СПОСОБ ОПТИМИЗАЦИИ УДЕЛЬНОГО РАСХОДА ТОПЛИВА ДВУХМОТОРНОГО ВЕРТОЛЕТА

Способ оптимизации удельного расхода топлива вертолета, оборудованного двумя газотурбинными двигателями (1, 2), каждый из которых содержит газогенератор (11, 21), оснащенный камерой (СС) сгорания, при этом каждый из этих газотурбинных двигателей (1, 2) выполнен с возможностью самостоятельно работать в постоянном полетном режиме, а другой газотурбинный двигатель (2, 1) находится при этом в так называемом режиме сверхмалого газа с нулевой мощностью и с включенной камерой (СС) сгорания, причем этот режим сверхмалого газа поддерживают посредством механического приведения во вращение вала (АЕ) газогенератора в этом режиме таким образом, чтобы снизить рабочую температуру и расход топлива этого газогенератора. 3 з.п. ф-лы, 1 ил.

Компрессор с приводом от установки для утилизации тепла с органическим циклом Ренкина и способ регулирования

Описана система преобразования энергии, содержащая источник (17) отходящего тепла и систему (5) с органическим циклом Ренкина. Система с органическим циклом Ренкина, в свою очередь, содержит по меньшей мере турбодетандер (21), содержащий регулируемые входные направляющие аппараты (57А, 57В), по меньшей мере вращающуюся нагрузку (29), механически соединенную с турбодетандером (21) и приводимую посредством этого в движение, и механическое соединение (31) с переменной скоростью между турбодетандером (21) и вращающейся нагрузкой (29). Обеспечивают возможность запуска с постепенным ускорением турбодетандера до скорости прогрева, чтобы обеспечить экономичный и безопасный прогрев указанного турбодетандера благодаря постепенному открытию пускового клапана и изменению скорости вращения механического соединения в результате воздействия на указанные пусковой клапан и механическое соединение. 2 н. и 16 з.п. ф-лы, 4 ил.

УСТРОЙСТВО С ВИНТАМИ ПРОТИВОПОЛОЖНОГО ВРАЩЕНИЯ, ИМЕЮЩЕЕ СРЕДСТВО ИЗМЕНЕНИЯ ШАГА ВИНТОВ

Изобретение относится к области авиации, в частности к устройствам регулирования шага винтов. Устройство с винтами противоположного вращения, имеющее средство изменения шага, содержит пару винтов противоположного вращения с изменяемым шагом. Шаг может изменяться посредством приводного механизма (31), расположенного в центральной полости валов. Система подвода энергии к приводному механизму (31) содержит линию (34) питания, проходящую в статический корпус (5) рядом с эпициклоидальной передачей (17), линии управления (38), которые подходят к приводному механизму (31) и проходят через держатель сателлитов (14) эпициклоидальной передачи (17). Соединение осуществляется посредством уплотнения для подвижных соединений с коллектором. Такой подвод позволяет избежать турбин с высокой температурой. Приводной механизм может быть электрическим, при этом уплотнения для подвижных соединений представляют собой уплотнения для электрических соединений. Система подвода содержит линии смазки (51), которые ...

ГАЗОТУРБИНЫЙ ДВИГАТЕЛЬ ДЛЯ ЛЕТАТЕЛЬНОГО АППАРАТА, ОСНАЩЕННЫЙ АВТОМАТИЧЕСКИ АКТИВИРУЕМЫМ ЦЕНТРУЮЩИМ ЭЛЕМЕНТОМ

Изобретение относится к авиационным газотурбинным двигателям, в частности к газотурбинным двигателям, содержащим свободную турбину, один из опорных подшипников которой вынесен в редуктор. Объектом изобретения является газотурбинный двигатель, содержащий картер, в котором расположены газогенератор и свободная турбина, установленная на силовом валу. Силовой вал выполнен с возможностью механического соединения/разъединения с редуктором. Газотурбинный двигатель содержит по меньшей мере один центрующий элемент, подвижный между положением, называемым активным положением, и положением, называемым пассивным положением. В активном положении он образует опорный подшипник упомянутого силового вала и оно соответствует механическому разъединению между упомянутым силовым валом и упомянутым редуктором. В пассивном положении он отходит от упомянутого силового вала и оно соответствует механическому соединению между упомянутым силовым валом и упомянутым редуктором. Изобретение позволяет исключить риск повреждения ...

ТУРБИННАЯ СИСТЕМА С ТРЕМЯ ПОДКЛЮЧЕННЫМИ К ЦЕНТРАЛЬНОМУ РЕДУКТОРУ ТУРБИНАМИ И СПОСОБ ЭКСПЛУАТАЦИИ РАБОЧЕЙ МАШИНЫ

... 1. Турбинная система (110, 310, 410), содержащаяпервую турбину (151, 251, 351),вторую турбину (152, 252, 352),третью турбину (153, 253, 353),центральный редуктор (130, 230, 330), который с входной стороны механически связан с тремя турбинами (151, 251, 351; 152, 252, 352; 153, 253, 353), а с выходной стороны имеет механический подключающий элемент (337), к которому подключается воспринимающая механическую энергию рабочая машина (120, 220),первую линию (161, 261) текучей среды для проведения рабочей текучей среды от первой турбины (151, 251, 351) ко второй турбине (152, 252, 352),вторую линию (162, 262) текучей среды для проведения рабочей текучей среды от второй турбины (152, 252, 352) к третьей турбине (153, 253, 353),первое подключающее устройство (171, 271), которое связано с первой линией (161, 261) текучей среды и которое выполнено таким образом, что возможен отбор первого частичного массового потока (171a) рабочей текучей среды из первой линии (161, 261) текучей среды или его подвод ...

СПОСОБ ЭКСПЛУАТАЦИИ ГАЗОТУРБИННОЙ УСТАНОВКИ И ГАЗОТУРБИННАЯ УСТАНОВКА ДЛЯ РЕАЛИЗАЦИИ ДАННОГО СПОСОБА

... 1. Способ эксплуатации газотурбинной установки, содержащей компрессор (1), который на впуске всасывающей стороны всасывает воздух (2) и нагнетает его для обеспечения подачи воздуха (3) на выходе компрессора на напорной стороне, камеру сгорания (4, 9), в которой сгорает топливо (5, 10), используя воздух (3) на выходе компрессора, образуя горячие газы (6, 11), а также турбину (7, 12), в которой горячие газы (6, 11) расширяются, осуществляя работу, характеризующийся тем, чтосжатый воздух, выходящий из компрессора (1), направляют в виде потока (20, 22, 23, 24) охлаждающего воздуха в камеру сгорания (4, 9) и/или в турбину (7, 12) для охлаждения термически нагруженных компонентов, при этом, по меньшей мере, одним из потоков (20, 22, 23, 24) охлаждающего воздуха управляют органом управления (15, 16, 17, 21) в зависимости от эксплуатационной задачи.2. Способ по п.1, отличающийся тем, что в качестве эксплуатационной задачи используют, по меньшей мере, одной из следующих- увеличение эксплуатационный ...

УЗЕЛ ГАЗОТУРБИННОГО ДВИГАТЕЛЯ И СПОСОБ ЕГО СБОРКИ

... 1. Узел (10) газотурбинного двигателя, содержащий ! газотурбинный двигатель (12) внутреннего контура, ! турбину (14) низкого давления, соединенную с газотурбинным двигателем внутреннего контура, ! вентиляторный узел (16) с противоположным вращением, соединенный с турбиной низкого давления, и ! вспомогательный компрессор (24), соединенный непосредственно с турбиной низкого давления таким образом, что обеспечивается вращение вспомогательного компрессора и турбины низкого давления в одном и том же направлении. ! 2. Узел (10) по п.1, дополнительно содержащий редуктор (100), соединенный между турбиной (14) высокого давления и вентиляторным узлом (16) с противоположным вращением. ! 3. Узел (10) по п.1, дополнительно содержащий вентиляторный каркас (13), причем вспомогательный компрессор (24) расположен ниже по потоку от вентиляторного каркаса. ! 4. Узел (10) по п.2, дополнительно содержащий приводной вал (34), соединенный между турбиной (14) низкого давления и редуктором (100). ! 5. Узел (10) ...

ПРИВОДНАЯ КОНСТРУКЦИЯ ДЛЯ ГАЗОТУРБИННОГО ДВИГАТЕЛЯ (ВАРИАНТЫ)

... 1. Приводная конструкция для газотурбинного двигателя, содержащая:вал вентилятора;раму, поддерживающую вал вентилятора и задающую репер поперечной жесткости (Kframe);зубчатую систему, приводящую во вращение вал вентилятора;гибкую несущую конструкцию, по меньшей мере частично поддерживающую зубчатую систему и имеющую поперечную жесткость (KFS), определяемую относительно поперечной жесткости (Kframe) рамы, ивходной узел зубчатой системы, имеющий поперечную жесткость (KIC), определяемую относительно поперечной жесткости (Kframe) рамы.2. Конструкция по п.1, в которой рама и гибкая несущая конструкция установлены на статической конструкции.3. Конструкция по п.1, в которой рама и гибкая несущая конструкция установлены на статической конструкции в составе газотурбинного двигателя.4. Конструкция по п.1, в которой рама и гибкая несущая конструкция установлены на переднюю центральную часть корпуса газотурбинного двигателя.5. Конструкция по п.1, в которой гибкая несущая конструкция связана с водилом ...

Getriebeturbomaschine

Getriebeturbomaschine (10), die ein Getriebe (11), ein Antriebsaggregat (12), und mehrere Abtriebsaggregate (13, 14, 15, 16) aufweist, die zu einem Maschinenstrang integriert sind; wobei das Getriebe (11) ein zentrales Großrad (17) mit einer Großradwelle (18) und mehrere in das Großrad (17) kämmende Ritzel (21, 23, 25) mit Ritzelwellen (22, 24, 26) umfasst; wobei das Antriebsaggregat (12) als Dampfturbine ausgebildet ist, in welcher zur Bereitstellung mechanischer Antriebsleistung ein erstens Prozessgas entspannt wird; wobei ein erstes Abtriebsaggregat (13) als mindestens zweistufiger Hauptkompressor ausgebildet ist, in welchem unter Nutzung der vom Antriebsaggregat (12) bereitgestellten mechanischen Antriebsleistung ein zweites Prozessgas verdichtet wird, wobei zwischen jeweils zwei Stufen des Hauptkompressors (13) jeweils ein Zwischenkühler (36) geschaltet ist, um das zweite Prozessgas zu kühlen; wobei mehrere zweite Abtriebsaggregate (14, 15, 16) als Getriebekompressoren ausgebildet ...

Geared turbomachine

A geared turbomachine comprises a gear unit 11 with a drive unit 12 such as a steam turbine, and at least two geared compressors 13, 14. There is a large central gear 17 mounted on a main shaft 18, and at least two pinion gears 21, 23 meshing with the main gear 17 and mounted on pinion shafts 22, 24. The drive unit 12 is coupled to a first pinion shaft 22 via a clutch 29, and a main compressor 13 is coupled to the opposite end of the first pinion shaft 22 via a clutch 30. At least one further compressor 14 is fixedly coupled to at least one further pinion shaft 24. The main compressor is therefore coupled to the drive unit without any gearing, which reduces mechanical losses.

Efficient gas turbine engine installation and operation

Remote shaft driven open rotor propulsion system with electrical power generation

Improvements in and relating to steam turbines for working with steam of high pressures and high degrees of superheat

... 210,456. Akt.-Ges. Brown, Boverie, et Cie. Jan. 29, 1923, [Convention date]. Motor power plant; transmission mechanism ; bearings ; balancing end-thrust.-In a compound turbine using steam at very high initial pressure and superheat, the high-pressure stage turbines are arranged to have small individual heat-drops as described in Specification 196,913, and their rotors are mounted in an overhung manner on pinion shafts which are disposed symmetrically around the main gear-wheel on the shaft of the low-pressure stage turbine or on the shaft of a dynamo. When two high-pressure stages are used they are disposed diametrically opposite one another in a horizontal plane. Fig. 2 shows two high-pressure stages a<1>, a<4> in series driving a main overhung wheel i. The casings are secured by flanges to the gear box and centering wedges or keys arranged crosswise are employed to locate the casings and permit expansion. The casings may be divided longitudinally in a plane inclined at 45‹ to the horizontal ...

Cryogenic air separation

A geared gas turbine engine and a gearbox

Centrifugal blower or - compressors

GAS TURBINE UNIT

System and method for oxidant compression in a stoichiometric exhaust gas recirculation gas turbine system

SYSTEM AND METHOD FOR OXIDANT COMPRESSION IN A STOICHIOMETRIC EXHAUST GAS RECIRCULATION GAS TURBINE SYSTEM [00212] A system includes a gas turbine system having a turbine combustor, a turbine driven by combustion products from the turbine combustor, and an exhaust gas compressor driven by the turbine. The exhaust gas compressor is configured to compress and supply an exhaust gas to the turbine combustor. The gas turbine system also has an exhaust gas recirculation (EGR) system. The EGR system is configured to recirculate the exhaust gas along an exhaust recirculation path from the turbine to the exhaust gas compressor. The system further includes a main oxidant compression system having one or more oxidant compressors. The one or more oxidant compressors are separate from the exhaust gas compressor, and the one or more oxidant compressors are configured to supply all compressed oxidant utilized by the turbine combustor in generating the combustion products. + 5/24 I-: C'", D ( Cn L U O ...

REMOTE SHAFT DRIVEN OPEN ROTOR PROPULSION SYSTEM WITH ELECTRICAL POWER GENERATION

A system for aircraft propulsion is disclosed herein. The system includes a power plant. The system also includes an open rotor module operable to rotate. The open rotor module has a plurality of variable-pitch blades. The system also includes a first linkage extending between the power plant and the open rotor module. The first linkage is operable to transmit rotational power to the open rotor module for rotating the plurality of variable-pitch blades. The system also includes an actuator operable to change a pitch of the plurality of variable-pitch blades. The system also includes a generator operable to generate electric power. The system also includes a second linkage extending between the power plant and the generator. The second linkage is operable to transmit rotational power to the generator. The generator is operable to convert the rotational power to electrical power. The system also includes a controller operably coupled to the actuator to vary a pitch of the plurality of variable-pitch ...

GAS TURBINE

Gas turbine (1), in particular for model aircraft, model helicopters and other small propulsion units, comprising a drive shaft (3), extending through an annular combustion chamber (2), rotatably mounted by means of two main bearings (13,13') and connected to a compressor rotor (4) and a turbine rotor (11) and a driven shaft (14) driven by the drive shaft (3). A device with a curvic gears is provided for torque transfer from the drive shaft (3) to the driven shaft (14), between the two main bearings (13, 13').

DISTRIBUTED PROPULSION SYSTEM POWER UNIT CONTROL

A propulsion system that includes a plurality of power units, a plurality of propulsors, where respective power units of the plurality of power units are controllably coupled to the plurality of propulsors, and a controller configured to receive a desired throttle value corresponding to a desired propulsive force, determine a number of power units of the plurality of power units to be coupled to the plurality of propulsors to achieve the desired propulsive force based on a respective power value associated with each respective power unit of the plurality of power units, and cause the number of power units of the plurality of power units to be coupled to the plurality of propulsors.

A TURBINE ENGINE COMPRISING A DRIVE SYSTEM FOR A DEVICE SUCH AS AN ACCESSORIES CASE

L'invention concerne une turbomachine (10) comprenant : au moins un bras de guidage (44) creux s'étendant radialement depuis un moyeu (38) jusqu'à un carter annulaire (42), et présentant une extrémité radialement extérieure qui débouche sur une ouverture (60) dudit carter annulaire, un arbre de renvoi radial (48) situé dans ledit bras de guidage (44) et destiné à entraîner en rotation au moins un équipement agencé en périphérie dudit carter annulaire (42), Selon l'invention, un boîtier de transfert (50) est disposé en regard de l'ouverture (60) du carter annulaire, et la turbomachine comprend des moyens de transfert (78, 52) de puissance en rotation entre l'arbre de renvoi radial (48) et l'équipement, le bras de guidage (44) étant muni, à son extrémité extérieure, d'une portion élargie (56) par laquelle il débouche sur l'ouverture (60) du carter annulaire et dans laquelle est logée une partie du boîtier de transfert (50).

INTERSHAFT INTEGRATED SEAL AND LOCK-NUT

An intershaft integrated seal and lock-nut assembly includes spaced apart forward and aft shafts, forward lock-nut threaded onto aft end of forward shaft, aft lock-nut threaded onto a forward end of aft shaft, and seal ring sealingly engaging and disposed between forward and aft lock-nuts operable to seal annular gap between forward and aft shafts. Seal ring may be disposed in annular ring groove extending radially inwardly into one of forward and aft lock-nuts and annular cylindrical inner surface on other one of lock-nuts operable to seal against seal ring. Seal ring may be carbon seal ring and split. Aft lock-nut may include forwardly extending annular arm having annular cylindrical inner surface. Retention tabbed rings may engage lock-nuts and shafts and snap rings may engage retention tabbed rings and lock-nuts. Seal ring may be in ring groove in aft lock-nut.

OIL TRANSFER UNIT FOR TRANSFERRING OIL BETWEEN A STATIONARY PART AND A ROTATING PART

An oil transfer unit (1) has a rotating part (19) extending along an axis (7), a stationary part (18) provided with an oil mouth (29), and a floating part (20) having a cylindrical surface (87) fitted onto an outer cylindrical surface (88) of the rotating part (19) in a non-contact configuration; an annular groove is provided between the floating part (20) and the rotating part (19) to put the oil mouth (29) into communication with an inner chamber of the rotating part (19); both sides of the groove are sealed by a hydrostatic seal defined by a radial gap between the cylindrical surfaces (87,88); the unit has at least one oil transfer tube (45), coupled to the stationary part (18) and the floating part (20) in a fluid-tight manner and with freedom of movement, and a connecting rod (60) to prevent rotation of the floating part (20); the opposite ends (61) of the connecting rod (60) are coupled to the stationary part (18) and floating part (20) by respective spherical joints (63).

PLANET GEARBOX WITH CYLINDRICAL ROLLER BEARING WITH HIGH DENSITY ROLLER PACKING

A planet gearbox is provided for connection to a carrier of an epicyclic gearing arrangement with a single input and single output and including a sun gear, a ring gear and at least one double helix planet gear rotatable on a cylindrical roller bearing with a cage having a cross-web thickness of 15% to 25% of the diameter of the cylindrical rollers and an L/D ratio exceeding 1Ø A gas turbine engine includes a fan and LP shaft, which couples a compressor to a turbine. An epicyclic gearing arrangement has a single input from the LP shaft coupled to a sun gear, a single output coupled to the fan's shaft, and a planet bearing cage having a cross-web thickness measuring 15% to 25% of the roller's diameter.

GEARED AXIAL MULTISTAGE EXPANDER DEVICE, SYSTEM AND METHOD

Method, system and axial multistage expander including a casing and a plurality of stages. A stage includes a stator part connected to the casing and having plural statoric airfoils, and a rotor part configured to rotate relative to the stator part and having plural rotoric airfoils. The axial multistage expander also includes a support mechanism connected to the casing and configured to rotatably support the rotor part. Rotoric airfoils of at least one stage of the plurality of stages are configured to rotate with a speed different from rotoric airfoils of the other stages. The stator part, the rotor part and the support mechanism of the plurality of stages are provided inside the casing.

Dampfturbinenanlage.

Mehrgehäusige Getriebedampfturbine.

Kraftanlage mit einer Dampf- bzw. Gasturbine.

Planetengetriebe, insbesondere für Schiffsantrieb

Arbeitsmaschine

Uber ein Untersetzungsgetriebe mit der anzutreibenden Maschine gleichachsig gekuppelte Dampf- oder Gasturbine

Mechanical drive architecturehybrid bearings and low density materials with low loss.

Eine mechanische Antriebsarchitektur (100) enthält eine Gasturbine (10), die einen Verdichterabschnitt (105), einen Turbinenabschnitt (115) und einen Brennkammerabschnitt (110) aufweist. Durch die Gasturbine (10) wird ein Ladungsverdichter (160) angetrieben. Eine Rotorwelle (125) erstreckt sich durch die Gasturbine (10) und den Ladungsverdichter (160) hindurch. Die Rotorwelle (125) weist rotierende Schaufeln (130, 135, 165) auf, die in einer Umfangsanordnung angeordnet sind, um mehrere Laufschaufelreihen in der Gasturbine (10) und dem Ladungsverdichter (160) zu definieren. Wenigstens eine der rotierenden Schaufeln (130, 135, 165) in der Gasturbine (10) oder dem Ladungsverdichter (160) enthält ein Material geringer Dichte. Lager (140) lagern die Rotorwelle (125) innerhalb der Gasturbine (10) und des Ladungsverdichters (160), wobei wenigstens eines der Lager (140) ein verlustarmes Lager (140) vom Hybridtyp ist.

Transmission turbomachine.

Getriebeturbomaschine (10), die ein Getriebe (11), mindestens ein Antriebsaggregat (12) und mindestens zwei Abtriebsaggregate (13) aufweist, die zu einem Maschinenstrang integriert sind; wobei das Getriebe (11) ein zentrales Grossrad mit einer Grossradwelle (18) und mindestens zwei in das Grossrad kämmende Ritzel mit Ritzelwellen (22, 24) umfasst; wobei das Antriebsaggregat (12) mit einer ersten Ritzelwelle (22) der Ritzelwellen des Getriebes (11) auf einer Seite desselben über eine erste Kupplung (29) gekoppelt ist; wobei ein erstes der Abtriebsaggregate (13) als Hauptkompressor ausgebildet ist, in welchem unter Nutzung der vom Antriebsaggregat (12) bereitgestellten mechanischen Antriebsleistung ein erstes Prozessgas verdichtbar ist, wobei das erste Abtriebsaggregat (13) mit der ersten Ritzelwelle (22) des Getriebes (11) auf der gegenüberliegenden Seite desselben über eine zweite Kupplung (30) derart gekoppelt ist, dass das erste Abtriebsaggregat (13) bei geschlossener erster Kupplung ...

transmission turbomachine.

Getriebeturbomaschine (10), die ein Getriebe (11), mindestens ein Antriebsaggregat (12), und mindestens zwei Abtriebsaggregate (13, 14) aufweist, die zu einem Maschinenstrang integriert sind; wobei das Getriebe (11) ein zentrales Grossrad mit einer Grossradwelle (18) und mindestens zwei in das Grossrad kämmende Ritzel mit mindestens zwei Ritzelwellen (22, 24) umfasst; wobei die Getriebeturbomaschine (10) ein Antriebsaggregat (12) umfasst, wobei das Antriebsaggregat (12) mit einer ersten Ritzelwelle (22) des Getriebes (11) auf einer Seite desselben über eine erste Kupplung (29) gekoppelt ist; wobei ein erstes Abtriebsaggregat (13) als Hauptkompressor ausgebildet ist, in welchem unter Nutzung der vom Antriebsaggregat (12) bereitgestellten mechanischen Antriebsleistung ein erstes Prozessgas verdichtet wird, wobei das erste Abtriebsaggregat (13) mit der ersten Ritzelwelle (22) des Getriebes (11) auf der gegenüberliegenden Seite desselben über eine zweite Kupplung (30) derart gekoppelt ist, ...

