или более штырей, входящих в упомянутую полость платформы, чтобы удерживать удерживающую пластину в полости платформы.6. Внутренняя платформа ...

Air-cooled gas turbine blade

A turbine blade for a gas turbine engine has internal passages which are opened at their roots to supply pressure so that there is a constant flow of cooling air inside the blade in use. The blade has an airfoil surface and a number of internal passages inside the walls of the blade, some being formed adjacent the mid-chord section and extend from the root section to the tip section to define feed channels open at the blade root section. A number of radially spaced film cooling holes in the airfoil surface communicate with each feed channel to flow a film of cooling air adjacent the airfoil surface. A number of replenishment holes are radially spaced in the wall to flow air from the mid-chord section to the feed channel(s) to replenish the cooling air in the feed channel(s) that is otherwise lost in supplying air to the film cooling holes. A device communicates cooling air from the root section via the feed channels to discharge from orifices in the airfoil surface at the tip section.

Gekühlte Schaufeln eines Gasturbinentriebwerks

Gekühlte Schaufel eines Gasturbinentriebwerks, umfassend: einen Kühlkanal, der innerhalb der gekühlten Schaufel ausgebildet ist und eine darin strömende Kühlluft aufweist; ausgebildete Filmkühllöcher, die von einer Innenwandfläche des Kühlkanals zu einer Außenwandfläche der Kühlschaufel durchdringen und einen Kühlfilm auf einer äußeren Fläche der Schaufel bilden, und ein darauf ausgebildetes Prallkühlelement, welches eine Vielzahl von die Kühlluft ausstoßenden kleinen Löchern aufweist, und gekennzeichnet ist durch: wobei das Prallkühlelement im Kühlkanal angeordnet ist und dabei einen vorbestimmten Zwischenraum entfernt von der Innenwandfläche lässt; wobei ein durch die Innenwandfläche und das Prallkühlelement ausgebildeter Zwischenraum einen darin angebrachten Dichtungsabschnitt aufweist, der den entsprechenden Zwischenraum in einer Schaufelsehnenrichtung teilt, und wobei der Dichtungsabschnitt zwischen in einer Schaufelsehnenrichtung benachbarten Filmkühllöchern angebracht ist.

GASTURBINE UND GASTURBINENBETRIEBSVERFAHREN

Eine Gasturbine wird mit einem Abgasdiffusor 5, in dem ein Abgasströmungsweg Pe zum Zirkulieren von Abgas von einer Turbine gebildet ist, und einer Kühlvorrichtung 6 zum Kühlen einer Struktur bereitgestellt, die zu dem Abgasströmungsweg Pe im Abgasdiffusor 5 gerichtet ist. Die Kühlvorrichtung 6 weist ein Führungsteil 7, in dem ein Führungsströmungsweg Pg zum Zirkulieren eines Kühlmediums gebildet ist, und das das Kühlmedium zu der Struktur führt, und ein Umschaltteil 8 auf, das in der Lage ist, zwischen einem ersten Zustand, bei dem eine Fließgeschwindigkeit des Kühlmediums, das durch den Führungsströmungsweg Pg strömt, eine erste Fließgeschwindigkeit ist, die einer Fließgeschwindigkeit während eines Nennbetriebs entspricht, und einem zweiten Zustand, bei dem die Fließgeschiwindigkeit des Kühlmediums eine zweite Fließgeschwindigkeit ist, die höher als die erste Fließgeschwindigkeit ist, umzuschalten.

STRÖMUNGSMASCHINE MIT EINER KÜHLBAREN ANORDNUNG VON WANDELEMENTEN UND VERFAHREN ZUR KÜHLUNG EINER ANORDNUNG VON WANDELEMENTEN

Gasturbine mit einer sequentiellen Verbrennung

Gekühlte Stator- oder Rotorschaufel für eine Turbomaschine

BOGENFOERMIGES WAND- UND DICHTSEGMENT FUER EINE AXIALSTROEMUNGSMASCHINE

Gasturbinenleitschaufel

Strömungsmaschinenschaufel mit aparter Kühlung der Anströmkante

Gasturbinenstatorschaufelsegment mit einem zweigeteilten Hohlraum

Gas turbine engine aerofoil

Turbine cooling flow modulation

A Gas Turbine Ducted Fan Engine.

... 1,161,186. Gas turbine jet propulion plant; axial flow fans. ROLLS-ROYCE Ltd. May 28, 1968, No.25484/68. Headings F1C, F1G and F1J. In a gas turbine ducted fan engine having a front fan surrounded by an annular casing at least a portion 22 of the casing is formed by two spaced apart walls 20, 24, the latter wall consisting of members 25 each having an integral flange 28 extending to the wall 20 and secured thereto, e.g. by welding. Circumferentially adjacent members 25 overlap at recesses 27 and are welded together. The flanges 28 define axially extending cavities 23. The portion 22 forms part of the radially inner wall of the annular fan casing (16) upstream of the tips of the fan blades (17, Fig. 1, not shown), the outer wall of the casing being formed by a wall 21 welded to the portion 22 via annular members 33, 34 and brackets 35. In operation, hot de-icing air supplied to the downstream end of the casing passes through apertures 36, 37 in the members 33, 34 to impinge on the leading ...

Gas turbine aerofoil cooling with pressure attenuation chambers

A gas turbine engine aerofoil 24 has a plurality of attenuation chambers 34 positioned between a cooling air passageway 26 and its leading edge 28. Cooling air passing from the passageway 26 to the exterior surface of the leading edge is attenuated in pressure by impingement on the opposing walls of the respective chambers 34, prior to leaving the chambers 34 via exit passageways 38 which straddle the leading edge. A plurality of input passageways 36 are provided for flow from passageway 26 into the chambers 34, and each chamber may have numerically less input passageways 36 than exit passageways 38. In alternative embodiments, passageways (36, fig 5) connect to upper and/or lower ends of each chamber 36 for pressure attenuation resulting from expansion therein. Hot spots due to blockage and effects due to pressure fluctuation are reduced.

A Blade for a rotor

COOLING OF WALL MEMBERS OF STRUCTURES

TURBINE SHROUDING

A turbine wheel for a turbine engine with cooling and anti-vibration means

A turbine wheel (fig 1, 10) comprises a disk (fig 1, 11), carrying blades 114; each blade having a platform 116, carrying an airfoil 115; and connected by a tang (fig 4, 117) to a root (fig 4, 113). Between the tangs are spaces where sealing and damping sheets 118 are located. The inner face of the platform has a line of studs 122, arranged along the longitudinal axis of the wheel; against which the sheets bear. This defines a radial clearance J or airflow space, supplied with cooling air for the platform by rows of holes 120 in the sheets, which are also arranged along the longitudinal axis of the wheel. The sheets are dish shaped with the concave side facing radially inwards, they are also of undulating cross-section, having a convex centre portion (fig 6, 131) which bears against the end portions of two adjacent platforms (fig 6, 116&116â ).

A turbine blade support assembly and a turbine assembly

Cooling system for turbine exhaust assembly

A cooling system for a turbine exhaust assembly 10 comprises an annular case 20, 21, a flow-path ring 12 and a splash plate or baffle 32. The flow-path ring 12 is coaxially disposed within the annular case 20, 21, and the splash plate or baffle 32 extends axially between the annular case 20, 21 and the flow-path ring 12. A plurality of cooling fluid apertures 31 are formed in the annular case 20, 21 in order to provide cooling fluid flow onto the splash plate or baffle 32. A plurality of impingement holes 33 are formed in the splash plate or baffle 32 in order to provide impingement cooling flow onto the flow-path ring 12. The flow-path ring may be provided with film cooling apertures 41 along the aft portion and/or the full axial length thereof. A blocker door 28 may be provided outside of the annular case 20, 21, downstream of the apertures 31, to increase a relative pressure of the cooling fluid flow. The annular case 20, 21 may comprise an inner turbine exhaust case or an outer turbine ...

Article, airfoil component and method for forming article

An article 100, which may be an airfoil or airfoil component, comprises manifold 102, article wall 104, post-impingement cavity 106 and a plurality of post-impingement partitions 118. The manifold includes impingement wall 108 defining plenum 110 and a plurality of impingement apertures 112. The article wall includes a plurality of external apertures 114. The post-impingement cavity is disposed between the manifold and the article wall, and is arranged to receive fluid 116 from the plenum through the impingement apertures and exhaust through the external apertures. Post-impingement partitions divide the post-impingement cavity into a plurality of hermetically separated sub-cavities 120. The impingement wall, article wall and post-impingement partitions are integrally formed as a single, continuous article. Impingement apertures may be arranged to distribute fluid to generate higher heat transfer coefficient in a sub-cavity exposed to higher temperatures. Apertures may be arranged to distribute ...

Improvements in and relating to turbo machines

An annular sleeve surrounds a turbine rotor and is fixed to a stationary part, the sleeve being elastically deformable in a radial direction. A wear ring is carried by the sleeve, the wear ring being composed of segments arranged end-to-end circumferentially. The adjacent ends of the segments are spaced apart at all operating conditions of the turbine. The thermal expansions of the turbine rotor and of the sleeve and ring are so related that under operating conditions of the turbine, when the engine of which it forms a part is running, the sleeve thermally expands radially. The rate of thermal expansion of the stationary part is directly related to the rate of thermal expansion of the turbine rotor. Air leaving the compressor is directed against the sleeve and the stationary part carrying the sleeve. The wear ring and portion carrying the sleeve may be insulated. The sleeve may be mounted on a combustion chamber surrounding the turbine.

Blade cooling

Aerofoil cooling arrangement

Heatshield for a gas turbine engine

Cooling of gas turbine engine accessories

GAS TURBINE ENGINE

... 1519590 Gas turbine engines; cooling ROLLS-ROYCE Ltd 1 Dec 1975 [11 Nov 1974] 48662/74 Heading F1G The invention relates to a method of cooling a wall of a gas turbine engine, the wall defining a duct 8 through which combustion gases flow. A chamber 12 is defined outwardly of the wall 11 and a partition 20 is disposed within the chamber to define an inlet region 12A, to which cooling air is supplied through inlet 16, and an outlet region 12B from which the cooling air finally discharges through outlet 17. The partition is in the form of a corrugated member which defines projections 20A which project towards the outer surface 11B of the wall 11 and discharge orifices 15 are formed in the projections through which cooling air discharges on to the surface 11B, the air then passing to the spaces 21 between the projections, finally discharging through outlets 17 which open into the inlet guide vanes 8 from which the air discharges through outlets along the trailing edge thereof. In Figs. 2 and ...

MEANS FOR COOLING A SURFACE BY THE IMPINGEMENT OF A COOLING FLUID

HIGH TEMPERATURE CELEBRATIONS SEALS ARRANGEMENT, IN PARTICULAR FOR GAS TURBINES

COOLING OF THE SIDE PANELS OF TUBINENLEITAPPARATSEGMENTEN

METHOD AROUND THE THICKNESS OF THE SIDE PANEL OF TURBINE CONTROL DEVICE SEGMENTS TO ADAPT AROUND THE COOLING TO IMPROVING

COOLING OF A TURBINE CONTROL DEVICE SEGMENT

SEALING MODULE FOR COMPONENTS OF A TURBO-ENGINE

SHROUD LEADING EDGE COOLING

A cooling device includes a plurality of passages extending through outer platforms of turbine vane segments for directing cooling air in a choked flow condition towards a downstream turbine shroud.

TURBINE BLADE AND GAS TURBINE

Holes 38 and 39 have upstream opening portions 38b and 39b and downstream opening portions 38a and 39a which have a larger cross-sectional area than upstream opening portions 38b and 39b, and are formed at top portion TP of each moving blade. Holes 38 and 39 have tapered shapes T1 and T2 or step portions, and preferably, downstream opening portions 38a and 39a are eccentrically formed toward the moving direction. When tip squealer 37 is formed, hole 38 is formed so that its opening portion is provided at the side surface of tip squealer 37. Without covering the holes for cooling which are formed at the top portion of the turbine blade due to rubbing or the like, the turbine blade is accurately cooled and stably driven.

COOLED TURBINE VANE

... 11 A hollow gas turbine vane is shown enclosing, in spaced relation, a vane insert for receiving cooling air. The insert has a plurality of apertures for selectively directing jets of the cooling air against the internal walls of the vane. A portion of the air is discharged from within the vane chamber through a slit in the trailing edge which contains cooling pins extending transversely thereacross to maintain the slit dimensionally stable and also induce turbulence in the exhausting cooling air to improve its cooling effectiveness. Certain apertures in the insert adjacent the trailing edge are selectively directed to cause jets of the cooling air to impinge at the base of certain of the pins in the inlet area of the slit to promote turbulence in the air entering the slit and adjacent the internal face, thereby maximizing heat transfer from the slit walls to the air.

COOLED TURBINE BLADE

IMPINGEMENT COOLING OF GAS TURBINE BLADES OR VANES

A turbine blade or vane (1) comprises an impingement tube (4) located gener~ally in a radial direction within the blade or vane aerofoil, the impingement tube comprising two parts (4a, 4b), extending into the blade or vane from opposite radial ends thereof and locating against a specially formed rib (51) which extends generally chordwise around the leading edge (2) of the aerofoil.

TURBINE AIRFOIL VANE STRUCTURE

TURBINE STATOR OUTER SHROUD COOLING FINS

A stator of a gas turbine engine is provided. A static shroud of the stator defines a platform having a radially outer surface. An impingement plate has impingement holes extending therethrough, and the impingement plate is radially spaced apart from the radially outer surface of the platform. The impingement holes are configured to direct a high-speed cooling air flow transversally to the radially outer surface of the stator shroud. A plurality of protrusions project away from the radially outer surface of the platform along a protrusion axis. The protrusions have a cruciate cross-sectional shape when viewed in a plane normal to the protrusion axis. A method of cooling a stator of a gas turbine engine is also provided.

HEATSHIELD FOR A GAS TURBINE ENGINE

A heat shield (60) for a gas turbine engine (10), the heat shield (10) comprising a main body (61) having a first surface (70) and a second surface (72), the first surface (70) being exposed to a hot working gas in use, a plurality of walls (74, 76, 78, 80) upstanding from the second surface (72) and an impingement plate (86). The impingement plate (86) is located on top of at least one wall of the plurality of walls (74, 76, 78, 80) and forms a chamber (88) with the second surface (72) and plurality of walls (74, 76, 78, 80) and comprises an array of impingement holes (90). At least one pair of divider walls (92, 94) comprising a first divider wall (92) and a second divider wall (94) formed within the chamber (88) and extending between the impingement plate (86) and the second surface (72). The first divider wall (92) having a length that extends from a first wall (74, 76, 78, 80) of the plurality of walls (74, 76, 78, 80) towards a second wall (74, 76, 78, 80), the second wall (74, 76 ...

VANE ASSEMBLY AND TEMPERATURE CONTROL ARRANGEMENT

COOLABLE WALL

TURBINE NOZZLE WITH INTEGRAL IMPINGEMENT BLANKET

A turbine nozzle segment includes: (a) an arcuate outer band segment (50); ( b) a hollow, airfoil-shaped turbine vane (30) extending radially inward from the outer band segment (50); (c) a manifold cover (54) secured to the outer band such that the manifold cover (54) and the outer band segment (50) cooperatively define an impingeme nt cavity; and (d) an impingement blanket (62) disposed in the impingement cavity, the impingement blanket (62) having at least one impingement hole formed therethrough which is arranged to direct cooling air at the outer band segment (50). A method is provided for impingement cooling the outer band segment (50).

RING SEAL SYSTEM WITH REDUCED COOLING REQUIREMENTS

Aspects of the invention are directed to systems for reducing the cooling requirements of a ring seal in a turbine engine. In one embodiment, the ring seal (54) can be made of a ceramic material, such as a ceramic matrix composite. The ceramic ring seal (54) can be connected to metal isolation rings (40, 42) by a plurality of pins (76, 78). The hot gas face (72) of the ring seal (54) can be coated with a thermal insulating material (74). In another embodiment, a ring seal (120) can be made of metal, but it can be operatively associated with a ceramic heat shield (122). The metal ring seal (120) can carry the mechanical loads imposed during engine operation, and the heat shield (122) can carry the thermal loads. By minimizing the amount of ring seal cooling, the ring seal systems according to aspects of the invention can result in improved engine performance and emissions ...

METHOD AND APPARATUS FOR ACTIVE CLEARANCE CONTROL

The turbomachine includes a compressor, an inner annular casing, and an outer annular casing. The inner annular casing and the outer annular casing define at least one cavity therebetween. The clearance control system includes a manifold system including at least one conduit disposed within the cavities and configured to channel a flow of cooling fluid between the cavities. The clearance control system also includes an impingement system including a header and at least one plenum configured to channel the flow of cooling fluid to the inner annular casing. The conduits configured to channel the flow of cooling fluid to the impingement system. The clearance control system further includes a channel system including at least one channels configured to channel the flow of cooling fluid to the turbomachine. The channels are configured to control the flow of cooling fluid to the manifold system.

TURBINE ENGINE SHUTDOWN TEMPERATURE CONTROL SYSTEM WITH NOZZLE INJECTION FOR A GAS TURBINE ENGINE

A turbine engine shutdown temperature control system (10) configured to limit thermal gradients from being created within an outer casing (12) surrounding a turbine blade assembly (14) during shutdown of a gas turbine engine (16) is disclosed. By reducing thermal gradients caused by hot air buoyancy within the mid-region cavities (18) in the outer casing (12), arched and sway-back bending of the outer casing (12) is prevented, thereby reducing the likelihood of blade tip rub, and potential blade damage, during a warm restart of the gas turbine engine (16). The turbine engine shutdown temperature control system (10) may operate during the shutdown process where the rotor (26) is still powered by combustion gases or during turning gear system operation after shutdown of the gas turbine engine, or both, to allow the outer casing (12) to uniformly, from top to bottom, cool down.

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.

TURBINE NOZZLE WITH IMPINGEMENT BAFFLE

A turbine nozzle apparatus includes: a vane (16) extending between inner and outer bands (12, 14), the interior of the vane (16) being open and communicating with an aperture (30) in the outer band (14), wherein the vane (16) and the bands (12, 14) are part of a monolithic whole of low-ductility material; a metallic baffle (42) inside the vane (16), the baffle (42) having upper (44) and lower ends (46) and a peripheral wall (48) including a plurality of impingement holes (56) defining an interior space, closed off by an end wall (50) at the lower end (46); and a metallic retainer (58) having a body (60) with a shape generally matching the shape of the aperture (30), the body (60) bearing against the upper end (44) of the impingement baffle (42) and being connected to the outer band (14) by a plurality of mechanical fasteners (86, 92).

SMALL EXIT DUCT FOR A REVERSE FLOW COMBUSTOR WITH INTEGRATED COOLING ELEMENTS

The described reverse flow combustor of a gas turbine engine includes inner and outer combustor liners defining a combustor chamber therewithin. A large exit duct and a small exit duct are disposed at downstream ends of the outer and inner liner respectively. The small exit duct includes an annular ring removably mounted to a support element of the gas turbine engine and includes a plurality of cooling elements integrally formed with the annular ring and projecting therefrom into impingement airflow. The cooling elements increase the effective surface area of the inner surface of the annular ring, which is adapted to be cooled by the impingement airflow.

TURBINE SHROUD HANGER WITH DEBRIS FILTER

A turbine shroud hanger (28) apparatus for a gas turbine engine includes: (a) an arcuate shroud hanger (28) having at least one cooling hole (52) passing therethrough, the cooling hole (52) having an inlet and an outlet; and (b) a filter (60) carried by the shroud hanger (28) positioned upstream of the inlet of the cooling hole (52), the filter (60) having a plurality of openings (62) formed therethrough which are sized to permit air flow through the cooling hole (52) while preventing the entn' of debris particles larger than a preselected size into the cooling hole (52).

STEAM AND AIR COOLING FOR STATOR STAGE OF A TURBINE

STEAM AND AIR COOLING FOR STATOR STAGE OF A TURBINE The second-stage nozzles include a plurality of stator vanes 10 having first, second, third, fourth and fifth passages 40, 42, 44, 46 and 48, respectively, for cooling the vanes. The first and fourth passages have a steam inlet along an outer sidewall 14 and a junction box 70 along the inner sidewall 12 for returning steam to the second passage 42. The third passage 44 has a contour corresponding to the contour of the leading edge and impingement steam is directed through openings in a partition 52, cooling the leading edge. Steam flows from the third passage 44 directly into the return passage 42 and also into a channel 64 for cooling the inner sidewall 12. Cooling air flows through fifth passage 48 radially inwardly through the inner sidewall 12 into a cavity 72 in the diaphragm 32 for flow axially outwardly into wheel cavities 86 and 88.

GAS TURBINE VANE WITH A COOLED INNER SHROUD

A gas turbine vane (17) having an inner shroud (26) that is cooled by a portion of the cooling air directed to a cavity (45) between two adjacent rows of discs (55, 56). A portion of the cooling air in the cavity flows through impingement plates (83, 84) and impinges on the inner (98) surface of the inner shroud (26). Another portion of the cooling air flows through a passage (88) in the leading edge (42) of the inner shroud that has a pin fin (89) array for enhanced cooling. The impingement plates form chambers that collect both the impingement air and the pin fin passage air and direct it through holes (92) in the trailing edge (43) of the inner shroud for cooling of the trailing edge. Longitudinal passages (93, 94) along the side of the inner shroud direct the cooling air from the pin fin passage to the trailing edge (43).

COOLED STATIONARY BLADE OF GAS TURBINE

There is provided a cooled stationary blade of a gas turbine in which the portions which can be cooled sufficiently by air is air-cooled, and the portions which is difficult to cool by air is steam-cooled, by which high temperatures can be overcome. In a stationary blade 1, there are formed a serpentine passage 3 in which cooling steam flows and an air passage 10 adjacent to the trailing edge portion and separated from the serpentine passage 3. Also, an outside shroud 4 is formed with an air cooling passage 16 at the outer edge portion and a steam impingement cooling portion 17 and an air impingement cooling portion 18 on the inside of the air cooling passage 16. An inside shroud 11 is provided with an air cooling passage 19 at the outer edge portion and shaped holes 20 formed on the inside of the air cooling passage 19. The air flowing out through the shaped holes 20 performs film cooling.

GAS TURBINE STATIONARY BLADE

Gas turbine stationary blade is improved in shapes of blade leading edge and fillets, in supporting of inserts and in blowing of cooling air, so that blade cooling efficiency is enhanced, insert supporting structure is simplified and clogging of cooling holes is prevented, thus reliability of the stationary blade is enhanced. Passages (23,24) are provided in stationary blade (10) . Front insert (2) is provided in the passage (23) and rear insert (5) in the passage (5) to be supported at two points of insert supporting portions (3a, 3b), (6a, 6b), respectively. Projection (1) is provided at blade leading edge so that portion where thermal load is high is made smaller in size and number of rows of cooling holes (11a) in this portion is lessened. Air blowing holes (4b) on dorsal side of the front insert (2) and film cooling holes (12) of the blade have diameters larger than those of other holes, so that dusts in cooling air are caused to flow out to prevent clogging of the holes. Curved surface ...

Schaufel für eine Strömungsmaschine

Turbinenschaufel

GAS TURBINE.

[...][...] PAR jET D'AIR ET PAR tRANSPIRATION.

AIR-COOLED TURBINE BLADE.

Turbine blade with a cooling device.

Turbinenlaufschaufel (14) mit einem Schaufelblatt (25), das eine Kühlungsanordnung beinhaltet, die mehrere längliche Strömungsdurchgänge (43, 44) zum Aufnehmen und Leiten eines Kühlmittels an einem Weg durch das Schaufelblatt (25) entlang hat. Die Kühlungsanordnung beinhaltet Folgendes: einen zentralen Strömungsdurchgang, der an jeder Seite von wandnahen Strömungsdurchgängen (43, 44) flankiert wird, zu denen ein druckseitiger wandnaher Strömungsdurchgang (43) und ein saugseitiger wandnaher Strömungsdurchgang (44) zählen, ein erstes Loch (46), das den mittleren Strömungsdurchgang mit dem druckseitigen wandnahen Strömungsdurchgang (43) in Strömungsverbindung setzt, ein zweites Loch (46), das den mittleren Strömungsdurchgang mit dem saugseitigen wandnahen Strömungsdurchgang (44) in Strömungsverbindung setzt, und Prallverbinder (48), die den mittleren Strömungsdurchgang mit einem Vorderkanten-Strömungsdurchgang (42) in Strömungsverbindung setzen.

Turbine blade with a chamber for receiving a coolant flow.

Eine Turbinenschaufel (16), umfassend ein Schaufelblatt, das von einer konkav geformten druckseitigen Aussenwand (26) und einer konvex geformten saugseitigen Aussenwand (27) definiert wird, die an Vorder- und Hinterkante (28, 29) entlang miteinander verbunden sind und dazwischen eine radial verlaufende Kammer zur Aufnahme, eines Kühlmittelstroms bilden. Die Turbinenschaufel (16) beinhaltet ferner Folgendes: eine Rippenanordnung (60, 62, 63, 64, 66, 67, 68, 69), die die Kammer in radial verlaufende Strömungsdurchgänge (40) unterteilt, die einen ersten Strömungsdurchgang und einen zweiten Strömungsdurchgang beinhalten, und einen Verbindungsdurchgang, der einen im ersten Strömungsdurchgang gebildeten Einlass mit einem im zweiten Strömungsdurchgang gebildeten Auslass in Strömungsverbindung setzt. Der Verbindungsdurchgang beinhaltet eine abgeschrägte Anordnung relativ zum zweiten Strömungsdurchgang.

Turbine blade with a chamber for receiving a coolant flow.

Eine Turbinenschaufel (16), die ein Schaufelblatt beinhaltet, das von einer konkav geformten druckseitigen Aussenwand (26) und einer konvex geformten saugseitigen Aussenwand (27) definiert wird, die an Vorder- und Hinterkante (28, 29) entlang miteinander verbunden sind und dazwischen eine radial verlaufende Kammer zur Aufnahme eines Kühlmittelstroms bilden. Die Turbinenschaufel (16) beinhaltet ferner eine Rippenanordnung (60, 62, 63, 64, 66, 67, 68, 69), die die Kammer des Schaufelblatts in radial verlaufende Strömungsdurchgänge (40) untergliedert. Ein erster Strömungsdurchgang beinhaltet eine erste Seite, an der Turbulenzerzeuger positioniert sind, wobei die Turbulenzerzeuger jeweils eine abgeschrägte Anordnung aufweisen.

Turbine blade with a chamber for receiving a coolant flow.

Turbinenschaufel (16) mit einem Schaufelblatt, das von einer konkav geformten druckseitigen Aussenwand (26) und einer konvex geformten saugseitigen Aussenwand (27) definiert wird, die an Vorder- und Hinterkanten (28, 29) entlang miteinander verbunden sind und dazwischen eine radial verlaufende Kammer zur Aufnahme eines Kühlmittelstroms bilden. Die Turbinenschaufel (16) beinhaltet eine Rippenanordnung (60, 62, 63, 64, 66, 67, 68, 49), die die Kammer unterteilt und einen Strömungsdurchgang mit einer ersten Seite und einer zweiten Seite definiert. Der Strömungsdurchgang beinhaltet ein Loch, das durch die erste Seite ausgebildet ist. Eine Projektion einer Mittelachse des Lochs durch den Strömungsdurchgang definiert einen Anprallpunkt auf der zweiten Seite des Strömungsdurchgangs, und auf der zweiten Seite des Strömungsdurchgangs ist eine Rückschlagaussparung zum Fassen des Anprallpunkts positioniert.

System with turbine case arrangement.

Ein System enthält eine Turbinengehäuseanordnung, die eine Aussenschale und eine im Wesentlichen konzentrisch in der Aussenschale positionierte Innenschale (30) enthält. Die Innenschale (30) enthält eine von der Aussenschale weg zeigende Innenoberfläche und eine zur Aussenschale hin zeigende Aussenoberfläche (74) und die Aussenoberfläche (74) hat einen oder mehrere Falschflansche (60). Wenigstens einer von dem einen oder den mehreren Falschflanschen (60) enthält eine aus der Aussenoberfläche (74) hervorstehende und zur Aussenschale hin zeigende erste Oberfläche (76) und einen sich zwischen der ersten Oberfläche (76) und der Aussenoberfläche (74) der Innenschale (70) erstreckenden Strömungsumlenkungsabschnitt (78). Der Strömungsumlenkungsabschnitt (78) enthält einen ersten Abschnitt (82); der in einer ersten Umfangsrichtung (84) zwischen der ersten Oberfläche (76) und der Aussenoberfläche (74) divergiert.

Turbine Blade.

Es ist hierin eine Turbinenschaufel (100) offenbart. Die Turbinenschaufel (100) enthält eine Plattform (102) und einen Schaftabschnitt (104), der sich von der Plattform (102) aus radial nach innen erstreckt. Der Schaftabschnitt (104) enthält eine Schlitzfläche (110), eine radiale Dichtungsstiftnut (118), die in der Schlitzfläche (110) ausgebildet ist, und wenigstens ein Kühlloch (122), das in der Schlitzfläche (110) an der radialen Dichtungsstiftnut (118) angeordnet ist.

[...] with improved cooling for a turbomachine.

Ein Turbinensystem enthält einen Überleitungskanal (50) mit einem Einlass und einem Auslass und einen sich zwischen dem Einlass und dem Auslass erstreckenden und eine longitudinale Achse (90), eine radiale Achse (94) und eine tangentiale Achse (92) definierenden Kanaldurchlass. Der Auslass des Überleitungskanals (50) ist gegenüber dem Einlass entlang der longitudinalen Achse (90) und der tangentialen Achse (92) versetzt. Der Kanaldurchlass enthält einen sich von dem Einlass aus erstreckenden stromaufwärts liegenden Abschnitt und einen sich von dem Auslass aus erstreckenden stromabwärts liegenden Abschnitt. Das Turbinensystem enthält ferner eine sich von einer Aussenoberfläche des Kanaldurchlasses aus erstreckende Rippe, wobei die Rippe den stromaufwärts liegenden Abschnitt und den stromabwärts liegenden Abschnitt teilt.

Zigzag cooling of the guide vane final wall.

Ein Gasturbinenleitschaufelabschnitt (100) einer Gasturbine weist eine innere Endwand (114) mit einer Vorderkante (115) auf. Ein Serpentinenkanal (116) ist im Wesentlichen innerhalb der Vorderkante (115) eingerichtet. Der Serpentinenkanal (116) weist einen Einlass und einen Auslass auf. Luft kann an dem Einlass (118) aufgenommen und unter Kühlung der Vorderkante (117) an dem Auslass (119) ausgegeben werden.

Enhanced serpentine cooling the guide vane final wall.

Ein Gasturbinenleitschaufelabschnitt (100) einer Gasturbine weist eine innere Endwand (114) mit einer Vorderkante (115) auf. Ein Serpentinenkanal (116) ist im Wesentlichen innerhalb der Vorderkante (115) eingerichtet. Der Serpentinenkanal (116) weist einen Einlass und einen Auslass auf. Luft kann an dem Einlass (118) aufgenommen und unter Kühlung der Vorderkante (117) an dem Auslass (119) ausgegeben werden.

Serpentine cooling the vane end wall.

Ein Gasturbinenleitschaufelabschnitt (110) einer Gasturbine weist eine innere Endwand (114) mit einer Vorderkante (115) auf. Ein Serpentinenkanal (116) ist im Wesentlichen innerhalb der Vorderkante (115) eingerichtet. Der Serpentinenkanal (116) weist einen Einlass und einen Auslass auf. Luft kann an dem Einlass (118) aufgenommen und unter Kühlung der Vorderkante (117) an dem Auslass (119) ausgegeben werden.

mould-hoop block segment for a gas turbine.

Ein Mantelringblocksegment (100) für eine Gasturbine weist einen Hauptkörper (102) mit einem vorderen Abschnitt (104), einem hinteren Abschnitt (106), einem ersten Seitenabschnitt und einem gegenüberliegenden zweiten Seitenabschnitt auf, die sich axial zwischen dem vorderen Abschnitt (104) und dem hinteren Abschnitt (106) erstrecken. Der Hauptkörper (102) weist ferner eine bogenförmige Verbrennungsgasseite, eine gegenüberliegende Rückseite (114) und eine Kühlkammer (118) auf, die in der Rückseite definiert ist. Ein Kühlplenum (130) und ein Auslassdurchgang (134) sind innerhalb des Hauptkörpers (102) definiert, wobei der Auslassdurchgang (134) eine Fluidübertragungsverbindung aus dem Kühlplenum (130) heraus schafft. Eine Einsatzöffnung (132) erstreckt sich innerhalb des Hauptkörpers (102) durch die Rückseite (114) hindurch in Richtung des Kühlplenums (130). Ein Kühlströmungseinsatz (148) ist innerhalb der Einsatzöffnung (132) angeordnet. Der Kühlströmungseinsatz (148) weist mehrere Kühlströmungsdurchgänge ...

Gas Turbines-Platform Cooling Of.

Ein Deckbandsegment (130) für ein Gehäuse einer Gasturbine enthält einen Körper, der zur Befestigung an dem Gehäuse in der Nähe einer lokalisierten kritischen Prozessstelle (144) in dem Gehäuse eingerichtet ist. Der Körper hat eine Vorderkante (132), eine Hinterkante (134) und zwei Seitenkanten (136, 138). Die kritische Prozessstelle (144) befindet sich zwischen der Vorderkante (132) und der Hinterkante (134), wenn der Körper an dem Gehäuse befestigt ist. Ein Kühlkanal (146, 148, 158, 160) ist in dem Körper entlang einer von den Seitenkanten (136, 138) mit einem Einlass (166, 154) oder einem Auslass (152, 164) in der Nähe der kritischen Prozessstelle (144) definiert. Der Kühlkanal (146, 148, 158, 160) ist gross genug dimensioniert, um die eine an den Kühlkanal (146, 148, 158, 160) angrenzende Seitenkante (136, 138) auf ein gewünschtes Niveau während des Betriebs der Gasturbine zu kühlen. Die kritischen Prozessstellen (144) können auf Temperaturen, Drücke oder andere messbare Merkmale bezüglich ...

High pressure compressor, in particular in a gas turbine.

Die Erfindung betrifft einen Hochdruck-Verdichter (12), insbesondere in einer Gasturbine, welcher einen Verdichter-Rotor (17) aufweist, der unter Ausbildung eines Haupt-Strömungskanals (25) von einem Stator (18, 19) umgeben ist und am Verdichter-Ausgang durch eine sich im Wesentlichen in radialer Richtung erstreckende Endfläche (30) begrenzt ist, an der zur Kühlung in radialer Richtung Kühlluft (21) entlanggeführt wird. Eine verlängerte Lebensdauer wird dadurch erreicht, dass die Endfläche (30) mit ersten Mitteln (23) zur Verbesserung des Wärmeübergangs zwischen der Kühlluft (21) und der Endfläche (30) versehen ist.

Turbine control device with side panel cooling plenum.

Ein Turbinenleitapparatsegment (10) enthält einen Aussen-bandabschnitt (12), einen Innenbandabschnitt (20) und wenigstens eine Leitschaufel, die sich zwischen den Bandabschnitten erstreckt. In einer Fügeseitenfläche (18) wenigstens eines der Bandabschnitte ist ein Kühlplenum (46) definiert, das sich in Querrichtung wenigstens teilweise durch den jeweiligen Bandabschnitt hindurch erstreckt. Erste und zweite Kühlkanäle verlaufen von dem Kühlplenum zu jeweiligen ersten (38) und zweiten (40) Kühlkammern.

System for cooling the wall of a gas turbine combustion chamber.

In einem Ausführungsbeispiel umfasst ein System eine Triebwerkswand (58). Die Triebwerkswand (58) weist eine Kaltseite (72) und eine Heissseite (74) auf. Die Triebwerkswand (58) ist mit einem oder mehreren Verdünnungslöchern (70) ausgebildet, wobei zu jedem der Verdünnungslöcher (70) gehört: eine erste Mündung auf der Kaltseite (72), eine zweite Mündung auf einer Heissseite (74), und ein Abschnitt einer Wärmebarrierenbeschichtung (TBC) (78), die auf der Kaltseite (72) angebracht ist und eine Öffnung aufweist, die die erste Mündung im Wesentlichen umgibt.

Turbine Blade Assembly.

Eine Turbinenschaufelanordnung weist ein Schaufelblatt (210) mit einer Innenwand (231), einer Aussenwand, einer Vorderkante und einer Hinterkante auf. Das Schaufelblatt weist eine oder mehrere Kammern (211217) auf, die sich im Wesentlichen in einer Sehnenrichtung des Schaufelblattes (210) erstrecken. Ein Einsatz (221) weist mehrere Pralllöcher (230) auf, und der Einsatz (221) ist konfiguriert, um in eine der Kammern (211217) eingesetzt zu werden. Der Einsatz (221) ist konfiguriert, um das Schaufelblatt (210) über die mehreren Pralllöcher (230) zu kühlen. Ein Kammerungselement (240) ist nur an dem Einsatz (221) befestigt, wobei das Kammerungselement (240) konfiguriert ist, um einen erhöhten Kühlgasdruck im Inneren eines Grenzbereiches (250), der durch das Kammerungselement (240) definiert ist, relativ zu einem Bereich ausserhalb des Grenzbereiches (250) zu erzielen. Ein Spalt (275) liegt zwischen der Innenwand (231) des Schaufelblattes (210) und dem Kammerungselement (240) vor, wobei der ...

Turbine blade for a turbine section of a gas turbine.

Eine Turbinenlaufschaufel (100) enthält einen Montageabschnitt (102), der teilweise einen Kühlkreislauf (138) im Innern der Turbinenlaufschaufel (100) definiert, und einen Schaufelblattabschnitt (104), der sich von dem Montageabschnitt (102) aus radial nach aussen erstreckt. Der Schaufelblattabschnitt (104) definiert weiter den Kühlkreislauf (138). Die Turbinenlaufschaufel (100) enthält ferner einen Plattformabschnitt (106), der radial zwischen dem Montageabschnitt (102) und dem Schaufelblatt angeordnet ist. Der Plattformabschnitt (106) enthält eine untere Wand (126), eine obere Wand (128), eine vordere Wand (122) eine hintere Wand (124) und ein Paar gegenüberliegender Seitenwände (130). Ein Kühlplenum, das wenigstens teilweise den Kühlkreislauf (138) definiert, ist innerhalb des Plattformabschnitts (106) definiert. Das Kühlplenum ist wenigstens teilweise zwischen der vorderen Wand (122), der hinteren Wand (124), und zwischen dem Paar gegenüberliegender Seitenwände (130) definiert.

[...]foot coolinghot gas path construction unit with.

Die vorliegende Anmeldung stellt eine Heissgaspfadkomponente zur Verwendung in einem Heissgaspfad einer Gasturbine bereit. Die Heissgaspfadkomponente enthält eine innere Wand (140), eine dem Heissgaspfad zugewandte äussere Wand (150), eine Prallwand (160), eine Anzahl von inneren Wandfüssen (180), die zwischen der inneren Wand (140) und der Prallwand (160) angeordnet sind, und eine Anzahl von äusseren Wandfüssen (190), die zwischen der äusseren Wand (150) und der Prallwand (160) angeordnet sind.

System for the cooling of turbine blades.

In einer Ausführungsform enthält ein System (10) eine Turbinenlaufschaufel (40) mit einer radialen Schaufelspitze (62). Das System (10) enthält weiterhin eine Hinterkantennut (94), die in der radialen Schaufelspitze (62) ausgebildet ist und sich zu einer Hinterkante (68) der Turbinenlaufschaufel (40) hin erstreckt. Die Hinterkantennut (94) enthält weiterhin eine erste Gruppe von Kühlkanälen, von denen jede eine erste Nut aufweist, die entlang einer ersten Seitenwand der Hinterkantennut ausgebildet ist, wobei die Nut mit einer zugehörigen ersten Öffnung verbunden ist, die sich durch einen Boden der Hinterkantennut hindurch erstreckt.

Platform cooling arrangement in a turbine rotor blade as well as procedure for the production of such.

Eine Plattformkühlanordnung einer Turbinenrotorschaufel weist eine Plattform auf, die einen in ihr ausgebildeten Kühlkanal (116) hat. Die Plattformkühlanordnung weist auf: Ein Hauptvolumen (132), das etwas innerbords bezüglich der planaren Oberseite angeordnet ist und sich innerhalb der Druckseite und/oder der Saugseite der Plattform von einer hinteren Position in eine vordere Position erstreckt, wobei das Hauptvolumen (132) eine Längsachse aufweist, die ungefähr parallel zu der planaren Oberseite ausgerichtet ist; ein Liefervolumen (134), das sich zwischen dem Hauptvolumen (132) und dem inneren Kühlkanal (116) erstreckt; sowie eine Anzahl von Kühlöffnungen (136), wobei sich jede Kühlöffnung (136) von der druckseitigen und/oder der saugseitigen Spaltfläche (126) weg zu einer Verbindung mit dem Hauptvolumen (132) erstreckt.

Turbine blade, Turbinenrotor, and procedure for the cooling of a turbine blade.

Die Erfindung betrifft eine Turbinenschaufel mit einem Blattkühlungsdurchgang, einem Deckband (120), einer Wandung mit Austrittsöffnung für Kühlfluid aus dem Blattkühlungsdurchgang, einer Deckband-Kühlkammer (142), verbunden mit der Austrittsöffnung, wobei die Austrittsöffnung auf eine Ziel-Wandfläche (134) der Deckband-Kühlkammer (142) gerichtet ist, wodurch die Austrittsöffnung eine Prallöffnung (132) zur Prallkühlung der Ziel-Wandfläche (134) als eine Prallzone definiert und zumindest eine Auslassöffnung (136, 138) zum Ausleiten von verbrauchtem Prallkühlungsfluid aus der Deckband-Kühlkammer.

Turbine blade with shovel point cooling.