Centrifugal compressor

Is used in the turbine rotor and steam turbine

DISPOSITIF DE LIMITATION DU DEPLACEMENT RADIAL D'UN ROTOR A PALIER DYNAMIQUE FORME PAR UN ENGRENAGE PLANETAIRE

DISPOSITIF DE LIMITATION DU DEPLACEMENT RADIAL D'UN ROTOR DE MACHINE TOURNANTE DONT L'ARBRE EST SUPPORTE D'UNE PART PAR UN PALIER ET D'AUTRE PART PAR UN PALIER DYNAMIQUE CONSTITUE PAR UN ENGRENAGE PLANETAIRE. LE PIGNON CENTRAL 4 ET AU MOINS LES PIGNONS SATELLITES 5 DONT L'AXE EST SITUE A UN NIVEAU PLUS ELEVE QUE CELUI DU PIGNON CENTRAL 4 SONT POURVUS, SUR L'UNE DE LEURS FACES, D'UN GALET DE ROULEMENT 10, 11 DE DIAMETRE LEGEREMENT INFERIEUR A CELUI DU CERCLE PRIMITIF DU PIGNON ASSOCIE, TOUS LES GALETS ETANT SITUES DANS LE MEME PLAN PERPENDICULAIRE AUX AXES DES PIGNONS.

Improvements with the turbines with gas fluid

MOTEUR A TURBINES S'ENTRAINANT MUTUELLEMENT

LA PRESENTE INVENTION CONCERNE UN MOTEUR FONCTIONNANT AVEC UNE ENERGIE DE BASE TRES FAIBLE, DONC PEU COUTEUSE; IL COMPORTE PLUSIEURS TURBINES A LA SUITE LES UNES DES AUTRES ABSORBANT AINSI L'ENERGIE PRIMAIRE AU MAXIMUM POUR LA TRANSFORMER EN MECANIQUE; IL EST ROTATIF, DONC EST SILENCIEUX ET N'EMET PAS DE VIBRATIONS. LE MOTEUR OBJET DE L'INVENTION COMPORTE UNE CHAMBRE A D'OU UN GAZ EST LIBERE ET QUI ACTIONNE A SON PASSAGE LES CINQ TURBINES S'ENTRAINANT MUTUELLEMENT GRACE AUX ENGRENAGES SUIVANTS, SOIT DIRECTIONNELS 6, 7, 8, 9, 10, 11, 12, 13, 14 ET 15, SOIT DEMULTIPLICATEURS, MULTIPLICATEURS ET INVERSEURS 16, 17, 18, 19, 20, 21, 22, 23, 24, 25, 26 ET 27, CE GAZ A SA SORTIE EN B A DONC ETE DANS DE TRES GRANDES PROPOSTIONS TRANSFORME EN ENERGIE MECANIQUE.

HYBRID AND PROCEEDED DRIVING INSTALLATION OF ORDERING Of SUCH a DRIVING INSTALLATION

La présente invention concerne une installation motrice (200) hybride comportant un moyen d'entraînement (204) apte à entraîner en rotation un élément mécanique (BTP, BTA). De plus, l'installation motrice hybride est remarquable en ce qu'elle comporte au moins un turbomoteur (253, 254) et au moins un moteur électrique (201) mécaniquement lié audit moyen d'entraînement (204) pour le mettre en rotation.

TURBINE ENGINE FOR AN AIRCRAFT EQUIPPED WITH A CENTRALIZER READING SYSTEM

ARM FOR A TURBOMACHINE CASING COMPRISING AN ADDITIONAL REMOVABLE

Pour permettre un démontage aisé d'un palier logé dans un bras de carter de turbomachine et assurant le guidage d'un arbre de transmission de la turbomachine, il est proposé un bras de carter (50) comprenant un corps (100) formé d'un montant (112) et d'une base (110) délimitant une partie d'un logement d'arbre (102), une pièce additionnelle (106) fixée de manière amovible sur le corps (100) et délimitant un logement de palier (104) débouchant dans le logement d'arbre (102), et une cale (140) amovible interposée entre la base (110) et la pièce additionnelle (106). Il est également proposé une turbomachine (10) comprenant un tel bras de carter (50) ainsi qu'un procédé de démontage du palier d'une telle turbomachine comprenant le retrait de la cale (140) et le déplacement de la pièce additionnelle (106) vers la base (110), permettant un accès au palier.

ASSEMBLY FOR PROPELLING AN AIRCRAFT EQUIPPED WITH A BLOWER OF AT LEAST ONE BLOWER OFF-SITE

La présente invention porte sur un ensemble de propulsion d'un aéronef comprenant : - au moins une turbine principale (5, 6) montée suivant un axe longitudinal (XX); - au moins une soufflante principale (8) disposée suivant l'axe longitudinal (XX) et entrainée en rotation par ladite turbine principale; - au moins une turbine auxiliaire (7) montée suivant l'axe longitudinal (XX), la turbine auxiliaire étant indépendante de la turbine principale, - au moins une soufflante auxiliaire (9, 9') d'axe (XY, XY') décalé par rapport à l'axe longitudinal (XX) et entrainée par la turbine auxiliaire.

METHOD FOR THE MAINTENANCE OF HOUSING ELEMENT CASING ACCESSORIES TURBOMACHINE

L'invention propose un procédé de maintenance d'une pièce de fonderie en alliage de magnésium, ladite pièce étant un carter de boitier d'accessoires de turbomachine ou un couvercle de carter, comprenant au moins une surface cylindrique apte à recevoir une cage de roulement, le procédé étant caractérisé en ce qu'il comprend une étape (200) consistant à réaliser un dépôt de molybdène sur la surface de la pièce par projection plasmatique.

원심 압축기

... 입력 드라이브를 갖는 임펠러(impeller)를 포함하는 원심 압축기로서, 상기 임펠러는 그 임펠러에서 배출되는 가스에 의해 구동되는 임펠러의 주변을 둘러싸도록 배치된 충동형(impulse) 터빈을 구비하며, 상기 터빈의 출력은 상기 임펠러의 드라이브에 연결되도록 구성된다.

2-axis gas turbine power plant

GAS TURBINE ENGINE WITH GEARED ARCHITECTURE

A gas turbine engine according to an exemplary aspect of the present disclosure includes a core nacelle defined about an engine centerline axis, a fan nacelle mounted at least partially around the core nacelle to define a fan bypass flow path for a fan bypass airflow, a fan variable area nozzle axially movable relative the fan nacelle to vary a fan nozzle exit area and adjust a pressure ratio of the fan bypass airflow during engine operation, and a gear system driven by a core engine within the core nacelle to drive a fan within the fan nacelle, the gear system defines a gear reduction ratio of greater than or equal to about 2.3.

GAS TURBINE

Gas turbine (1), in particular for model aircraft, model helicopters and other small propulsion units, comprising a drive shaft (3), extending through an annular combustion chamber (2), rotatably mounted by means of two main bearings (13,13') and connected to a compressor rotor (4) and a turbine rotor (11) and a driven shaft (14) driven by the drive shaft (3). A device with a curvic gears is provided for torque transfer from the drive shaft (3) to the driven shaft (14), between the two main bearings (13, 13').

TURBINE ENGINE GEARBOX

A gas turbine engine according to an example of the present disclosure includes, among other things, a fan section, a compressor section, and a turbine section including a fan drive turbine that drives the fan through a gear reduction. The gear reduction includes at least two double helical gears in meshed engagement, each of the at least two double helical gears having a first plurality of gear teeth separated from a second plurality of gear teeth such that a first end of the first plurality of gear teeth and a first end of the second plurality of gear teeth are spaced apart by an axial distance. Each of the first plurality of gear teeth is offset a first circumferential offset distance in relation to the next gear tooth of the second plurality of gear teeth when moving in a circumferential direction relative to respective axes. 1. A turbofan engine comprising:a fan section including a fan and an outer housing surrounding the fan to define a bypass duct;a compressor section including a low pressure compressor and a high pressure compressor;a gear reduction; anda turbine section including a fan drive turbine and a high pressure turbine, wherein the fan drive turbine drives the fan through the gear reduction;wherein the gear reduction is an epicyclic gear system that includes a carrier and a plurality of gears, and the plurality of gears including a sun gear, a ring gear, and a plurality of intermediate gears that engage the sun gear and the ring gear;wherein at least two of the plurality of gears are double helical gears in meshed engagement, each of the double helical gears disposed to rotate about respective axes, each of the double helical gears having a first plurality of gear teeth separated from a second plurality of gear teeth such that a first end of the first plurality of gear teeth and a first end of the second plurality of gear teeth are spaced apart by an axial distance;wherein each of the first plurality of gear teeth is offset a first circumferential ...

Geared Architecture for High Speed and Small Volume Fan Drive Turbine

A gas turbine engine includes a gear system that provides a speed reduction between a fan drive turbine and a fan rotor. Aspects of the gear system are provided with defined flexibility. The fan drive turbine has a first exit area and rotates at a first speed. A second turbine section has a second exit area and rotates at a second speed, which is faster than said first speed. A performance quantity can be defined for both turbine sections as the products of the respective areas and respective speeds squared. A performance quantity ratio of the performance quantity for the fan drive turbine to the performance quantity for the second turbine section is between 0.5 and 1.5.

Multishaft geared multishaft turbocompressor with return channel stages and radial expaner

A geared multishaft turbocompressor with impellers arranged in series in terms of flow, are attached to two or more pinion shafts (6), which are arranged in parallel to one another and are driven directly via a central gear (5) or indirectly via pinion shafts at the circumference of the central gear (5). A plurality of impellers (8, 8a) are arranged in series, via the interstage diaphram of a disk-type diffuser (9) and of a return ring (10), on at least one pinion shaft end (6) in high-pressure stages following low-pressure stages (first pinion shaft or first and second pinion shafts) after the second or third pinion shaft (6). Reversing the direction of flow, i.e., admission of the gas on the high-pressure side and discharge of the gas on the low-pressure side, as well as with simultaneous reversal of the direction of rotation, a radial expander (turbine) is obtained with the same basic design.

SYSTEM AND APPARATUS FOR DIVERSIFIED GEARBOX

A gas turbine engine assembly comprising, a gearbox including a first housing that includes a first auxiliary gear drive on a first portion thereof, a second housing that includes a second auxiliary gear drive on a second portion thereof, and a third housing that includes a third auxiliary gear drive on a third portion thereof, the housings being interconnected so that the first portion of the first housing, the second portion of the second housing and the third portion of the third housing form a substantially triangular polyhedron shape, with the second portion of the second housing disposed between the first portion of the first housing and the third portion of the third housing. The first auxiliary gear drive, the second auxiliary gear drive and the third auxiliary gear drive project outwardly in mutually divergent directions.

LOW PRESSURE RATIO FAN ENGINE HAVING A DIMENSIONAL RELATIONSHIP BETWEEN INLET AND FAN SIZE

According to an example embodiment, a gas turbine engine assembly includes, among other things, a fan section including a fan, the fan including a plurality of fan blades, a diameter of the fan having a dimension D that is based on a dimension of the fan blades, each fan blade having a leading edge, and a forward most portion on the leading edges of the fan blades in a first reference plane, a geared architecture, a turbine section including a high pressure turbine and a low pressure turbine, the low pressure turbine driving the fan through the geared architecture, a nacelle surrounding the fan, the nacelle including an inlet portion forward of the fan, a forward edge on the inlet portion in a second reference plane, and a length of the inlet portion having a dimension L measured along an engine axis between the first reference plane and the second reference plane. A dimensional relationship of L/D is between 0.20 and 0.40. 1. A gas turbine engine assembly comprising:a fan section including a fan, the fan including a plurality of fan blades, a diameter of the fan having a dimension D that is based on a dimension of the fan blades, each fan blade having a leading edge, and a forward most portion on the leading edges of the fan blades in a first reference plane;a compressor section including a low pressure compressor and a high pressure compressor;a geared architecture including an epicyclical gear train that drives the fan at a lower speed than an input speed in the geared architecture;a turbine section including a high pressure turbine and a low pressure turbine, the low pressure turbine driving the fan through the geared architecture;a low spool and a high spool, the low spool comprising the low pressure compressor and the low pressure turbine, the high spool comprising the high pressure compressor and the high pressure turbine, the low spool and the high spool rotatable about an engine central longitudinal axis;a nacelle surrounding the fan, the nacelle including ...

EMBEDDED AUXILIARY OIL SYSTEM FOR GEARBOX PROTECTION

A gearbox is provided for use in a gas turbine engine. The gearbox includes a plurality of gears, a casing surrounding the plurality of gears, and a lubricant re-circulation system. The casing defines a sump positioned beneath the plurality of gears to retain lubricant. The lubricant re-circulation system includes a re-circulation pump, a re-circulation inlet positioned within the sump, and a re-circulation outlet positioned to supply lubricant to the plurality of gears.

Adjustable flange material and torque path isolation for splined fan drive gear system flexible support

A gear assembly support for a gas turbine engine includes a first portion engageable to a case of the gas turbine engine and a second portion configured for supporting a gear assembly. The support includes a torque reacting portion for transferring torque from the second portion to the first portion, a forward flange disposed forward of the torque reacting portion, the forward flange defining a first interface to the case and an aft flange disposed aft of the torque reacting portion, the aft flange defining a second interface to the case.

ROTOR BLADE WITH L-SHAPED FEATHER SEAL

A turbine engine assembly arranged relative to an axis includes a rotor blade and a feather seal with an L-shaped geometry. The rotor blade includes an airfoil and a base. The airfoil extends axially between an upstream leading edge and a downstream trailing edge, and radially out from the base. The base includes a neck and a pocket. The neck extends axially to a downstream surface. The pocket extends laterally into the neck. The feather seal extends laterally into the pocket, and includes a down stream leg that extends substantially along the downstream surface.

Near zero velocity lubrication system for a turbine engine

A system is provided for a turbine engine. This turbine engine system includes a rotating assembly, a bearing and a lubrication system. The bearing is configured with the rotating assembly. The lubrication system is configured to lubricate the bearing. The lubrication system includes a lubricant pump and a lubricant reservoir. The lubricant pump is mechanically coupled with and driven by the rotating assembly. The lubricant pump is configured with the lubricant reservoir so as to be at least partially submersed in lubricant contained within the lubricant reservoir.

Lubrication system for a gear system of a gas turbine

The present disclosure is directed to lubrication systems for gear systems of a gas turbine engine. In one embodiment, a lubrication system includes an auxiliary reservoir configured to store lubricant and a lubrication collection device. The auxiliary reservoir is contained within an annular structure of the gas turbine engine. The lubricant collection device can collect lubricant from a gear system of the gas turbine engine and direct collected lubricant to the auxiliary reservoir. The auxiliary reservoir is configured to receive collected lubricant by way of a channel within the annular structure. The lubrication system may compliment a main lubrication system, and in particular, when the main lubrication system is temporarily unable to supply lubricant to the gear system.

Lightweight journal support pin

A journal support pin to support intermediate gears for use in gas turbine engine comprises a titanium body, and an outer surface outside of the titanium body having a surface hardness that is harder than the body. A gas turbine engine and a method of forming a journal support pin to support intermediate gears for use in gas turbine engine are also disclosed.

Accessory gearbox for gas turbine engine with compressor drive

A gas turbine engine has a low speed input shaft drives a first plurality of accessories. A high speed input shaft drives a second plurality of accessories. The first plurality of accessories rotating about a first set of rotational axes perpendicular to a first plane. The second plurality of accessories rotating about a second set rotational axes perpendicular to a second plane. The first and second planes extending in opposed directions away from a drive input axis. Compressed air is tapped and passes through a heat exchanger, then to a boost compressor, and then to at least one rotatable components in a main compressor section and a main turbine section. The boost compressor driven on a boost axis, which is non-parallel to the first set of rotational axes and the second set of rotational axes.

INTERMEDIATE GEAR DEVICE FOR MOUNTING GENERATOR TO EXPANDER

Lubrication device for a turbine engine

Geared turbomachine

Gas turbine engine with core mount

Electric turbomachine

Decoupler assembly for engine starter

A gas turbine architecture

GAS EXPANSION MACHINE

GASTURBINE

GAS TURBINE

Gas-turbine unit

A centrifugal compressor

A centrifugal compressor comprising an impeller having an input drive, the impeller having an impulse turbine positioned around the periphery of the impeller to be driven by the gas exiting the impeller, the output of the turbine being coupled to the drive of the impeller.

Coupling of a turbopump for molten salts

A device comprising at least one vertical pump (3) and at least one associated turbine (4), for transporting a heat-transfer fluid carried at high temperature over a difference in level, characterised in that the device comprises a mechanical device for coupling the turbine (4) with the pump (3), comprising a gearbox (21) with a gimbal coupling (41) located on the turbine (4) side, allowing the mechanical energy produced by the turbine (4) to be reused for actuating the pump (3).

MULTI-ENGINE SYSTEM AND METHOD

A method of operating a multi-engine system of a rotorcraft includes, during a cruise flight segment of the rotorcraft, controlling a first engine to provide sufficient power and/or rotor speed demands of the cruise flight segment; and controlling a second engine to by providing a fuel flow to the second engine that is between 70% and 99.5% less than a fuel flow provided to the first engine. A turboshaft engine for a multi-engine system configured to drive a common load is also described.

DOUBLE ROW CYLINDRICAL ROLLER BEARING WITH HIGH LENGTH TO DIAMETER RATIO ROLLERS

A planet gearbox is provided for connection to a carrier of an epicyclic gearing arrangement with a single input and single output and including a sun gear, a ring gear and at least one double helix planet gear rotatable on a cylindrical roller bearing wherein the ratio of each cylindrical roller's length to each cylindrical roller's diameter exceeds 1Ø A gas turbine engine includes a shaft coupling a compressor of a compressor section to a turbine of a turbine section. An epicyclic gearing arrangement has a single input from the shaft coupled to a sun gear, a single output from the carrier that is coupled to the shaft of a fan and includes a planet gearbox with cylindrical rollers having a length-to-diameter ratio exceeding 1Ø ...

TURBOMACHINE WITH A GEARBOX AND INTEGRATED ELECTRIC MACHINE ASSEMBLY

A turbomachine includes a turbine section including a turbine. The turbine includes a first plurality of turbine rotor blades and a second plurality of turbine rotor blades, the first plurality of turbine rotor blades and second plurality of turbine rotor blades alternatingly spaced along the axial direction. The turbomachine also includes a gearbox. The first plurality of turbine rotor blades and the second plurality of turbine rotor blades are each coupled to one of a ring gear, a planet gear, or a sun gear of the gearbox such that the first plurality of turbine rotor blades is rotatable with the second plurality of turbine rotor blades through the gearbox. The turbomachine also includes an electric machine assembly including a rotor coupled to one of the ring gear, the planet gear, or the sun gear of the gearbox such that the rotor rotates relative to a stator during operation.

DEVICE FOR TRANSFERRING OIL BETWEEN TWO REPOSITORIES ROTATING RELATIVE TO EACH OTHER, AND PROPELLER TURBOMACHINE FOR AN AIRCRAFT WITH SUCH A DEVICE

Le dispositif (20) comporte deux bagues concentriques externe et interne (22, 23) dont l'une est raccordée à une alimentation d'huile issue de l'un des référentiels, et dont l'autre bague est liée à l'autre des référentiels, l'huile circulant entre lesdites bagues, et des paliers entre les bagues pour réaliser le changement de référentiels entre celles-ci. Selon l'invention, le dispositif (20) comporte, de plus, un moyen souple (31 ) formant amortisseur, prévu entre une première desdites bagues et une bague intermédiaire (41 ) qui est séparée d'une seconde desdites bagues par lesdits paliers (25), ledit moyen souple (31 ) définissant une chambre étanche déformable (32) dans laquelle transite l'huile entre les deux référentiels.

AUXILLIARY POWER AND THRUST UNIT DRIVE SYSTEM

An aircraft auxiliary power and thrust unit includes at least one blade mounted to a fan shaft which are mounted to a tail cone of the aircraft and rotatable relative to the tail cone. Also included is an air intake assembly which includes an opening defined by the tail cone and a channel in fluid communication with the opening and the at least one blade. A first drive shaft positioned to extend in a direction transverse to the fan shaft and engagageable to rotate with the fan shaft and first drive shaft engaged and the first drive shaft engagageable to a second drive shaft positioned transverse to the first drive shaft such that the second drive shaft rotates with the first drive shaft with the first and second drive shafts engaged. The second drive shaft is positioned outside of the channel of the air intake.

TURBINE/MOTOR (GENERATOR) DRIVEN BOOSTER COMPRESSOR

A compressor assembly for cryogenic gas separation wherein the assembly comprises a compressor, an expansion turbine and an electric motor integrally connected via a gear drive, and processes for using the compressor assembly.

Improved turbine used for reducing hard particle damage and application method of improved turbine

Improvement to the housings for mounting planetary gear set

FORMATION GEAR V ON A TURBOMACHINE

Boîte d'engrenages (10) pour l'entraînement d'équipements d'une turbomachine, comprenant un carter de forme sensiblement en V et comportant deux bras (20) reliés entre eux par une partie de jonction (22), les bras contenant des lignes d'engrenages reliées entre elles au niveau de la partie de jonction, et des moyens de fixation à la turbomachine, caractérisée en ce que ces moyens de fixation comprennent des moyens (38) d'encastrement et/ou de fixation de la partie de jonction et des moyens (40) de suspension des bras.

LUBRICATING DEVICE FOR A TURBOMACHINE

L'invention concerne un dispositif de lubrification pour une turbomachine, comportant une conduite (23) d'arrivée d'huile équipée d'une pompe (24) d'alimentation en huile et de moyens de régulation (30) situés en aval de la pompe d'alimentation (24), une conduite d'alimentation (26) destinée à alimenter en huile un organe à lubrifier et une conduite de recyclage (27) débouchant en amont de la pompe d'alimentation (24), les moyens de régulation (30) permettant de diriger tout ou partie du débit d'huile issu de la conduite d'arrivée (23) vers la conduite d'alimentation (26) et/ou vers la conduite de recyclage (27).

TURBOMACHINE COMPRISING A DECOUPLING MEANS A BLOWER

OIL TRANSFER DEVICE BETWEEN TWO REPOSITORIES IN ROTATION RELATIVE TO EACH OTHER, AND TURBOMACHINE SCROLL FOR AIRCRAFT WITH SUCH A DEVICE

Le dispositif (20) comporte deux bagues concentriques externe et interne (22, 23) dont l'une est raccordée à une alimentation d'huile issue de l'un des référentiels, et dont l'autre bague est liée à l'autre des référentiels, l'huile circulant entre lesdites bagues, et des paliers entre les bagues pour réaliser le changement de référentiels entre celles-ci. Selon l'invention, le dispositif (20) comporte, de plus, un moyen souple (31) formant amortisseur, prévu entre les bagues, et définissant une chambre étanche déformable (32) dans laquelle transite l'huile entre les deux référentiels.

Multiple stage turbo machine - has one high pressure and two low pressure stages all geared to output shaft

AIR INLET APPARATUS OF COMPRESSOR AND GAS TURBINE COMPRISING SAME

The present invention provides an air inlet apparatus of a compressor, comprising: an IGV assembly including a plurality of inlet vanes; at least one VGV assembly including a plurality of variable vanes and arranged at the back of the IGV assembly; and a single actuator supported by a compressor casing and selectively rotating the inlet vanes of the IGV assembly or variable vanes of each VGV assembly. According to the present invention, the IGV assembly and the VGV assembly can be selectively controlled with the single actuator such that the number of parts can be reduced, and manufacturing and maintenance of the air inlet apparatus can be facilitated. COPYRIGHT KIPO 2018 ...