Die Erfindung betrifft eine Turbinenlaufschaufel (40) mit einer Schaufelspitze (62). Die Schaufelspitze (62) enthält weiterhin eine Schaufelspitzennut (94), die in der radialen Schaufelspitze (62) ausgebildet ist und sich zu einer Hinterkante (68) der Turbinenlaufschaufel (40) hin erstreckt. Die Schaufelspitzennut (94) enthält weiterhin erste Kühlkanäle, welche sich durch den Nutboden und zur Hinterkante (68) geneigt erstrecken. In Verlängerung der Kühlkanäle können Kühlkanalnuten in den Nutseitenwänden vorgesehen sein.

Heat shield.

Es ist ein Hitzeschild (100) offenbart. Der Hitzeschild (100) enthält eine Basisschicht (102) und eine Abstandsschicht (101). Die Abstandsschicht (101) ist mit der Basisschicht (102) gekoppelt. Die Abstandsschicht (101) definiert mehrere Strömungskanäle (49). Die Basisschicht (102) und die Abstandsschicht (101) ist zur Verknüpfung mit einer Heissgaspfadkomponente (34) konfiguriert.

COOLINGCASH WALL SEGMENT FOR GAS-TURBINE ENGINES.

Gas turbine system

Es ist ein Gasturbinensystem (100) vorgestellt. Das Gasturbinensystem (100) umfasst eine Anzahl von Einlass-Leitschaufeln (150), einen Verdichter (110), eine Turbine (140) und ein Luftbewegungsmittel (155), welches ausgelegt ist, eine CO-Emission aus dem Gasturbinensystem (100) zu reduzieren.

Apparatus and method for cooling a combustor cap

A combustor includes an end cap having a perforated downstream plate and a combustion chamber downstream of the downstream plate. A plenum is in fluid communication with the downstream plate and supplies a cooling medium to the combustion chamber through the perforations in the downstream plate. A method for cooling a combustor includes flowing a cooling medium into a combustor end cap and impinging the cooling medium on a downstream plate in the combustor end cap. The method further includes flowing the cooling medium into a combustion chamber through perforations in the downstream plate.

Annular flow channel section for a turbomachine

An annular flow channel section for a turbomachine is provided. The annular flow channel includes a guide vane ring having a number of guide vanes which are arranged next to each other in the circumferential direction, each guide vane including a vane base, a platform and an airfoil that projects into the flow channel in a radiating pattern. The flow channel is delimited on the platform side by shielding elements, each of which being disposed between two immediately adjacent airfoils, wherein the shielding elements are arranged on the platforms for creating a particularly space-saving flow channel section while creating a gap, and impingement-cooling openings are provided in the platform.

Gas turbine of the axial flow type

In an axial flow gas turbine ( 30 ), a reduction in cooling air mass flow and leakage in combination with an improved cooling and effective thermal protection of critical parts within the turbine stages of the turbine is achieved by providing, within a turbine stage (TS), devices ( 43 - 48 ) to direct cooling air that has already been used to cool, especially the airfoils of the vanes ( 31 ) of the turbine stage (TS), into a first cavity ( 41 ) located between the outer blade platforms ( 34 ) and the opposed stator heat shields ( 36 ) for protecting the stator heat shields ( 36 ) against the hot gas and for cooling the outer blade platforms ( 34 ).

Fluid impingement arrangement

A fluid impingement arrangement comprising a supply manifold and at least one nozzle exit coupled to the supply manifold. The nozzle exit is arranged as a Coanda surface having a restriction and has at least one static pressure tapping that cross-connects two regions of the restriction to induce passive oscillation in a fluid jet passing through the nozzle exit.

Gas Turbine Nozzle Arrangement and Gas Turbine

A sealing element is provided for sealing a leak path between a radial outer platform of a turbine nozzle and a carrier ring for carrying said radial outer platform. The carrier ring has an axially facing carrier ring surface and the radial outer platform has an axially facing platform surface. The carrier ring surface forms a first sealing surface and the platform surface forming a second sealing surface. The first and second sealing surfaces is aligned in a plane with a radial gap between them. The sealing element includes a leaf seal adapted to cover the gap between the first and second sealing surfaces, and an impingement plate for allowing impingement cooling of a radial outer surface of the radial outer platform. The impingement plate is adapted to be fixed to the turbine nozzle. The sealing element may be part of a nozzle arrangement of a gas turbine.

Gas turbine engine

A gas turbine engine including a segment of an annular guide vane assembly is provided. When the engine is used, the segment directs hot combustion gases onto rotor blades of the engine. The segment includes a platform disposed at a side of the segment radially inward/outward with respect to the axis of rotation of the engine. The platform has a trailing edge portion downstream with respect to the flow of hot combustion gases through the segment, the trailing edge portion includes a rail that extends radially inwardly/outwardly from the trailing edge portion. The engine also includes a support and cooling arrangement for supporting the segment and directing a cooling fluid to cool the segment. The arrangement is located radially inward/outward of the platform, and includes a flange part that extends radially outwardly/inwardly from the arrangement. The arrangement further includes a leaf seal and a retaining pin.

Method and apparatus for cooling combustor liner in combustor

A method and apparatus for cooling a combustor liner in a combustor are disclosed. In one embodiment, a combustor is disclosed. The combustor includes a transition piece, and an impingement sleeve at least partially surrounding the transition piece and at least partially defining a generally annular flow path therebetween. The combustor further includes an injection sleeve mounted to one of the transition piece or the impingement sleeve and positioned radially outward of the impingement sleeve, the injection sleeve at least partially defining a flow channel configured to flow working fluid to the flow path. In another embodiment, a method for cooling a combustor liner in a combustor is disclosed. The method includes flowing a working fluid through a flow channel at least partially defined by an injection sleeve, and exhausting the working fluid from the flow channel into a flow path adjacent the combustor liner.

System for measuring parameters of fluid flow in turbomachinery

A system, including, a boundary layer rake, including a rake body, a coolant path extending through the rake body, and a first probe coupled to the rake body, wherein the first probe is configured to measure a first parameter of a first boundary layer flow along a first wall.

Turbine combustion system cooling scoop

A scoop ( 54 ) over a coolant inlet hole ( 48 ) in an outer wall ( 40 B) of a double-walled tubular structure ( 40 A, 40 B) of a gas turbine engine component ( 26, 28 ). The scoop redirects a coolant flow ( 37 ) into the hole. The leading edge ( 56, 58 ) of the scoop has a central projection ( 56 ) or tongue that overhangs the coolant inlet hole, and a curved undercut ( 58 ) on each side of the tongue between the tongue and a generally C-shaped or generally U-shaped attachment base ( 53 ) of the scoop. A partial scoop ( 62 ) may be cooperatively positioned with the scoop ( 54 ).

Turbine wheel for a turbine engine

A turbine wheel for a turbine engine, comprising a disk carrying blades, each having a platform carrying an airfoil and connected by a tang to a root, and sealing and damping sheets housed in the inter-tang spaces, the platforms including projections on their radially internal faces against which the sheets bear radially in operation, in order to define radial clearance and create at least one space between the sheets and the platforms, and the sheets including holes for feeding air to the or each space.

Turbine shroud segment

A turbine shroud segment is metal injection molded (MIM) about a core to provide a composite structure. In one aspect, the core is held in position in an injection mold and then the MIM material is injected in the mold to form the body of the shroud segment about the core. Any suitable combination of materials can be used for the core and the MIM shroud body, each material selected for its own characteristics. The core may be imbedded in the shroud platform to provide a multilayered reinforced platform, which may offer resistance against crack propagation.

Sealing arrangement

A seal assembly for a bearing chamber of a gas turbine engine. The seal assembly includes a seal land having a sealing surface and a non-sealing surface, and at least one non-contact seal member having a sealing surface and a non-sealing surface. The opposing sealing surfaces define a fluid flow path for a sealing fluid such as air from a compressor of the gas turbine engine. The seal assembly includes a sealing fluid cooling arrangement comprising an oil jet configured to provide cooling oil to one or both of the seal member and the seal runner non-sealing surfaces.

Cooled turbine blade for gas turbine engine

A cooled turbine blade for a gas turbine engine including a pressure surface wall, a suction surface wall and a distal wall connecting the pressure surface wall and the suction surface wall, arranged so as to create in the region of the distal end of the blade at least one external cavity forming a bathtub-shaped cavity and at least one internal cavity separated by the distal wall, the blade having at least one opening for the introduction of a flow of cooling air into the external cavity, wherein, on the one hand, at least one part of the distal wall is inclined relative to the verticals of the pressure surface wall and, on the other hand, the opening is created in the vicinity of the distal wall so that the flow of cooling air is directed towards the distal end of the pressure surface wall.

Impingement Cooling System for Use with Contoured Surfaces

The present application provides an impingement cooling system for use with a contoured surface. The impingement cooling system may include an impingement plenum and an impingement plate with a linear shape facing the contoured surface. The impingement surface may include a number of projected area thereon with a number of impingement holes having varying sizes and varying spacings.

Turbine bucket with squealer tip

A turbine bucket having an airfoil is disclosed. The airfoil may include a pressure side wall and a suction side wall extending between a leading edge and a trailing edge. In addition, the airfoil may include a base and a tip disposed opposite the base. The tip may include a tip floor and pressure and suction side tip walls extending outwardly from the tip floor. Moreover, the tip may include an intermediate tip wall extending outwardly from the tip floor between the pressure and suction side tip walls. The intermediate tip wall may define a height that is less than a height of the pressure and/or suction side tip walls.

Gas turbine engine component having platform trench

A component for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a platform that axially extends between a leading edge and a trailing edge and circumferentially extends between a first mate face and a second mate face and a trench disposed on at least one of the first mate face and the second mate face. A plurality of cooling holes are axially disposed within the trench.

Cooled turbine airfoil structures

In accordance with an exemplary embodiment, disclosed is an air-cooled turbine blade having an airfoil shape, including a convex suction side wall, a concave pressure side wall, the walls including an interior surface that defines an interior with the blade, a suction side flow circuit formed within the blade interior, a pressure side flow circuit formed within the blade interior; and a trailing edge pin bank positioned aft of the suction side and pressure side flow circuits. The turbine blade includes a wishbone-shaped architecture at a transition point between the suction side flow circuit and the pressure side flow circuit and the trailing edge pin bank.

INSERTS FOR AIRFOILS OF GAS TURBINE ENGINES

Baffle inserts for airfoils of gas turbine engines are described. The baffle inserts include a baffle insert body having a first side portion and a second side portion, wherein each side portion has a respective end, a first set of vortex generation elements is arranged at the end of the first side portion, and a second set of vortex generation elements is arranged at the end of the second side portion. The first set of vortex generation elements and the second set of vortex generation elements are arranged at an aft end of the baffle insert body. 1. A baffle insert for an airfoil of a gas turbine engine , the baffle insert comprising:a baffle insert body having a first side portion and a second side portion, wherein each side portion has a respective end;a first set of vortex generation elements arranged at the end of the first side portion; anda second set of vortex generation elements arranged at the end of the second side portion,wherein the first set of vortex generation elements and the second set of vortex generation elements are arranged at an aft end of the baffle insert body,wherein the end of the first side portion is joined with the end of the second side portion and the first set of vortex generation elements and the second set of vortex generation elements are arranged in an alternating and overlapping pattern where the first side portion joins with the second side portion.2. The baffle insert of claim 1 , wherein a gap is defined at the aft end of the baffle insert body to allow air to flow aftward through the gap.3. The baffle insert of claim 1 , wherein the baffle insert body is formed from sheet metal.4. The baffle insert of claim 1 , wherein each vortex generation element of at least one of the first set of vortex generation elements and the second set of vortex generation elements has a squared-shape geometry.5. The baffle insert of claim 1 , wherein each vortex generation element of at least one of the first set of vortex generation elements and ...

GAS TURBINE BLADE

A gas turbine blade includes a blade root and a blade aerofoil, a cooling fluid plenum extending inside the gas turbine blade through the blade root, the blade aerofoil and the blade tip, a blade root impingement plate in the cooling fluid plenum inside the blade root and a blade tip impingement plate in the cooling fluid plenum inside the blade tip, the blade tip impingement plate having at least one cooling fluid hole configured and arranged to enable a cooling fluid to flow from the blade tip into the blade aerofoil via the cooling fluid hole or holes, and a pipe extending in the cooling fluid plenum from the blade root impingement plate to the blade tip impingement plate. The blade root impingement plate can direct the cooling fluid from the blade root to the pipe. 1. A gas turbine blade comprising:a blade root and a blade aerofoil, the blade root being attached to a first end of the blade aerofoil;a blade tip attached to a second end of the blade aerofoil;a cooling fluid plenum extending inside the gas turbine blade through the blade root, the blade aerofoil and the blade tip;a blade root impingement plate in the cooling fluid plenum inside the blade root and a blade tip impingement plate in the cooling fluid plenum inside the blade tip, the blade tip impingement plate having at least one cooling fluid hole configured and arranged to enable a cooling fluid to flow from the blade tip into the blade aerofoil via the cooling fluid hole or holes; anda pipe extending in the cooling fluid plenum from the blade root impingement plate to the blade tip impingement plate, and the pipe being configured and arranged to transport the cooling fluid from the blade root to the blade tip; andthe blade root impingement plate being configured and arranged to direct the cooling fluid from the blade root to the pipe.2. The gas turbine blade of claim 1 , wherein the pipe is attached to the blade tip impingement plate and slidably attached to the blade root impingement plate claim 1 ...

Guide vane of a gas turbine engine, in particular of an aircraft engine

A guide vane of a gas turbine engine, in particular of an aircraft engine, which has a pressure-side wall, a suction-side wall, a guide vane root, a guide vane tip, a guide vane leading edge area that is impinged by a cooling air flow of a cooling system, a guide vane trailing edge area that is facing away from the guide vane leading edge area, and at least one channel for conducting a fluid to be cooled arranged in an internal space of the guide vane. At that, during operation of the gas turbine engine, a first part of the cooling air flow flows around a pressure-side wall, and a second part of the cooling air flow flows around the suction-side wall, and a third part of the cooling air flow flows through the internal space including the channel. What is further suggested is a gas turbine engine with at least one such static guide vane.

GAS TURBINE ENGINE COMPONENT HAVING TRANSVERSELY ANGLED IMPINGEMENT RIBS

A component for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a body portion that includes a first wall spaced apart from a second wall and disposed about a centerline axis. At least one rib extends between the first wall and the second wall. The at least one rib extends along a rib axis that is transversely angled relative to the centerline axis. At least one impingement hole extends through the at least one rib. 1. A component for a gas turbine engine , comprising:a body portion that includes a first wall spaced apart from a second wall and disposed about a centerline axis;at least one rib that extends between said first wall and said second wall, wherein said at least one rib extends along a rib axis that is transversely angled relative to said centerline axis; andat least one impingement hole that extends through said at least one rib.2. The component as recited in claim 1 , wherein said body portion is an airfoil of one of a blade and a vane.3. The component as recited in claim 1 , wherein said body portion is part of one of a blade outer air seal (BOAS) and a combustor liner.4. The component as recited in claim 1 , comprising a cooling circuit disposed within said body portion and including at least a first cavity and a second cavity in fluid communication with said first cavity.5. The component as recited in claim 1 , wherein said first wall is a suction side wall and said second wall is a pressure side wall.6. The component as recited in claim 1 , wherein said at least one impingement hole is oriented toward said first wall.7. The component as recited in claim 1 , wherein said at least one impingement hole is oriented toward said second wall.8. The component as recited in claim 1 , wherein said at least one rib includes a first impingement hole that is oriented toward said first wall and a second impingement hole that is oriented toward said second wall.9. The component as recited in claim 1 , ...

GAS TURBINE ENGINE STATOR VANE BAFFLE ARRANGEMENT

A stator vane for a gas turbine engine includes an airfoil that has an exterior wall that provides a cooling cavity. The exterior surface has an interior surface that has multiple pin fins extending therefrom. A baffle is arranged in the cooling cavity and is supported by the pin fins. 1. A stator vane for a gas turbine engine comprising:an airfoil having an exterior wall providing a cooling cavity, the exterior surface has an interior surface having multiple pin fins extending therefrom; anda baffle arranged in the cooling cavity and supported by the pin fins.2. The stator vane according to claim 1 , wherein the baffle is sheet steel.3. The stator vane according to claim 2 , wherein the exterior wall provides pressure and suction sides joined at leading and trailing edges claim 2 , and the baffle includes impingement holes configured to provide impingement cooling fluid onto the exterior wall at the leading edge.4. The stator vane according to claim 2 , wherein the baffle includes a generally smooth outer contour free of protrusions.5. The stator vane according to claim 4 , wherein the outer contour is provided by plastically deformation.6. The stator vane according to claim 4 , wherein cooling holes are provided by at least one of drilling claim 4 , laser drilling claim 4 , or electro discharge machining.7. The stator vane according to claim 1 , wherein a perimeter cavity is provided between the baffle and the exterior wall claim 1 , the pin fins arranged in the perimeter cavity.8. The stator vane according to claim 7 , wherein the perimeter cavity circumscribes the baffle.9. The stator vane according to claim 8 , wherein the pin fins provide the sole support for the baffle in the perimeter cavity.10. The stator vane according to claim 1 , wherein the pin fins are arranged in rows.11. The stator vane according to claim 1 , wherein the pin fins are radially spaced from one another.12. The stator vane according to claim 1 , wherein a rib separates the cooling cavity ...

TURBINE CASE COOLING SYSTEM

A turbine case cooling system including a turbine assembly having an inlet and an outlet and surrounded by a turbine casing. The turbine case cooling system is arranged to selectively impingement cool at least part of the turbine casing. The system includes an annular structure that is radially spaced from the turbine casing and includes a downstream end. The system includes an annular duct that is spaced radially outwardly from the turbine casing and radially inwardly from the annular structure. The duct is sealingly coupled to the turbine casing at a first end towards the turbine inlet, and a second end extends axially towards the downstream end of the annular structure and the turbine outlet. 1. A turbine case cooling system comprising:a turbine assembly having an inlet and an outlet and surrounded by a turbine casing, wherein the turbine assembly comprises two or three turbine stages in axial series; the turbine case cooling system arranged to selectively impingement cool at least part of the turbine casing;an annular structure that is radially spaced from the turbine casing and comprises a downstream end; andan annular duct that is spaced radially outwardly from the turbine casing and radially inwardly from the annular structure; the duct sealingly coupled to the turbine casing at a first end towards the inlet of the axially first turbine stage of the turbine assembly, and a second end of the duct extending axially towards, to or beyond the downstream end of the annular structure and the outlet of the axially last of the turbine stages of the turbine assembly.2. A turbine case cooling system as claimed in wherein the second end of the duct extends axially to or beyond the downstream end of the annular structure.3. A turbine case cooling system as claimed in wherein the second end of the duct extends axially to or beyond the outlet of the turbine assembly.4. A turbine case cooling system as claimed in wherein the annular duct comprises at least one manifold ...

Gas turbine engine combustor liner panel with synergistic cooling features

A liner panel for a combustor of a gas turbine engine includes a multiple of heat transfer augmentors. At least one of the multiple of heat transfer augmentors includes a cone shaped pin.

TURBINE TIP SHROUD ASSEMBLY WITH PLURAL SHROUD SEGMENTS HAVING INTER-SEGMENT SEAL ARRANGEMENT

A shroud assembly for a gas turbine engine includes a plurality of shroud segments that are attached to a shroud support with an inter-segment joint defined between shroud segments. The shroud assembly also includes a cooling flow path cooperatively defined by the shroud support and the first shroud segment. The cooling flow path includes an internal cooling passage within the shroud segments. The cooling flow path includes an outlet chamber configured to receive flow from the internal cooling passage. The shroud assembly additionally includes a seal arrangement that extends across the inter-segment joint. The seal arrangement, the first shroud segment, and the second shroud segment cooperatively define a seal chamber that is enclosed. 1. A shroud assembly for a gas turbine engine comprising:a shroud support that extends arcuately about an axis;a plurality of shroud segments that are attached to the shroud support and that are arranged annularly about the axis at different circumferential positions with respect to the axis, the plurality of shroud segments including a first shroud segment and a second shroud segment, an inter-segment joint defined circumferentially between the first and second shroud segments;a seal arrangement that extends circumferentially across the inter-segment joint; andthe seal arrangement, the first shroud segment, and the second shroud segment cooperatively defining a seal chamber that is enclosed.2. The shroud assembly of claim 1 , wherein the intersegment joint includes a leading edge and a trailing edge that are separated apart at a distance along the axis; andwherein the seal chamber is disposed proximate the trailing edge and is spaced apart at a distance from the leading edge.3. The shroud assembly of claim 1 , wherein the seal arrangement includes a first sealing member and a second sealing member that are arranged in-series with the seal chamber separating the first sealing member and the sealing member apart at a distance.4. The ...

VANE ANGLE SYSTEM ACCURACY IMPROVEMENT

A stator vane angle system includes an engine case, a plurality of stator vanes located at an interior of the engine case. Each stator vane is rotatable about a stator vane axis. A synchronization ring is located at an exterior of the engine case. The synchronization ring is operably connected to each stator vane of the plurality of stator vanes such that movement of the synchronization ring urges rotation of each stator vane of the plurality of stator vanes about their respective stator vane axes. A plurality of impingement openings extend through the engine case from the interior of the engine case to the exterior of the engine case. The plurality of impingement openings are configured to direct flowpath gases from the interior of the engine case to impinge on the synchronization ring, thereby reducing a thermal mismatch between the engine case and the synchronization ring. 1. A stator vane angle system , comprising:an engine case;a plurality of stator vanes disposed at an interior of the engine case, each stator vane rotatable about a stator vane axis;a synchronization ring disposed at an exterior of the engine case, the synchronization ring operably connected to each stator vane of the plurality of stator vanes such that movement of the synchronization ring urges rotation of each stator vane of the plurality of stator vanes about their respective stator vane axes; anda plurality of impingement openings extending through the engine case from the interior of the engine case to the exterior of the engine case, the plurality of impingement openings configured to direct flowpath gases from the interior of the engine case to impinge on the synchronization ring.2. The stator vane angle system of claim 1 , wherein the plurality of impingement openings each have an impingement opening outlet disposed at a same axial location as the synchronization ring.3. The stator vane angle system of claim 1 , wherein the plurality of impingement openings each extend perpendicular to ...

FLUID COOLING SYSTEMS FOR A GAS TURBINE ENGINE

A heat exchanger includes an airfoil configured to be positioned in a coolant stream. The airfoil includes a pressure sidewall and a suction sidewall coupled to the pressure sidewall. The suction sidewall and the pressure sidewall define a leading edge and a trailing edge opposite the leading edge. The leading edge defines an impingement zone wherein the coolant stream is configured to impinge the airfoil. The heat exchanger also includes at least one channel defined within the airfoil between the pressure sidewall and the suction sidewall. The at least one channel is at least partially defined within the impingement zone proximate the leading edge. 1. A heat exchanger comprising: a pressure sidewall; and', 'a suction sidewall coupled to said pressure sidewall, said suction sidewall and said pressure sidewall define a leading edge and a trailing edge opposite said leading edge, said leading edge defines an impingement zone wherein the coolant stream is configured to impinge said airfoil; and, 'an airfoil configured to be positioned in a coolant stream, said airfoil comprisingat least one channel defined within said airfoil between said pressure sidewall and said suction sidewall, said at least one channel at least partially defined within the impingement zone proximate said leading edge.2. The heat exchanger in accordance with claim 1 , wherein said at least one channel is configured to receive a fluid stream such that heat is removed from the fluid stream at least in part through the coolant stream impinging on said leading edge.3. The heat exchanger in accordance with claim 1 , wherein said suction sidewall and said pressure sidewall further define a root portion and a tip portion opposite said root portion claim 1 , said at least one channel comprises:an inlet section extending from said root portion to adjacent said tip portion proximate said leading edge; andan outlet section extending from adjacent said tip portion to said root portion such that said at least ...

Manufacturing of single or multiple panels

A method of manufacturing of a structured cooling panel includes cutting of desized 2D ceramic into tissues; slurry infiltration in the tissues by at least one knife blade coating method; laminating the tissues in a multi-layer panel, with slurry impregnation after each layer, wherein the tissue has combined fibres and/or pre-build cooling holes; drying; de-moulding; sintering the multi-layer panel, wherein part of the combined fibres burns out during the sintering process leaving a negative architecture forming the cooling structure and/or the pre-build cooling holes define the cooling structure; finishing, using of i) post-machine, and/or ii) surface smoothening/rework, and/or iii) coating application, and/or other procedures.

Stepped heat shield for a turbine engine combustor

An assembly is provided for a turbine engine. This turbine engine assembly includes a combustor wall with a shell and a heat shield. The combustor wall defines a quench aperture therethrough. The combustor wall also defines a cavity between the shell and the heat shield. The shell defines a first aperture through which air is directed into the cavity. The heat shield includes a rail that at least partially defines a second aperture configured to direct at least some of the air within the cavity out of the combustor wall and towards the quench aperture.

FILM COOLING A COMBUSTOR WALL OF A TURBINE ENGINE

An assembly is provided for a turbine engine. This turbine engine assembly includes a combustor wall including a shell and a heat shield. The combustor wall defines first and second cavities between the shell and the heat shield. The heat shield defines a first outlet and an elongated second outlet. The first outlet is fluidly coupled with the first cavity. The second outlet is fluidly coupled with the second cavity. The combustor wall defines one of the cavities with a tapered geometry. 1. An assembly for a turbine engine , the assembly comprising:a combustor wall including a shell and a heat shield, the combustor wall defining first and second cavities between the shell and the heat shield;the heat shield defining a first outlet and an elongated second outlet, the first outlet fluidly coupled with the first cavity, and the second outlet fluidly coupled with the second cavity;wherein the combustor wall defines one of the cavities with a tapered geometry.2. The assembly of claim 1 , whereinthe combustor wall defines a quench aperture through the shell and the heat shield; andthe heat shield is configured to direct cooling air from the second cavity through the second outlet and towards the quench aperture.3. The assembly of claim 2 , wherein the heat shield defines a third outlet fluidly coupled with the second cavity claim 2 , and is configured to direct additional cooling air from the second cavity through the third outlet and towards the quench aperture.4. The assembly of claim 3 , wherein the second and the third outlets are staggered.5. The assembly of claim 1 , wherein the heat shield defines an elongated aperture therethrough that at least partially forms the second outlet.6. The assembly of claim 1 , wherein the heat shield defines a plurality of apertures therethrough that at least partially form the second outlet.7. The assembly of claim 1 , wherein the shell and the heat shield converge towards one another thereby at least partially defining the ...

COOLED SEAL ASSEMBLY FOR ARRANGING BETWEEN A STATOR AND A ROTOR

A seal assembly is provided that extends along an axial centerline, and includes a rotor seal element that engages with a stator seal element. The rotor seal element extends radially between an inner element side and an outer element side, and includes a channel, a plurality of first passages and a plurality of second passages. The channel extends radially into the rotor seal element from the inner element side. The first and the second passages are fluidly coupled with the channel. Each of the first passages extends radially through the rotor seal element to a respective first passage outlet. Each of the second passages extends axially through the rotor seal element to a respective second passage outlet. 1. A seal assembly extending along an axial centerline , comprising:an annular stator seal element; and an annular channel extending radially into the rotor seal element from the inner element side;', 'a plurality of first passages fluidly coupled with the channel, the first passages extending radially through the rotor seal element to a respective first passage outlet; and', 'a plurality of second passages fluidly coupled with the channel, the second passages extending axially through the rotor seal element to a respective second passage outlet., 'an annular rotor seal element engaged with the stator seal element, the rotor seal element extending radially between an inner element side and an outer element side, and including'}2. The seal assembly of claim 1 , whereinthe rotor seal element includes a sleeve and a seal seat flange that extends radially from the sleeve to the outer element side;each of the first passages extends radially through the seal seat flange to the respective first passage outlet; andeach of the second passages extends axially through the sleeve to the respective second passage outlet.3. The seal assembly of claim 2 , whereinthe sleeve extends axially between a first element end and a second element end; andthe seal seat flange is axially ...

COMBUSTOR ASSEMBLIES FOR USE IN TURBINE ENGINES AND METHODS OF ASSEMBLING SAME

A combustor assembly for use in a gas turbine engine includes a combustor liner that defines a combustion chamber and includes an axial combustion portion and a curved transition portion. The combustion liner also includes an inner surface and an outer surface and a first plurality of cooling channels defined between the inner and outer surfaces. The combustor assembly also includes a sleeve substantially circumscribing the combustor liner such that an annular cavity is defined between the combustor liner and the sleeve. The sleeve includes a second plurality of cooling channels defined therethrough that are configured to channel a fluid against the combustor liner outer surface. 1. A combustor assembly for use in a gas turbine engine , said combustor assembly comprising:a combustor liner defining a combustion chamber and comprising an axial combustion portion and a curved transition portion, wherein said combustion liner comprises an inner surface and an outer surface and a first plurality of cooling channels defined between said inner and outer surfaces; anda sleeve substantially circumscribing said combustor liner such that an annular cavity is defined between said combustor liner and said sleeve, wherein said sleeve comprises a second plurality of cooling channels defined therethrough, said second plurality of cooling channels configured to channel a fluid against said combustor liner outer surface.2. The combustor assembly in accordance with claim 1 , wherein said combustor liner comprises an inlet and an outlet that define a first length therebetween claim 1 , and wherein said sleeve comprises an inlet and an outlet that define a second length therebetween that is substantially equal to the first length.3. The combustor assembly in accordance with claim 1 , wherein said combustor portion comprises a plurality of dilution openings defined therethrough.4. The combustor assembly in accordance with claim 1 , wherein said first plurality of cooling channels are ...

FLUID COOLING SYSTEMS FOR A GAS TURBINE ENGINE

A heat exchanger includes an airfoil configured to be positioned in a coolant stream. The airfoil includes a pressure sidewall and a suction sidewall coupled to the pressure sidewall. The suction sidewall and the pressure sidewall define a leading edge and a trailing edge opposite the leading edge. The leading edge defines an impingement zone wherein the coolant stream is configured to impinge the airfoil. The heat exchanger also includes at least one channel defined within the airfoil between the pressure sidewall and the suction sidewall. The at least one channel is at least partially defined within the impingement zone proximate the leading edge. 1. A heat exchanger comprising: a pressure sidewall; and', 'a suction sidewall coupled to said pressure sidewall, said suction sidewall and said pressure sidewall define a leading edge and a trailing edge opposite said leading edge, said leading edge defines an impingement zone wherein the coolant stream is configured to impinge said airfoil; and, 'an airfoil configured to be positioned in a coolant stream, said airfoil comprisingat least one channel defined within said airfoil between said pressure sidewall and said suction sidewall, said at least one channel at least partially defined within the impingement zone proximate said leading edge.2. The heat exchanger in accordance with claim 1 , wherein said at least one channel is configured to receive a fluid stream such that heat is removed from the fluid stream at least in part through the coolant stream impinging on said leading edge.3. The heat exchanger in accordance with claim 1 , wherein said suction sidewall and said pressure sidewall further define a root portion and a tip portion opposite said root portion claim 1 , said at least one channel comprises:an inlet section extending from said root portion to adjacent said tip portion proximate said leading edge; andan outlet section extending from adjacent said tip portion to said root potion such that said at least ...

Oxidation activated cooling flow

A flow regulating system for increasing a flow of cooling fluid supplied to a cooling system of a component of a gas turbine system is provided. The flow regulating system includes: a pneumatic circuit embedded within a section of the component, the pneumatic circuit including a set of interconnected pneumatic passages; and a pressure-actuated switch fluidly coupled to the pneumatic circuit. The pressure-actuated switch is activated in response to a formation of a breach in the section of the component and an exposure of at least one of the pneumatic passages of the pneumatic circuit embedded in the section of the component. The activation of the pressure-actuated switch increases the flow of cooling fluid supplied to the cooling system of the component.

HOT SECTION DUAL WALL COMPONENT ANTI-BLOCKAGE SYSTEM

A system for a hot section dual wall component in a gas turbine engine is used to avoid blockage by minimizing particulate deposits. The system includes impingement apertures formed in first wall of a cooling passageway of the dual wall component, and posts included on a second wall of the cooling passageway. The impingement apertures and the posts are respectively aligned opposite each other in the cooling passageway in operative cooperation so that working fluid exhausting into the cooling passageway from the impingement apertures in a first direction is directed to flow in a second direction along the cooling passageway to provide a laminar flow of the working fluid through the cooling passageway in order to minimize deposition of particles. 1. A system comprising:a hot section component of a gas turbine;a dual wall included in the hot section component, the dual wall including a first wall and a second wall, the first wall and the second wall disposed adjacently to define a cooling passageway therebetween;the first wall formed to include a series of impingement apertures providing fluid communication between the cooling passageway and a source of cooling fluid external to the cooling passageway; anda series of posts, each of the respective posts extending from the second wall toward a respective one of the series of impingement apertures, the posts sized and positioned to receive and direct a flow of fluid into the cooling passageway.2. The system of claim 1 , wherein a cross-sectional area of a base of each of the posts is greater than a cross sectional area of a tip of each of the posts claim 1 , the base coupled with second wall claim 1 , and one of the tip and a sidewall of the post aligned with the respective one of the series of impingement apertures.3. The system of claim 2 , wherein the sidewall is tapered between the tip and the base.4. The system of claim 3 , wherein the sidewall is a planar surface between the tip and the base.5. The system of claim 3 ...

COMPARTMENTALIZATION OF COOLING AIR FLOW IN A STRUCTURE COMPRISING A CMC COMPONENT

A structure in a gas turbine engine comprises a spar and a CMC component adjoining the spar and separated from the spar by a cavity supplied by cooling air. At least one rope seal is installed in the cavity within a groove made in the spar to thus compartmentalize the cavity and control the flow of cooling air. 1. A gas turbine engine structure comprising:a static metal component;a CMC component spaced apart from the static metal component and separated therefrom by a cavity receiving cooling air; andat least one rope seal located between the static metal component and the CMC component,the rope seal dividing the cavity into sections to thereby compartmentalize the cavity and cooling air flow is ensured.2. The structure as claimed in claim 1 , wherein the at least one rope seal is pressed into a groove formed in the static metal component.3. The structure as claimed in claim 2 , wherein the groove is formed between two raised landings in the static metal component.4. The structure as claimed in claim 1 , wherein the seal is made with a single rope.5. The structure as claimed in claim 1 , wherein the seal is made with multiple ropes dividing the cavity into a plurality of sections.6. The structure as claimed in claim 1 , wherein the rope seal is made of thermo stable material.7. The structure as claimed in claim 1 , wherein the static metal component includes a vane spar.8. The structure as claimed in claim 1 , wherein the static metal component includes a flow path wall.9. The structure as claimed in claim 1 , wherein the static metal component includes a combustor liner.10. A method of compartmentalizing cooling airflow in a gas turbine engine structure claim 1 , the method comprising the steps of:providing a static metal component, preferably a spar,providing a CMC component adjoining the spar and separated therefrom by a cavity carrying the cooling airflow, anddividing the cavity into at least two sections by installing a rope seal between the spar and CMC ...

IMPINGEMENT COOLING OF TURBINE BLADES OR VANES

A turbine assembly having a hollow aerofoil and impingement device, the aerofoil having a first side wall from leading to trailing edge and a cavity arranged a distance to an inner surface of the cavity for impingement cooling and a flow channel for cooling medium from the leading to trailing edge, the impingement device has first and second pieces arranged side by side, the second piece downstream of the first forming a first flow passage providing passage from one side of the aerofoil towards an opposite side. A blocking element is arranged in the flow channel between the second piece and first side wall at a suction side for blocking flow of cooling medium from leading to trailing edge denying access to a section of the flow channel downstream of the blocking element while directing cooling medium in the first flow passage away from the suction side towards pressure side. 1. A turbine assembly comprising: wherein the hollow aerofoil has at least a first side wall extending from a leading edge towards a trailing edge of the hollow aerofoil and at least a cavity in which in an assembled state of the at least one impingement device in the hollow aerofoil the at least one impingement device is arranged with a predetermined distance in respect to an inner surface of the cavity for impingement cooling of the at least one inner surface and to form a flow channel for a cooling medium extending from the leading edge towards the trailing edge and', 'wherein the at least one impingement device comprises a first piece and a second piece being arranged side by side in an axial direction with the second piece being located viewed in the axial direction downstream of the first piece and with an axial distance in respect to each other forming a first flow passage providing a passage from one side of the aerofoil towards an opposite side of the aerofoil, and, 'a basically hollow aerofoil and at least an impingement device,'}at least a first blocking element, which is arranged in ...

Airfoil platform impingement cooling holes

An airfoil structure for a gas turbine engine includes an airfoil which includes a leading edge and a trailing edge. A platform is located adjacent a first end of the airfoil and includes a core passage that extends through the platform, a mate-face for engaging an adjacent airfoil structure and a set of impingement cooling holes in communication with the core passage that extend through the mate-face adjacent the trialing edge of the airfoil.

ANNULAR RING ASSEMBLY FOR SHROUD COOLING

A gas turbine engine includes an annular casing and a plurality of shroud segments forming an annular shroud. The annular shroud forms with the annular casing an annular cavity therebetween. The annular cavity includes an inlet and an outlet. An annular ring assembly is disposed in the annular cavity between the casing and the shroud and cooperating therewith to provide a first annular chamber and a second annular chamber. The annular ring assembly and a first portion of the shroud form the first annular chamber. The annular ring assembly and a second portion of the shroud form the second annular chamber. The annular ring assembly forms an intermediate annular chamber disposed between the first annular chamber and the second annular chamber. A flow path for coolant air is sequentially defined through the inlet, the first annular chamber, the intermediate annular chamber, the second annular chamber and the outlet.

HEAT TRANSFER DEVICE AND RELATED TURBINE AIRFOIL

Various embodiments include a heat transfer device, while other embodiments include a turbine component. In some cases, the device can include: a body portion having an inner surface and an outer surface, the inner surface defining an inner region; at least one aperture in the body portion, the at least one aperture positioned to direct fluid from the inner region through the body portion; and at least one fluid receiving feature formed in the outer surface of the body portion, the at least one fluid receiving feature positioned to receive post-impingement fluid from the at least one aperture, wherein the at least one aperture does not define any portion of the at least one fluid receiving feature, and the at least one fluid receiving feature segregates relatively higher velocity post-impingement fluid from relatively lower velocity fluid within an impingement cross-flow region. 1. A device , comprising:a body portion having an inner surface and an outer surface, the inner surface defining an inner region;at least one aperture in the body portion, the at least one aperture positioned to direct fluid from the inner region through the body portion; andat least one fluid receiving feature formed in the outer surface of the body portion, the at least one fluid receiving feature positioned to receive post-impingement fluid from the at least one aperture;wherein the at least one aperture does not define any portion of the at least one fluid receiving feature, and the at least one fluid receiving feature segregates relatively higher velocity post-impingement fluid from relatively lower velocity fluid within an impingement cross-flow region.2. The device of claim 1 , wherein the at least one fluid receiving feature further comprises a fluid directing feature.3. The device of claim 2 , wherein the fluid directing feature is formed within the fluid receiving feature.4. The device of claim 2 , wherein the fluid directing feature comprises a turning vane positioned within an ...

TURBOMACHINE COMPONENT HAVING IMPINGEMENT HEAT TRANSFER FEATURE, RELATED TURBOMACHINE AND STORAGE MEDIUM

Various aspects include a turbomachine component, along with a turbomachine and related storage medium. In some cases, the turbomachine component includes: a body defining an inner cavity, the body having an outer surface and an inner surface opposing the outer surface, the inner surface facing the inner cavity; and a mount coupled with the inner surface of the body, the mount including: an impingement baffle coupled with and separated from the inner surface of the body, the impingement baffle including a set of apertures configured to permit flow of a heat transfer fluid therethrough to contact the inner surface of the body; and a reclamation channel connected with the impingement baffle for reclaiming the heat transfer fluid. 1. A turbomachine component comprising:a body defining an inner cavity, the body having an outer surface and an inner surface opposing the outer surface, the inner surface facing the inner cavity;and an impingement baffle coupled with and separated from the inner surface of the body, the impingement baffle including a set of apertures configured to permit flow of a heat transfer fluid therethrough to contact the inner surface of the body; and', 'a reclamation channel connected with the impingement baffle for reclaiming the heat transfer fluid., 'a mount coupled with the inner surface of the body, the mount including2. The turbomachine component of claim 1 , further comprising a set of connectors extending between the inner surface and the mount.3. The turbomachine component of claim 1 , wherein the body defines a portion of an airfoil or a platform.4. The turbomachine component of claim 1 , wherein the mount and the inner surface define a heat transfer region therebetween claim 1 , wherein the body includes at least one film cooling hole extending therethrough.5. The turbomachine component of claim 4 , wherein the reclamation channel is adjacent the impingement baffle and fluidly coupled with the heat transfer region.6. The turbomachine ...

AIRFOIL WITH TIP RAIL COOLING

An apparatus and method for cooling an airfoil tip for a turbine engine can include a blade, such as a cooled turbine blade, having a tip rail extending beyond a tip wall () enclosing an interior for the airfoil at the tip. A plurality of film-holes can be provided in the tip rail. A flow of cooling fluid can be provided through the film-holes from the interior of the airfoil to cool the tip of the airfoil. 1. An airfoil for a turbine engine , the airfoil comprising:a body defining an interior, and extending axially between a leading edge and a trailing edge to define a chord-wise direction and radially between a root and a tip to define a span-wise direction, which terminates in a tip wall and a tip rail extending from the tip wall;at least one cooling passage formed in the interior;at least one cooling cavity provided within the tip rail and comprising at least one cooling conduit defining a flow path having a centerline intersecting with a first surface of the cooling cavity and fluidly coupled to the cooling passage; andat least one film-hole non-aligned in the chord-wise direction with the at least one cooling conduit having an inlet fluidly coupled to the at least one cooling cavity at a second surface opposite the first surface and an outlet provided on an exterior surface of the tip rail.2. The airfoil of wherein the at least one cooling cavity comprises multiple cooling cavities.3. The airfoil of wherein the at least one cooling conduit comprises multiple cooling conduits.4. The airfoil of wherein the at least one cooling conduit comprises a curved cooling conduit.5. The airfoil of wherein the at least one cooling conduit comprises multiple cooling conduits.6. The airfoil of wherein the at least one cooling conduit comprises multiple cooling conduits.7. The airfoil of further comprising a plurality of film-holes provided along a distal end of the tip rail.8. The airfoil of wherein the exterior surface comprises an outer wall and the outlet is fluidly ...