ROTOR FOR A STEAM TURBINE

The invention relates to a rotor (3) for a turbomachine, wherein the rotor (3) comprises a planetary gear (1) on which low-pressure end stage blades (17) are arranged, wherein the planetary gear (1) is designed in such a way that the frequency of the rotating blades (17) is lower than the frequency of the rotor (3).

System and method for oxidant compression in a stoichiometric exhaust gas recirculation gas turbine system

A system includes a gas turbine system having a turbine combustor, a turbine driven by combustion products from the turbine combustor, and an exhaust gas compressor driven by the turbine. The exhaust gas compressor is configured to compress and supply an exhaust gas to the turbine combustor. The gas turbine system also has an exhaust gas recirculation (EGR) system. The EGR system is configured to recirculate the exhaust gas along an exhaust recirculation path from the turbine to the exhaust gas compressor. The system further includes a main oxidant compression system having one or more oxidant compressors. The one or more oxidant compressors are separate from the exhaust gas compressor, and the one or more oxidant compressors are configured to supply all compressed oxidant utilized by the turbine combustor in generating the combustion products.

FLEXIBLE SUPPORT STRUCTURE FOR A GEARED ARCHITECTURE GAS TURBINE ENGINE

A gas turbine engine according to an example of the present disclosure includes, among other things, a fan shaft driving a fan having fan blades, a fan shaft support that supports the fan shaft and defines a fan shaft support transverse stiffness. A gear system connected to the fan shaft and includes a gear mesh defining a gear mesh transverse stiffness, and a reduction ratio greater than 2.3. A flexible support supports said gear system and defines a flexible support transverse stiffness. The flexible support transverse stiffness is less than 20% of the fan shaft support transverse stiffness.

Aircraft propulsion assembly equipped with a main fan and with at least one offset fan

An aircraft propulsion assembly comprising at least a main turbine mounted along a longitudinal axis, at least one main fan arranged upstream of the main turbine along the longitudinal axis and driven in rotation by the said main turbine, the said main fan being ducted by a main fan casing, an auxiliary turbine mounted along the longitudinal axis, the auxiliary turbine being independent of the main turbine, an auxiliary fan of axis offset with respect to the longitudinal axis and driven by the auxiliary turbine, the auxiliary fan being ducted by an auxiliary fan casing, the main casing being separate and distinct from the auxiliary casing so as respectively to generate a main secondary flow and an auxiliary secondary flow which remain independent of one another until they are discharged into the atmosphere.

Geared turbofan with three turbines all counter-rotating

A gas turbine engine has a fan rotor, a first compressor rotor and a second compressor rotor. The second compressor rotor compresses air to a higher pressure than the first compressor rotor. A first turbine rotor drives the second compressor rotor and a second turbine rotor. The second turbine drives the compressor rotor. A fan drive turbine is positioned downstream of the second turbine rotor. The fan drive turbine drives the fan through a gear reduction. The first compressor rotor and second turbine rotor rotate as an intermediate speed spool. The second compressor rotor and first turbine rotor together as a high speed spool. The high speed spool and the fan drive turbine configured to rotate in the same first direction. The intermediate speed spool rotates in an opposed, second direction.

TURBOMACHINE COMPRISING A MEANS OF UNCOUPLING A FAN

A turbomachine includes a fan shaft driven by a turbine shaft via a device for reducing a speed of rotation. The turbomachine includes an uncoupling device interposed between the reduction device and the turbine shaft. The uncoupling device is configured to uncouple the reduction device and the turbine shaft in response to the exceeding of a determined resistant torque exerted by the reduction device on the turbine shaft.

Fan drive gear system with improved misalignment capability

An epicyclic gear assembly according to an exemplary embodiment includes, among other things, a carrier including a first plate axially spaced from a second plate by a radially outer connector. A first set of epicyclic gears supported adjacent the first plate include a first set of circumferentially offset intermediate gears meshing with a first sun gear and a first ring gear. A second set of epicyclic gears are axially spaced from the first set of epicyclic gears and supported adjacent the second plate, and include a second set of circumferentially offset intermediate gears meshing with a second sun gear and a second ring gear. The first epicyclic gear set and the second epicyclic gear set maintain relative intermeshing alignment during flexure induced deformation of the carrier.

Lubrication and scavenge system

A lubrication system comprises a lubricant feed tank, a lubricant feed pump, one or more lubricant nozzles, a scavenge pump drive motor, and one or more scavenge pumps. The lubricant feed pump takes suction from the lubricant feed tank and pumps a lubricant feed as lubricant jets exiting the one or more lubricant nozzles. The lubricant feed pump has a rotating feed pump shaft coupled to a power source. The scavenge pump drive motor drives a rotating scavenge pump shaft that is coupled to the one or more scavenge pumps. The scavenge pumps return lubricant to the lubricant feed tank. The feed pump shaft and scavenge pump shaft rotate independently of each other.

REVERSE FLOW ENGINE ARCHITECTURE

A reverse flow gas turbine engine has a low pressure (LP) spool and a high pressure (HP) spool arranged sequentially in an axial direction. The LP spool comprises an LP compressor disposed forward of an LP turbine and drivingly connected thereto via an LP compressor gear train. The HP spool comprises an HP compressor in flow communication with the LP compressor, and an HP turbine disposed forward of the HP compressor and drivingly connected thereto via an HP shaft.

TURBINE ENGINE ASSEMBLY AND METHOD OF MANUFACTURING THEREOF

A turbine engine assembly is provided. The assembly includes a stationary component, a drive shaft, and a gearbox coupled along the drive shaft, and coupled to the stationary component. The assembly also includes a vibration-reducing mechanism coupled between the stationary component and the gearbox. The vibration-reducing mechanism is configured to isolate a vibratory response of the gearbox from the stationary component.

PLANETARY GEAR

A planetary gear with a sun wheel, a hollow wheel and a planetary carrier on which a planetary wheel is rotatably mounted. In the axial direction on a first side of the planetary carrier, the sun wheel and hollow wheel include connection areas for coupling the sun wheel and hollow wheel to rotatable or torque-proof areas of an engine. The planetary carrier has a connection area for attaching to rotatable or torque-proof areas of the engine on its opposite second side. The structural component stiffnesses of the sun wheel, planetary carrier, the hollow wheel and the planetary wheel are adjusted to each other such that, during operation they have twistings in the axial direction of the planetary gear that respectively qualitatively correspond to each other between the connection areas and side areas facing away from the connection areas due to the respectively applied torques.

GAS TURBINE ENGINE OIL BUFFERING

A turbine engine includes a shaft, a fan, at least one bearing mounted on the shaft and rotationally supporting the fan, a fan drive gear system coupled to drive the fan, a bearing compartment around the at least one bearing and a source of pressurized air in communication with a region outside of the bearing compartment. 1. A turbine engine comprising:a shaft;a fan;at least one bearing mounted on the shaft and rotationally supporting the fan;a fan drive gear system coupled to drive the fan;a bearing compartment around the at least one bearing; anda source of pressurized air in communication with a region outside of the bearing compartment.2. The turbine engine as recited in claim 1 , wherein the fan drive gear system includes an epicyclic gear train.3. The turbine engine as recited in claim 2 , wherein the epicyclic gear train has a gear reduction ratio of greater than or equal to about 2.3.4. The turbine engine as recited in claim 2 , wherein the epicyclic gear train has a gear reduction ratio of greater than or equal to 2.3.5. The turbine engine as recited in claim 2 , wherein the epicyclic gear train has a gear reduction ratio of greater than or equal to about 2.5.6. The turbine engine as recited in claim 2 , wherein the epicyclic gear train has a gear reduction ratio of greater than or equal to 2.5.7. The turbine engine as recited in claim 1 , wherein the fan defines a bypass ratio of greater than about ten (10) with regard to a bypass airflow and a core airflow.8. The turbine engine as recited in claim 1 , wherein the fan defines a bypass ratio of greater than 10.5:1 with regard to a bypass airflow and a core airflow.9. The turbine engine as recited in claim 1 , wherein the fan defines a bypass ratio of greater than ten (10) with regard to a bypass airflow and a core airflow.10. The turbine engine as recited in claim 1 , wherein the fan defines a pressure ratio that is less than about 1.45.11. The turbine engine as recited in claim 1 , wherein the fan defines ...

Gearbox in a turbine engine

A gearbox in a turbine engine for imparting rotary drive to at least one piece of rotary equipment, the gearbox including a transmission shaft guided in rotation in bearings and carrying a toothed wheel meshing with at least one rotary drive gearwheel. One of the bearings is a rolling bearing mounted inside the toothed wheel in a radial plane containing the toothed wheel and the drive gearwheel, and another of the bearings is a smooth bearing for taking up forces tending to tilt the transmission shaft.

Power Plant Line Having a Variable-Speed Pump

The invention relates to a power plant line, comprising a steam turbine and/or gas turbine that rotates at a constant speed in order to drive an electric generator; a variable-speed pump for conveying and/or compressing a working medium in order to drive and/or supply the process of the steam turbine and/or the gas turbine or to pump and/or compress an exhaust gas produced in the process supply or in the gas turbine. The invention is characterized in that the variable-speed pump is driven by the steam turbine and/or gas turbine and a speed-controllable gear train is arranged in the driving connection, said gear train having a power split, which comprises a mechanical main branch and a hydrodynamic secondary branch, wherein driving power is branched off from the mechanical main branch via a hydrodynamic coupling or a hydrodynamic converter by means of the hydrodynamic secondary branch and fed back to the mechanical main branch at the output side of the gear train in a variable-speed manner by means of a superimposing gear train.

Gas turbine engine with fan variable area nozzle for low fan pressure ratio

A gas turbine engine includes a fan section with twenty (20) or less fan blades and a fan pressure ratio less than about 1.45.

GAS EXPANDER SYSTEM

A gas expander system suitable for use in a turbomachine, the gas expander system comprising: a gas expander provided with a moveable part; a magnetic gear arrangement; and a shaft; the moveable part of the gas expander being connectable to a load via the magnetic gear arrangement and the shaft, and movement of the moveable part of the gas expander being arranged to cause movement of the shaft, wherein the magnetic gear arrangement is used in a closed loop heat recovery system, with the inner and outer rotors of the magnetic gear separated by a wall that contains the stator. 1. A turbocharger system comprising:a turbocharger, comprising a turbine and a compressor;a gas expander system comprising:a gas expander provided with a moveable part;a magnetic gear arrangement; anda shaft;the moveable part of the gas expander being connectable to a load via the magnetic gear arrangement and the shaft, and movement of the moveable part of the gas expander being arranged to cause movement of the shaft,wherein the turbocharger system comprises a source of heat arranged to heat a working fluid provided in a closed-loop system to cause expansion of that working fluid to move the moveable part of the gas expanderwherein the magnetic gear comprises a first rotor and a second rotor, the first rotor and second rotor being separated from one another by at least a part of a wall forming part of the closed loop system.2. The system of wherein the at least a part of a wall forming part of the closed loop system that separates the first rotor from the second rotor comprises a stator of the magnetic gear.3. The system of wherein the wall encloses the closed loop system.4. The system of wherein the source of heat is provided by a part of the turbocharger system claim 1 , an engine to which the turbocharger is connected claim 1 , or a fluid flowing into or out of the turbocharger or the engine.5. The system of wherein the moveable part of the gas expander is located within the closed loop ...

Geared Architecture for High Speed and Small Volume Fan Drive Turbine

A gas turbine engine includes a flex mount for a fan drive gear system. A very high speed fan drive turbine drives the fan drive gear system.

Geared Architecture for High Speed and Small Volume Fan Drive Turbine

A gas turbine engine includes a flex mount for a fan drive gear system. A very high speed fan drive turbine drives the fan drive gear system. 1. A gas turbine engine comprising:a fan shaft driving a fan;a frame which supports said fan shaft;a plurality of gears to drive said fan shaft;a flexible support which at least partially supports said plurality of gears, said flexible support having a lesser stiffness than said frame;a first turbine section providing a drive input into said plurality of gears; anda second turbine section,wherein said first turbine section has a first exit area at a first exit point and rotates at a first speed,wherein said second turbine section has a second exit area at a second exit point and rotates at a second speed, which is faster than the first speed,wherein a first performance quantity is defined as the product of the first speed squared and the first area,wherein a second performance quantity is defined as the product of the second speed squared and the second area, andwherein a ratio of the first performance quantity to the second performance quantity is between about 0.5 and about 1.5.2. The turbine section as set forth in claim 1 , wherein said ratio is above or equal to about 0.8.3. The turbine section as set forth in claim 1 , wherein said first turbine section has at least 3 stages.4. The turbine section as set forth in claim 1 , wherein said first turbine section has up to 6 stages.5. The turbine section as set forth in claim 1 , wherein said second turbine section has 2 or fewer stages.6. The turbine section as set forth in claim 1 , wherein a pressure ratio across the first turbine section is greater than about 5:1.7. The gas turbine engine as set forth in claim 1 , including a ratio of a thrust provided by said engine claim 1 , to a volume of a turbine section including both said high pressure turbine and said low pressure turbine being greater than or equal to about 1.5 and less than or equal to about 5.5 lbf/inch.8. The ...

GAS TURBINE ENGINE FAN DRIVE GEAR SYSTEM DAMPER

A gas turbine engine includes a fan section. A turbine section is coupled to the fan section via a geared architecture. The geared architecture includes a torque frame and a flex support spaced apart from one another at a location. A gear train is supported by the torque frame. A viscous damper is provided between the torque frame and the flex support at the location. 1. A gas turbine engine comprising:first and second members spaced apart from one another at a location;a gear train supported by the first member; anda damper provided between the first member and the second member at the location.2. The gas turbine engine according to claim 1 , wherein the first member is a torque frame claim 1 , and the second member is a flex support having a bellow claim 1 , the flex support grounded to a static structure.3. The gas turbine engine according to claim 2 , wherein the torque frame and flex support are secured to one another by fasteners in an area spaced radially inward from the location.4. The gas turbine engine according to claim 3 , wherein multiple viscous dampers are arranged circumferentially between the torque frame and the flex support claim 3 , and the bellow is provided between the fasteners and the viscous dampers.5. The gas turbine engine according to claim 4 , wherein the torque frame supports a carrier to which star gears are mounted claim 4 , a sun gear is arranged centrally relative to and intermeshing with the star gears claim 4 , and a ring gear circumscribing and intermeshing with the star gears.6. The gas turbine engine according to claim 5 , comprising a fan coupled to the ring gear claim 5 , and a low speed spool coupled to the sun gear.7. The gas turbine engine according to claim 2 , wherein the first and second members respectively include first and second apertures aligned with one another in an axial direction claim 2 , and wherein the damper is a viscous damper extending between and received in the first and second apertures.8. The gas ...

ELECTRIC ACTUATOR FOR ENGINE CONTROL

An electric actuator for control of an engine includes an electric motor coupled to a drive shaft that extends to align a gear interface of the electric actuator with a variable geometry adjustment interface of the engine. A position feedback shaft extends coaxially with respect to the drive shaft. The position feedback shaft is coupled to an output shaft of the gear interface at a gear interface end of the position feedback shaft. A rotational position sensor is coupled to a motor end of the position feedback shaft proximate the electric motor. The drive shaft and the position feedback shaft are sized to position an output ring gear of the output shaft in contact with the variable geometry adjustment interface within a casing of the engine and to further position the electric motor and the rotational position sensor external to the casing of the engine. 1. An electric actuator for control of an engine , the electric actuator comprising:an electric motor coupled to a drive shaft that extends to align a gear interface of the electric actuator with a variable geometry adjustment interface of the engine;a position feedback shaft that extends coaxially with respect to the drive shaft, wherein the position feedback shaft is coupled to an output shaft of the gear interface at a gear interface end of the position feedback shaft; anda rotational position sensor coupled to a motor end of the position feedback shaft proximate the electric motor, wherein the drive shaft and the position feedback shaft are sized to position an output ring gear of the output shaft in contact with the variable geometry adjustment interface within a casing of the engine and to further position the electric motor and the rotational position sensor external to the casing of the engine.2. The electric actuator according to claim 1 , further comprising a retracting mechanism configured to selectively retract the drive shaft and a portion of the gear interface to decouple the drive shaft from the ...

FAN-DRIVE GEAR SYSTEM WITH SEPARATE SCAVENGE PUMP

A lubrication circulation system may comprise a first oil pump configured to retrieve oil from a bearing compartment, and a second oil pump located separately from the first oil pump and configured to retrieve the oil from a fan-drive gear system compartment. The first oil pump may be configured to provide the oil to the bearing compartment and the fan-drive gear system compartment. 1. A lubrication circulation system , comprising:a first oil pump configured to retrieve oil from a bearing compartment; anda second oil pump located separate from the first oil pump and configured to retrieve the oil from a fan-drive gear system compartment, wherein the first oil pump is configured to provide the oil to the bearing compartment and the fan-drive gear system compartment.2. The lubrication circulation system of claim 1 , further comprising an epicyclic gear system in the fan-drive gear system compartment claim 1 , wherein the second oil pump is configured to retrieve the oil from the epicyclic gear system.3. The lubrication circulation system of claim 1 , wherein the second oil pump is located in the fan-drive gear system compartment.4. The lubrication circulation system of claim 1 , further comprising a scavenge line fluidly coupled to the fan-drive gear system compartment and the second oil pump.5. The lubrication circulation system of claim 4 , wherein the scavenge line comprises a diameter less than 2 inches.6. The lubrication circulation system of claim 1 , further comprising an oil tank claim 1 , wherein the second oil pump and the first oil pump urge the oil to the oil tank.7. The lubrication circulation system of claim 6 , wherein the first oil pump is configured to pump the oil from the oil tank into the bearing compartment and the fan-drive gear system compartment.8. The lubrication circulation system of claim 1 , wherein the second oil pump is located closer to the fan-drive gear system compartment than the first oil pump.9. A gas turbine engine claim 1 , ...

FAN DRIVE GEAR SYSTEM SPLINE OIL LUBRICATION SCHEME

An input coupling for a fan drive gear system includes features for maintaining lubricant within a splined interface. The fan drive gear system includes a gear rotatable about an axis that includes an inner spline. The input coupling includes an outer spline engaged to the inner spline of the gear. The input coupling includes an aft oil dam for maintaining lubricant within an interface between the outer spline and the inner spline. 1. A fan drive gear system comprising:a gear rotatable about an axis, the gear including an inner spline; andan input coupling including an outer spline engaged to the inner spline of the gear, the input coupling including an aft oil dam for maintaining lubricant within an interface between the outer spline and the inner spline.2. The fan drive gear system as recited in claim 1 , wherein the gear includes a forward tab extending radially inward from an inner surface of the gear forward of the inner spline.3. The fan drive gear system as recited in claim 1 , wherein the gear includes an aft tab extending radially inward from an inner surface of the gear aft of the inner spline and forward of the aft oil dam.4. The fan drive gear system as recited in claim 1 , wherein the aft oil dam includes a retaining ring supported within an annular channel of the input coupling.5. The fan drive gear system as recited in claim 4 , wherein the retaining ring extends radially outward into contact with an inner surface of the gear.6. The fan drive gear system as recited in claim 1 , wherein the gear comprises a sun gear.7. The fan drive gear system as recited in claim 1 , wherein the input coupling includes at least one U-shaped portion for accommodating relative movement and misalignment with the gear.8. A geared turbofan engine comprising:a fan rotatable about an engine axis;a core engine including a turbine section driving a turbine shaft;a gearbox including a sun gear driven by the turbine shaft; andan input coupling transferring power between the ...

Fan Axial Containment System

A gas turbine engine () and method for containing a fan inside an engine after a fan thrust bearing assembly failure. The engine () may comprise a fan (), a housing () including a compartment (), a fan shaft () inside the compartment () and comprising a bowl (), a support structure () inside the compartment (), a speed sensor pickup () mounted on the outer surface () of the bowl (), a speed sensor () mounted on the support structure (), and a fan thrust bearing assembly () disposed forward of the bowl (). The fan thrust bearing assembly () including a bearing (). The speed sensor () and the sensor pickup () define a defining a sensor gap (). The bearing () and the outer surface () defining a fan thrust bearing gap (), wherein the sensor gap () is less than the fan thrust bearing gap. 1. A gas turbine engine disposed about a longitudinal engine axis (A) , the engine comprising:a fan;an exterior housing including an interior compartment disposed adjacent to the fan;a fan shaft disposed inside the compartment, the fan shaft configured to drive the fan, the fan shaft comprising an elongated pole and a bowl, the bowl including an outer surface and an inner surface;a fan bearing support structure disposed inside the compartment;a speed sensor pickup mounted on the outer surface of the bowl;a speed sensor mounted on the fan bearing support structure, the speed sensor and the sensor pickup defining a sensor gap in the axial direction; anda fan thrust bearing assembly disposed forward of the bowl, the fan thrust bearing assembly including a bearing, the bearing and the outer surface of the bowl defining in the axial direction a fan thrust bearing gap, wherein the sensor gap is less than the fan thrust bearing gap.2. The gas turbine engine of claim 1 , in which the fan thrust bearing assembly includes mounting hardware claim 1 , wherein the fan thrust bearing assembly is disposed adjacent to the pole and is mounted below the speed sensor to the fan bearing support structure ...

GAS TURBINE ENGINE

A gas turbine engine comprises a gearbox comprising a sun gear, an annulus gear, a plurality of planet gears and a planet gear carrier. The sun gear meshes with the planet gears and the planet gears mesh with the annulus gear. Each planet gear is rotatably mounted in the planet gear carrier. The planet gear carrier comprises a plurality of axles arranged parallel to the axis of the gearbox. The axially spaced ends of each axle are secured to the planet gear carrier. Each planet gear is rotatably mounted on a corresponding one of the axles by a bearing arrangement. Each bearing arrangement comprises a journal bearing and a rolling element bearing and each planet gear is rotatably mounted on a journal bearing and each journal bearing is rotatably mounted on an axle by at least one rolling element bearing. 1. A gas turbine engine comprising a bearing arrangement for first and second relatively rotatable members , the bearing arrangement comprising a journal bearing and a rolling element bearing , the second member being arranged coaxially around the first member , the second member having a cylindrical inner surface , the rolling element bearing and the journal bearing being arranged radially between the first member and the second member , the rolling element bearing being positioned radially between the first member and the journal bearing and the journal bearing being positioned radially between the rolling element bearing and the second member , the journal bearing comprising a tubular member , the tubular member having a cylindrical outer surface arranged to cooperate with the cylindrical inner surface of the second member , the tubular member and the second member being relatively rotatable , the bearing arrangement comprising a lubricant supply to supply lubricant to the rolling element bearing and the journal bearing having at least one passage extending radially there-through to supply lubricant from the rolling element bearing to the journal bearing.2. A gas ...