TURBOMACHINE IMPINGEMENT COOLING INSERT

The present disclosure is directed to an impingement insert for a turbomachine. The impingement insert includes an insert body having an inner surface, an outer surface spaced apart from the inner surface, and a thickness extending from the inner surface to the outer surface. The insert body defines a first depression extending from one of the inner surface or the outer surface into the insert body. The first depression has a diameter. The insert body further defines an impingement aperture extending from the first depression through the insert body. The impingement aperture has a length and a diameter. The thickness of the insert body is greater than the length of the impingement aperture and the diameter of the first depression is greater than the diameter of the impingement aperture. 1. An impingement insert for a turbomachine , comprising:an insert body including an inner surface, an outer surface spaced apart from the inner surface, and a thickness extending from the inner surface to the outer surface, the insert body defining a first depression extending from one of the inner surface or the outer surface into the insert body, the first depression having a diameter, the insert body further defining an impingement aperture extending from the first depression through the insert body, the impingement aperture having a length and a diameter,wherein the thickness of the insert body is greater than the length of the impingement aperture and the diameter of the first depression is greater than the diameter of the impingement aperture.2. The impingement insert of claim 1 , wherein the length of the impingement aperture is less than or equal to the diameter of the impingement aperture.3. The impingement insert of claim 1 , wherein the diameter of the first depression is between two times and four times greater than the diameter of the impingement aperture.4. The impingement insert of claim 1 , wherein the diameter of the first depression is at least three times greater ...

IMPINGEMENT TUBES FOR GAS TURBINE ENGINE ASSEMBLIES WITH CERAMIC MATRIX COMPOSITE COMPONENTS

A turbine shroud adapted for use in a gas turbine engine includes a plurality of metallic carrier segments and a plurality of blade track segments mounted to corresponding metallic carrier segments. Impingement tubes direct cooling air onto the blade track segments to cool the blade track segments when exposed to high temperatures in a gas turbine engine. 1. A turbine engine assembly , the assembly comprisinga carrier component comprising metallic material,a supported component comprising ceramic-matrix composite material coupled to the carrier component, each supported component including a runner that extends partway around a central axis and at least two attachment features extending radially outward relative to the central axis from the runner, the at least two attachment features axially spaced apart from one another and circumferentially extending along the runner around the central axis, anda plurality of impingement tubes, each impingement tube extending into one of the carrier components and configured to direct a flow of cooling air toward a radially-outward facing side of the supported component, each impingement tube positioned between one of the at least two attachment features and the runner of the supported component.2. The assembly of claim 1 , wherein each impingement tube includes an elongated body defining an internal plenum claim 1 , an opening formed through one end of the elongated body and extending into the internal plenum claim 1 , and an impingement hole formed through the elongated body and in fluid communication with the internal plenum.3. The assembly of claim 2 , wherein the impingement hole is configured to direct the flow of cooling air toward the radially-outward facing side of the supported component at an angle relative to the radially-outward facing side such that the flow of cooling air is not normal to the radially-outward facing side.4. The assembly of claim 2 , wherein the impingement hole is configured to direct the flow of ...

CMC BOAS ARRANGEMENT

A blade outer air seal assembly includes a blade outer air seal that has a plurality of segments that extend circumferentially about an axis and are mounted in a support structure via a carrier. At least one of the segments have a first hook circumferentially spaced from a second hook. A base portion extends from the first hook to the second hook. The carrier has a cavity on a radially inner surface between the carrier and the base portion. 1. A blade outer air seal assembly , comprising:a blade outer air seal having a plurality of segments extending circumferentially about an axis and mounted in a support structure via a carrier, at least one of the segments having a first hook circumferentially spaced from a second hook, and a base portion extending from the first hook to the second hook; andthe carrier having a cavity on a radially inner surface between the carrier and the base portion.2. The blade outer air seal assembly of claim 1 , wherein the cavity extends at least 50% of a circumferential width of the base portion.3. The blade outer air seal assembly of claim 1 , wherein a plurality of grooves are arranged in the cavity.4. The blade outer air seal assembly of claim 1 , wherein an impingement plate is arranged between the carrier and the base portion.5. The blade outer air seal assembly of claim 4 , wherein the impingement plate has a plurality of orifices.6. The blade outer air seal assembly of claim 5 , wherein the plurality of orifices are configured to communicate cooling air to the at least one segment.7. The blade outer air seal assembly of claim 4 , wherein the impingement plate wraps around a flat portion of the carrier.8. The blade outer air seal assembly of claim 4 , wherein a coating is on a surface between the carrier and the impingement plate.9. The blade outer air seal assembly of claim 4 , wherein the impingement plate is a metallic material.10. The blade outer air seal assembly of claim 4 , wherein the impingement plate is welded to the ...

MODULATED TURBINE COMPONENT COOLING

Features and methods for modulating a flow of cooling fluid to gas turbine engine components are provided. In one embodiment, an airfoil is provided having a flow modulation insert for modulating a flow of cooling fluid received in a cavity of a body of the airfoil. In another embodiment, a shroud is provided comprising a cooling channel for a flow of cooling fluid and an insert that varies in position to modulate the flow of cooling fluid through the cooling channel. In yet another embodiment, a method for operating a gas turbine engine having a cooling circuit for cooling one or more components of the gas turbine engine comprises increasing power provided to the engine and decreasing power provided to the engine to modulate a position of a flow modulation insert located in the cooling circuit and thereby modulate the flow of cooling fluid through the cooling circuit. 119.-. (canceled)20. A method for operating a gas turbine engine , the gas turbine engine including a cooling circuit for providing a flow of cooling fluid to one or more components of the gas turbine engine , the method comprising:increasing power provided to the gas turbine engine; anddecreasing power provided to the gas turbine engine,wherein increasing and decreasing the power provided to the gas turbine engine modulates a position of a flow modulation insert located in the cooling circuit to modulate the flow of cooling fluid through the cooling circuit.21. The method of claim 20 , wherein the flow modulation insert deflects toward an open position when increasing power provided to the gas turbine engine claim 20 , and wherein the flow modulation insert deflects toward a closed position when decreasing power provided to the gas turbine engine.22. The method of claim 20 , wherein the cooling circuit comprises a shroud positioned radially adjacent a plurality of turbine rotor blades claim 20 , and wherein the shroud comprises:a cooling channel for a flow of cooling fluid, the cooling channel having ...

BLADE ASSEMBLY FOR A TURBOMACHINE ON THE BASIS OF A MODULAR STRUCTURE

A blade assembly of a power plant having a modular structure, wherein blade elements include at least one blade airfoil, and at least one footboard mounting part. Blade elements can each have at its one ending a configuration for an interchangeable connection among each other. The connection of the airfoil with respect to other elements can be based on a fixation in radially or quasi-radially extension relative gas turbine axis, wherein the assembling of the blade airfoil in connection with the footboard mounting part is based on a friction-locked bonding actuated by adherence interconnecting, or on use of a metallic and/or ceramic surface fixing blade elements to each other, or on closure configuration with a detachable, permanent or semi-permanent fixation. 1. A rotor blade assembly for a turbomachine having a modular structure , wherein the blade assembly comprises:blade elements each having at least one blade airfoil, and at least one footboard mounting part, wherein the blade elements each have one ending means for interchangeable connection among each other, wherein the connection of the airfoil with respect to other elements is based on a fixation in radial or quasi-radial extension compared to an axis of the rotor of a turbomachine, wherein assembling of the blade airfoil in connection with the footboard mounting part is based on a friction-locked bonding actuated by adherence interconnecting, or the assembling of the blade airfoil in connection with the footboard mounting part is based on the use of a metallic and/or ceramic surface fixing blade elements to each other, or the assembling of the blade airfoil in connection with the footboard mounting part is based on closure means with a detachable, permanent or semi-permanent fixture, wherein the footboard mounting part includes at least two-folded elements, wherein the assembly of separated footboard mounting parts with respect to a foot-side elongated portion of the blade airfoil is conducted with a ...

BLADE WITH PARALLEL CORRUGATED SURFACES ON INNER AND OUTER SURFACES

A blade includes an airfoil body defined by a concave pressure side outer wall and a convex suction side outer wall that connect along leading and trailing edges and, therebetween, form a radially extending chamber for receiving the flow of a coolant. The airfoil body has an inner surface facing the radially extending chamber and an outer surface, a first corrugated surface on a portion of the outer surface, and a second corrugated surface on the inner surface paralleling the first corrugated surface. The corrugated surface on the outer surface of the airfoil provides wake mixing. The blade may also include an integrally formed impingement cooling structure having a third corrugated surface parallel to the second corrugated surface, which is made possible through additive manufacturing. The impingement cooling structure so formed provides improved cooling of the blade. 1. A blade comprising:an airfoil body defined by a concave pressure side outer wall and a convex suction side outer wall that connect along leading and trailing edges and, therebetween, form a radially extending chamber for receiving the flow of a coolant, the airfoil body having an outer surface and an inner surface facing the radially extending chamber;a first corrugated surface on a portion of the outer surface;a second corrugated surface on the inner surface paralleling the first corrugated surface; andan impingement cooling structure positioned within the radially extending chamber, the impingement cooling structure including a portion of an outer surface thereof having a third corrugated surface paralleling the second corrugated surface on the inner surface of the airfoil body.2. The blade of claim 1 , further comprising a plurality of internal supports positioning the impingement cooling structure relative to the radially extending chamber claim 1 , and wherein the airfoil body and the impingement sleeve include a plurality of integral material layers that also include the plurality of internal ...

BLADE WITH INTERNAL RIB HAVING CORRUGATED SURFACE(S)

A blade includes an airfoil body defined by a concave pressure side outer wall and a convex suction side outer wall that connect along leading and trailing edges and, therebetween, form a radially extending chamber for receiving the flow of a coolant, the airfoil body having an outer surface and an inner surface facing the radially extending chamber. A first corrugated surface is on at least a portion of the outer surface of the airfoil body; and a first rib partitions the radially extending chamber, the first rib including a first side and an opposing second side. A second corrugated surface is on at least a portion of at least one of the first and second sides of the first rib. 1. A blade comprising:an airfoil body defined by a concave pressure side outer wall and a convex suction side outer wall that connect along leading and trailing edges and, therebetween, form a radially extending chamber for receiving the flow of a coolant, the airfoil body having an outer surface and an inner surface facing the radially extending chamber;a first corrugated surface on at least a portion of the outer surface of the airfoil body;a first rib partitioning the radially extending chamber, the first rib including a first side and an opposing second side; anda second corrugated surface on at least a portion of at least one of the first and second sides of the first rib.2. The blade of claim 1 , further comprising a pin bank positioned on the first side of the first rib between the first rib and one of the concave pressure side outer wall and the convex suction side outer wall.3. The blade of claim 2 , wherein the second corrugated surface is on at least a portion of the second side of the first rib.4. The blade of claim 3 , further comprising at least one of:a third corrugated surface on the inner surface of the airfoil body, wherein the third corrugated surface parallels the first corrugated surface and the pin bank couples to the third corrugated surface; anda fourth corrugated ...

RING SEGMENT SYSTEM FOR GAS TURBINE ENGINES

A ring segment system () for a gas turbine engine () is disclosed. The ring segment system () may be formed from ring segments () that circumferentially surround a rotor assembly (). The ring segments () may each include a carrier portion () that is coupled to a vane carrier (), and a heat shielding portion () that is detachably coupled to the carrier portion (). The detachable coupling may allow the heat shielding portion () to be uncoupled from the carrier portion () and removed from the gas turbine engine () axially. The ring segments () may further include cooling fluid supply channels () that allow cooling fluid to flow from a radially outward facing backside () of the ring segments () to a radially inward facing front side (). Additionally, the ring segments () may also include ingestion prevention channels () that allow cooling fluid to create a barrier over the gap () between the ring segments () and the adjacent vane (). 110. A turbine engine () , characterized in that:{'b': 40', '20, 'a rotor assembly () having at least one circumferentially aligned row of turbine blades () extending radially outward therefrom;'}{'b': 28', '40', '28', '18, 'a vane carrier () positioned circumferentially around at least a portion of the rotor assembly (), the vane carrier () having at least one circumferentially aligned row of vanes () extending radially inward therefrom; and'}{'b': 50', '20', '28', '50, 'claim-text': [{'b': 34', '28, 'a carrier portion () coupled to the vane carrier (); and'}, {'b': 38', '34', '38', '34', '38', '34', '10, 'a heat shielding portion () positioned radially inward from the carrier portion (), the heat shielding portion () being detachably coupled to the carrier portion (), characterized in that the detachable coupling is configured to allow the heat shielding portion () to be uncoupled from the carrier portion () and removed from the turbine engine () axially.'}], 'one or more ring segments () positioned radially outward from the ...

AIR COOLED COMPONENT FOR A GAS TURBINE ENGINE

An air cooled component for a turbine stage of a gas turbine engine, comprising: a main body having radially inner main gas path wall and a cooling chamber, the main gas path wall separating the main gas path of the turbine stage and the cooling chamber; at least one flange extending from the main body; a cooling cavity enclosed within the flange; and, an inlet conduit extending between and fluidically connecting the cavity and cooling chamber. 1. An air cooled component for a turbine stage of a gas turbine engine , comprising:a main body having radially inner main gas path wall and a first cooling chamber, the main gas path wall separating the main gas path of the turbine stage and the cooling chamber;an attachment system providing radial retention of the component, the attachment system comprising at least one flange extending from the main body;a cooling cavity within the at least one flange; and,an inlet conduit extending between and fluidically connecting the cavity and cooling chamber.2. An air cooled component as claimed in claim 1 , further comprising: at least one outlet conduit extending between and fluidically connecting the cavity and a second cooling chamber.3. An air cooled component as claimed in claim 1 , wherein the first cooling chamber includes first and second sub-chambers claim 1 , the first and second sub-chambers being separated by a partitioning wall having one or more pressure reducing apertures such that the operating pressure of the first and second sub-chambers is different;further comprising: at least one outlet conduit extending between and fluidically connecting the cavity and second sub-chamber.4. An air cooled component as claimed in claim 1 , wherein the cavity is defined in part by the radially inner main gas path wall.51. An air cooled component as claimed in claim 1 , wherein the flange forms part of a coupling for receiving another part of the turbine stage.6. An air cooled component as claimed in claim 1 , wherein the cavity is ...

TURBINE AIRFOIL HAVING FLOW DISPLACEMENT FEATURE WITH PARTIALLY SEALED RADIAL PASSAGES

A turbine airfoil () includes a flow displacement element (A-B, A′-B′) positioned in an interior portion () of an airfoil body () between a pair of adjacent partition walls () and comprising a radially extending elongated main body (). The main body () is spaced from the pressure and suction side walls () and further spaced from one or both of the adjacent partition walls (), whereby a first near wall passage () is defined between the main body () and the pressure side wall (), a second near wall passage () is defined between the main body () and the pressure side wall () and a central channel () is defined between the main body () and a respective one of the adjacent partition walls (). The central channel () is connected to the near wall passages () along a radial extent. One or more radial ribs () are positioned in the central channel () that extend partially across the central channel () between the main body () and the respective adjacent partition wall (). 1. A turbine airfoil comprising:a generally hollow airfoil body formed by an outer wall extending span-wise along a radial direction, the outer wall comprising a pressure side wall and a suction side wall joined at a leading edge and a trailing edge, wherein a chordal axis is defined extending generally centrally between the pressure side wall and the suction side wall,a plurality of radially extending partition walls positioned in an interior portion of the airfoil body and connecting the pressure and suction side walls, the partition walls being spaced along the chordal axis, anda flow displacement element positioned in a space between a pair of adjacent partition walls and comprising a radially extending elongated main body which is spaced from the pressure and suction side walls and spaced from one or both of the adjacent partition walls, whereby a first near wall passage is defined between the main body and the pressure side wall, a second near wall passage is defined between the main body and the ...

TURBINE BLADE COOLING

A gas turbine engine vane or blade comprises a plurality of micro-structures disposed on an internal surface of the outer wall. The micro-structures may comprise an array of fins having a thickness of less than 0.002″, for example. Cooling air may be impinged upon the surface and routed through at least one flow channel to convectively cool the outer wall. Additionally or alternatively, micro-channels may be disposed along a suction side wall and pressure side wall of the vane or blade to convectively cool the respective walls. An embodiment of a vane or blade in accordance with the present invention may be constructed from a plurality of thin metal foils, stacked and bonded together. 1. An air impingement cooled turbine vane , having a plurality of stacked and bonded metal foils defining the turbine vane , the turbine vane having an internal air inlet plenum for receiving an inlet airflow into the turbine vane , at least a first foil of the stacked and bonded metal foils , comprising:an outer wall having an airfoil shape with a leading edge region, a pressure-side, a suction-side and a trailing edge, wherein said pressure-side and said suction-side extend between said leading edge region and said trailing edge;at least a first inner wall spaced from an inside surface of said outer wall and extending along at least a portion of said leading edge region including a stagnation point of the turbine vane for said first foil, wherein said inside surface of said outer wall and an outside surface of said first inner wall at least partially define a first flow channel therebetween; anda plurality of micro-structures extending from said inside surface of said outer wall into said flow channel, wherein air from said internal air inlet plenum passes into a first inlet of said first flow channel and across said plurality of micro-structures.2. The turbine vane of claim 1 , wherein said plurality of micro-structures are integrally formed with said inside surface of said outer ...

COMPONENT FOR A TURBINE ENGINE WITH A COOLING HOLE

An apparatus and method relating an airfoil for a turbine engine with an outer wall bounding an interior and defining a pressure side and a suction side extending between a leading edge and a trailing edge to define a chord-wise direction and extending between a root and a tip to define a span-wise direction. The airfoil includes a first cooling passage extending in the span-wise direction within the interior and a second cooling passage defining an impingement surface and located proximate the first cooling passage, an interior wall separating the first cooling passage from the second cooling passage, and at least one cooling hole passing through the interior wall. 1. An airfoil for a turbine engine , the airfoil comprising:an outer wall bounding an interior and defining a pressure side and a suction side extending between a leading edge and a trailing edge to define a chord-wise direction and extending between a root and a tip to define a span-wise direction;a first cooling passage extending in the span-wise direction within the interior and defining a corner;a second cooling passage defining an impingement surface and located proximate the first cooling passage;an interior wall separating the first cooling passage from the second cooling passage; andat least one cooling hole passing through the interior wall and having an inlet at the corner, an outlet at the second cooling passage, and a connecting passage extending between the inlet and the outlet, with the connecting passage having a curve defined by at least a first portion extending in a first direction, and a second portion extending in a second direction, different from the first direction.2. The airfoil of claim 1 , wherein the connecting passage further comprises a restriction such that a cross-sectional area at the restriction is smaller than a cross-sectional area upstream of the restriction.3. The airfoil of claim 2 , wherein the restriction is at the outlet.4. The airfoil of claim 3 , wherein the ...

Airfoil with Tunable Cooling Configuration

Airfoils, additively manufactured airfoils, and methods of manufacturing airfoils are provided. For example, an airfoil comprises opposite pressure and suction sides that extend axially from a leading edge to a trailing edge and radially spaced apart inner and outer ends. The airfoil also comprises an outer wall defining the pressure and suction sides and leading and trailing edges. A rib extends within the airfoil from the pressure side to the suction side of the outer wall and radially from the inner to the outer end. The airfoil further comprises a first pre-impingement chamber surrounded by a first post-impingement chamber and a first dividing wall segment separating the first pre-impingement and first post-impingement chambers and having a plurality of cooling holes defined therein. The outer wall, rib, and first dividing wall segment are integrally formed as a single monolithic component. 1. An airfoil , comprising:a concave pressure side opposite a convex suction side and an inner end radially spaced apart from an outer end, the pressure side and the suction side extending axially from a leading edge to a trailing edge;an outer wall defining the pressure side, suction side, leading edge, and trailing edge;a rib extending within the airfoil from the pressure side of the outer wall to the suction side of the outer wall, the rib further extending radially from the inner end to the outer end;a first pre-impingement chamber;a first post-impingement chamber surrounding the first pre-impingement chamber;a first dividing wall segment separating the first pre-impingement chamber from the first post-impingement chamber; anda plurality of cooling holes defined in the first dividing wall segment, wherein the outer wall, rib, and first dividing wall segment are integrally formed as a single monolithic component.2. The airfoil of claim 1 , wherein the plurality of cooling holes are defined at a plurality of radial and axial locations along the dividing wall.3. The airfoil ...

Airfoil having angled trailing edge slots

Airfoils for gas turbine engines are described. The airfoils include an airfoil body having a leading edge and a trailing edge extending in a radial direction, a cooling cavity defined within the airfoil body at the trailing edge, and a plurality of angled pedestals arranged along the trailing edge, wherein the plurality of angled pedestals define a plurality of angled trailing edge slots therebetween. Adjacent angled pedestals of the plurality of angled pedestals define a meter section of a respective angled trailing edge slot and a diffuser section of the respective angled trailing edge slot, wherein the meter section is defined by parallel sides of the adjacent angled pedestals, wherein the parallel sides of the adjacent angled pedestals are oriented at a bleed direction that is less than 90° with respect to a feed direction through the cooling cavity.

ANGLED IMPINGEMENT INSERTS WITH COOLING FEATURES

An engine component assembly for impingement cooling. The engine component assembly includes an engine first component having a cooled surface. The engine first component having a flow path on one side of the cooled surface. A second component is a disposed adjacent to the engine first component between the flow path and the engine first component, and has a plurality of openings forming an array through the second component. The cooling flow path passes through the plurality of openings to cool the cooled surface. The second component having a surface facing the cooled surface of the engine first component. A plurality of discrete cooling features that have at least one wall that has a curved cross-section extend from the second component surface into a gap between and toward the cooled surface of the engine first component and defining an array. 1. An engine component assembly for impingement cooling , comprising:an engine first component having a cooled surface;said engine first component having a flow path on one side of said cooled surface;a second component disposed adjacent to said engine first component between said flow path and said engine first component, said second component having a plurality of openings forming an array through said second component, said cooling flow path passing through said plurality of openings to cool said cooled surface;said second component having a surface facing said cooled surface of said engine first component;a plurality of discrete cooling features extending from said second component surface into a gap between and toward said cooled surface of said engine first component and defining an array, wherein a forward wall of at least one cooling feature of the plurality of cooling features has a curved cross-section; andsaid openings extending through said second component at a non-orthogonal angle to said second component surface; andwherein the array comprises a staggered array, the staggered array comprising a first row and ...

TURBINE SHROUD SEGMENT WITH LOAD DISTRIBUTION SPRINGS

A turbine shroud adapted for use in a gas turbine engine includes a plurality of metallic carrier segments and a plurality of blade track segments mounted to corresponding metallic carrier segments. Cooling air is directed onto the blade track segments to cool the blade track segments when exposed to high temperatures in a gas turbine engine. 1. A turbine shroud segment adapted for use in a gas turbine engine having a central axis , the turbine shroud segment comprisinga carrier segment comprising metallic materials, the carrier segment formed to include an attachment-receiving space,a blade track segment comprising ceramic matrix composite materials, the blade track segment formed to include a runner shaped to extend at least partway around the central axis and an attachment portion that extends radially outward from the runner into the attachment-receiving space formed by the carrier segment,an attachment assembly including a first attachment post that extends from the carrier segment through an attachment hole formed in the attachment portion of the blade track segment, a first attachment support arranged inside a cavity formed by the attachment portion of the blade track segment that is shielded by the runner of the blade track segment from the central axis and coupled to the first attachment post to block withdrawal of the attachment post through the attachment hole, and a first spring member arranged outside of the attachment-receiving space and configured to pull the first attachment support radially outward away from the central axis, wherein the first attachment support contacts the attachment portion of the blade track segment at locations spaced apart from the attachment hole formed in the attachment portion of the blade track segment.2. The turbine shroud segment of claim 1 , wherein the first spring member is arranged radially outward of the carrier segment.3. The turbine shroud segment of claim 2 , wherein the first spring member is a coil spring that ...

TURBINE SHROUD INCLUDING PLURALITY OF COOLING PASSAGES

Turbine shrouds for turbine systems are disclosed. The turbine shrouds may include a unitary body including a forward and aft end, an outer surface facing a cooling chamber formed between the unitary body and a turbine casing of the turbine system, and an inner surface facing a hot gas flow path. The shrouds may also include a first cooling passage extending within the unitary body, and a plurality of impingement openings formed through the outer surface of the unitary body to fluidly couple the first cooling passage to the cooling chamber. Additionally, the shrouds may include a second cooling passage and/or a third cooling passage. The second cooling passage may extend adjacent the forward end and may be in fluid communication with the first cooling passage. The third cooling passage may extend adjacent the aft end, and may be in fluid communication with the first cooling passage. 1. A turbine shroud coupled to a turbine casing of a turbine system , the turbine shroud comprising: a forward end;', 'an aft end positioned opposite the forward end;', 'an outer surface facing a cooling chamber formed between the unitary body and the turbine casing; and', 'an inner surface facing a hot gas flow path for the turbine system;, 'a unitary body includinga first cooling passage extending within the unitary body, the first cooling passage including a forward part positioned adjacent the forward end of the unitary body, an aft part positioned adjacent the aft end of the unitary body, and a central part positioned between the forward part and the aft part;a plurality of impingement openings formed through the outer surface of the unitary body to fluidly couple the first cooling passage to the cooling chamber; and a second cooling passage extending within the unitary body adjacent the forward end, the second cooling passage in fluid communication with the first cooling passage, or', 'a third cooling passage extending within the unitary body adjacent the aft end, the third cooling ...

ATTACHMENT BLOCK FOR BLADE OUTER AIR SEAL PROVIDING IMPINGEMENT COOLING

A gas turbine engine includes a compressor section and a turbine section. The turbine section includes at least one turbine rotor having a radially extending turbine blade. The turbine section is rotatable about an axis of rotation. A blade outer air seal is positioned radially outwardly of a radially outer tip of the at least one turbine blade. The blade outer air seal has axially spaced forward and aft portions and a central web between the axially spaced portions. An attachment block is supported on structure within the engine. The attachment block mounts the blade outer air seal. A passage extends into a central chamber within the attachment block, and communicates with cooling holes through a radially inner face of the attachment block to direct cooling air at the central web of the blade outer air seal. A blade outer air seal is also disclosed. 1. A gas turbine engine comprising:a compressor section and a turbine section;said turbine section including at least one turbine rotor having a radially extending turbine blade, and said turbine section being rotatable about an axis of rotation, and a blade outer air seal positioned radially outwardly of a radially outer tip of said at least one turbine blade, said blade outer air seal having axially spaced forward and aft portions and a central web between said axially spaced portions; andan attachment block supported on structure within said engine, and said attachment block mounting said blade outer air seal, and a passage extending into a central chamber within said attachment block, and communicating with cooling holes through a radially inner face of said attachment block to direct cooling air at said central web of said blade outer air seal.2. The gas turbine engine as set forth in claim 1 , wherein said central chamber including circumferentially extending fingers communicating with said passage claim 1 , such that air may pass into said passage claim 1 , and then circumferentially into said circumferentially ...

Attachment block for blade outer air seal providing convection cooling

A gas turbine engine includes a compressor section and a turbine section. The turbine section includes at least one turbine rotor having a radially extending turbine blade. The turbine section is rotatable about an axis of rotation. A blade outer air seal is positioned radially outwardly of a radially outer tip of the at least one turbine blade. The blade outer air seal has axially spaced forward and aft portions and a central web between the axially spaced portions. An attachment block is supported on structure within the engine. The attachment block mounts the blade outer air seal. A passage extends into a chamber within the attachment block, and communicates to circumferential edges of the attachment block to direct cooling air along the central web of the blade outer air seal. A blade outer air seal assembly is also disclosed.

SELF-REGULATING BACK-SIDE PRESSURIZATION SYSTEM FOR THERMAL INSULATION BLANKETS

High-pressure fan duct bleed air is used to pressurize a cavity between the fan duct inner wall and the inner wall thermal insulation blankets. The cavity is pressurized to prevent hot air from the nacelle core compartment from flowing under the insulation blankets and degrading the fan duct inner wall. Pressure regulating valves (PRV) allow better control of the cavity pressure during different phases of the flight profile and under different levels of insulation blanket seal degradation by passively controlling exit area from the cavity based on an established pressure limit. Moreover, the pressurization system can be implemented as a passive cooling system by increasing the mass flow rate into the cavity and then the core compartment to a level suitable for core compartment cooling. The cooling air can be vented at the forward end of the insulation blanket assembly to provide core compartment ventilation flow, or vented through dedicated ports in the insulation blanket for targeted core compartment component cooling. 1400. An aircraft engine () , comprising:{'b': '402', 'an engine core ();'}{'b': 404', '202, 'a fan duct () including an inner wall ();'}{'b': 1', '202, 'a first orifice (A) through the inner wall ();'}{'b': 206', '202', '202', '402, 'an insulation blanket () coupled to the inner wall () so as to shield the inner wall () from heat generated in the engine core ();'}{'b': 2', '206, 'a second orifice (A) through the insulation blanket ();'}{'b': 206', '202', '206', '206', '406', '206', '1', '404, 'a cavity () bounded by the inner wall () and the insulation blanket (), the cavity () receiving air () inputted into the cavity () through the first orifice (A) from the fan duct ();'}{'b': 210', '404', '402', '210', '408', '402', '410', '206, 'a core compartment () within the fan duct () and housing the engine core (), the core compartment () having a first boundary () with the engine core () and second boundary () with the insulation blanket (); and'}{'b': ...

Dual-wall impingement, convection, effusion combustor tile

A gas turbine engine includes a combustor having a dual-wall impingement convention effusion combustor tile assembly. The dual-wall tile assembly provides a cooling air flow channel and attachments for securing the tile to the cold skin liner of the combustor. Cooling is more efficient in part due to the dual wall construction and in part due to reduced parasitic leakage, and the design is less sensitive to attachment features which operate at lower temperatures.

TURBINE SHROUD WITH CERAMIC MATRIX COMPOSITE SEAL SEGMENTS MOUNTED TO METALLIC CARRIERS

A turbine shroud assembly of a gas turbine engine includes a carrier, a blade track segment and a mounting assembly. The carrier is made from metallic materials. The blade track segment is made from ceramic matrix composite materials. The mounting assembly is configured to couple the blade track segment to the carrier. 1. A turbine shroud segment adapted to extend part-way around a turbine wheel in a gas turbine engine , the turbine shroud segment comprisinga carrier made from metallic materials and adapted to be coupled to a turbine case,a blade track segment made from ceramic matrix composite materials, the blade track segment shaped to include a runner that extends partway around a central reference axis to define in-part a primary gas path and an attachment feature that extends radially-outwardly from the runner, anda mounting assembly configured to couple the blade track segment to the carrier, the mounting assembly including (i) at least one hanger with a bracket that engages the attachment feature of the blade track segment and a shaft that extends through an aperture in the carrier, and (ii) a heat shield coupled to the bracket of the hanger and having a shield panel arranged radially between the runner of the blade track segment and the bracket to resist the transmission of heat from the runner to the bracket.2. The turbine shroud segment of claim 1 , wherein the mounting assembly includes a plurality of hangers each having a bracket that engages the attachment feature of the blade track segment and a shaft that extends through an aperture in the carrier claim 1 , and wherein the shield panel of the heat shield is arranged radially between the runner and the brackets of each of the plurality of hangers to resist the transmission of heat from the runner to the brackets of each of the plurality of hangers.3. The turbine shroud segment of claim 2 , wherein the heat shield includes a plurality of attachment clips that extend from the heat shield to engage with ...

METHOD FOR COOLING A TURBO-ENGINE COMPONENT AND TURBO-ENGINE COMPONENT

Disclosed is a turbo-engine component and a method for cooling a turbo-engine component. The method includes guiding a working fluid flow along a hot gas side surface of a wall of the component and in a main working fluid flow direction, discharging a coolant discharge flow at the hot gas side surface from a coolant discharge duct provided in the wall, and supplying a coolant supply flow to the coolant discharge duct and through a coolant supply path. The method also includes discharging the coolant supply flow into the coolant discharge duct as a free jet oriented across a cross section of the coolant discharge duct, and directing the free jet onto an inner surface section of the coolant discharge duct, thus effecting impingement cooling of the inner surface section. 1. A method for cooling a turbo-engine component , the method comprising:guiding a working fluid flow along a hot gas side surface of a wall of the component and in a main working fluid flow direction,discharging a coolant discharge flow at the hot gas side surface from a coolant discharge duct provided in the wall,supplying a coolant supply flow to the coolant discharge duct and through a coolant supply path,discharging the coolant supply flow into the coolant discharge duct as a free jet oriented across a cross section of the coolant discharge duct, anddirecting the free jet onto an inner surface section of the coolant discharge duct, thus effecting impingement cooling of the inner surface section.2. The method according to claim 1 , comprising:guiding the coolant supply flow through a means for generating a free jet and discharging the free jet from said means for generating.3. The method according to claim 1 , comprising:discharging the coolant discharge flow in a direction inclined with respect to a normal of the hot gas side surface at: discharge location, whereby the coolant discharge duct is inclined with respect to said normal thus having a first inner surface section disposed towards the hot ...

TURBINE AIRFOILS WITH MICRO COOLING FEATURES

A blade used in a gas turbine engine includes a pair of pedestals and an airfoil coupled between the pedestals. The airfoil includes cooling features to cool the airfoil. 1. An airfoil for use in a gas turbine engine and having a pressure side and a suction side , the airfoil comprisinga spar formed to define a cooling air plenum adapted to receive a flow of cooling air, anda skin coupled to an exterior surface of the spar and positioned to at least partially cover the spar along the pressure side and the suction side,wherein at least one axially extending groove is formed in the exterior surface of the spar on the pressure side that defines at least one cooling passageway between the spar and the skin, at least one inlet port is formed in the spar adjacent a trailing edge of the spar, the at least one inlet port is in fluid communication with the cooling air plenum and the at least one cooling passageway to pass the flow of cooling air into the at least one cooling passageway from the cooling air plenum, at least one outlet port is formed through the skin on the pressure side and axially forward of the at least one inlet port, the at least one outlet port is configured to pass the flow of cooling air from the at least one cooling passageway to an exterior of the airfoil, and at least one turbulator is positioned within the at least one cooling passageway.2. The airfoil of claim 1 , wherein the at least one axially extending groove includes a plurality of axially extending grooves formed in the exterior surface of the spar on the pressure side and radially spaced apart from one another to define a plurality of stand-offs therebetween claim 1 , wherein the plurality of axially extending grooves define a plurality of cooling passageways between the spar and the skin claim 1 , and wherein the at least one inlet port includes a plurality of inlet ports formed in the spar adjacent a trailing edge of the spar.3. The airfoil of claim 2 , wherein the skin is bonded to at ...

Article, airfoil component and method for forming article

An article is disclosed including a manifold, an article wall, a post-impingement cavity and a plurality of post-impingement partitions. The manifold includes an impingement wall defining a plenum and a plurality of impingement apertures. The article wall includes a plurality of external apertures. The post-impingement cavity is disposed between the manifold and the article wall, and is arranged to receive a fluid from the plenum through the plurality of impingement apertures and exhaust the fluid through the plurality of external apertures. The plurality of post-impingement partitions divide the post-impingement cavity into a plurality of sub-cavities, and hermetically separate the plurality of sub-cavities from one another. The impingement wall, article wall and plurality of post-impingement partitions are integrally formed as a single, continuous article. The article may be an airfoil component. A method for forming the article includes forming a single, continuous object by an additive manufacturing technique.

TURBINE VANE REAR INSERT SCHEME

An internally cooled turbine vane for a gas turbine engine has coolant flow channels between the interior walls of the vane and an insert, where the channels serve to convey a portion of the cooling air flow from a pressure side chamber to a suction side chamber. The turbine vane defines a radially extending passage with a dividing wall defining a front section and a rear section; the rear section having interior walls spaced apart from an insert to define the pressure side chamber and the suction side chamber. The insert may receive cooling air and conveys the cooling air into the pressure side chamber and the suction side chamber. A front surface of the insert or a rear surface of the dividing wall may have a clearance gap and an air flow channel communicating between the pressure side chamber and the suction side chamber. 1. A turbine vane comprising:a pressure side; a suction side; and a hollow front section separated from a hollow rear section by a dividing wall;the hollow rear section having interior walls spaced apart from a hollow insert by protrusions to define a pressure side chamber and a suction side chamber;the hollow insert adapted to be in fluid communication with a source of pressurized cooling air and having openings for conveying cooling air into the pressure side chamber and the suction side chamber;a front surface of the hollow insert and a rear surface of the dividing wall being spaced apart defining a gap; andat least one of: the front surface of the insert; and the rear surface of the dividing wall, including a channel communicating between the pressure side chamber and the suction side chamber.2. The turbine vane according to claim 1 , wherein the channel comprises a recess formed within the rear surface of the dividing wall.3. The turbine vane according to claim 1 , wherein the channel comprises a dimple within the front surface of the insert.4. The turbine vane according to claim 1 , comprising two channels radially spaced apart.5. The ...

TURBINE COMPONENTS AND METHOD FOR FORMING TURBINE COMPONENTS

Turbine components are disclosed including a component wall defining a constrained portion, a manifold having an impingement wall, and a post-impingement cavity disposed between the manifold and the component wall. The impingement wall includes a wall thickness and defines a plenum and a tapered portion. The tapered portion tapers toward the constrained portion and includes a plurality of impingement apertures and a wall inflection. The wall inflection is disposed proximal to the constrained portion, and the tapered portion is integrally formed as a single, continuous object. The wall inflection may include an inflection radius of less than about 3 times the wall thickness of the impingement wall, or the tapered portion may include a consolidated portion with the impingement wall extending across the plenum. A method for forming the turbine component is also disclosed, including forming the tapered portion as a single, continuous tapered portion by an additive manufacturing technique. 1. A turbine component , comprising:a component wall including a plurality of external apertures and defining a constrained portion;a manifold disposed within the component wall, the manifold including an impingement wall, the impingement wall including a wall thickness and defining a plenum and a tapered portion, the tapered portion tapering toward the constrained portion and including a plurality of impingement apertures and a wall inflection, the wall inflection being disposed proximal to the constrained portion; anda post-impingement cavity disposed between the manifold and the component wall, the post-impingement cavity arranged to receive a fluid from the plenum through the plurality of impingement apertures and exhaust the fluid through the plurality of external apertures, the post-impingement cavity including an enervated zone disposed between the tapered portion and the constrained portion,wherein the tapered portion is integrally formed as a single, continuous object, and the ...

IMPACT-COOLING TUBULAR INSERT FOR A TURBOMACHINE DISTRIBUTOR

A tubular ventilation sleeve for a turbomachine distributor, in particular for an aircraft, the sleeve having a generally elongate shape along an axis (A-A) and including a perforated tubular wall around said axis, one of the axial ends of the sleeve being open and the other being closed by a bottom wall, wherein it further includes support beams when the sleeve is made by additive manufacturing, the beams extending inside the sleeve between the tubular wall and the bottom wall and having a longitudinal cross-section with a generally triangular shape, two sides of which are respectively connected to the tubular wall and the bottom wall and the last side of which is free and extends inside the sleeve, perforations in the tubular wall being provided between the support beams. 1. A tubular ventilation sleeve for a turbomachine distributor , in particular for an aircraft , the sleeve having a generally elongated shape along an axis and comprising a tubular wall perforated around this axis , one of the axial ends of the sleeve being open and the other being closed by a bottom wall , wherein it furthermore comprises support beams when the sleeve is made by additive manufacturing , these beams extending inside the sleeve , between the tubular wall and the bottom wall and having a longitudinal cross-section with a generally triangular shape , two sides of which are respectively connected to the tubular wall and to the bottom wall , and the last side of which is free and extends inside the sleeve , perforations in the tubular wall being provided between the support beams.2. The sleeve according to claim 1 , wherein each beam has claim 1 , at the level of its side connected to the bottom wall claim 1 , a greater thickness of material than the rest of the beam.3. The sleeve according to any claim 1 , wherein the beams cover substantially the entire internal surface of the bottom wall.4. The sleeve according to any claim 1 , wherein the beams are separated into two series which ...

INTERIOR COOLING CONFIGURATIONS IN TURBINE BLADES AND METHODS OF MANUFACTURE RELATING THERETO

A rotor blade for use in combustion turbine engine. The rotor blade may include: an airfoil assembled from two radially stacked non-integral sections in which a body section resides inboard of a cap section; an outboard tip of the airfoil that is enclosed by a tip plate having a tip rail; and a cooling configuration that includes cooling channels for directing a coolant through the rotor blade. Each of the cooling channels may include segments, in which: a supply segment extends radially through the airfoil, the supply segment being radially defined between a floor and ceiling; a rail segment extends through an interior of the tip rail; and a connecting segment extends between the supply segment and rail segment. For each of the one or more cooling channels, the ceiling of the supply segment may be defined within the cap section of the airfoil. 1. A rotor blade for use in combustion turbine engine , the rotor blade comprising:an airfoil assembled from two radially stacked non-integral sections in which a body section resides inboard of a cap section;an outboard tip of the airfoil that is enclosed by a tip plate and, formed along a periphery of the tip plate, a tip rail; and a supply segment extends radially through the airfoil, the supply segment being radially defined between a floor, which comprises an inboard boundary, and a ceiling, which comprises an outboard boundary;', 'a rail segment extends through an interior of the tip rail; and', 'a connecting segment extends between the supply segment and the rail segment, the connecting segment comprising an upstream port, which connects to the supply segment, and a downstream port, which connects to the rail segment;, 'a cooling configuration that includes one or more cooling channels for receiving and directing a coolant through an interior of the rotor blade, each of the one or more cooling channels comprising fluidly communicative segments, in whichwherein, for each of the one or more cooling channels, the ceiling ...