GAS TURBINE

A gas turbine engine, in particular an aircraft engine, including: a turbine connected via an input shaft device to a gearbox device having a sun gear, a planet carrier having a plurality of planet gears attached thereto, and a ring gear, the sun gear is connected to the input shaft device, the planet carrier or the ring gear is connected to a propulsive fan via an output shaft device of the gearbox device, with an inter-shaft bearing system being positioned radially between the input shaft device and the planet carrier of the gearbox device. 1. A gas turbine engine , in particular an aircraft engine , comprising:a turbine connected via an input shaft device to a gearbox device having a sun gear, a planet carrier having a plurality of planet gears attached thereto, and a ring gear,the sun gear is connected to the input shaft device,the planet carrier or the ring gear is connected to a propulsive fan via an output shaft device of the gearbox device, withan inter-shaft bearing system being positioned radially between the input shaft device and the planet carrier of the gearbox device.2. The gas turbine of claim 1 , wherein the inter-shaft bearing system is located axially within or in front of a low-pressure compressor or an intermediate compressor.3. The gas turbine of claim 1 , wherein the inter-shaft bearing system is axially adjacent to the gearbox device on the input and/or the output side claim 1 , in particular with an axial distance measured from the centreline of the gearbox between 0.001 and 4 times the inner radius of the inter-shaft bearing system.4. The gas turbine of claim 1 , wherein the inter-shaft bearing device comprises at least one ball bearing.5. The gas turbine of claim 1 , wherein a fan shaft bearing system is radially located between a fan shaft as part of the output shaft device and a static structure 1 , in particular a static front cone structure 1 , in particular the fan shaft bearing system being axially positioned within the width of the ...

GAS TURBINE

A gas turbine engine, in particular an aircraft engine, includes: a turbine connected via an input shaft device to a gearbox device having a sun gear, a planet carrier having a plurality of planet gears attached thereto, and a ring gear, the sun gear is connected to the input shaft device, the planet carrier or the ring gear is connected to a propulsive fan via an output shaft device of the gearbox device, with a rear carrier bearing device radially between the planet carrier and a static structure on the input side of the gearbox device, an inter-shaft bearing system being positioned radially between the input shaft device and the planet carrier of the gearbox device. The input shaft device having a high rigidity or the input shaft device having a means for decreasing the rigidity, in particular a diaphragm section. 1. A gas turbine engine , comprising:a turbine connected via an input shaft device to a gearbox device having a sun gear, a planet carrier having a plurality of planet gears attached thereto, and a ring gear,the sun gear is connected to the input shaft device,the planet carrier or the ring gear is connected to a propulsive fan via an output shaft device of the gearbox device, witha rear carrier bearing device radially between the planet carrier and a static structure on the input side of the gearbox device,an inter-shaft bearing system being positioned radially between the input shaft device and the planet carrier of the gearbox device.the input shaft device having a high rigidity.2. A gas turbine engine , comprising:a turbine connected via an input shaft device to a gearbox device having a sun gear, a planet carrier having a plurality of planet gears attached thereto, and a ring gear,the sun gear is connected to the input shaft device,the planet carrier or the ring gear is connected to a propulsive fan via an output shaft device of the gearbox device, witha rear carrier bearing device radially between the planet carrier and a static structure on the ...

Gas turbine

A gas turbine engine, in particular an aircraft engine, includes: a turbine connected via an input shaft device to a gearbox device having a sun gear, a planet carrier having a plurality of planet gears attached thereto, and a ring gear, the sun gear is connected to the input shaft device, the planet carrier or the ring gear is connected to a propulsive fan via an output shaft device of the gearbox device, with a rear carrier bearing device radially between the planet carrier and a part on the input shaft on the input side of the gearbox device.

Low pressure ratio fan engine having a dimensional relationship between inlet and fan size

A gas turbine engine assembly may include, among other things, a fan section including a fan, the fan including a plurality of fan blades, a diameter of the fan having a dimension D, each fan blade having a leading edge, and a forward most portion on the leading edges of the fan blades in a first reference plane, a geared architecture, a turbine section including a high pressure turbine and a low pressure turbine, the low pressure turbine driving the fan through the geared architecture, a nacelle surrounding the fan, the nacelle including an inlet portion forward of the fan, a forward edge on the inlet portion in a second reference plane, and a length of the inlet portion having a dimension L between the first reference plane and the second reference plane. A dimensional relationship of L/D may be between 0.30 and 0.40.

Fan drive thrust balance

A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a fan section, a shaft including a bearing system, a turbine section in communication with the shaft, a speed change mechanism coupling the fan section to the turbine section and a biasing device configured to apply a biasing force against the shaft.

MANIFOLD FOR GAS TURBINE

In various embodiments, a manifold assembly () for conducting one or more fluids to a gear assembly () in a gas turbine engine () is provided. The manifold assembly () may comprise a first plate () and a second plate () that rotatably couple together. The manifold assembly () may be retained and/or held together by a channel () and engagement member () arrangement. 1. A manifold assembly , comprising:a first manifold comprising an engagement member;a second manifold having a groove defined therein, wherein the engagement member is installable in the groove; andthe manifold assembly configured to conduct a fluid to a gear assembly through the first manifold and the second manifold.2. The manifold assembly of claim 1 , further comprising an anti-rotation element.3. The manifold assembly of claim 2 , wherein the anti-rotation element is at least one of a fastener claim 2 , a pin claim 2 , an adhesive claim 2 , a tensioning device and a detent assembly.4. The manifold assembly of claim 1 , wherein the groove comprises a receiving portion and a retention portion.5. The manifold assembly of claim 1 , wherein the engagement member is a tongue.6. The manifold assembly of claim 1 , wherein the first manifold is rotatably coupled to the second manifold.7. A turbine engine claim 1 , comprising;a gear assembly; a first portion having a first groove and a second groove, wherein the first groove and the second groove are defined along a diameter of the first portion; and', 'a second portion having a first engagement member installable in the first groove and a second engagement member installable in the second groove., 'a manifold operatively coupled to and in fluid communication with the gear assembly, the manifold comprising8. The turbine engine of claim 7 , wherein the first groove further comprises a first groove portion that is defined radially outward from a centerline of the turbine engine.9. The turbine engine of claim 7 , wherein the engagement member comprises a shaft ...

GEARED TURBOFAN WITH INTEGRAL FRONT SUPPORT AND CARRIER

A gas turbine engine includes a fan section including a fan hub. A speed reduction device includes a star gear system. A turbine section is connected to the fan section through the speed reduction device. A first fan bearing for supporting rotation of the fan hub is connected forward of the speed reduction device. A second fan bearing for supporting rotation of the fan hub is connected aft of the speed reduction device. A first outer race of the first fan bearing is attached to the fan hub. 1. A gas turbine engine comprising:a fan section including a fan hub;a speed reduction device including a star gear system;a turbine section connected to the fan section through the speed reduction device;a first fan bearing for supporting rotation of the fan hub connected forward of the speed reduction device;a second fan bearing for supporting rotation of the fan hub connected aft of the speed reduction device; anda first outer race of the first fan bearing attached to the fan hub.2. The gas turbine engine of including a compressor section configured to rotate with the fan section.3. The gas turbine engine of including a first inner race of the first fan bearing connected to a static structure and a second inner race of the second fan bearing connected to a static structure.4. The gas turbine engine of wherein the first bearing and the second bearing include at least one of roller bearings claim 1 , ball bearings claim 1 , or tapered bearings.5. The gas turbine engine of including a high pressure compressor with a compression ratio of at least 20:1.6. The gas turbine engine of including a low pressure compressor with a compression ratio of at least 2:1.7. The gas turbine engine of including a fan by pass ratio greater than 10.8. The gas turbine engine of claim 1 , wherein the star gear system includes a sun gear claim 1 , star gears claim 1 , a ring gear mechanically attached to the fan section claim 1 , and a carrier fixed from rotation.9. The gas turbine engine of claim 8 , ...

GAS TURBINE ENGINE WITH FAN VARIABLE AREA NOZZLE FOR LOW FAN PRESSURE RATIO

A gas turbine engine includes a fan section with twenty (20) or less fan blades and a fan pressure ratio less than about 1.45. 1. A gas turbine engine comprising:a core nacelle defined about an engine centerline axis;a core engine at least partially disposed within the core nacelle;a fan section with twenty (20) or less fan blades;a gear system driven by the core engine to drive said fan section;a fan nacelle mounted at least partially around said fan section and said core nacelle to define a fan bypass flow path for a fan bypass airflow, said fan bypass airflow having a fan pressure ratio of the fan bypass airflow during engine operation, said fan pressure ratio less than about 1.45;a variable fan nozzle axially movable relative to the fan nacelle, the variable fan nozzle including at least two sectors; anda controller for independently adjusting each of the at least two sectors.2. (canceled)3. The engine as recited in claim 1 , wherein the controller is operable to reduce said fan nozzle exit area at a cruise flight condition.4. The engine as recited in claim 1 , wherein said controller is operable to control said fan nozzle exit area to reduce a fan instability.5. The engine as recited in claim 1 , wherein said fan variable area nozzle defines a trailing edge of said fan nacelle.6. The engine as recited in claim 1 , wherein said fan variable area nozzle is axially movable relative to said fan nacelle.7. (canceled)8. The engine as recited in claim 1 , wherein said fan section defines a corrected fan tip speed less than about 1150 ft/second.9. The engine as recited in claim 1 , wherein said core engine includes a low pressure turbine which defines a low pressure turbine pressure ratio that is greater than about five (5).10. The engine as recited in claim 7 , wherein said core engine includes a low pressure turbine which defines a low pressure turbine pressure ratio that is greater than five (5).11. The engine as recited in claim 1 , further comprising a gear system ...

PINNED MECHANICAL FUSE FOR ENGINE MOTORING SYSTEM

A motoring system for a gas turbine engine having: a reduction gear train having an input and output; a motor operably connected to the input; a clutch operably connected to the output, the clutch in operation engages and disengages the reduction gear train; and a pinned mechanical fuse operably connecting the output to the clutch, the pinned mechanical fuse having at least one shear pin. The pinned mechanical fuse having: an outer sleeve having a first section, second section, inner chamber, outer wall, and at least one through hole connecting the inner chamber to the outer wall within the first section; and an inner sleeve having a first portion, second portion, outer surface, and at least one blind hole located in the outer surface within the second portion. The second portion being located within the inner chamber and operably connected to the outer sleeve through at least one shear pin. 1. A motoring system for a gas turbine engine comprising:a reduction gear train having an input and an output;a motor operably connected to the input;a clutch operably connected to the output, the clutch in operation engages and disengages the reduction gear train; and an outer sleeve having a first section, a second section, an inner chamber, an outer wall, and at least one through hole connecting the inner chamber to the outer wall within the first section; and', 'an inner sleeve having a first portion, a second portion, an outer surface, and at least one blind hole located in the outer surface within the second portion, the second portion being located within the inner chamber and operably connected to the outer sleeve through the at least one shear pin,', 'wherein the at least one through hole is aligned with the at least one blind hole,', 'wherein the at least one shear pin is secured within the at least one through hole and the at least one blind hole, and', 'wherein the at least one shear pin in operation shears when torque on the pinned mechanical fuse is greater than or ...

MECHANICAL SHEAR FUSE FOR ENGINE MOTORING SYSTEM

A motoring system for a gas turbine engine having: a reduction gear train having an input and an output; an electric motor operably connected to the input; a clutch operably connected to the output, the clutch in operation engages and disengages the reduction gear train; and a mechanical shaft fuse operably connecting the output to the clutch, the mechanical shaft fuse in operation shears when torque on the mechanical shaft fuse is greater than or equal to a selected value. The mechanical shaft fuse includes a plurality of through holes. 1. A motoring system for a gas turbine engine comprising:a reduction gear train having an input and an output;an electric motor operably connected to the input;a clutch operably connected to the output, the clutch in operation engages and disengages the reduction gear train; anda mechanical shaft fuse operably connecting the output to the clutch, the mechanical shaft fuse in operation shears when torque on the mechanical shaft fuse is greater than or equal to a selected value,wherein the mechanical shaft fuse includes a plurality of through holes.2. The motoring system of claim 1 , wherein:the plurality of through holes are oriented around an approximate axial center point of the mechanical shaft fuse.3. The motoring system of claim 1 , wherein:each of the holes has a diameter of about 0.187 inches (0.475 centimeters).4. The motoring system of claim 1 , wherein:the plurality of holes comprises six holes.5. The motoring system of claim 1 , wherein:the selected value is about 64 foot-pounds (87 newton-meters).6. The motoring system of claim 1 , wherein:the mechanical shaft fuse includes a first outer diameter of about 0.63 inches (1.6 centimeters).7. The motoring system of claim 1 , wherein:the mechanical shaft fuse is hollow and includes a thickness of about 0.09 inches (0.229 centimeters).8. The motoring system of claim 1 , wherein:the mechanical shaft fuse has a hexagonal cross-sectional shape.9. The motoring system of claim 6 , ...

GAS TURBINE ENGINE ELECTRICAL GENERATOR

An aircraft gas turbine engine () comprises a main engine shaft () arranged to couple a turbine () and a compressor (), the main engine shaft () defining an axial direction (). The gas turbine engine () further comprises at least one radially extending offtake shaft () coupled to the main engine shaft (), and a radially extending electric machine () coupled to the radially extending offtake shaft (). 1. An aircraft gas turbine engine comprising:a main engine shaft arranged to couple a turbine and a compressor, the main engine shaft defining an axial direction;at least one radially extending offtake shaft coupled to the main engine shaft; anda radially extending electric machine coupled to the radially extending offtake shaft.2. A gas turbine engine according to claim 1 , wherein the electric machine is directly coupled to the radially extending offtake shaft.3. A gas turbine engine according to claim 1 , wherein the electric machine is coupled to the radially extending offtake shaft by a reduction gearbox.4. A gas turbine engine according to claim 1 , wherein the electric machine comprises at least one of an electric motor configured to provide motive power to start the gas turbine engine in a starting mode claim 1 , and a generator configured to generate electrical power when in a running mode.5. A gas turbine engine according to claim 1 , wherein the electric machine comprises an axial flux electric machine in which a stator of the electric machine is axially offset relative to a rotor of the electric machine.6. A gas turbine engine according to claim 1 , wherein the electric machine comprises a radial flux electric machine claim 1 , in which a stator of the electric machine is radially inward or radially outward of a rotor of the electric machine.7. A gas turbine engine according to claim 1 , wherein the gas turbine engine comprises a plurality of radially extending offtake shafts circumferentially arrayed around the main engine shaft.8. A gas turbine engine ...

GEARED TURBOFAN WITH FOUR STAR/PLANETARY GEAR REDUCTION

A turbofan engine assembly includes a nacelle and a turbofan engine. The turbofan engine includes a fan, which includes a fan rotor having fan blades, and a nacelle enclosing the fan rotor and the blades. A turbine rotor drives the fan rotor. An epicyclic gear reduction is positioned between the fan rotor and the turbine rotor. The epicyclic gear reduction includes a ring gear, a sun gear, and four intermediate gears that engage the sun gear and the ring gear. A gear ratio of the gear reduction is greater than 3.06. The fan drive turbine drives the sun gear to, in turn, drive the fan rotor. 1. A turbofan engine comprising:a fan including a fan rotor having fan blades, and housing enclosing said fan rotor and said blades;a turbine rotor driving said fan rotor;an epicyclic gear reduction positioned between said fan rotor and said turbine rotor, said epicyclic gear reduction including a ring gear, a sun gear, and four intermediate gears that engage said sun gear and said ring gear, a gear ratio of said gear reduction is greater than 3.06;wherein said fan drive turbine drives said sun gear to, in turn, drive said fan rotor;a bypass ratio is defined as a volume of air delivered by said fan into a bypass duct inward of said housing compared to a volume of air delivered into a compressor, said bypass ratio is greater than or equal to 12.0; andwherein there is a primary oil supply supplying oil to journal bearings that support said intermediate gears.2. The turbofan engine as set forth in claim 1 , wherein said gear ratio is greater than or equal to 4.0.3. The turbofan engine as set forth in claim 2 , wherein said gear ratio is greater than or equal to 4.2.4. The turbofan engine as set forth in claim 2 , wherein said gear ratio is less than or equal to 4.4.5. The turbofan engine as set forth in claim 2 , wherein there is an auxiliary oil supply supplying oil to said journal bearings when there is windmilling of said fan rotor.6. The turbofan engine as set forth in claim 1 , ...

GEARED TURBOFAN WITH INTEGRAL FRONT SUPPORT AND CARRIER

A gas turbine engine includes a nacelle, and a bypass flow path in a bypass duct within the nacelle of the turbofan engine. A fan section includes a fan with fan blades. The fan section drives air along the bypass flow path. A fan shaft drives a fan that has fan blades and the fan rotates about a central longitudinal axis of the turbofan engine. A speed reduction device includes an epicyclic gear system. A turbine section is connected to the fan section through the speed reduction device and the turbine section rotates about the central longitudinal axis. A first fan bearing for supporting rotation of the fan hub is located axially forward of the speed reduction device. A second fan bearing for supporting rotation of the fan hub is located axially aft of the speed reduction device. A first outer race of the first fan bearing is fixed relative to the fan hub. 1. A gas turbine engine comprising:a fan section including a fan having fan blades extending from a fan hub, wherein said fan section drives air along a bypass flow path;a fan shaft driving said fan and said fan rotates about a central longitudinal axis of said gas turbine engine;a speed reduction device including an epicyclic gear system including a plurality of intermediate gears supported on a corresponding one of a plurality of flexible posts on a static carrier of the epicyclic gear system;a turbine section connected to the fan section through the speed reduction device and said turbine section rotates about said central longitudinal axis;a first fan bearing for supporting rotation of the fan hub located axially forward of the speed reduction device;a second fan bearing for supporting rotation of the fan hub located axially aft of the speed reduction device; anda first outer race of the first fan bearing is fixed relative to the fan hub.2. The gas turbine engine of claim 1 , wherein each of the plurality of flexible posts extend from the static carrier at a proximal end to a distal free end.3. The gas ...

Gas turbine engine

A gas turbine engine includes a spool which rotates in use, a static structure and an air bearing, the air bearing being provided at an interface between the spool and the static structure.

GEARED ARCHITECTURE TO PROTECT CRITICAL HARDWARE DURING FAN BLADE OUT

A turbofan engine including a fan section including a plurality of fan blades rotatable about an axis, a compressor including a plurality of compressor blades, a turbine including a plurality of turbine blades and a geared architecture driven by the turbine for driving the fan section at a speed and direction different than the turbine is disclosed. A rub strip proximate at least one of the compressor blades, the turbine blades and the fan blades slows rotation when engaged. The rub strip generates a torque opposing rotation when in an engaged condition that is between 2 and 6 times a torque encountered in a non-engaged condition.

Turbofan Engine Assembly Methods

A method for assembling a turbofan engine () comprises coupling a bearing assembly () and a shaft () as a unit to a bearing support (). A transmission () and the fan shaft () are installed to a front frame assembly (). The bearing assembly and shaft are installed as a unit so that the shaft engages a central gear () of the transmission and the bearing support engages the front frame assembly.

GEARED GAS TURBINE ENGINE AND A GEARBOX

A gas turbine engine comprises a gearbox comprises a sun gear, an annulus gear, a plurality of planet gears and a carrier. The sun gear meshes with the planet gears and the planet gears mesh with the annulus gear. The planet gear carrier comprising a first ring, a second ring spaced axially from the first ring and a plurality of circumferentially spaced axles extending axially between the first ring and the second ring. Each planet gear is rotatably mounted on a respective one of the axles and the axles are arranged at a first radius. At least one of the first ring and the second ring comprises a metal matrix composite material ring and the metal matrix composite ring comprises a ring of reinforcing fibres and the ring of reinforcing fibres having a second radius greater than the first radius. 1. A gas turbine engine comprising a gearbox , the gearbox comprising a sun gear , an annulus gear , a plurality of planet gears and a carrier , the sun gear meshing with the planet gears and the planet gears meshing with the annulus gear , the carrier comprising a first ring , a second ring spaced axially from the first ring and a plurality of circumferentially spaced axles extending axially between the first ring and the second ring , each planet gear being rotatably mounted on a respective one of the axles , the axles being arranged at a first radius , at least one of the first ring and the second ring comprising a metal matrix composite material , the metal matrix composite material comprising a ring of reinforcing fibres and the ring of reinforcing fibres having a second radius greater than the first radius.2. A gas turbine engine as claimed in wherein the first ring comprises a first metal matrix composite material and the second ring comprises a second metal matrix composite material claim 1 , the first metal matrix composite material comprises a first ring of reinforcing fibres claim 1 , the first ring of reinforcing fibres having a second radius greater than the first ...

Geared Turbocharged Engine

A geared turbocharged engine, having a gearbox, a drive and output assembly, and an oil supply system. The gearbox has a housing. The oil supply system has a reservoir, a supply line, an oil pump to convey oil from the reservoir, and a return line, to return oil from the gearbox to the reservoir. The housing is mounted on a support plate of reservoir a supply-line opening and a return-line opening are introduced into the housing and the support plate such that the supply-line opening of the support plate and the supply-line opening of the housing and the return-line opening of the gearbox and the return-line opening of the support plate are flush, so that the inflow of oil to the gearbox and the return flow of oil from the gearbox take place via the gearbox housing. 1. A geared turbocharged engine , comprising: at least one drive and output assembly; a gearbox housing;', 'a central gear wheel with a gear-wheel shaft; and', 'at least one pinion, with at least one pinion shaft, meshing into the central gear wheel; and, 'a gearbox comprising a support plate;', 'an oil storage reservoir in which oil is held;', 'an oil supply line by which oil can be drawn from the oil storage reservoir and guided toward the gearbox to be lubricated;', 'an oil pump configured to convey the oil out of the oil storage reservoir, and', 'an oil return line by which oil emanating from the gearbox flows back into the oil storage reservoir;', 'wherein:', 'the gearbox housing is mounted standing on the support plate by a gearbox housing base;', 'a supply-line opening and a return-line opening are introduced into each of the gearbox housing base and the support plate such that the supply-line opening of the support plate and the supply-line opening of the gearbox housing base are flush and the return-line opening of the gearbox housing base and the return-line opening of the support plate,', 'whereby inflow of the oil to the gearbox and return flow of the oil from the gearbox take place via the ...

COUNTER ROTATING POWER TURBINE WITH REDUCTION GEARBOX

The present disclosure is directed to a turbine engine defining a longitudinal direction, a radial direction, and a circumferential direction. The turbine engine includes a power turbine including a first turbine rotor assembly interdigitated with a second turbine rotor assembly along the longitudinal direction; a gear assembly coupled to the first turbine rotor assembly and the second turbine rotor assembly, wherein the gear assembly includes a first input interface coupled to the first turbine rotor assembly, a second input interface coupled to the second turbine rotor assembly, and one or more third gears coupled to the first input interface and the second input interface therebetween; and a first output shaft and a second output shaft, wherein each of the first output shaft and the second output shaft are configured to couple to an electrical load device. 1. A turbine engine defining a longitudinal direction , a radial direction , and a circumferential direction , the turbine engine comprising:a power turbine comprising a first turbine rotor assembly interdigitated with a second turbine rotor assembly along the longitudinal direction;a gear assembly coupled to the first turbine rotor assembly and the second turbine rotor assembly, wherein the gear assembly comprises a first input interface coupled to the first turbine rotor assembly, a second input interface coupled to the second turbine rotor assembly, and one or more third gears coupled to the first input interface and the second input interface therebetween; anda first output shaft and a second output shaft, wherein each of the first output shaft and the second output shaft are configured to couple to an electrical load device.2. The turbine engine of claim 1 , wherein the gear assembly further comprises an output interface claim 1 , and wherein the first output shaft is coupled to the output interface of the gear assembly.3. The turbine engine of claim 1 , wherein the second turbine rotor assembly is coupled ...