SHELL AND TILED LINER ARRANGEMENT FOR A COMBUSTOR

A combustor adapted for use in a gas turbine engine is disclosed. The combustor includes a metallic shell forming a cavity and a ceramic liner arranged in the cavity of the metallic shell. The ceramic liner defines a combustion chamber in which fuel is burned during operation of a gas turbine engine. The ceramic liner includes a plurality of ceramic tiles mounted to the metallic shell and arranged to shield the metallic shell from heat generated in the combustion chamber. 1. A combustor for use in a gas turbine engine , the combustor comprisingan annular metallic shell forming an annular cavity around a central reference axis, andan annular liner arranged in the annular cavity of the annular metallic shell along an annular combustion chamber inside the annular metallic shell, the annular liner including a plurality of ceramic tiles arranged around the central reference axis and located to shield an axially-extending wall of the annular metallic shell from burning fuel in the combustion chamber,wherein a first ceramic tile of the plurality of ceramic tiles includes a first shelf that cooperates with a second shelf in a circumferentially adjacent second ceramic tile of the plurality of ceramic tiles to form a ship lapped joint that provides a labyrinth-like seal between the first ceramic tile and the second ceramic tile.2. The combustor of claim 1 , wherein the first shelf is arranged in direct confronting relation to the combustion chamber and the second shelf is shielded by the first shelf from the combustion chamber.3. The combustor of claim 2 , wherein the first shelf has a radial thickness that is less than a maximum radial thickness of the first ceramic tile and the second shelf has a radial thickness that is less than a maximum radial thickness of the second ceramic tile.4. The combustor of claim 3 , wherein the radial thickness of the first shelf added to the radial thickness of the second shelf is substantially equal to the maximum radial thickness of the ...

IMPINGEMENT RING ELEMENT ATTACHMENT AND SEALING

A turbine includes a supporting structure which extends along a circumferential direction of the turbine, wherein the supporting structure has a groove to guide cooling air. The groove extends along the circumferential direction. The turbine has a first impingement ring element having a plurality of first cooling holes and a second impingement ring element having a plurality of second cooling holes, wherein the first and second impingement ring elements are mounted one after another along the circumferential direction to the groove such that the groove is covered by the first and second impingement ring elements. A first coupling ring element is arranged between the first and second impingement ring elements such that the groove is covered by the first coupling ring element between the first impingement ring element and the second impingement ring element. The first coupling ring element forms a sliding contact with the second impingement ring element. 1. A turbine comprisinga supporting structure which extends along a circumferential direction of the turbine, wherein the supporting structure comprises a groove through which cooling air is guidable, wherein the groove extends along the circumferential direction,a first impingement ring element comprising a plurality of first cooling holes,a second impingement ring element comprising a plurality of second cooling holes,wherein the first impingement ring element and the second impingement ring element are mounted one after another along the circumferential direction to the groove such that the groove is covered by the first impingement ring element and the second impingement ring element, anda first coupling ring element which is arranged between the first impingement ring element and the second impingement ring element such that the groove is covered by the first coupling ring element between the first impingement ring element and the second impingement ring element, andwherein the first coupling ring element forms a ...

TURBOMACHINE COMPONENT WITH FLOW GUIDES FOR FILM COOLING HOLES IN FILM COOLING ARRANGEMENT

A turbomachine component having film cooling arrangement includes a cooling passage, an external wall having an outer surface to be positioned in a hot gas path and an inner surface forming a part of the cooling passage, film cooling holes formed through the external wall, and a flow guide arrangement having a flow guide corresponding to one of the film cooling holes. Each film cooling hole has an inlet at the inner surface and an outlet at the outer surface. The inlet receives a cooling fluid from the cooling passage and the outlet releases it over the outer surface. The flow guide positioned at the inlet of the corresponding film cooling hole on the inner surface redirects within the cooling passage a flow of the cooling fluid such that the flow makes a U-turn before entering the inlet of the corresponding film cooling hole in a reversed flow. 1. A turbomachine component having film cooling arrangement for a gas turbine engine , the turbomachine component comprising:a cooling passage defined within the turbomachine component;an external wall of the turbomachine component, wherein the external wall comprises an outer surface adapted to be positioned in a hot gas path of the gas turbine engine and an inner surface forming a part of the cooling passage;a plurality of film cooling holes formed through the external wall of the turbomachine component, the film cooling holes being spaced apart over at least a part of the external wall, wherein each of the film cooling holes has an inlet and an outlet, and wherein the inlet is positioned on the inner surface of the external wall in the cooling passage and is adapted to receive a cooling fluid flowing through the cooling passage and to direct the cooling fluid towards the outlet, and wherein the outlet is positioned on the outer surface of the external wall and is adapted to release the cooling fluid over the outer surface of the external wall to form a cooling film over at least a part of the outer surface of the external ...

GAS TURBINE ENGINE STATOR VANE BAFFLE ARRANGEMENT

A method of flowing cooling fluid through a stator vane in a gas turbine engine includes the step of providing an airfoil that has an exterior wall that provides a cooling cavity. The exterior surface has an interior surface that has multiple pin fins that extend therefrom. A baffle is arranged in the cooling cavity and supported by the pin fins. A perimeter cavity is provided between the baffle and the exterior wall. The pin fins are arranged in the perimeter cavity. Cooling fluid flows through a region in the perimeter cavity. The pin fins are arranged in the region having a low Reynolds number and through which the cooling fluid 1. A method of flowing cooling fluid through a stator vane in a gas turbine engine , comprising the steps of:providing an airfoil having an exterior wall providing a cooling cavity, the exterior surface has an interior surface having multiple pin fins extending therefrom;providing a baffle arranged in the cooling cavity and supported by the pin fins, wherein a perimeter cavity is provided between the baffle and the exterior wall, the pin fins arranged in the perimeter cavity; andflowing cooling fluid through a region in the perimeter cavity, wherein the pin fins are arranged in the region having a low Reynolds number and through which the cooling fluid flows.2. The method according to claim 1 , wherein the baffle is sheet steel.3. The method according to claim 2 , wherein the exterior wall provides pressure and suction sides joined at leading and trailing edges claim 2 , and the baffle includes impingement holes configured to provide impingement cooling fluid onto the exterior wall at the leading edge.4. The method according to claim 2 , wherein the baffle includes a generally smooth outer contour free of protrusions.5. The method according to claim 4 , wherein the outer contour is provided by plastic deformation.6. The method according to claim 4 , wherein cooling holes are provided by at least one of drilling claim 4 , laser drilling ...

FLARED CENTRAL CAVITY AFT OF AIRFOIL LEADING EDGE

A blade includes an airfoil defined by a pressure side outer wall and a suction side outer wall connecting along leading and trailing edges and form a radially extending chamber for receiving a coolant flow. A rib configuration may include: a leading edge transverse rib connecting the pressure side outer wall and the suction side outer wall and partitioning the radially extending chamber into a leading edge passage within the leading edge of the airfoil and a central passage adjacent to the leading edge passage. One or both camber line ribs connect to a corresponding pressure side outer wall and suction side outer wall at a point aft of the leading edge transverse rib causing the central passage to extend towards one or both of the pressure side outer wall and the suction side outer wall, resulting in a flared center cavity aft of the leading edge. 1. A blade comprising an airfoil defined by a concave pressure side outer wall and a convex suction side outer wall that connect along leading and trailing edges and , therebetween , forming a radially extending chamber for receiving a flow of a coolant , the blade further comprising:{'claim-text': ['a leading edge transverse rib connecting the pressure side outer wall and the suction side outer wall and partitioning the radially extending chamber into a leading edge passage within the leading edge of the airfoil and a central passage adjacent to the leading edge passage;', 'a suction side camber line rib connected to the suction side outer wall at a point aft of the leading edge transverse rib causing the central passage to extend towards the suction side outer wall, wherein a leading end of the suction side camber line rib connects directly to the suction side outer wall; and', 'a pressure side camber line rib directly connected to the leading edge transverse rib creating a pressure side flow passage opposite the central passage extension towards the suction side outer wall.'], '#text': 'a rib configuration including:'} ...

GAS TURBINE ENGINES INCLUDING EMBEDDED ELECTRICAL MACHINES AND ASSOCIATED COOLING SYSTEMS

A gas turbine engine includes a fan located at a forward portion of the gas turbine engine. A compressor section and a turbine section are arranged in serial flow order. The compressor section and the turbine section together define a core airflow path. A rotary member is rotatable with at least a portion of the compressor section and with at least a portion of the turbine section. An electrical machine is coupled to the rotary member and is located at least partially inward of the core airflow path in a radial direction. An enclosure at least partially encloses the electrical machine. The enclosure at least partially defines a first cooling airflow path within the enclosure that at least partially defines a first cooling airflow buffer cavity at least partially around the electrical machine. The first cooling airflow path is in communication with a second cooling airflow path located outside the enclosure that at least partially defines a second cooling airflow buffer cavity at least partially around the enclosure. A cooling duct provides pressurized air to the first cooling airflow path such that the air flows along both the first cooling airflow path and the second cooling airflow path providing the first cooling airflow buffer cavity and the second cooling airflow buffer cavity. 1. A gas turbine engine comprising:a fan located at a forward portion of the gas turbine engine;a compressor section and a turbine section arranged in serial flow order, the compressor section and the turbine section together defining a core airflow path;a rotary member rotatable with at least a portion of the compressor section and with at least a portion of the turbine section;an electrical machine coupled to the rotary member and located at least partially inward of the core airflow path in a radial direction;an enclosure that at least partially encloses the electrical machine, the enclosure at least partially defining a first cooling airflow path within the enclosure that at least ...

BLADE AND GAS TURBINE PROVIDED WITH SAME

A blade includes a blade body in which a cooling flow passage through which a cooling medium flows is formed. The blade body includes a top plate, and a thinning which is formed on a top surface of the top plate, closer to a suction side than a camber line, and which protrudes and extends along the camber line. A top plate flow passage is formed inside the top plate. The top plate flow passage includes an inlet flow passage which is formed closer to the suction side than the camber line and into which the cooling medium flows, a main flow passage which extends in a direction intersecting the camber line along the top surface, and an outlet flow passage through which the cooling medium is discharged to an outside of the blade body from a position closer to a pressure side than the camber line. 1. A blade , comprising a blade body inside which a cooling flow passage through which a cooling medium flows is formed , a top plate formed on a blade end portion of the blade body in a blade height direction, and', 'a thinning which is formed on a top surface of the top plate facing an outside in the blade height direction, closer to a suction side of the blade body than a camber line of the blade body, and which protrudes toward the outside and extends along the camber line,, 'wherein the blade body includes'}wherein a top plate flow passage through which the cooling medium from the cooling flow passage flows is formed inside the top plate, an inlet flow passage which is formed closer to the suction side than the camber line and into which the cooling medium flows from the cooling flow passage,', 'a main flow passage connected to the inlet flow passage and extending in a direction intersecting the camber line along the top surface, and', 'an outlet flow passage which is connected to the main flow passage and through which the cooling medium is discharged to an outside of the blade body from a position closer to a pressure side of the blade body than the camber line,, ' ...

Turbomachine Cooling Trench

A component for a gas turbine engine. The component includes a body. The body has an exterior surface abutting a flowpath for the flow of a hot combustion gas through the gas turbine engine. Further, the body defines a cooling passageway within the body to supply cool air to the component. The component includes a leading face and a trailing face defining a trench therebetween on the exterior surface. The body defines a plurality of cooling holes extending between the cooling passageway and a plurality of outlets defined in the trench such that the trench is fluidly coupled to the cooling passageway. Additionally, the leading face and trailing face are each tangent to at least one of the plurality of outlets. The trench directs the cool air along a contour of the component. 1. A component for a gas turbine engine , comprising:a body, an exterior surface of the body abutting a flowpath for the flow of a hot combustion gas through the gas turbine engine, wherein the body defines a cooling passageway within the body to supply cool air to the component; anda leading face and a trailing face defining a trench therebetween on the exterior surface, wherein the body defines a plurality of cooling holes extending between the cooling passageway and a plurality of outlets defined in the trench such that the trench is fluidly coupled to the cooling passageway, and wherein at least one of the leading face or the trailing face is tangent to at least one of the plurality of outlets,wherein the trench directs the cool air along a contour of the component.2. The component of claim 1 , wherein the leading face defines a first radius of curvature claim 1 , and the trailing face defines a second radius of curvature less than the first radius of curvature claim 1 , and wherein the cool air impinges on the trailing face such that the second radius of curvature directs the cool air along the contour of the component.3. The component of claim 1 , wherein the body is an airfoil claim 1 , ...

BURNER ARRANGEMENT AND METHOD FOR OPERATING A BURNER ARRANGEMENT

The invention relates to a burner arrangement for using in a single combustion chamber or in a can-combustor comprising a center body burner located upstream of a combustion zone, an annular duct with a cross section area, intermediate lobes which are arranged in circumferential direction and in longitudinal direction of the center body. The lobes being actively connected to the cross section area of the annular duct, wherein a cooling air is guided through a number of pipes within the lobes to the center body and cools beforehand at least the front section of the center body based on impingement cooling. Subsequently, the impingement cooling air cools the middle and back face of the center body based on convective and/or effusion cooling. At least the back face of the center body includes on the inside at least one damper. 1. A burner arrangement for using in a single combustion chamber or in a can-combustor comprising a center body burner located upstream of a combustion zone , an annular duct with a cross section area , intermediate lobes which are arranged in circumferential direction and in longitudinal or quasi-longitudinal direction of the center body burner , wherein the lobes being actively connected to the cross section area of the annular duct , wherein a cooling air is guided through a number of pipes within the lobes to the center body burner and based on impingement cooling cools beforehand at least the front section of the center body burner and in a subsequent flow the impingement cooling air based on convective and/or effusion cooling cools the middle and back face of the center body burner , wherein at least the back face of the center body burner includes on the inside at least one damper.2. The burner arrangement according to claim 1 , wherein the damper is operatively designed as low frequency damper.3. The burner arrangement according to claim 1 , wherein the front section of the center body burner having a impingement cooling cavity disposed ...

AIRFOIL WITH TUNABLE COOLING CONFIGURATION

Airfoils, additively manufactured airfoils, and methods of manufacturing airfoils are provided. For example, an airfoil comprises opposite pressure and suction sides that extend axially from a leading edge to a trailing edge and radially spaced apart inner and outer ends. The airfoil also comprises an outer wall defining the pressure and suction sides and leading and trailing edges. A rib extends within the airfoil from the pressure side to the suction side of the outer wall and radially from the inner to the outer end. The airfoil further comprises a first pre-impingement chamber surrounded by a first post-impingement chamber and a first dividing wall segment separating the first pre-impingement and first post-impingement chambers and having a plurality of cooling holes defined therein. The outer wall, rib, and first dividing wall segment are integrally formed as a single monolithic component. 120.-. (canceled)21. An airfoil , comprising:a concave pressure side opposite a convex suction side and an inner end radially spaced apart from an outer end, the pressure side and the suction side extending axially from a leading edge to a trailing edge;an outer wall defining the pressure side, suction side, leading edge, and trailing edge;a rib extending linearly within the airfoil from the pressure side of the outer wall to the suction side of the outer wall, the rib further extending radially from the inner end to the outer end;a pre-impingement chamber;a post-impingement chamber surrounding the pre-impingement chamber; anda dividing wall separating the pre-impingement chamber from the post-impingement chamber.22. The airfoil of claim 21 , wherein the outer wall claim 21 , the rib claim 21 , and the dividing wall are integrally formed as a single monolithic component.23. The airfoil of claim 21 , further comprising:a plurality of cooling holes defined in the dividing wall, the plurality of cooling holes including a first plurality of cooling holes defined at a first radial ...

TURBINE ASSEMBLY AND CORRESPONDING METHOD OF OPERATION

A turbine assembly having a basically hollow aerofoil with at least one cavity spanning the aerofoil in span wise direction of the aerofoil, an outer platform and an inner platform, each comprising at least one cavity, which are in flow communication with each other over at least one jumper tube, which extends in span wise direction along a whole length of the cavity of the aerofoil, and with a sealed gap being arranged between an outer surface of the jumper tube and an inner surface of a cavity wall of the aerofoil. A corresponding method operates a turbine assembly. 1. A turbine assembly , comprising:a hollow aerofoil formed by a cavity wall defining at least one cavity spanning the aerofoil in span wise direction of the aerofoil,an outer platform and an inner platform, each comprising at least one cavity, which are in flow communication with each other through at least one jumper tube, which extends in span wise direction along a length of the aerofoil,a gap is arranged between an outer surface of the jumper tube and an inner surface of the cavity wallwherein the jumper tube has a main inlet and a main outlet for a main part of a cooling medium andwherein the turbine assembly has at least one inlet aperture to the gap located within 0.2L of one of the inner and outer platforms and at least one outlet aperture located within 0.2L of the other inner and outer platforms for passing a fraction of the cooling medium through the gap.2. The turbine assembly according to claim 1 ,wherein the fraction of cooling medium of the cooling medium flowing in span wise direction along the gap provides an insulation for the jumper tube to prevent a heat transfer between the jumper tube and the cavity wall of the aerofoil.3. The turbine assembly according to claim 1 ,wherein at least 80% of the span wise length of the aerofoil is travelled by the cooling medium.4. The turbine assembly according to claim 1 ,wherein the jumper tube is arranged in the cavity of the aerofoil the gap ...

TURBINE SHROUD SEGMENT SEALING

An integrated shroud structure surrounds a circumferential array of stator vanes and a circumferential array of rotor blades of a gas turbine engine. The shroud structure includes a plurality of vane shroud segments and a plurality of blade shroud segments. The blade shroud segments integrally extend downstream from the vane shroud segments and each pair of circumferentially adjacent blade shroud segments defines an inter-segment gap. At least one slot extends axially from a location downstream of the vane shroud segments to an aft end of the blade shroud segment. The inter-segment gaps and slots are sealed by a sealing band mounted around the full circumference of the integrated shroud structure. 1. A shroud structure integrated to a circumferential array of stator vanes for surrounding a circumferential array of rotor blades of a gas turbine engine , the circumferential array of stator vanes positioned axially upstream of the circumferential array of rotor blades , the shroud structure comprising:a plurality of blade shroud segments disposed circumferentially one adjacent to another and configured to surround the circumferential array of rotor blades, the blade shroud segments extending integrally from the circumferential array of stator vanes, each pair of circumferentially adjacent blade shroud segments defining an inter-segment gap, at least one of the plurality of blade shroud segments having a radially inner gas path surface and an opposed radially outer surface and at least one slot extending axially from a location downstream of the circumferential array of stator vanes to a downstream end of the at least one of the plurality of the blade shroud segments between the radially inner gas path surface and the opposed radially outer surface thereof; anda sealing band mounted around the radially outer surface of the blade shroud segments and extending across the inter-segment gaps and the at least one slot around the full circumference of the integrated shroud ...

IMPINGEMENT STRUCTURE FOR JET ENGINE MID-RUTBINE FRAME

A mid-turbine frame (MTF”) for a jet engine is disclosed and comprises a duct that extends between a high pressure turbine (“HPT”) and a low pressure turbine (“LPT”), the duct comprising a plurality of segments that together form an outer annular structure and an inner annular structure, the inner annular structure situated radially inward of the outer annular structure, and/or a plurality of vanes that extend radially outward from the inner annular structure toward the outer annular structure, each vane comprising a channel. Each segment may be coupled to an adjacent segment by a seal. 1. An apparatus comprising:a radially inward segment portion;a radially outward segment portion; anda radially outward perforated structure coupled to the radially outward segment portion.2. The apparatus of claim 1 , wherein a vane extends from the radially inward segment portion to the radially outward segment portion.3. The apparatus of claim 2 , further comprising a radially inward perforated structure coupled to the radially inward segment portion.4. The apparatus of claim 2 , wherein the vane defines a channel.5. The apparatus of claim 2 , wherein the radially inward segment portion and the radially outward segment portion comprise a segment.6. The apparatus of claim 5 , wherein the segment comprise a first tenon that defines a first axial terminus of the segment and a second tenon that defines a second axial terminus of the segment.7. The apparatus of claim 5 , wherein the segment at least partially defines a hot gas path between a high pressure turbine (“HPT”) and a high pressure turbine (“LPT”).8. The apparatus of claim 1 , further comprising a slot proximate to a seal that clamps a second segment to the segment.9. The apparatus of claim 8 , wherein the slot is configured to relay cooling air through the seal and into a hot gas path.10. The apparatus of claim 9 , further comprising a tortuous gutter.11. The apparatus of claim 10 , wherein the tortuous gutter comprises a ...