GAS TURBINE ENGINE WITH GEARBOX HEALTH FEATURES

A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a fan and a braking system. The braking system is configured to selectively engage the fan during ground windmilling to apply a first level of braking to slow rotation of the fan. Further, when the rotation of the fan sufficiently slows, the braking system is further configured to apply a second level of braking more restrictive than the first level of braking. 1. A gas turbine engine , comprising:a fan; anda lubrication system configured to pump lubricant into a fan drive gearbox when the fan is windmilling at any rotational speed and direction.2. The gas turbine engine as recited in claim 1 , wherein the lubrication system is configured to pump lubricant to the fan drive gearbox when the fan rotates below 1 claim 1 ,000 rpm.3. The gas turbine engine as recited in claim 1 , wherein:the lubrication system includes a main pump and a main reservoir fluidly coupled to the main pump, the main pump configured to pump lubricant from the main reservoir to the fan drive gearbox during normal operating conditions; andthe lubrication system further includes a secondary pump and a secondary reservoir fluidly coupled to the secondary pump, the secondary pump configured to pump lubricant from the secondary reservoir to the fan drive gearbox when the main pump does not provide adequate lubricant to the fan drive gearbox.4. The gas turbine engine as recited in claim 3 , wherein the secondary pump is one of (1) an accessory pump claim 3 , (2) a rotary-shaft driven pump claim 3 , (3) an electrically-driven pump claim 3 , and (4) an aircraft hydraulic system-powered pump.5. A gas turbine engine claim 3 , comprising:a fan;a geared architecture coupled to the fan;at least one sensor configured to generate signals indicative of a condition of the geared architecture; anda control unit electrically coupled to the at least one sensor, the control unit configured to interpret ...

Air inlet for a gas turbine engine

A radial air inlet for a gas turbine engine. The air inlet has an inlet duct defined between two axially-spaced radially-extending annular walls and has a plurality of circumferentially-spaced axially-extending struts extending between the annular walls adjacent a radially-outer portion of the air inlet. At least one of the struts has an internal passage extending between a first opening in a forward end of the strut and a second opening in an aft end of the strut, the first and second openings being axially spaced apart. A transmission shaft extends through the internal passage of said strut.

COMPRESSOR OPERABILITY CONTROL FOR HYBRID ELECTRIC PROPULSION

A hybrid electric propulsion system includes a gas turbine engine having a low speed spool and a high speed spool. The low speed spool includes a low pressure compressor and turbine, and the high speed spool includes a high pressure compressor and turbine. The hybrid electric propulsion system includes an electric generator configured to extract power from the low speed spool, an electric motor configured to augment rotational power of the high speed spool, and a controller. The controller is operable to determine a target operating condition of the low pressure compressor to achieve a compressor stability margin in the gas turbine engine, determine a current operating condition of the low pressure compressor, and control a power transfer between the electric generator of the low speed spool and the electric motor of the high speed spool to adjust the current operating condition based on the target operating condition. 1. A hybrid electric propulsion system comprising:a gas turbine engine comprising a low speed spool and a high speed spool, the low speed spool comprising a low pressure compressor and a low pressure turbine, and the high speed spool comprising a high pressure compressor and a high pressure turbine;an electric generator configured to extract power from the low speed spool;an electric motor configured to augment rotational power of the high speed spool; and determine a target operating condition of the low pressure compressor to achieve a compressor stability margin in the gas turbine engine;', 'determine a current operating condition of the low pressure compressor; and', 'control a power transfer between the electric generator of the low speed spool and the electric motor of the high speed spool to adjust the current operating condition based on the target operating condition., 'a controller operable to2. The hybrid electric propulsion system of claim 1 , wherein the target operating condition of the low pressure compressor is determined by the ...

GAS TURBINE ENGINE WITH LOW STAGE COUNT LOW PRESSURE TURBINE

A gas turbine engine according to an example of the present disclosure includes, among other things, a fan section, a first compressor section including three (3) or more stages, a second compressor section including between eight (8) and thirteen (13) stages, and a first turbine section operable for driving the first compressor section, the first turbine section including between three (3) and six (6) stages, a second turbine section operable for driving the second compressor section, and a gear train defined along an engine centerline axis. One of the first turbine section and the second turbine section is operable to drive the fan section through the gear train. 1. A gas turbine engine comprising:a fan section;a first compressor section including three (3) or more stages;a second compressor section including between eight (8) and thirteen (13) stages;a first turbine section operable for driving the first compressor section, the first turbine section including between three (3) and six (6) stages;a second turbine section operable for driving the second compressor section; anda gear train defined along an engine centerline axis, wherein one of the first turbine section and the second turbine section is operable to drive the fan section through the gear train.2. The engine as recited in claim 1 , wherein the first turbine section includes three (3) stages.3. The engine as recited in claim 2 , wherein the first turbine defines a pressure ratio that is greater than about 5.0.4. The engine as recited in claim 1 , including a fan case circumscribing the fan section that defines a fan bypass airflow claim 1 , and wherein a bypass ratio of the engine is greater than about 10.0.5. The engine as recited in claim 1 , wherein said gear train defines a gear reduction ratio of greater than about 2.3.6. The gas turbine engine as set forth in claim 1 , further comprising a fan variable area nozzle to vary a fan nozzle exit area and adjust a fan pressure ratio of fan bypass ...

GAS TURBINE ENGINE WITH LOW STAGE COUNT LOW PRESSURE TURBINE

A gas turbine engine according to an example of the present disclosure includes, among other things, a fan section, a low spool that includes a low pressure compressor section, the low pressure compressor section including three (3) or more stages, and a high spool including a high pressure compressor section. The high pressure compressor section includes between eight to thirteen (8-13) stages. A gear train is defined along an engine axis. The low spool is operable to drive the fan section through the gear train. 1. A gas turbine engine comprising:a fan section;a low spool that includes a low pressure compressor section, said low pressure compressor section includes three (3) or more stages;a high spool that includes a high pressure compressor section, said high pressure compressor section including between eight to thirteen (8-13) stages; anda gear train defined along an engine axis, said low spool operable to drive said fan section through said gear train.2. The engine as recited in claim 1 , wherein said low pressure compressor includes three (3) stages.3. The engine as recited in claim 1 , wherein said low pressure compressor includes four (4) stages.4. The engine as recited in claim 1 , wherein said high pressure compressor includes eight (8) stages.5. The engine as recited in claim 1 , wherein said low spool includes a low pressure turbine with three to six (3-6) stages.6. The engine as recited in claim 5 , wherein said low pressure turbine defines a low pressure turbine pressure ratio that is greater than about five (5).7. The engine as recited in claim 1 , wherein said low spool includes a low pressure turbine with three to six (3-6) stages claim 1 , and said low pressure turbine defines a low pressure turbine pressure ratio that is greater than five (5) claim 1 , said low pressure compressor includes three (3) or four (4) stages.8. The engine as recited in claim 7 , wherein said gear train defines a gear reduction ratio of greater than about 2.3.9. The ...

GAS TURBINE ENGINE WITH AIR-OIL COOLER OIL TANK

A thermal management system according to one disclosed non-limiting embodiment of the present disclosure includes an at least partially annular oil tank; and a fan airflow diverter upstream of the cooling fins. 1. An air-oil cooler oil tank system comprising:an at least partially annular oil tank; anda fan airflow diverter upstream of said at least partially annular oil tank.2. The system as recited in claim 1 , wherein said at least partially annular oil tank is in fluid communication with a geared architecture.3. The system as recited in claim 1 , further comprising an array of cooling fins defined about an outer diameter of said at least partially annular oil tank claim 1 , said fan airflow diverter in communication with said array of cooling fins.4. The system as recited in claim 3 , wherein said array of cooling fins include a shroud that is flush along a fan bypass flowpath.5. The system as recited in claim 3 , wherein said array of cooling fins include a shroud that is flush with a core nacelle.6. The system as recited in claim 3 , wherein said array of cooling fins are downstream of a Fan Exit Guide Vane array.7. The system as recited in claim 3 , wherein said array of cooling fins are defined about an outer periphery of said at least partially annular oil tank.8. The system as recited in claim 1 , further comprising a bleed compartment which contains said fan airflow diverter and said at least partially annular oil tank.9. The system as recited in claim 1 , further comprising a shield radially inboard of said at least partially annular oil tank.10. The system as recited in claim 1 , wherein said at least partially annular oil tank is manufactured of a multiple of arcuate segments.11. A gas turbine engine comprising:an at least partially annular oil tank with an array of cooling fins that include a shroud adjacent to a fan bypass flowpath12. The gas turbine engine as recited in claim 11 , further comprising a geared architecture claim 11 , wherein said at ...

Geared Architecture Turbofan Engine Thermal Management System and Method

A method of sizing a heat exchanger for a geared architecture gas turbine engine includes sizing a minimum frontal area of at least one heat exchanger located in communication with a fan bypass airflow such that a ratio of waste heat area to horsepower generation characteristic area is between 1.6 to 8.75. 1. A method of sizing a heat exchanger for a geared architecture gas turbine engine comprising:sizing a minimum frontal area of at least one heat exchanger located in communication with a fan bypass airflow such that a ratio of waste heat area to horsepower generation characteristic area is between 1.6 to 17.5.2. The method as recited in claim 1 , wherein the waste heat area is defined by the minimum frontal area of the HEX.3. The method as recited in claim 1 , wherein the horsepower generation characteristic area is defined by an exit area of a high pressure compressor.4. The method as recited in claim 1 , further comprising locating the at least one heat exchanger within a fan bypass airflow path such that the ratio of waste heat area to horsepower generation characteristic area is between 1.6 to 8.75.5. The method as recited in claim 1 , further comprising locating the at least one heat exchanger with respect to a fan duct total pressure profile.6. A method of sizing a heat exchanger for a geared architecture gas turbine engine comprising:determining an efficiency of a geared architecture;determination a temperature requirement of the oil at a particular flight condition;determining a fan pressure ratio; andsizing a minimum frontal area of the at least one heat exchanger in response to the efficiency of the geared architecture, the temperature requirements of the oil at a particular flight condition, and the fan pressure ratio.7. The method as recited in claim 6 , further comprising sizing the minimum frontal area of the at least one heat exchanger such that a ratio of waste heat area to horsepower generation characteristic area is between 1.6 to 17.5.8. The ...

PIVOT DOOR THRUST REVERSER

A geared turbofan engine includes a fan, a low pressure turbine, a geared architecture, and a nacelle. The fan and low pressure turbine are capable of rotation about an axial centerline of the gas turbine engine. The geared architecture connects the fan to be driven by the low pressure turbine. The nacelle is disposed circumferentially around the fan and defines a portion of a bypass flow duct. The nacelle includes a thrust reverser assembly with one or more doors that pivot to block at least a portion of the bypass flow duct in a deployed position. 1. A geared turbofan engine with a bypass ratio that is greater than six , the engine comprising:a first spool capable of rotation about an axial centerline of the gas turbine engine;a second spool capable of rotation about the axial centerline;a fan capable of rotation about the axial centerline;a geared architecture, wherein the fan is coupled to the low pressure compressor and the low pressure turbine through the geared architecture; and 'a thrust reverser assembly with one or more doors that pivot to block at least a portion of the bypass flow duct in a deployed position, wherein the thrust reverser has an effective flow area that is greater than a bypass flow duct exit area of the bypass flow duct.', 'a nacelle arranged circumferentially around the axial centerline and defining a portion of a bypass flow duct, the nacelle comprising2. The geared turbofan engine of claim 1 , wherein the one or more doors pivot on hinges.3. The geared turbofan engine of claim 1 , wherein an actuator drives the one or more pivoting doors between a stowed position and a deployed position.4. The geared turbofan engine of claim 3 , wherein the actuator includes a rod that is extensible and retractable.5. The geared turbofan engine of claim 1 , wherein the first and second spools include a fan claim 1 , a low pressure compressor claim 1 , a high pressure compressor claim 1 , a low pressure turbine claim 1 , and a high pressure turbine.6. ...

TURBINE SHAFT POWER TAKE-OFF

A multi-spool gas turbine engine comprises a low pressure (LP) spool and a high pressure (HP) spool independently rotatable of one another about an engine axis. The LP spool comprises an LP turbine, an LP compressor and an LP shaft. The HP pressure spool comprises an HP turbine, an HP compressor and an HP shaft. The LP turbine is in fluid flow communication with the HP turbine and disposed downstream therefrom. The HP compressor is in fluid flow communication with the LP compressor and disposed downstream therefrom. The LP shaft has an upstream shaft portion extending upstream of the LP turbine to a location upstream of the LP compressor to provide a first power take-off at an upstream end of the engine and a downstream shaft portion extending downstream of the LP turbine to provide a second power take-off at a downstream end of the engine, thereby allowing mounting of a reduction gear box at either end of the engine. 1. A multi-spool gas turbine engine comprising: a low pressure (LP) spool and a high pressure (HP) spool independently rotatable of one another about an engine axis , the LP spool comprising an LP turbine , an LP compressor and an LP shaft , the HP pressure spool comprising an HP turbine , an HP compressor and an HP shaft; the LP turbine being in fluid flow communication with the HP turbine and disposed downstream therefrom , the HP compressor being in fluid flow communication with the LP compressor and disposed downstream therefrom , the LP shaft having an upstream shaft portion extending upstream of the LP turbine to a location upstream of the LP compressor to provide a first power take-off at an upstream end of the engine and a downstream shaft portion extending downstream of the LP turbine to provide a second power take-off at a downstream end of the engine.2. The multi spool gas turbine engine defined in claim 1 , further comprising an accessory gearbox (AGB) disposed at the upstream end of the engine upstream of the LP compressor claim 1 , the ...

AIRCRAFT TURBOMACHINE WITH REDUCTION GEARSET

A turbomachine includes a single ducted fan including a first shaft rotated by a second shaft via a speed reduction gearset, the second shaft being rotated by a third shaft of a turbine, the first shaft being guided in rotation with respect to a fixed structure via a first bearing and a second bearing placed upstream of the speed reduction gearset. The second shaft is guided in rotation with respect to the first shaft via a rolling bearing placed upstream of the speed reduction gearset, the rolling bearing comprising an outer ring housed in the first shaft, an inner ring connected to the second shaft and rolling elements arranged between the inner and outer rings. 1. A turbomachine with an axis of rotation , comprising: a single ducted fan comprising a first shaft carrying blades and driven in rotation by a second shaft via a speed reducer , said second shaft being driven in rotation by a third shaft of a turbine , said first , second and third shafts sharing the axis of rotation , said first shaft being guided in rotation with respect to a fixed structure of the turbomachine via a first bearing and a second bearing located upstream of said speed reducer , wherein the second shaft is guided in rotation with respect to the first shaft via a third rolling bearing located upstream of said speed reducer , said third rolling bearing comprising an outer ring housed in said first shaft , an inner ring attached to said second shaft , and rolling elements disposed between said inner and outer rings.2. The turbomachine according to claim 1 , further comprising sealing means located upstream of said third bearing and configured to ensure tightness between said first shaft and said second shaft.3. The turbomachine according to claim 2 , wherein said sealing means comprise a first member and a second member claim 2 , said first member comprising at least one annular lip in radial contact with an abradable ring of said second member.4. The turbomachine according to claim 3 , ...

FLEXIBLE SUPPORT STRUCTURE FOR A GEARED ARCHITECTURE GAS TURBINE ENGINE

A gas turbine engine includes a fan shaft that drives a fan that has fan blades. A fan shaft support that supports the fan shaft defines a fan shaft support transverse stiffness. A gear system is connected to the fan shaft and includes a gear mesh that defines a gear mesh transverse stiffness. A flexible support supports the gear system and defines a flexible support transverse stiffness. The flexible support transverse stiffness is less than 11% of the fan shaft support transverse stiffness and less than 8% of the gear mesh transverse stiffness. 1. A gas turbine engine , comprising:a fan shaft driving a fan having fan blades;a fan shaft support that supports said fan shaft defining a fan shaft support transverse stiffness;a gear system connected to said fan shaft, wherein said gear system includes a gear mesh defining a gear mesh transverse stiffness;a flexible support supporting said gear system defining a flexible support transverse stiffness; andwherein said flexible support transverse stiffness is less than 11% of said fan shaft support transverse stiffness and less than 8% of said gear mesh transverse stiffness.2. The gas turbine engine of claim 1 , wherein said gear mesh defines a gear mesh lateral stiffness and said gear system includes a ring gear defining a ring gear lateral stiffness and said ring gear lateral stiffness is less than 12% of said gear mesh lateral stiffness.3. The gas turbine engine of claim 2 , wherein said flexible support defines a flexible support lateral stiffness and said fan shaft support defines a fan shaft support lateral stiffness and said flexible support lateral stiffness is less than 11% of said fan shaft support lateral stiffness.4. The gas turbine engine of claim 3 , wherein said gear system is an epicyclic gear system claim 3 , said gear mesh defines a gear mesh lateral stiffness and said flexible support lateral stiffness is less than 8% of said gear mesh lateral stiffness.5. The gas turbine engine of claim 4 , wherein said ...

FLEXIBLE SUPPORT STRUCTURE FOR A GEARED ARCHITECTURE GAS TURBINE ENGINE

A gas turbine engine according to an example of the present disclosure includes, among other things, a fan shaft driving a fan having fan blades. A fan shaft support supports the fan shaft and defines a support transverse stiffness. A gear system is connected to the fan shaft and includes a gear mesh defining a gear mesh transverse stiffness and a reduction ratio greater than 2.3. A gear system input is connected to the gear system and defines a gear system input lateral stiffness. A flexible support supports the gear system and defines a flexible support transverse stiffness. The gear system input lateral stiffness is less than 5% of the gear mesh lateral stiffness and the flexible support transverse stiffness is less than 20% of the fan shaft support transverse stiffness. 1. A gas turbine engine , comprising:a fan shaft drivingly connected to a fan, said fan having fan blades;a nacelle, and a bypass flow path in a bypass duct within said nacelle;a fan shaft support that supports said fan shaft;a gear system connected to said fan shaft, said gear system includes a ring gear defining a ring gear lateral stiffness and a ring gear transverse stiffness, a gear mesh defining gear mesh lateral stiffness and a gear mesh transverse stiffness, and a reduction ratio greater than 2.3; andwherein said ring gear lateral stiffness and said ring gear transverse stiffness are each less than 12% of a respective one of said gear mesh lateral stiffness and said gear mesh transverse stiffness.2. The gas turbine engine of claim 1 , further comprising a flexible support which supports said gear system relative to a static structure and defines a flexible support lateral stiffness and a flexible support transverse stiffness claim 1 , where at least one of said flexible support lateral stiffness and said flexible support transverse stiffness is less than 8% of a respective one of said gear mesh lateral stiffness and said gear mesh transverse stiffness.3. The gas turbine engine of claim 2 ...

GEARED TURBOFAN WITH INTEGRAL FRONT SUPPORT AND CARRIER

A gas turbine engine includes a nacelle, and a bypass flow path in a bypass duct within the nacelle of the turbofan engine. A fan section includes a fan with fan blades. The fan section drives air along the bypass flow path. A fan shaft drives a fan that has fan blades and the fan rotates about a central longitudinal axis of the turbofan engine. A speed reduction device includes an epicyclic gear system. A turbine section is connected to the fan section through the speed reduction device and the turbine section rotates about the central longitudinal axis. A first fan bearing for supporting rotation of the fan hub is located axially forward of the speed reduction device. A second fan bearing for supporting rotation of the fan hub is located axially aft of the speed reduction device. A first outer race of the first fan bearing is fixed relative to the fan hub. 1. A gas turbine engine comprising:a nacelle, and a bypass flow path in a bypass duct within said nacelle of said turbofan engine;a fan section including a fan with fan blades, wherein said fan section drives air along said bypass flow path a fan shaft driving a fan having fan blades and said fan rotates about a central longitudinal axis of said turbofan engine;a speed reduction device including an epicyclic gear system;a turbine section connected to the fan section through the speed reduction device and said turbine section rotates about said central longitudinal axis;a first fan bearing for supporting rotation of the fan hub located axially forward of the speed reduction device;a second fan bearing for supporting rotation of the fan hub located axially aft of the speed reduction device; anda first outer race of the first fan bearing fixed relative to the fan hub.2. The gas turbine engine of claim 1 , wherein the second fan bearing includes an outer race and the outer race of the first fan bearing and the outer race of the second fan bearing are fixed relative to the fan hub and rotate with the fan hub in the ...

V-SHAPED GEARBOX FOR DRIVING TURBOMACHINE EQUIPMENT

A gearbox for driving equipment of a turbomachine, including a substantially V-shaped box having two arms joined together by a joining part, the arms containing gear trains joined together at the joining part, and an attachment to the turbomachine. The attachment include a mechanism for insetting and/or for attaching the joining part and a suspension device of the arms. 111-. (canceled)12. A turbine engine comprising a gearbox comprising a substantially V-shaped casing and comprising two arms which are interconnected by a joining part which extends over part of the length of the arms , the arms being formed by two parts of said casing which are separate outside said joining part , the arms containing gear lines which are located in non-parallel planes and are interconnected in the region of the joining part , the gearbox further comprising means for attachment to the turbine engine , the attachment means comprising both means for embedding and/or attaching the joining part and means for the suspension of the arms , the gearbox being mounted downstream of a fan in the space located between a casing of a compressor and an inner cylindrical wall of an intermediate casing , the arms being located symmetrically on either side of a plane passing through the longitudinal axis of the turbine engine.13. The turbine engine according to claim 12 , wherein the joining part of the gearbox is oriented towards the upstream end and the arms extend towards the downstream end.14. The turbine engine according to claim 12 , wherein the joining part of the gearbox is oriented towards the downstream end and the arms extend towards the upstream end.15. The turbine engine according to claim 12 , wherein the casing comprises a tubular member through which a shaft for driving the gear lines is intended to pass claim 12 , said tubular member comprising an end which is connected to the part for joining the arms and an opposite free end part comprising the embedding and/or attachment means ...

LOW NOISE TURBINE FOR GEARED TURBOFAN ENGINE

A gas turbine engine comprises a fan and a turbine having a fan drive rotor. There is also a second turbine rotor. A gear reduction effects a reduction in the speed of the fan relative to an input speed from the fan drive rotor. The fan drive rotor has a number of turbine blades in at least one of a plurality of rows of the fan drive rotor, and the turbine blades operate at least some of the time at a rotational speed. The number of turbine blades in the at least one row and the rotational speed are such that the following formula holds true for the at least one row of the fan drive turbine: (number of blades×speed)/60≧5500 Hz. The rotational speed is in revolutions per minute. A method of designing a gas turbine engine, and a turbine module are also disclosed. 1. A gas turbine engine comprising:a fan and a turbine having a fan drive rotor, there also being a second turbine rotor;a gear reduction effecting a reduction in the speed of said fan relative to an input speed from said fan drive rotor;said fan drive rotor having a number of turbine blades in at least one of a plurality of rows of said fan drive rotor, and said turbine blades operating at least some of the time at a rotational speed, and said number of turbine blades in said at least one row and said rotational speed being such that the following formula holds true for said at least one row of the fan drive turbine(number of blades×speed)/60≧5500 Hz; andsaid rotational speed being in revolutions per minute.2. The gas turbine engine as set forth in claim 1 , wherein the formula results in a number greater than or equal to 6000 Hz.3. The gas turbine engine as set forth in claim 2 , wherein said gas turbine engine is rated to produce 15 claim 2 ,000 pounds of thrust or more.4. The gas turbine engine as set forth in claim 1 , wherein the formula holds true for the majority of blade rows of the fan drive rotor.5. The gas turbine engine as set forth in claim 1 , wherein said rotational speed being an approach ...

Self Cleaning Debris Filter for Fan Drive Gear System

A pump system for a gas turbine engine has at least one pump. At least one valve has an outlet and at least one inlet is fluidly connected to the at least one pump. A geared architecture is positioned within a bearing compartment. The geared architecture is configured to receive lubricating fluid from the outlet of the at least one valve, a self-cleaning filter is positioned downstream of the at least one valve and upstream of the geared architecture. A gas turbine engine and a method are also disclosed. 1. A pump system for a gas turbine engine comprising:at least one pump;at least one valve having an outlet and at least one inlet fluidly connected to the at least one pump;a geared architecture positioned within a bearing compartment, wherein the geared architecture is configured to receive lubricating fluid from the outlet of the at least one valve; anda self-cleaning filter positioned downstream of the at least one valve and upstream of the geared architecture.2. The pump system according to wherein the self-cleaning filter comprises a filter screen positioned within a filter housing claim 1 , and wherein the filter housing has a filter inlet fluidly connected to the outlet of the at least one valve claim 1 , a first outlet fluidly connected to the geared architecture claim 1 , and a second outlet fluidly connected to the bearing compartment.3. The pump system according to wherein fluid flows into the filter inlet at a first flow rate and flows out of the first outlet at a flow rate that is less than the first flow rate.4. The pump system according to wherein fluid flows out of the second outlet at a third flow rate that is less than the second flow rate.5. The pump system according to wherein the second flow rate is approximately 95% of the first flow rate.6. The pump system according to wherein the filter screen is set to filter to a level comprising approximately 75 microns.7. The pump system according to wherein the at least one inlet of the at least one ...

CIRCULATING LUBRICANT THROUGH A TURBINE ENGINE COMPONENT WITH PARALLEL PUMPS

A system for a turbine engine includes a turbine engine component, a lubricant collection device and a plurality of lubricant circuits. The lubricant collection device is fluidly coupled with the turbine engine component. The lubricant circuits are fluidly coupled between the lubricant collection device and the turbine engine component. The lubricant circuits include a first circuit and a second circuit configured in parallel with the first circuit. Each of the lubricant circuits includes a lubricant pump. The first and the second circuits receive lubricant from the lubricant collection device, and direct the received lubricant to the turbine engine component. 1. A system for a turbine engine , comprising:a turbine engine component;a lubricant collection device fluidly coupled with the turbine engine component; anda plurality of lubricant circuits fluidly coupled between the lubricant collection device and the turbine engine component, and including a first circuit and a second circuit configured in parallel with the first circuit;each of the lubricant circuits including a lubricant pump;wherein the first and the second circuits receive lubricant from the lubricant collection device and direct the received lubricant to the turbine engine component.2. The system of claim 1 , wherein the first and the second circuits extend to the turbine engine component.3. The system of claim 2 , whereinthe turbine engine component includes a bearing and a manifold; andthe manifold respectively receives the lubricant from the first and the second circuits, and directs the received lubricant to the bearing.4. The system of claim 1 , whereinthe first circuit directs the received lubricant to the turbine engine component at a first flow rate; andthe second circuit directs the received lubricant to the turbine engine component at a second flow rate that is different than the first flow rate.5. The system of claim 1 , wherein one of the lubricant circuits further includes a heat ...

SYSTEM AND METHOD FOR REDUCING EDGE CONTACT STRESS CONCENTRATIONS IN A PRESS-FIT

An apparatus and method for reducing edge contact stress concentrations in a press-fit. The apparatus and method of the present disclosure specifically provide for a press-fit collar having a channel circumscribing a collar axis. The channel having an asymmetrical cross-sectional profile in a radial face. The asymmetrical cross-sectional profile being configured to reduce an edge contact pressure. 1. An epicyclic gearing for a gas turbine aviation engine , the epicyclic gearing comprising:a plurality of planet gears circumferentially disposed about a transmission axis and operably coupled to a plurality of planet pins;{'claim-text': ['a side plate comprising a coupling portion for connecting the side plate to a rotating member or to a static structure, and', 'a central ring coaxial to the side plate along the transmission axis, each planet pin of the plurality of planet pins being coupled to the central ring via a press-fit collar; and'], '#text': 'a planet-carrier, the planet-carrier comprising:'}{'claim-text': ['a first radial contact face defining a collar outer diameter, the first radial contact face being configured to interface with a pin opening defined by the central ring,', 'a second radial contact face disposed radially inward of the first radial contact face and defining a collar inner diameter centered about a collar axis, the second radial contact face being configured to accept one of the plurality of planet pins, and', 'an axial face extending between the first radial contact face and the second radial contact face, the axial face facing the plurality of planet gears, the axial face defining a channel having an asymmetrical cross-sectional profile, the channel circumscribing the collar axis, wherein the asymmetrical cross-sectional profile is configured to reduce an edge contact pressure.'], '#text': 'the press-fit collar defining a collar axis and comprising an annular body having:'}2. The epicyclic gearing of claim 1 , wherein the axial face further ...

FLEXIBLE SUPPORT STRUCTURE FOR A GEARED ARCHITECTURE GAS TURBINE ENGINE

A gas turbine engine according to an example of the present disclosure includes, among other things, a fan shaft configured to drive a fan, a support configured to support at least a portion of the fan shaft, the support defining a support transverse stiffness and a support lateral stiffness, a gear system coupled to the fan shaft, and a flexible support configured to at least partially support the gear system. The flexible support defines a flexible support transverse stiffness with respect to the support transverse stiffness and a flexible support lateral stiffness with respect to the support lateral stiffness. The input defines an input transverse stiffness with respect to the support transverse stiffness and an input lateral stiffness with respect to the support lateral stiffness. 1. A gas turbine engine , comprising:a fan having fan blades;an outer housing surrounding the fan, and a bypass flow path within said outer housing;a fan shaft drivingly connected to said fan;a frame supporting said fan shaft and defining a frame lateral stiffness and a frame transverse stiffness;a gear system connected to said fan shaft and driven through an input defining an input lateral stiffness and an input transverse stiffness, said gear system includes a gear mesh defining a gear mesh lateral stiffness and a gear mesh transverse stiffness; anda gear system flex mount arrangement, wherein said flex mount arrangement accommodates misalignment of said fan shaft and said input during operation and includes a flexible support which supports said gear system relative to a static structure and defines a flexible support lateral stiffness and a flexible support transverse stiffness;wherein at least one of said flexible support lateral stiffness and said flexible support transverse stiffness is less than 8% of a respective one of said gear mesh lateral stiffness and said gear mesh transverse stiffness; andwherein at least one of said input lateral stiffness and said input transverse ...

GEAR ASSEMBLY FOR AERONAUTICAL ENGINE WITH LUBRICANT STORING POCKETS

A gear assembly for an aeronautical engine includes a first gear disposed at a centerline axis of the gear assembly, a second gear coupled to the first gear in adjacent radial arrangement to form a first mesh between the first gear and the second gear, a static portion coupled to the second gear in adjacent circumferential arrangement, the static portion defining a pocket, and a spraybar disposed within the static portion such that a supply opening of the spraybar is directed at the first mesh between the first gear and the second gear. The supply opening provides a flow of lubricant to the first mesh between the first gear and the second gear and at least a portion of the flow of lubricant is collected by the pocket. The flow of lubricant is continuously released from the pocket to the gear system. 1. A gear assembly , comprising:a first gear disposed at a centerline axis of the gear assembly;a second gear coupled to the first gear in adjacent radial arrangement to form a first mesh between the first gear and the second gear;a static portion coupled to the second gear in adjacent circumferential arrangement, the static portion defining a pocket; anda spraybar disposed within the static portion such that a supply opening of the spraybar is directed at the first mesh between the first gear and the second gear,wherein the supply opening is configured to provide a flow of lubricant to the first mesh between the first gear and the second gear and at least a portion of the flow of lubricant is collected by the pocket.2. The gear assembly of claim 1 , further comprising:a third gear coupled to the second gear in adjacent radial arrangement to form a second mesh between the second gear and the third gear,wherein the flow of lubricant is released from the pocket to the second mesh between the second gear and the third gear during an interruption condition of a primary lubrication system in which the flow of lubricant is interrupted from being provided to the supply opening, ...

AIR TURBINE STARTER

An air turbine starter for starting an engine, comprising a housing defining an inlet, an outlet, and a flow path extending between the inlet and the outlet for communicating a flow of gas there through. A turbine member is journaled within the housing and disposed within the flow path for rotatably extracting mechanical power from the flow of gas. A gear box includes a gear train coupled with the turbine member. A gear box housing at least partially defines an interior having an open face and a body at least partially closes the open face. 1. An air turbine starter , comprising:a housing;at least one turbine member journaled within the housing;a gear box at least partially defining an interior having an open face and where the interior at least partially houses a gear train that is drivingly coupled with the at least one turbine member;a retainer having a body at least partially closing the open face; andan output shaft operably coupled with the gear train and having an output end.2. The air turbine starter of wherein the retainer includes a plate that spans at least a portion of the open face.3. The air turbine starter of wherein the plate is non-planar.4. The air turbine starter of wherein the plate includes a peripheral portion mounted between the housing and the gear box.5. The air turbine starter of wherein the plate includes a central portion that extends into the housing.6. The air turbine starter of wherein the plate includes an aperture within the central portion.7. The air turbine starter of claim 6 , further comprising a drive shaft operably coupled between the at least one turbine member and the gear train and passing through the aperture.8. The air turbine starter of wherein the body further includes a peripheral portion mounted to at least one of the housing or the gear box.9. The air turbine starter of wherein the body further includes a peripheral portion mounted between the housing and the gear box.10. The air turbine starter of claim 1 , further ...

Air turbine starter

An air turbine starter for starting an engine, comprising a housing defining an inlet, an outlet, and a flow path extending between the inlet and the outlet for communicating a flow of gas there through. A turbine member is journaled within the housing and disposed within the flow path for rotatably extracting mechanical power from the flow of gas. A gear train is drivingly coupled with the turbine member, a drive shaft is operably coupled with the gear train, and an output shaft is selectively operably coupled to rotate with the engine.

AIR TURBINE STARTER

An air turbine starter for starting an engine, comprising a housing defining an inlet, an outlet, and a flow path extending between the inlet and the outlet for communicating a flow of gas there through. A turbine member is journaled within the housing and disposed within the flow path for rotatably extracting mechanical power from the flow of gas. 1. An air turbine starter for starting an engine , comprising:a housing defining an inlet, an outlet, and a flow path extending between the inlet and the outlet for communicating a flow of gas there through, the housing having a diameter that is less than 15.63 cm (6.15 inches); anda set of turbine members journaled within the housing and disposed within the flow path for rotatably extracting mechanical power from the flow of gas.2. The air turbine starter of wherein the set of turbine members comprises a first rotor and a second rotor defining a first stage and a second stage claim 1 , respectively.3. The air turbine starter of claim 2 , further comprising a first stator upstream of the first rotor and wherein the first stator includes 16 nozzles.4. The air turbine starter of claim 3 , further comprising a second stator upstream of the second rotor and wherein the second stator includes 25 nozzles.5. The air turbine starter of wherein the outlet comprises a plurality of apertures located downstream of the set of turbine members.6. The air turbine starter of wherein the plurality of apertures are circumferentially spaced about the housing.7. The air turbine starter of wherein the diameter of the housing is 14.6 cm (5.75 inches).8. The air turbine starter of claim 1 , further comprising an inlet assembly mounted to a first end of the housing and a gear box mounted to a second end of the housing claim 1 , the gear box at least partially housing a gear train drivingly coupled with the set of turbine members.9. The air turbine starter of claim 8 , further comprising a drive shaft operably coupled with the gear train and ...

GAS TURBINE STARTING DEVICE AND GAS TURBINE SYSTEM

Provided is a gas turbine starting device which has a sun gear, a planetary carrier, an internal gear, and a planetary gear, the sun gear, the planetary carrier and the internal gear serving as rotating bodies rotating about an axis; and a variable speed power source which has a rotor connected to one of the rotating bodies in the planetary gear mechanism. A rotating shaft of the gas turbine is connected to one of the rotating bodies other than the rotating body to which the variable speed power source is connected, and the rotating shaft of the compressor is connected to the remaining one of the rotating bodies other than the rotating bodies to which the rotor of the variable speed power source and the rotating shaft of the gas turbine are connected. 1. A gas turbine starting device comprising:a planetary gear mechanism which is provided in a gas turbine system having a gas turbine, and a fluid machine rotationally driven by a driving force of the gas turbine on the same axis as a rotation center of a rotating shaft of the gas turbine, and has a sun gear, a planetary carrier disposed on an outer circumferential side of the sun gear, an internal gear having a gear unit and disposed on an outer circumferential side of the planetary carrier, and a planetary gear which is supported by the planetary carrier to mesh with the sun gear and the internal gear, the sun gear, the planetary carrier and the internal gear serving as rotating bodies rotating about the axis; anda variable speed power source which has a driving shaft connected to rotate one of the rotating bodies, which is the sun gear, the planetary carrier, or the internal gear, in the planetary gear mechanism about the axis,wherein the rotating shaft of the gas turbine is connected to rotate one of the rotating bodies, which is the sun gear, the planetary carrier, or the internal gear, other than the rotating body, to which the variable speed power source is connected, about the axis, anda rotating shaft of the ...

TURBINE SYSTEM WITH THREE TURBINES COUPLED TO A CENTRAL GEARBOX AND METHOD FOR OPERATING A WORK MACHINE

A turbine system and a method for operating a work machine are provided herein. The turbine system including: a first, a second, and a third turbine, a central gearbox mechanically coupled on the input side to the three turbines and having a mechanical connection on the output side for connecting a work machine, a first fluid line for conveying a working fluid from the first turbine to the second turbine, a second fluid line for conveying the working fluid from the second turbine to the third turbine, a first connecting unit designed such that a first partial mass flow of the working fluid can be removed from the first fluid line or supplied to the first fluid line, and a second connecting unit designed such that a second partial mass flow of the working fluid can be removed from the second fluid line or supplied to the second fluid line.

GEARBOX EFFICIENCY RATING FOR TURBOMACHINE ENGINES

A turbomachine engine can include a fan assembly, a vane assembly, a core engine, a gearbox, and a gearbox efficiency rating. The fan assembly can include a plurality of fan blades. The vane assembly can include a plurality of vanes, and the vanes can, in some instances, be disposed aft of the fan blades. The core engine can include one or more compressor sections and one or more turbine sections. The gearbox includes an input and an output. The input is coupled to the one or more turbine sections of the core engine and comprises a first rotational speed, the output is coupled to the fan assembly and has a second rotational speed, and a gear ratio of the first rotational speed to the second rotational speed is within a range of 4.1-14.0. The gearbox efficiency rating is 0.10-1.8. 2. The turbomachine engine of claim 1 , wherein the gearbox efficiency rating is 0.10-1.01.3. The turbomachine engine of claim 1 , wherein the gearbox efficiency rating is 0.19-1.8.4. The turbomachine engine of claim 1 , wherein the gear ratio is within a range of 4.5-12.0.5. The turbomachine engine of claim 1 , wherein the gear ratio is within a range of 6.0-11.0.6. The turbomachine engine of claim 1 , wherein Q is within a range of 5-55 gallons per minute.7. The turbomachine engine of claim 1 , wherein Q is within a range of 6-36 gallons per minute.8. The turbomachine engine of claim 1 , wherein D is 120-192 inches.9. The turbomachine engine of claim 1 , wherein T is within a range of 10 claim 1 ,000-100 claim 1 ,000 pounds force.10. The turbomachine engine of claim 1 , wherein T is within a range of 12 claim 1 ,000-30 claim 1 ,000 pounds force.12. The turbomachine engine of claim 11 , wherein the gearbox efficiency rating is 0.10-1.01.13. The turbomachine engine of claim 11 , wherein the gearbox efficiency rating is 0.19-1.8.14. The turbomachine engine of claim 11 , wherein the gear ratio is within a range of 4.5-12.0.15. The turbomachine engine of claim 11 , wherein the gear ratio is ...

WASTE HEAT UTILIZATION APPARATUS

A waste heat utilization apparatus is provided with a Rankine cycle and a power transmission mechanism that transmits power regenerated by an expander to an engine. The power transmission mechanism includes an expander clutch that interrupts or permits The transmission of the power from to expander to the engine. The expander includes a rotational speed sensor that detects a rotational speed of the expander. An increase in friction of the expander is detected on the basis of an increase in the rotational speed of the expander detected by the rotational speed sensor when the expander clutch is disconnected. 1. A waste heat utilization apparatus comprising:a Rankine cycle including a heat exchanger that recovers waste heat from an engine with a refrigerant, an expander that produces power by using the refrigerant that exits the heat exchanger, a condenser that condenses the refrigerant that exits the expander, and a refrigerant pump that supplies the refrigerant that exits the condenser to the heat exchanger; anda power transmission mechanism that transmits the power regenerated by the expander to the engine, wherein:the power transmission mechanism includes a disconnection/connection means that interrupts or permits the transmission of the power from the expander to the engine; andthe expander includes a rotational speed detection means that detects a rotational speed of the expander,the waste heat utilization apparatus further comprising a friction increase detection means that detects an increase in friction of the expander on the basis of an increase in the rotational speed of the expander detected by the rotational speed defection means when the disconnection/connection means is disconnected.2. The waste heat utilization apparatus according to claim 1 , wherein the Rankine cycle includes a bypass passage that bypasses the refrigerant introduced into the expander claim 1 , and a bypass valve that interrupts or permits conduction of the refrigerant into the bypass ...

GAS TURBINE ENGINE WITH LOW STAGE COUNT LOW PRESSURE TURBINE

A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a spool along an engine axis which drives a gear train, the spool including a low stage count low pressure turbine. 2. The engine as recited in claim 1 , wherein said first turbine section includes three (3) or four (4) stages.3. The engine as recited in claim 1 , wherein said first turbine section includes five (5) stages.4. The engine as recited in claim 1 , wherein said first turbine section includes six (6) stages.5. The engine as recited in claim 1 , further includinga fan variable area nozzle configured to vary a fan nozzle exit area and adjust a pressure ratio of the fan bypass airflow during engine operation.6. The engine as recited in claim 5 , further comprising:a controller operable to control the fan variable area nozzle to vary the fan nozzle exit area and adjust the pressure ratio of the fan bypass airflow.7. The engine as recited in claim 1 , wherein said gear train defines a gear reduction ratio of greater than or equal to about 2.5.8. The engine as recited in claim 7 , wherein said first turbine section defines a pressure ratio that is greater than about five (5.0).9. The engine as recited in claim 1 , wherein a fan pressure ratio across the fan blades is less than about 1.45.10. The engine as recited in claim 9 , wherein the bypass ratio is between about six and about ten (6.0-10.0).11. The engine as recited in claim 1 , wherein the first turbine section is configured to drive the fan through the gear train to provide a low corrected fan tip speed of less than about 1150 feet/second.12. The engine as recited in claim 1 , wherein the bypass ratio is between about six and about ten (6.0-10.0).13. The engine as recited in claim 1 , wherein said first turbine section is one of three turbine rotors claim 1 , while said second turbine section and another one of said turbine rotors each drives a compressor section.14. The engine as recited in ...

MECHANICALLY DRIVEN AIR VEHICLE THERMAL MANAGEMENT DEVICE

The present disclosure is directed to an aircraft power generation system including a reverse Brayton cycle system, a gas turbine engine, and a gearbox. The gas turbine engine includes a compressor section, a turbine section, and an engine shaft. The compressor section is arranged in serial flow arrangement with the turbine section. The engine shaft is rotatable with at least a portion of the compressor section and with at least a portion of the turbine section. The reverse Brayton cycle system includes a compressor, a driveshaft, a turbine, and a first exchanger. The driveshaft is rotatable with the compressor or the turbine, and the compressor, the first heat exchanger, and the turbine are in serial flow arrangement. The gearbox is configured to receive mechanical energy from the engine shaft and transmit mechanical energy to the reverse Brayton cycle system through the driveshaft. 1. An aircraft power generation system , comprising:a gas turbine engine including a compressor section, a turbine section, and an engine shaft, the compressor section arranged in serial flow arrangement with the turbine section, and the engine shaft rotatable with at least a portion of the compressor section and with at least a portion of the turbine section;a reverse Brayton cycle system, including a compressor, a driveshaft, a turbine, and a first exchanger, the driveshaft rotatable with the compressor or the turbine, and the compressor, the first heat exchanger, and the turbine in serial flow arrangement; anda gearbox, wherein the gearbox is configured to receive mechanical energy from the engine shaft and transmit mechanical energy to the reverse Brayton cycle system through the driveshaft.2. The system in claim 1 , further comprising:a thermal management system; anda working fluid, wherein the working fluid is in the reverse Brayton cycle system, and wherein the working fluid is in fluid communication with the thermal management system.3. The system in claim 2 , wherein the ...

A VANE ASSEMBLY OF A GAS TURBINE ENGINE

A first stage vane array of a high pressure turbine that may be for a geared turbofan engine includes a plurality of airfoils circumferentially spaced from one-another and orientated about an engine axis. Each airfoil has a leading edge and a trailing edge with the trailing edge being circumferentially separated by the next adjacent trailing edge by a pitch distance. The leading a trailing edges of each one of the plurality of airfoils are axially separated by an axial chord length. A pitch-to-chord ratio of the pitch distance over the axial chord length is equal to or greater than 1.7. 1. A first stage vane assembly of a high pressure turbine of a gas turbine engine comprising:a first airfoil configured to be circumferentially spaced from an adjacent second airfoil and orientated about an engine axis, wherein the first airfoil has a leading edge and a trailing edge with the trailing edge configured to be circumferentially separated from the trailing edge of the second airfoil by a pitch distance, and the leading and trailing edges of the first airfoil are axially separated by an axial chord length, and wherein a pitch-to-chord ratio of pitch distance over axial chord length is equal to or greater than 1.7.2. The first stage vane assembly set forth in claim 1 , wherein the first airfoil has a thickness-to-axial chord ratio that is greater than forty percent.3. The first stage vane assembly set forth in claim 2 , wherein the thickness-to-axial chord ratio is about fifty-three percent.4. The first stage vane assembly set forth in claim 1 , wherein the trailing edge has an angle that is greater than seventy-five degrees.5. The first stage vane assembly set forth in claim 1 , wherein the pitch-to-chord ratio is within a range of about 1.7 to 2.0.6. The first stage vane assembly set forth in claim 4 , wherein the pitch-to-chord ratio is within a range of about 1.7 to 2.0.7. The first stage vane assembly set forth in claim 1 , wherein the pitch-to-chord ratio is about 1.8 ...