Impingement system for an airfoil

An airfoil includes an exterior wall, a triple trailing edge pin bank, and an impingement system. The exterior wall defines a first interior space. The exterior wall also includes a pressure sidewall and a suction sidewall. The impingement system is disposed within the first interior space. The impingement system includes an interior wall which defines a second interior space a plurality of impingement holes configured to channel a flow of coolant from the second interior space to the first interior space. The interior wall has an impingement hole density with a varying hole density pattern. The impingement system also includes dividing walls extending from the interior wall to the exterior wall. The interior wall, exterior wall, and dividing walls define a first and second zone coupled in flow communication. The impingement hole density is configured to separately meter flow to the first and second zones.

IMPINGEMENT SYSTEM FOR AN AIRFOIL

An airfoil includes an exterior wall, a triple trailing edge pin bank, and first and second impingement systems. The exterior wall defines a first interior space. A segmenting wall divides the first interior space into a second and a first interior space. The first impingement system is disposed within the first interior space and the second impingement system is disposed within the first interior space. First impingement system includes impingement holes with a first impingement hole density. Second impingement system includes impingement holes with a second impingement hole density. The impingement systems also include dividing walls extending from a first and second interior wall to the exterior wall. The interior walls, exterior wall, and dividing walls define a first and second zone. The first and second impingement hole density is configured to separately meter flow to the first and second zones. 1. An airfoil comprising: a pressure sidewall; and', 'a suction sidewall coupled to said pressure sidewall, wherein said suction sidewall and said pressure sidewall define a leading edge and an trailing edge opposite said leading edge;, 'an exterior wall comprising an inner surface, an outer surface, and a plurality of exterior wall regions, said exterior wall defining a first interior space, said exterior wall further comprisinga root portion;a tip portion opposite said root portion;a triple trailing edge pin bank disposed within said first interior space;a first impingement system disposed within said first interior space and a second impingement system disposed within said first interior space, said first and second impingement systems configured to channel a coolant stream to said exterior wall, said first impingement system comprising a first interior wall substantially parallel to said exterior wall and defining a second interior space, said first interior wall further defines a plurality of first impingement holes configured to channel a flow of coolant from ...

COMPONENTS HAVING OUTER WALL RECESSES FOR IMPINGEMENT COOLING

A component configured for impingement cooling includes an inner wall defining a plurality of apertures extending therethrough. Each aperture of the plurality of apertures is configured to emit a cooling fluid therethrough. The component also includes an outer wall that includes an exterior surface, an opposite interior surface, and a thickness defined therebetween. The component further includes a plurality of recesses defined in the outer wall. Each recess of the plurality of recesses extends from a recess first end to an opposite recess second end. The second recess end is defined at the interior surface, and the recess first end is positioned within the outer wall at a depth less than the thickness. Each recess is aligned with a corresponding aperture of the plurality of apertures to receive the cooling fluid therefrom. 1. A component configured for impingement cooling , said component comprising:an inner wall defining a plurality of apertures extending therethrough, each aperture of said plurality of apertures configured to emit a cooling fluid therethrough;an outer wall comprising an exterior surface, an opposite interior surface, and a thickness defined therebetween; anda plurality of recesses defined in said outer wall, each recess of said plurality of recesses extending from a recess first end to an opposite recess second end, wherein said recess second end is defined at said interior surface and said recess first end is positioned within said outer wall at a depth less than said thickness, wherein said each recess is aligned with a corresponding aperture of said plurality of apertures to receive the cooling fluid therefrom.2. The component of claim 1 , further comprising at least one plenum that is at least partially defined by said inner wall and interior thereto claim 1 , said at least one plenum configured to supply the cooling fluid to said apertures.3. The component of claim 1 , further comprising:at least one chamber defined between said inner wall ...

IMPINGEMENT SYSTEM FOR AN AIRFOIL

An airfoil includes an exterior wall, a trailing edge pin bank, and an impingement system. The exterior wall includes an inner surface and an outer surface and defines a first interior space. The impingement system is disposed within the first interior space and is configured to channel a coolant stream to the exterior wall. The coolant stream has a velocity. The impingement system includes an interior wall which defines a second interior space and a plurality of impingement holes having an impingement hole density. The impingement system also includes dividing walls extending from the interior wall to the exterior wall. The interior wall, exterior wall, and dividing walls define a first and second zone. A first dividing wall is coupled to the trailing edge pin bank and separates the first and second zones. The impingement hole density configured to separately meter flow to the first and second zones. 1. An airfoil comprising: a pressure sidewall; and', 'a suction sidewall coupled to said pressure sidewall, wherein said suction sidewall and said pressure sidewall define a leading edge and a trailing edge opposite said leading edge;, 'an exterior wall comprising an inner surface, an outer surface, and a plurality of exterior wall regions, said exterior wall defining a first interior space, said exterior wall further comprisinga root portion;a tip portion opposite said root portion;a trailing edge pin bank disposed within said first interior space; and an interior wall substantially parallel to said exterior wall, said interior wall defining a second interior space, said interior wall further defines a plurality of impingement holes configured to channel a flow of coolant from said second interior space to said first interior space, said interior wall having an impingement hole density having a varying hole density pattern; and', 'a plurality of dividing walls extending from said interior wall to said exterior wall, said interior wall, said exterior wall, and said ...

SEQUENTIAL COMBUSTION ARRANGEMENT WITH COOLING GAS FOR DILUTION

A gas turbine with a sequential combustor arrangement as disclosed includes a first combustor with a first burner for admitting a first fuel into a combustor inlet gas during operation and a first combustion chamber for burning the first fuel, a dilution gas admixer for admixing a dilution gas to the first combustor combustion products leaving the first combustion chamber, a second burner for admixing a second fuel and a second combustion chamber. To assure a temperature profile after the dilution gas admixer and to increase the gas turbine's power and efficiency a vane and/or blade of the turbine has a closed loop cooling. The outlet of the closed loop cooling is connected to the dilution gas admixer for admixing the heated cooling gas leaving the vane and/or blade into the first combustor combustion products. 1. A gas turbine with a compressor , a turbine , and sequential combustor arrangement comprising:a first combustor with a first burner for admitting a first fuel into a combustor inlet gas during operation and a first combustion chamber for burning the first fuel;a dilution gas admixer for admixing a dilution gas to the first combustor combustion products leaving the first combustion chamber;a second burner for admixing a second fuel and a second combustion chamber, wherein the first combustor, the dilution gas admixer, the second burner and second combustion chamber are arranged sequentially in a fluid flow connection; anda vane and/or blade of the turbine having a closed loop cooling which is connected to a compressor plenum for feeding compressed cooling fluid into the vane and/or blade and an outlet of the closed loop cooling being connected to the dilution gas admixer for admixing the heated cooling gas leaving the vane and/or blade into the first combustor combustion products.2. A gas turbine as claimed in claim 1 , comprising:a cooling gas feed connecting the compressor plenum to the closed loop cooling.3. A gas turbine as claimed in claim 2 , wherein ...

IMPINGEMENT PANEL SUPPORT STRUCTURE AND METHOD OF MANUFACTURE

An integrated combustor nozzle includes a combustion liner that extends between an inner liner segment and an outer liner segment along a radial direction. The combustion liner including a forward end portion, an aft end portion, a first side wall, and a second side wall. An impingement panel having an impingement plate disposed along an exterior surface of one of the inner liner segment or the outer liner segment. The impingement plate defines a plurality of impingement apertures that direct coolant in discrete jets towards the exterior surface of the inner liner segment or the outer liner segment. The impingement panel includes an inlet portion that extends from the impingement plate to a collection duct. The impingement panel further includes a plurality of supports spaced apart from one another. The plurality of supports extend between, and are coupled to, the inlet portion, the collection duct, and the impingement plate

AIRFOIL WITH RIBS HAVING CONNECTOR ARMS AND APERTURES DEFINING A COOLING CIRCUIT

An airfoil includes an airfoil wall that defines a leading end, a trailing end, and first and second sides joining the leading end and the trailing end. First and second ribs each connect the first and second sides of the airfoil wall. Each of the first and second ribs define a tube portion that circumscribes a rib passage and includes cooling apertures, and first and second connector arms that solely join the tube portion to, respectively, the first and second sides of the airfoil wall. The airfoil wall and the first and second ribs bound a cooling channel there between. The cooling apertures of the first and second ribs open at the cooling channel such that there is a cooling circuit in which the rib passages of the first and second ribs are fluidly connected through the cooling apertures and the cooling channel 1. An airfoil comprising:an airfoil wall defining a leading end, a trailing end, and first and second sides joining the leading end and the trailing end; a tube portion circumscribing a rib passage and including cooling apertures, and', 'first and second connector arms solely joining the tube portion to, respectively, the first and second sides of the airfoil wall,, 'first and second ribs each connecting the first and second sides of the airfoil wall, each of the first and second ribs defining'}the airfoil wall and the first and second ribs bounding a cooling channel there between, the cooling apertures of the first and second ribs opening at the cooling channel such that there is a cooling circuit in which the rib passages of the first and second ribs are fluidly connected through the cooling apertures and the cooling channel2. The airfoil as recited in claim 1 , wherein the tube portion includes forward and aft walls and first and second side walls joining the forward and aft walls claim 1 , the first connector arm projects from the first side wall and the second connector arm projects from the second side wall claim 1 , and one or more of the cooling ...

GEARED GAS TURBINE ENGINE

A rotor assembly includes a plurality of rotor disks that each have a rim portion. The plurality of rotor disks includes a first rotor disk and at least one first bleed air passage that extends through a forward rim portion of the first rotor disk. At least one second bleed air passage that extends through an aft rim portion of the first rotor disk. 1. A rotor assembly comprising:a plurality of rotor disks each having a rim portion, wherein the plurality of rotor disks includes a first rotor disk; andat least one first bleed air passage extending through a forward rim portion of the first rotor disk and at least one second bleed air passage extending through an aft rim portion of the first rotor disk.2. The rotor assembly of claim 1 , wherein the plurality of rotor disks are configured to rotate with a shaft having an internal cooling passage.3. The rotor assembly of claim 2 , wherein the at least one first bleed air passage and the at least one second bleed air passage are in fluid communication with the internal cooling passage in the shaft.4. The rotor assembly of claim 3 , further comprising a second rotor disk claim 3 , wherein a rim of a second rotor disk includes at least one turbine bleed air passage.5. The rotor assembly of claim 4 , wherein the at least one turbine bleed air passage is in fluid communication with the internal cooling passage.6. The rotor assembly of claim 5 , wherein the second rotor disk is located upstream of the first rotor.7. The rotor assembly of claim 6 , further comprising at least one rotor disk of the plurality of rotor disks located between the first rotor and the second rotor.8. The rotor assembly of claim 5 , wherein a cross-sectional area of the at least one turbine bleed air passage is larger than a sum of a cross-sectional area of all of the at least one first bleed air passage and the at least one second bleed air passage.9. A gas turbine engine comprising: a plurality of rotor disks each having a rim portion, wherein the ...

Combustor heat shield and method of cooling same

A combustor for a gas turbine engine includes an annular shell, an annular bulkhead connected to the shell, and a heat shield panel. The heat shield panel has a first surface facing a combustion chamber and a second surface opposite the first surface. The heat shield panel is mounted to the bulkhead and defines a cooling chamber between the bulkhead and the heat shield panel. The heat shield panel has a wall extending from the heat shield panel toward the bulkhead around at least a portion of a periphery of the heat shield panel. The wall includes a circumferential wall portion including at least one cooling air passage extending between the cooling chamber and a cavity defined between the circumferential wall portion and the shell. The at least one cooling air passage is configured to purge the cavity by directing a first cooling air stream from the cooling chamber into the cavity.

COOLING CIRCUIT FOR A MULTI-WALL BLADE

A cooling circuit according to an embodiment includes: a cooling circuit for a multi-wall blade, the cooling circuit including: a pressure side cavity with a surface adjacent a pressure side of the multi-wall blade; a suction side cavity with a surface adjacent a suction side of the multi-wall blade; a central cavity disposed between the pressure side and suction side cavities, the central cavity including no surfaces adjacent the pressure and suction sides of the multi-wall blade; a first leading edge cavity with surfaces adjacent the pressure and suction sides of the multi-wall blade; and at least one impingement opening for fluidly coupling the first leading edge cavity with a second leading edge cavity. 1. A cooling circuit for a multi-wall blade , comprising:a pressure side cavity with a surface adjacent a pressure side of the multi-wall blade;a suction side cavity with a surface adjacent a suction side of the multi-wall blade;a central cavity disposed between the pressure side and suction side cavities, the central cavity including no surfaces adjacent the pressure and suction sides of the multi-wall blade;a first leading edge cavity with surfaces adjacent the pressure and suction sides of the multi-wall blade; andat least one impingement opening for fluidly coupling the first leading edge cavity with a second leading edge cavity.2. The cooling circuit of claim 1 , further including at least one leading edge film hole for fluidly coupling the second leading edge cavity to a leading edge of the multi-wall blade.3. The cooling circuit of claim 1 , further including at least one channel for fluidly coupling the second leading edge cavity to a tip of the multi-wall blade.4. The cooling circuit of claim 1 , further comprising:a flow of cooling air split between the pressure and suction side cavities, wherein a first portion of the flow of cooling air is directed into the pressure side cavity, and wherein a second portion of the flow of cooling air is directed into ...

System for cooling a hot gas component for a combustor of a gas turbine

A system for cooling a hot gas path component for a combustor generally includes an impingement sleeve that circumferentially surrounds an outer surface of the hot gas path component. A first cooling chamber is defined between the impingement sleeve and a first portion of the outer surface of the hot gas path component. A second cooling chamber is disposed downstream from the first cooling chamber. The second cooling chamber is defined between the impingement sleeve and a second portion of the outer surface of hot gas path component. An inlet extends through the impingement sleeve so as to define a first flow path into the first cooling chamber. An outlet defines a second flow path between the first cooling chamber and the second cooling chamber.

AIR-FILM COOLED COMPONENT FOR A GAS TURBINE ENGINE

A component for a gas turbine engine that separates a cooling air plenum from a heated gas environment. The component defines a hot section surface adjacent to the heated gas environment having a plurality of cooling apertures fluidically connecting the cooling air plenum to the heated gas environment to allow a cooling air to flow from the cooling air plenum to the heated gas environment through the plurality of cooling apertures. The plurality of cooling apertures each have an aperture diameter of less than about millimeters (mm) and an average surface roughness of less than about 1 micrometer (1 μm). 1. An article for a gas turbine engine comprising:a component separating a cooling air plenum from a heated gas environment, wherein the component defines a hot section surface adjacent to the heated gas environment, wherein the hot section surface defines a plurality of cooling apertures fluidically connecting the cooling air plenum to the heated gas environment to allow a cooling air to flow from the cooling air plenum to the heated gas environment through the plurality of cooling apertures, wherein the plurality of cooling apertures each comprise an aperture diameter of less than about 3 millimeters (mm) and an aperture surface having average surface roughness of less than about 1 micrometer (1 μm).2. The article of claim 1 , wherein the component comprises:a turbine airfoil comprising an exterior surface and defining an internal chamber, wherein the exterior surface comprises the hot section surface, and wherein the internal chamber comprises the cooling air plenum.3. The article of claim 1 , wherein the component comprises a combustor component that separates the cooling air plenum from a combustion chamber that comprises the heated gas environment.4. The article of claim 1 , wherein the component comprises a single-walled structure separating the cooling air plenum from the heated gas environment.5. The article of article of claim 1 , wherein the component ...

INJECTOR NOZZLE CONFIGURATION FOR SWIRL ANTI-ICING SYSTEM

An anti-icing system for annular gas turbine engine inlet housings includes a substantially closed annular housing at a leading edge of the gas turbine engine inlet housing, the annular housing containing a quantity of air and a conduit extending from a source of high-pressure hot bleed air to the annular housing. The system also includes an injector connected to the end of the conduit and extending into the annular housing and one or more nozzles extending outwardly from the injector in a direction that the quantity of air circulates in the annular housing. The system may include one or both of an airfoil on an upstream side of the injector and an air direction element disposed one the injector that causes the quantity of air to be directed toward an outlet of the one or more nozzles. 1. An anti-icing system for annular gas turbine engine inlet housings comprising:a substantially closed annular housing at a leading edge of the gas turbine engine inlet housing, the annular housing containing a quantity of air;a conduit extending from a source of high-pressure hot bleed air to the annular housing;an injector connected to the end of the conduit and extending into the annular housing, the injector including an airfoil on an upstream side; andone or more nozzles extending outwardly from the injector in a direction that the quantity of cooler air circulates in the annular housing.2. The system of claim 1 , further comprising a gas turbine engine compressor claim 1 , wherein the gas turbine engine compressor is the source of the high-pressure hot bleed air.3. The system of claim 1 , further comprising:a first air directional element disposed on an end of the injector that causes the quantity of air to be directed toward an outlet of the one or more nozzles.4. The system of claim 1 , further comprising:a second air direction element disposed on either a top or a bottom the injector that causes the quantity of air to be directed toward an outlet of the one or more nozzles.5 ...

TAIL CONE EJECTOR FOR POWER CABLE COOLING SYSTEM IN A GAS TURBINE ENGINE

An ejector assembly for a cooling system of a gas turbine engine may comprise: a tail cone having a tail cone outlet in fluid communication with a cooling air flow of the cooling system; an ejector body defining a mixing section, a constant area section, and a diffuser section; and a nozzle section in fluid communication with an exhaust air flow of the gas turbine engine, the ejector assembly configured to entrain the cooling air flow via the exhaust air flow. 1. An ejector assembly for a cooling system of a gas turbine engine , the ejector assembly comprising:a tail cone having a tail cone outlet in fluid communication with a cooling air flow of the cooling system;an ejector body defining a mixing section, a constant area section, and a diffuser section; anda nozzle section in fluid communication with an exhaust air flow of the gas turbine engine, the ejector assembly configured to entrain the cooling air flow via the exhaust air flow.2. The ejector assembly of claim 1 , wherein the nozzle section is defined by a nozzle portion of the ejector body and the tail cone.3. The ejector assembly of claim 1 , wherein the ejector body further comprises a scoop defining an inlet to the nozzle section.4. The ejector assembly of claim 1 , wherein a throat of the nozzle section is disposed forward of the tail cone outlet.5. The ejector assembly of claim 1 , wherein the cooling air flow is from an external air source disposed radially outward from a bypass air flow path of the gas turbine engine.6. The ejector assembly of claim 1 , wherein:the tail cone comprises a scoop configured to divert the exhaust air flow of the gas turbine engine internal to the tail cone;the tail cone further comprises a channel extending through the tail cone and in fluid communication with the exhaust air flow, the channel defining a throat at a nozzle portion of the channel.7. The ejector assembly of claim 6 , wherein:a cooling air flow path is defined radially outward from the channel and between ...

Fuel injector assembly for a turbine engine

An apparatus is provided for a turbine engine. This apparatus includes a fuel conduit and a fuel nozzle. The fuel conduit includes a supply passage. The fuel nozzle includes a nozzle passage, an end wall and a nozzle orifice. The nozzle passage has a longitudinal centerline and extends longitudinally through the fuel nozzle along the longitudinal centerline from the end wall to the nozzle orifice. The nozzle passage is configured with a convergent portion and a throat portion. The nozzle passage converges radially inward towards the longitudinal centerline as the convergent portion extends longitudinally along the longitudinal centerline away from the end wall and towards the throat portion. The supply passage is fluidly coupled to the nozzle passage by a fuel aperture in the end wall. A centerline of the fuel aperture is angularly and laterally offset from the longitudinal centerline.

COOLANT FLOW REDIRECTION COMPONENT

A gas turbine engine includes a compressor section, a combustor fluidly connected to the compressor section, and a turbine section fluidly connected to the combustor and mechanically connected to the compressor section via a shaft. Multiple rotors are disposed in one of the compressor section and the turbine section. Each of the rotors includes a rotor disk portion having a radially inward bore, and is static relative to the shaft. Each rotor is axially adjacent at least one other rotor and a gap is defined between each rotor and an adjacent rotor. A cooling passage for a cooling flow is defined between the shaft and the rotors, and a cooling flow redirection component is disposed at the gap and is operable to redirect the cooling flow in the cooling passage into the gap. 1. A gas turbine engine comprising:a compressor section;a combustor fluidly connected to the compressor section;a turbine section fluidly connected to the combustor and mechanically connected to said compressor section via a shaft;a plurality of rotors disposed in one of said compressor section and said turbine section, each of said rotors including a rotor disk portion including a radially inward bore, and each of said rotors being static relative to said shaft;each rotor in said plurality of rotors being axially adjacent at least one other of said rotors in said plurality of rotors and defining a gap between each of said rotors and said axially adjacent rotors;a cooling passage for a cooling flow defined between said shaft and said rotors; anda cooling flow redirection component disposed at said gap and operable to redirect said cooling flow in said cooling passage into said gap.2. The gas turbine engine of claim 1 , wherein said gap is defined between radially aligned surfaces of adjacent rotor disks.3. The gas turbine engine of claim 1 , wherein said cooling flow redirection component includes a radially outward protrusion from said shaft.4. The gas turbine engine of claim 3 , wherein cooling ...