DECOUPLER ASSEMBLY FOR ENGINE STARTER

An air turbine starter for starting an engine, comprising a housing defining an inlet, an outlet, and a flow path extending between the inlet and the outlet for communicating a flow of gas there through. A turbine member is journaled within the housing and disposed within the flow path for rotatably extracting mechanical power from the flow of gas. A gear train is drivingly coupled with the turbine member, a drive train is operably coupled with the gear train, and an output shaft is selectively operably coupled to rotate with the engine via a decoupler. 1. An air turbine starter for starting an engine , comprising:a housing defining an inlet, an outlet, and a flow path extending between the inlet and the outlet for communicating a flow of gas there through;a turbine member journaled within the housing and disposed within the flow path for rotatably extracting mechanical power from the flow of gas;a gear train drivingly coupled with the turbine member;a drive train operably coupled with the gear train and configured to provide a rotational output; and a drive hub operably coupled to the drive train and having a first set of teeth;', 'an output shaft having a first end having a second set of teeth configured to mate with the first set of teeth and a second end selectively operably coupled to the engine, the second set of teeth allow for driving torque transfer from the drive hub to the output shaft and the second set of teeth slide on the first set of teeth when back driving torque is transmitted such that the output shaft is moved axially away from the drive hub;', 'a shear pin operably coupled at a first end to the drive hub and operably coupled at a second end to the output shaft and having a shear fuse; and', 'a sheath surrounding at least a portion of the shear pin and axially moveable along a portion of the shear pin;', 'a load path for torque transmission through the drive hub and output shaft occurs during normal operation; and, 'a decoupler assembly, ...

COOPERATIVE SHAPE MEMORY ALLOY TORQUE TUBES FOR CONTINUOUS-ACTION TURNING MOTOR

An engine turning “clock work” motor including two shape memory alloy (SMA) torque tube actuators, ratcheting mechanisms, and gearing. The gearing communicates the SMA torque tube actuators with a common gear that applies torque to a shaft, so that while one torque tube is heated and applying torque, the other torque tube is relaxed (using a cooling mechanism). The ratchet prevents the relaxing torque tube from applying torque in the incorrect direction. 1200. A motor () , comprising:{'b': 224', '202', '202', '204', '202', '202', '208', '202', '202', '216', '218, 'i': a', 'b', 'a', 'b', 'c', 'a', 'b, 'cooperatively connected () torque tubes (), () each comprising a shape memory alloy (), wherein the cooperatively connected torque tubes (), () generate a continuous torque output () when the torque tubes (), () sequentially change shape () in response to heat ().'}2204. The motor of claim 1 , wherein the shape memory alloy () comprises Nickel and Titanium.3204. The motor of claim 1 , wherein the shape memory alloy () comprises an alloy including at least two metals selected from nickel claim 1 , titanium claim 1 , zinc claim 1 , copper claim 1 , gold claim 1 , palladium claim 1 , platinum claim 1 , and iron.4200204. The motor () of claim 1 , wherein the SMA () consists essentially of NiTiHf having an Hf content in a range of 10%-30%.5. The motor of claim 1 , wherein the torque tubes each have a length (L) in a range of 5-50 inches and a diameter (OD) claim 1 , (ID) in a range of 0.1-2 inches.6204202202202202208babac. The motor of claim 1 , wherein the shape memory alloy () comprises a composition and the torque tubes () claim 1 , () have dimensions (L) claim 1 , (OD) claim 1 , (ID) claim 1 , such that the torque tubes () claim 1 , () output at least 200 inch pounds of torque () in response to the SMA changing temperature by 300 degrees Celsius.7. The motor of claim 1 , wherein:{'b': 224', '202', '202', '202', '202', '206, 'i': a', 'b', 'a,', 'b, 'the cooperatively ...

GAS TURBINE ENGINE WITH MOUNT FOR LOW PRESSURE TURBINE SECTION

A turbine section of a gas turbine engine according to an example of the present disclosure includes, among other things, a first turbine section and a second turbine section. The first turbine section has a first exit area and rotates at a first speed. The second turbine section has a second exit area and rotates at a second speed. A first performance quantity is defined as the product of the first speed squared and the first exit area. A second performance quantity is defined as the product of the second speed squared and the second exit area. 1. A turbine section of a gas turbine engine comprising:a first turbine section; anda second turbine section,wherein said first turbine section has a first exit area and rotates at a first speed,wherein said second turbine section has a second exit area and rotates at a second speed, which is faster than the first speed,wherein a first performance quantity is defined as the product of the first speed squared and the first exit area,wherein a second performance quantity is defined as the product of the second speed squared and the second exit area,wherein a ratio of the first performance quantity to the second performance quantity is between about 0.5 and about 1.5;wherein said first turbine section is supported on two bearings, with a first bearing mounted in a mid-turbine frame that is positioned intermediate said first turbine section and said second turbine section, and a second bearing mounting said first turbine section, with said second bearing having a support extending downstream of said first turbine section; andwherein said second turbine has an inlet and is supported on a third bearing aft of said inlet.2. The turbine section as set forth in claim 1 , wherein said ratio is above or equal to about 0.8.3. The turbine section as set forth in claim 1 , wherein said third bearing is situated in said mid-turbine frame.4. The turbine section as set forth in claim 1 , wherein said first and third bearings are situated ...

GAS TURBINE ENGINE WITH HIGH SPEED LOW PRESSURE TURBINE SECTION AND BEARING SUPPORT FEATURES

A turbine section of a gas turbine engine according to an example of the present disclosure includes, among other things, a fan drive turbine section and a second turbine section. The fan drive turbine section has a first exit area at a first exit point and is configured to rotate at a first speed. The second turbine section has a second exit area at a second exit point and is configured to rotate at a second speed, which is faster than the first speed. 1. A turbine section of a gas turbine engine , comprising:a fan drive turbine section; anda second turbine section;wherein said fan drive turbine section has a first exit area at a first exit point and is configured to rotate at a first speed,wherein said second turbine section has a second exit area at a second exit point and is configured to rotate at a second speed, which is faster than the first speed, wherein a first performance quantity is defined as the product of the first speed squared and the first area,wherein a second performance quantity is defined as the product of the second speed squared and the second area,wherein a ratio of the first performance quantity to the second performance quantity is between about 0.5 and about 1.5, andwherein said second turbine section is configured to drive a first shaft supported on a bearing, with said bearing being situated between said first exit area and said second exit area.2. The turbine section as set forth in claim 1 , wherein said ratio is above or equal to about 0.8.3. The turbine section as set forth in claim 1 , wherein:said fan drive turbine section has between three and six stages;said second turbine section has two or fewer stages; anda pressure ratio across said fan drive turbine section is greater than about 5:1.4. The turbine section as set forth in claim 1 , wherein said fan drive and second turbine sections are configured to rotate in opposed directions.5. The turbine section as set forth in claim 1 , wherein each of said fan drive turbine section ...

FLEXIBLE SUPPORT STRUCTURE FOR A GEARED ARCHITECTURE GAS TURBINE ENGINE

A gas turbine engine according to an example of the present disclosure includes, among other things, a fan shaft configured to drive a fan, a support configured to support at least a portion of the fan shaft, the support defining a support transverse stiffness and a support lateral stiffness, a gear system coupled to the fan shaft, and a flexible support configured to at least partially support the gear system. The flexible support defines a flexible support transverse stiffness with respect to the support transverse stiffness and a flexible support lateral stiffness with respect to the support lateral stiffness. The input defines an input transverse stiffness with respect to the support transverse stiffness and an input lateral stiffness with respect to the support lateral stiffness. 1. A gas turbine engine , comprising:a fan shaft configured to drive a fan;a support configured to support at least a portion of the fan shaft, the support defining a support transverse stiffness and a support lateral stiffness;a gear system coupled to the fan shaft;a flexible support configured to at least partially support the gear system, the flexible support defining a flexible support transverse stiffness with respect to the support transverse stiffness and a flexible support lateral stiffness with respect to the support lateral stiffness; andan input to the gear system, the input defining an input transverse stiffness with respect to the support transverse stiffness and an input lateral stiffness with respect to the support lateral stiffness.2. The gas turbine engine as recited in claim 1 , wherein the support and the flexible support are mounted to a static structure.3. The gas turbine engine as recited in claim 2 , wherein the static structure is a front center body of the gas turbine engine.4. The gas turbine engine as recited in claim 1 , wherein the flexible support is mounted to a planet carrier of the gear system claim 1 , and the input is mounted to a sun gear of the gear ...

PROPULSION ENGINE FOR AN AIRCRAFT

A propulsion system for an aircraft includes an electric propulsion engine. The electric propulsion engine includes an electric motor and a fan rotatable about a central axis of the electric propulsion engine by the electric motor. The electric propulsion engine also includes a bearing supporting rotation of the fan and a thermal management system. The thermal management system includes a lubrication oil circulation assembly and a heat exchanger thermally connected to the lubrication oil circulation assembly. The lubrication oil circulation assembly is configured for providing the bearing with lubrication oil. Such an electric propulsion engine may be a relatively self-sufficient engine. 1. A propulsion system for an aircraft having an aft end , the propulsion system comprising: an electric motor;', 'a fan rotatable about the central axis of the electric propulsion engine by the electric motor;', 'a bearing supporting rotation of the fan; and', a lubrication oil circulation assembly for providing the bearing with lubrication oil; and', 'a heat exchanger thermally connected to the lubrication oil circulation assembly., 'a thermal management system comprising'}], 'an electric propulsion engine defining a central axis, the electric propulsion engine comprising'}2. The propulsion system of claim 1 , wherein the lubrication oil circulation assembly comprises a lubrication oil supply pump and a lubrication oil scavenge pump.3. The propulsion system of claim 1 , further comprising:a sump enclosing the bearing, wherein the lubrication oil circulation assembly is fluidly connected to the sump.4. The propulsion system of claim 1 , further comprising:an accessory gear box dedicated to the electric propulsion engine.5. The propulsion system of claim 4 , wherein the accessory gear box is driven by the electric motor.6. The propulsion system of claim 4 , wherein the lubrication oil circulation assembly comprises a lubrication oil supply pump and a lubrication oil scavenge pump ...

MULTI-ENGINE SYSTEM AND METHOD

A method of operating a multi-engine system of a rotorcraft includes, during a cruise flight segment of the rotorcraft, controlling a first engine to provide sufficient power and/or rotor speed demands of the cruise flight segment; and controlling a second engine to by providing a fuel flow to the second engine that is between 70% and 99.5% less than a fuel flow provided to the first engine. A turboshaft engine for a multi-engine system configured to drive a common load is also described. 1. A method of operating a multi-engine system of a rotorcraft , comprising:during a cruise flight segment of the rotorcraft, controlling a first engine to provide sufficient power and/or rotor speed demands of the cruise flight segment; andcontrolling a second engine to provide a fuel flow to the second engine that is between 70% and 99.5% less than a fuel flow provided to the first engine.2. The method of claim 1 , wherein the fuel flow to the second engine is between 70% to 90% less than a fuel flow provided to the first engine.3. The method of claim 2 , wherein the fuel flow to the second engine is between 80% to 90% less than a fuel flow provided to the first engine.4. The method of claim 1 , wherein the step of controlling the second engine includes using the fuel flow rate to the second engine as a control input variable to a controller of the multi-engine system.5. The method of claim 4 , further comprising a step of decoupling the second engine from the gearbox.6. The method of claim 1 , wherein the step of controlling the first engine is performed by using the power or rotor speed demand as a control input variable to the first engine and includes driving a rotor of the multi-engine rotorcraft via a common gearbox claim 1 , and the step of controlling the second engine includes controlling the fuel flow rate to the second engine so that a power output of the second engine to the common gearbox remains between 0% to 1% of a rated full-power of the second engine.7. The ...

Mechanical Gearbox for an Aircraft Turbomachine

A mechanical gearbox for aircraft 1. A mechanical gearbox for a turbomachine , comprising:a sun gear having an axis (X) of rotation and comprising an external toothing,a ring gear extending around the sun gear and comprising an internal toothing,planet gears meshed with the sun gear and the ring gear, each planet gear comprising a first toothing with a first average diameter meshed with the toothing of the sun gear, and a second toothing with a second average diameter different from the first diameter of the first toothing, the second toothing with a second diameter being meshed with the internal toothing of the ring gear, the planet gears being guided by hydrodynamic bearings carried by a planet carrier,wherein the hydrodynamic bearing for guiding each planet gear comprises a first smooth guiding surface extending about an axis (Y) of rotation of the planet gear, at least partly under the first toothing, and a second smooth guiding surface, different from said first surface and extending about the axis (Y) of rotation of the planet gear, at least partly under the second toothing, and in that the first surface is located on a first axial portion of a body of the hydrodynamic bearing, and the second surface is located on a second axial portion of the body of the hydrodynamic bearing, the first and second axial portions being connected together by a first annular web of the body of the hydrodynamic bearing.2. The gearbox according to claim 1 , wherein the first surface has a third diameter or average third diameter smaller than the first diameter of the first toothing claim 1 , and the second surface has a fourth diameter or average fourth diameter smaller than the second diameter of the second toothing.3. The gearbox according to claim 1 , wherein the first diameter of the first toothing is greater than the second diameter of the second toothing.4. The gearbox according to claim 1 , wherein the first annular web comprises a cylindrical rim axially supporting the ...

GEARED TURBOFAN WITH THREE TURBINES WITH HIGH SPEED FAN DRIVE TURBINE

A gas turbine engine according to an example of the present disclosure includes a fan section includes, among other things, a fan, a first compressor section, and a second compressor section configured to compress air to a higher pressure than the first compressor section. A first turbine section is configured to drive the second compressor section. A second turbine section is configured to drive the first compressor section. A fan drive turbine section is positioned downstream of the second turbine section. The fan drive turbine section is configured to drive the fan section through a gear reduction. The fan drive turbine section has a first exit area at a first exit point and is configured to rotate at a first speed. The second turbine section has a second exit area at a second exit point and is configured to rotate at a second speed, which is faster than the first speed. A first performance quantity is defined as the product of the first speed squared and the first area. A second performance quantity is defined as the product of the second speed squared and the second area. 1. A gas turbine engine comprising:a fan section including a fan, a first compressor section, and a second compressor section configured to compress air to a higher pressure than said first compressor section;a first turbine section configured to drive said second compressor section;a second turbine section configured to drive said first compressor section;a fan drive turbine section positioned downstream of said second turbine section, said fan drive turbine section configured to drive said fan section through a gear reduction;wherein said first compressor section and said second turbine section are configured to rotate as an intermediate speed spool, and said second compressor section and said first turbine section are configured to rotate together as a high speed spool, with said high speed spool, said intermediate speed spool, and said fan drive turbine section each configured to rotate in ...

GAS TURBINE ENGINE WITH HIGH SPEED LOW PRESSURE TURBINE SECTION

A gas turbine engine according to an example of the present disclosure includes, among other things, a fan having one or more fan blades. A compressor section is in fluid communication with the fan. The compressor section includes a first compressor section and a second compressor section. A turbine section is in fluid communication with the compressor section. The turbine section includes a first turbine section and a second turbine section. The first turbine section has a first exit area at a first exit point and rotates at a first speed. The second turbine section has a second exit area at a second exit point and rotates at a second speed. A first performance quantity is defined as the product of the first speed squared and the first area. A second performance quantity is defined as the product of the second speed squared and the second area. 1. A gas turbine engine comprising:a fan having one or more fan blades, the fan defining a pressure ratio less than about 1.45;a compressor section in fluid communication with the fan, the compressor section including a first compressor section and a second compressor section;a turbine section in fluid communication with the compressor section;wherein the turbine section includes a first turbine section and a second turbine section, the first turbine section and the first compressor section are configured to rotate in a first direction, and wherein the second turbine section and the second compressor section are configured to rotate in a second direction, opposed to said first direction;wherein a pressure ratio across the first turbine section is greater than about 5:1;wherein said first turbine section has a first exit area at a first exit point and rotates at a first speed;wherein said second turbine section has a second exit area at a second exit point and rotates at a second speed, which is faster than the first speed;wherein a first performance quantity is defined as the product of the first speed squared and the first ...

LOW PRESSURE RATIO FAN ENGINE HAVING A DIMENSIONAL RELATIONSHIP BETWEEN INLET AND FAN SIZE

According to an example embodiment, a gas turbine engine assembly includes, among other things, a fan that has a plurality of fan blades. A diameter of the fan has a dimension D that is based on a dimension of the fan blades. Each fan blade has a leading edge. An inlet portion is situated forward of the fan. A length of the inlet portion has a dimension L between a location of the leading edge of at least some of the fan blades and a forward edge on the inlet portion. A dimensional relationship of L/D is between about 0.2 and 0.45. 1. A gas turbine engine assembly , comprising:a fan including a plurality of fan blades, a diameter of the fan having a dimension D that is based on a dimension of the fan blades, each fan blade having a leading edge, the fan being, configured to deliver a portion of air into a compressor section and a portion of air into a bypass duct, wherein a bypass ratio is greater than about 6, the bypass ratio being a ratio of a volume of air passing to the bypass duct compared to a volume of air passing into the compressor section; andan inlet portion forward of the fan, a length of the inlet portion having a dimension L between a location of the leading edge of at least some of the fan blades and a forward edge on the inlet portion,wherein a dimensional relationship of L/D is between about 0.2 and 0.45.2. The assembly of claim 1 , wherein the dimensional relationship of L/D is between about 0.25 and 0.45.3. The assembly of claim 2 , wherein the dimensional relationship of L/D is between about 0.30 and about 0.40.4. The assembly of claim 2 , wherein the dimensional relationship of L/D is about 0.35.5. The assembly of claim 1 , whereinthe dimension L is different at a plurality of locations on the fan case;a greatest value of L corresponds to a value of L/D that is at most 0.45; anda smallest value of L corresponds to a value of L/D that is at least 0.20.6. The assembly of claim 1 , whereinthe dimension L varies; andthe dimensional relationship of ...

Method of increasing the safety of a power plant, and a power plant suitable for implementing the method

The present invention relates to a method of increasing the safety of a power plant provided with at least one heat engine and a gearbox (BTP), the engine driving the gearbox (BTP), the gearbox (BTP) having a lubrication system implemented using an aqueous medium stored in a reserve, in which method a fluid comprising water is injected into the heat engine to increase the power developed by the heat engine without increasing the temperature of a member of the heat engine or to decrease the temperature without modifying the power developed by the engine, the fluid being taken from the reserve.

LUBRICATING APPARATUS FOR TURBO COMPOUND SYSTEM

Disclosed is a lubricating apparatus including an oil supply pipe connected to a gearbox of a gearing device for transmitting turning force of a blowdown turbine in the turbo compound system to a crankshaft to supply oil into the gearbox, a first oil supply opening formed to be transpierced at the bearing housing of a first bearing rotatably supporting a shaft of an output gear for outputting turning force in the turbo compound system to supply oil supplied from the gearbox to the first bearing, a sub-supply pipe connected to a gear case in which the output gear is embedded to supply oil to the gear case, and a second oil supply opening formed to be transpierced at the bearing housing of a second bearing rotatably supporting the shaft of the output gear to supply oil supplied through the sub-supply pipe to the second bearing. 1. A lubricating apparatus of a turbo compound system comprising:an oil supply pipe connected to a gearbox of a gearing device for transmitting turning force of a blowdown turbine in the turbo compound system to a crankshaft to supply oil into the gearbox;a first oil supply opening formed to be transpierced at a bearing housing of a first bearing rotatably supporting a shaft of an output gear for outputting turning force in the turbo compound system to supply oil supplied from the gearbox to the first bearing;a sub-supply pipe connected to a gear case in which the output gear is embedded to supply oil to the gear case; anda second oil supply opening formed to be transpierced at the bearing housing of a second bearing rotatably supporting the shaft of the output gear to supply oil supplied through the sub-supply pipe to the second bearing.2. The lubricating apparatus according to claim 1 , wherein the oil supply pipe is connected from a cylinder block of an engine to the gearbox in which reduction gears of the gearing device are embedded to supply oil from the engine.3. The lubricating apparatus according to claim 1 , wherein the sub-supply pipe ...

SYSTEM AND APPARATUS FOR DIVERSIFIED GEARBOX

A gas turbine engine assembly comprising, a gearbox including a first housing that includes a first auxiliary gear drive on a first portion thereof, a second housing that includes a second auxiliary gear drive on a second portion thereof, and a third housing that includes a third auxiliary gear drive on a third portion thereof, the housings being interconnected so that the first portion of the first housing, the second portion of the second housing and the third portion of the third housing form a substantially triangular polyhedron shape, with the second portion of the second housing disposed between the first portion of the first housing and the third portion of the third housing. The first auxiliary gear drive, the second auxiliary gear drive and the third auxiliary gear drive project outwardly in mutually divergent directions. 1. An intermediate housing portion of a gearbox comprising:a generally triangular polyhedron shape;a coupling with a first housing portion, wherein the first housing portion comprises a first housing portion first face, and wherein a first auxiliary gear drive is arranged within the first housing portion; anda coupling with a second housing portion, wherein the second housing portion comprises a second housing portion first face, and wherein a second auxiliary gear drive is arranged within the second housing portion,wherein the first housing portion first face and the second housing portion first face intersect along a common edge distal the intermediate housing portion.2. The intermediate housing portion of the gearbox of claim 1 , wherein at least one of the first housing portion first face or the second housing portion first face includes a removable cover.3. The intermediate housing portion of the gearbox of claim 1 , further including a third auxiliary gear drive arranged within the intermediate housing portion.4. The intermediate housing portion of the gearbox of claim 3 , wherein a first set of bevel gears interconnects the first ...

TURBINE INTER-SPOOL ENERGY TRANSFER SYSTEM

An inter-spool energy transfer system is provided and includes a first spool, a second spool, which includes components that are rotatable at a different speed as compared to components of the first spool and an inter-spool interface system coupled to at least one of the components of the first spool and at least one of the components of the second spool. The inter-spool interface system includes a controller, which is configured to supply power to one of the first and second spools and to draw power from the other of the first and second spools. 1. An inter-spool energy transfer system , comprising:a first spool;a second spool, which includes components that are rotatable at a different speed as compared to components of the first spool; andan inter-spool interface system coupled to at least one of the components of the first spool and at least one of the components of the second spool, the inter-spool interface system comprising:a controller, which is configured to supply power to one of the first and second spools and to draw power from the other of the first and second spools.2. The inter-spool energy transfer system according to claim 1 , further comprising a power connection between the controller and one or more of the aircraft electrical systems claim 1 , hydraulic systems claim 1 , compressed air systems claim 1 , and mechanical systems.3. The inter-spool energy transfer system according to claim 1 , wherein the first spool is a core spool comprising a gas generator claim 1 , a gas generator turbine and a compressor shaft and the second spool is a power spool comprising a power shaft and a power turbine.4. An inter-spool energy transfer system for use in a vehicle including an electrical system claim 1 , comprising:a first spool;a second spool, which includes components that are rotatable at a lesser speed as compared to components of the first spool; andan inter-spool interface system coupled to at least one of the components of the first spool and at ...

Gas turbine engine with low stage count low pressure turbine

A gas turbine engine includes, among other things, a fan section including a fan rotor, a gear train defined about an engine axis of rotation, a first nacelle which at least partially surrounds a second nacelle and the fan rotor, the fan section configured to communicate airflow into the first nacelle and the second nacelle, a first turbine, and a second turbine followed by the first turbine. The first turbine is configured to drive the fan rotor through the gear train. A static structure includes a first engine mount location and a second engine mount location.

WINDAGE SHIELD

A windage shield (″) for mounting on a fan disc (″) of a fan () of a gas turbine engine (), the windage shield (″) comprising: a fan disc contacting portion (″) adapted to contact and structurally support a rear portion of the fan disc (″); wherein the fan disc contacting portion (″) includes one or more stiffening elements (″) to locally increase the hoop stiffness of the windage shield (″). 1. A windage shield for mounting on a fan disc of a fan of a gas turbine engine , the windage shield comprising: a fan disc contacting portion adapted to contact and structurally support a rear portion of the fan disc; wherein the fan disc contacting portion includes one or more stiffening elements to locally increase the hoop stiffness of the windage shield.2. A windage shield according to claim 1 , wherein at least one of the one or more stiffening elements is integral to the fan disc contacting portion.3. A windage shield according to claim 1 , wherein at least one of the one or more stiffening elements is provided as an insert receivable partially or fully on or in the fan disc contacting portion.4. A windage shield according to claim 1 , wherein at least one of the one or more stiffening elements comprises claim 1 , or consists essentially of claim 1 , a composite material such as a metal matrix composite or a ceramic matrix composite.5. A windage shield according to claim 1 , wherein at least one of the one or more stiffening elements comprises claim 1 , or consists essentially of claim 1 , a metal or alloy such as nickel alloy claim 1 , titanium claim 1 , a titanium alloy claim 1 , aluminium or an aluminium alloy.6. A fan for a gas turbine engine claim 1 , the fan comprising:a fan disc; and{'claim-ref': {'@idref': 'CLM-00001', 'claim 1'}, 'at least one windage shield according to mounted on the fan disc.'}7. A fan according to claim 6 , wherein the windage shield is mounted on a rear portion of the fan disc.8. A fan according to claim 7 , the fan disc further comprising ...

ROLLER BEARINGS FOR HIGH RATIO GEARED TURBOFAN ENGINE

A gear system for a geared turbofan engine is disclosed. The gear system includes a sun gear driven by a low spool shaft. The sun gear defines a sun gear diameter. A rotating carrier drives a fan. The carrier defines an outer carrier diameter and an inner carrier diameter. A non-rotating ring gear is also included. The ring gear defines a ring gear diameter and the ring gear diameter is smaller than the outer carrier diameter. A set of planet gears are mounted on corresponding rolling element bearing assemblies. Each roller element bearing assembly is supported within the carrier within a space defined between the carrier outer diameter and the carrier inner diameter. Each of the sun gear, ring gear and planet gears are substantially centered along a gearbox centerline transverse to an engine longitudinal axis and the gear system provides a speed reduction ratio between an input to the sun gear and an output from the carrier between 3:1 and 5:1. A method of creating a gear system for a geared turbofan engine and a geared turbofan system are also disclosed. 1. A gear system for a geared turbofan engine , the gear system comprising:a sun gear driven by a low spool shaft, the sun gear defining a sun gear diameter;a rotating carrier that drives a fan, the carrier defining an outer carrier diameter and an inner carrier diameter;a non-rotating ring gear, the ring gear defining a ring gear diameter, the ring gear diameter is smaller than the outer carrier diameter; anda set of planet gears including 3, 4 or 5 planet gears; androlling element bearing assemblies supporting each of the set of planet gears, each roller element bearing assembly is supported within the carrier within a space defined between the carrier outer diameter and the carrier inner diameter, wherein a diameter of each of the bearing elements of the roller element bearing assembly is less than a sum of the carrier outer diameter and the carrier inner diameter multiplied by √ and divided by twice the number ...

Geared turbofan engine with power density range

A turbofan gas turbine engine includes a fan section having a fan, a compressor section including a low pressure compressor and a high pressure compressor, a geared architecture including an epicyclic gear train, a turbine section including a low pressure turbine and a high pressure turbine, the fan driven by the low pressure turbine through the geared architecture, and a power density between 4.84 lbf/in3 and 5.5 lbf/in3.

GAS TURBINE ENGINE FOR AN AIRCRAFT COMPRISING AN AIR INTAKE

A gas turbine engine for an aircraft includes an engine core, fan, air intake and gearbox. The engine core includes a turbine, compressor, and core shaft connecting the turbine and compressor. The fan is located upstream of the core and includes a plurality of fan blades, the fan having a diameter greater than 2.0 m. The air intake is located upstream of the fan and has ratio of intake length to fan diameter of 0.20 to 0.60 and defines highlight, throat and diffuser areas. The gearbox receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft. The gas turbine engine has a bypass ratio greater than 10; and the air intake has an intake length and the ratio of the intake length to the fan diameter is from 0.20 to 0.60. 1. A gas turbine engine for an aircraft , the gas turbine engine comprising:an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor;a fan located upstream of the engine core, the fan comprising a plurality of fan blades and having a fan diameter greater than 2.0 m;an air intake located upstream of the fan, the air intake having a ratio of intake length to fan diameter of from 0.20 to 0.60 and defining a highlight area, a throat area and a diffuser area; anda gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft;wherein the gas turbine engine has a bypass ratio greater than 10; and the air intake has an intake length and the ratio of the intake length to the fan diameter is from 0.20 to 0.60.2. The gas turbine engine of claim 1 , wherein the ratio of the intake length to the fan diameter is from 0.20 to 0.50.3. The gas turbine engine claim 1 , wherein the ratio of the intake length to the fan diameter is from 0.25 to 0.45.4. The gas turbine engine of claim 1 , wherein the ratio of the intake length to the fan diameter is from 0.30 ...

GAS TURBINE ENGINE AIRFOIL

An airfoil for a turbine engine includes an airfoil having pressure and suction sides extending in a radial direction from a 0% span position at an inner flow path location to a 100% span position at an airfoil tip. The airfoil has a curve corresponding to a relationship between a trailing edge sweep angle and a span position. The trailing edge sweep angle is in a range of 0° to 10° in a range of 10-20% span position. The trailing edge sweep angle is in a range of the trailing edge sweep angle is positive from 0% span to at least 95% span. 1. An airfoil for a turbine engine comprising:an airfoil having pressure and suction sides and extending in a radial direction from a 0% span position at an inner flow path location to a 100% span position at an airfoil tip, wherein the airfoil has a curve corresponding to a relationship between a trailing edge sweep angle and a span position, wherein the trailing edge sweep angle is in a range of 0° to 10° in a range of 10-20% span position, and the trailing edge sweep angle is positive from 0% span to at least 95% span.2. The airfoil according to claim 1 , wherein the trailing edge sweep angle is in a range of 10° to 20° in a range of 40-70% span position.3. The airfoil according to claim 2 , wherein the trailing edge sweep angle is about 15°.4. The airfoil according to claim 1 , wherein the trailing edge sweep angle is positive from 0%-95% span.5. The airfoil according to claim 4 , wherein the trailing edge sweep angle transitions from less positive to more positive at greater than an 80% span position.6. The airfoil according to claim 4 , wherein a positive-most trailing edge sweep angle is at a greater than 50% span position.7. The airfoil according to claim 1 , wherein a trailing edge sweep angle at the 0% span position and a trailing edge sweep angle at the 100% span position are about the same.8. The airfoil according to claim 7 , wherein a positive-most trailing edge sweep angle is at about a 70% span position.9. The ...

Turbine engine assembly and method of maufacturing thereof

A turbine engine assembly includes: a fan assembly; a turbine coupled to the fan assembly through a gearbox; a stationary component; and an assembly extending between the gearbox and the stationary component to couple the gearbox to the stationary component, wherein the assembly includes at least one vibration-reducing mechanism configured to isolate a vibratory response of the gearbox from the stationary component.

Gas turbine engine airfoil

An airfoil of a turbine engine according to an example of the present disclosure includes, among other things, pressure and suction sides extending in a radial direction from a 0% span position to a 100% span position. The airfoil has a relationship between a gap/chord ratio and span position that defines a curve with a gap/chord ratio having a portion with a negative slope.

OIL TRANSFER BEARING

A transfer bearing assembly includes a transfer bearing shaft. An oil transfer bearing surrounds the transfer bearing shaft. A radially extending tube is attached to the transfer bearing shaft and is configured to engage a carrier on a speed reduction device. 1. A transfer bearing assembly including:a transfer bearing shaft;an oil transfer bearing surrounding the transfer bearing shaft; anda radially extending tube attached to the transfer bearing shaft configured to engage a carrier on a speed reduction device.2. The assembly of claim 1 , wherein the oil transfer bearing includes a stationary outer bearing and a rotating inner bearing attached to the transfer bearing shaft.3. The assembly of claim 2 , wherein the stationary outer bearing includes at least one race aligned with at least one opening in the rotating inner bearing configured to transfer oil from the stationary bearing to the transfer bearing shaft through the rotating inner bearing.4. The assembly of claim 3 , wherein the transfer bearing shaft includes at least one axially extending passage in communication with one of the at least one opening in the rotating inner bearing.5. The assembly of claim 4 , wherein a radially inner end of the tube engages the transfer bearing shaft and a radially outer end of the tube engages a carrier.6. The assembly of claim 5 , wherein the tube is configured to move in a radial direction relative to the transfer bearing shaft.7. The assembly of claim 5 , wherein the transfer bearing shaft includes at least one radially extending opening in communication with the axially extending passage and the tube is located within the radially extending opening.8. A gas turbine engine section including:a planetary gear set including a rotating carrier; and a transfer bearing shaft;', 'an oil transfer bearing surrounding the transfer bearing shaft; and', 'a radially extending tube attached to the transfer bearing shaft configured to engage a carrier on a speed reduction device., 'a ...

LOW PRESSURE RATIO FAN ENGINE HAVING A DIMENSIONAL RELATIONSHIP BETWEEN INLET AND FAN SIZE

A gas turbine engine assembly includes a fan. A diameter of the fan has a dimension D. The fan has a pressure ratio of greater than 1.20 and less than 1.45. A leading edge on an inlet portion of a nacelle is within a first reference plane oriented at an oblique angle. A forward most portion on the fan blade leading edges is in a second reference plane. A length of the inlet portion has a dimension L different at a plurality of locations on the inlet portion. A geared architecture has a gear reduction ratio of greater than 2.3, a bypass ratio is greater than 10, and a low pressure turbine includes a pressure ratio greater than 5:1. A dimensional relationship of UD is between 0.25 and 0.45. The leading edge on the inlet portion is further from the second reference plane near the top of the assembly. 1. A gas turbine engine assembly comprising:an engine central longitudinal axis;a fan and a fan case surrounding the fan, the fan including a plurality of fan blades having circumferentially outermost edges, a diameter of the fan having a dimension D extending between the circumferentially outermost edges of the fan blades, each fan blade further having a leading edge, wherein a portion of the fan case is forward of the leading edges of the fan blades, and the fan has a pressure ratio of greater than 1.20 and less than 1.45 across the fan blade alone;a nacelle surrounding the fan, the nacelle including an inlet portion forward of the fan, and a leading edge on the inlet portion is within a first reference plane;a forward most portion on the leading edges of the fan blades is in a second reference plane perpendicular to the engine central longitudinal axis, and the first reference plane is oriented at an oblique angle relative to the second reference plane and to the engine central longitudinal axis, and a length of the inlet portion has a dimension L measured in a direction parallel to the engine central longitudinal axis between a location of the leading edges of the fan ...

PUMP GEAR

An integrated drive generator includes an input shaft and an input drive gear connected to the input shaft. The integrated drive generator also includes an input driven gear meshed with the input drive gear. The integrated drive generator also includes a hydraulic speed trimming device. The hydraulic speed trimming device includes an input shaft connected to the input driven gear, an output shaft, an accessory drive gear connected to the output shaft, and an output ring gear connected to the output shaft. The input driven gear, the accessory drive gear, and the output ring gear are coaxial and disposed proximate a first end of the hydraulic speed trimming device. The integrated drive generator also includes a pump assembly with a pump drive shaft and a pump gear connected to the pump drive shaft and meshed with the accessory drive gear. 1. An integrated drive generator for a gas turbine engine , the integrated drive generator comprising:an input shaft;an input drive gear connected to the input shaft;an input driven gear meshed with the input drive gear; an input shaft connected to the input driven gear;', 'an output shaft;', 'an accessory drive gear connected to the output shaft; and', 'an output ring gear connected to the output shaft,', 'wherein the input driven gear, the accessory drive gear, and the output ring gear are coaxial and disposed proximate a first end of the hydraulic speed trimming device;, 'a hydraulic speed trimming device, the hydraulic speed trimming device comprisinga pump assembly comprising a pump drive shaft; anda pump gear connected to the pump drive shaft and meshed with the accessory drive gear.2. The integrated drive generator of claim 1 , wherein the pump gear is a spur gear.3. The integrated drive generator of claim 2 , wherein the pump gear comprises:an outside diameter; anda pitch diameter,wherein a ratio of the outside diameter and the pitch diameter is about 1.027 to about 1.029.4. The integrated drive generator of claim 2 , wherein ...

AUXILLIARY POWER AND THRUST UNIT DRIVE SYSTEM

An aircraft auxiliary power and thrust unit includes at least one blade mounted to a fan shaft which are mounted to a tail cone of the aircraft and rotatable relative to the tail cone. Also included is an air intake assembly which includes an opening defined by the tail cone and a channel in fluid communication with the opening and the at least one blade. A first drive shaft positioned to extend in a direction transverse to the fan shaft and engageable to rotate with the fan shaft and first drive shaft engaged and the first drive shaft engageable to a second drive shaft positioned transverse to the first drive shaft such that the second drive shaft rotates with the first drive shaft with the first and second drive shafts engaged. The second drive shaft is positioned outside of the channel of the air intake. 110. An improved aircraft auxiliary power and thrust unit (′) for an aircraft , comprising:{'b': 56', '59', '61', '58', '12, 'at least one blade (, , ) mounted to a fan shaft () wherein the at least one blade and the fan shaft are mounted to a tail cone () of the aircraft wherein the at least one blade and the fan shaft are rotatable relative to the aircraft;'}{'b': 38', '40', '16, 'an air intake assembly () comprising an opening () defined by the tail cone and a channel () in fluid communication with the opening and the at least one blade;'}{'b': 60', '62, 'a first drive shaft () positioned to extend in a direction transverse to the fan shaft with a first end portion () of the first drive shaft engageable with the fan shaft such that the fan shaft and the first drive shaft rotate together with the first end portion of the first drive shaft and the shaft engaged; and'}{'b': 64', '66', '68, 'a second drive shaft () positioned to extend in a direction transverse to the first drive shaft with a first end portion () of the second drive shaft engageable with a second end portion () of the first drive shaft such that rotation of the first drive shaft imparts rotation ...

FAN DRIVE GEAR SYSTEM WITH IMPROVED MISALIGNMENT CAPABILITY

An epicyclic gear assembly according to an exemplary aspect of the present disclosure includes, 60 among other things, a carrier including a first plate axially spaced from a second plate by a radially outer connector. A first set of epicyclic gears supported adjacent the first plate include a first set of circumferentially offset intermediate gears meshing with a first sun gear and a first ring gear. A second set of epicyclic gears are axially spaced from the first set of epicyclic gears and supported adjacent the second plate, and include a second set of circumferentially offset intermediate gears meshing with a second sun gear and a second ring gear. The first epicyclic gear set and the second epicyclic gear set maintain relative intermeshing alignment during flexure induced deformation of the carrier. 1. An epicyclic gear assembly comprising:a carrier including a first plate axially spaced from a second plate by a radially outer connector;a first set of epicyclic gears supported adjacent the first plate, including a first set of circumferentially offset intermediate gears meshing with a first sun gear and a first ring gear; anda second set of epicyclic gears axially spaced from the first set of epicyclic gears and supported adjacent the second plate, and including a second set of circumferentially offset intermediate gears meshing with a second sun gear and a second ring gear, whereby the first epicyclic gear set and the second epicyclic gear set maintain relative intermeshing alignment during flexure induced deformation of the carrier.2. The assembly of claim 1 , including at least one intermediate gear opening in the carrier claim 1 , wherein one of the first set of intermediate gears and one of the second set of intermediate gears are located within the at least one intermediate gear opening.3. The assembly of claim 1 , wherein the first set of intermediate gears move independently of the second set of intermediate gears.4. The assembly of claim 1 , wherein ...

AIRCRAFT TURBINE ENGINE WITH PLANETARY OR EPICYCLIC GEAR TRAIN

Aircraft turbine engine comprising a low-pressure spool that comprises a low-pressure shaft (), means () for taking off power from said low-pressure shaft, and a fan () that is driven by said low-pressure shaft by means of a reduction gear (), said reduction gear comprising at least one first element () that is connected to said low-pressure shaft for conjoint rotation, at least one second element () that is connected to said fan for conjoint rotation, and at least one third element () that is connected to a stator casing of the turbine engine, characterised in that said at least one third element is connected to said stator casing by disengageable connection means (), and comprising at least one member that can move from a first position in which said at least one third element is fixedly connected to said stator casing into a second position in which said at least one third element is separated from said stator casing and is free to rotate about said longitudinal axis. 1. An aircraft turbine engine comprising a low-pressure spool that comprises a low-pressure shaft that connects a rotor of a low-pressure compressor to a rotor of a low-pressure turbine , and a high-pressure spool that comprises a high-pressure shaft that connects a rotor of a high-pressure compressor to a rotor of a high-pressure turbine , the low-pressure and high-pressure shafts extending along the same longitudinal axis (A) , the turbine engine further comprising means for taking off power from said low-pressure shaft , and a fan that is driven by said low-pressure shaft by means of a planetary or epicyclic reduction gear , said reduction gear comprising at least one first element that is connected to said low-pressure shaft for conjoint rotation , at least one second element that is connected to said fan for conjoint rotation , and at least one third element that is connected to a stator casing of the turbine engine , characterised in that said at least one third element is connected to said ...

FLEXIBLE SUPPORT STRUCTURE FOR A GEARED ARCHITECTURE GAS TURBINE ENGINE

A gas turbine engine includes a fan that has fan blades. A fan shaft is drivingly connected to the fan. A gear system is connected to the fan shaft and driven through an input. A gear system flex mount arrangement accommodates misalignment of the fan shaft and the input during operation. The gear system flex mount arrangement includes a gear mesh that defines a gear mesh lateral stiffness and a flexible support that defines a flexible support lateral stiffness that is less than 8% of the gear mesh lateral stiffness. 1. A gas turbine engine , comprising:a fan having fan blades;a fan shaft drivingly connected to said fan;a gear system connected to said fan shaft and driven through an input; anda gear system flex mount arrangement, wherein said gear system flex mount arrangement accommodates misalignment of said fan shaft and said input during operation and said gear system flex mount arrangement includes a gear mesh defining a gear mesh lateral stiffness and a flexible support defining a flexible support lateral stiffness that is less than 8% of said gear mesh lateral stiffness.2. The gas turbine engine of claim 1 , wherein said gear system flex mount arrangement further includes a ring gear defining a ring gear lateral stiffness that is less than 12% of said gear mesh lateral stiffness.3. The gas turbine engine of claim 2 , wherein said gear system has a gear reduction ratio of greater than 2.3 claim 2 , and further comprising a bypass ratio greater than ten (10) claim 2 , a fan pressure ratio of less than 1.45 measured across said fan blades alone claim 2 , and a low pressure turbine with an inlet claim 2 , an outlet claim 2 , and a low pressure turbine pressure ratio greater than 5:1 claim 2 , wherein said low pressure turbine pressure ratio is a ratio of a pressure measured prior to said inlet as related to a pressure at said outlet prior to any exhaust nozzle.4. The gas turbine engine of claim 3 , wherein said gear mesh defines a gear mesh transverse stiffness ...

Auxiliary Journal Oil Supply System

A fluid circulation system may comprise a main pump configured to provide oil to a journal bearing. An auxiliary system including a pump system may be configured to provide oil to the journal bearing. A manifold may be configured to mix the oil from the main pump with oil from the auxiliary system. A journal delivery line may be configured to deliver the oil from the manifold to the journal bearing. 1. A fluid circulation system , comprising:a main pump configured to provide oil to a journal bearing;an auxiliary system including a pump system configured to provide oil to the journal bearing;a manifold configured to mix the oil from the main pump with oil from the auxiliary system; anda journal delivery line configured to deliver the oil from the manifold to the journal bearing.2. The fluid circulation system of claim 1 , further comprising:a main journal supply line coupled between the main pump and the manifold; andan auxiliary journal supply line coupled between the auxiliary system and the manifold, the journal delivery line coupled to the main journal supply line and the auxiliary journal supply line.3. The fluid circulation system of claim 2 , further comprising a check valve disposed on the main journal supply line between the main pump and the auxiliary journal supply line.4. The fluid circulation system of claim 2 , further comprising a filter disposed on the auxiliary journal supply line between the auxiliary system and the manifold.5. The fluid circulation system of claim 2 , further comprising an auxiliary oil pressure sensor disposed on the auxiliary journal supply line between the pump system and the manifold.6. The fluid circulation system of claim 1 , wherein the pump system of the auxiliary system includes:a first pump configured to retrieve oil from a gutter of a fan drive gear system, anda second pump configured to retrieve oil from a bearing compartment.7. The fluid circulation system of claim 1 , wherein the auxiliary system is configured to pump ...

Turbomachine

A turbomachine defining a radial direction and an axial direction, the turbomachine including: a structural assembly comprising a frame and a static structure; a gearbox coupled to the structural assembly through the static structure; a turbine coupled to the gearbox and having a plurality of turbine rotor blades spaced apart from one another in the axial direction, each of the turbine rotor blades extending in the radial direction; and a thrust bearing disposed in a load path from the turbine to the static structure, the load path extending through the gearbox to transmit axial loads from the turbine to the static structure.

GAS TURBINE ENGINE WITH A GEARED TURBOFAN ARRANGEMENT

A gas turbine engine with a geared turbofan arrangement with a gearbox in a drive shaft assembly driven by a turbine is provided. A driving side of the gearbox being driveably connected with at least one propulsive fan, with at least one mechanical fuse in the drive shaft assembly enabling a controlled disengagement of at least one engine part from the drive shaft assembly in case of a mechanical failure of the gas turbine engine or a part thereof and at least one load stop for bearing a load, in particular an axial or radial load in case of the mechanical failure of the gas turbine or a part thereof. A first mechanical fuse is positioned in a torque carrying shaft or a torque carrying part of a shaft, in particular in a torque bearing coupling between the shaft and the gearbox. 1. A gas turbine engine with a geared turbofan arrangement with a gearbox in a drive shaft assembly driven by a turbine , a driving side of the gearbox being driveably connected with at least one propulsive fan , withat least one mechanical fuse in the drive shaft assembly enabling a controlled disengagement of at least one engine part from the drive shaft assembly in case of a mechanical failure of the gas turbine engine or a part thereof andat least one load stop for bearing a load, in particular, an axial or radial load in case of the mechanical failure of the gas turbine or a part thereof,wherein a first mechanical fuse is positioned in a torque carrying shaft or a torque carrying part of a shaft, in particular in a torque bearing coupling between the shaft and the gearbox.2. The gas turbine engine according to claim 1 , wherein at least one mechanical fuse comprises a defined thinning claim 1 , a structuring of a load bearing material and/or a structure with a defined deformable zone.3. The gas turbine engine according to claim 1 , wherein the first mechanical fuse comprises a spline joint.4. The gas turbine engine according to claim 1 , wherein a second mechanical fuse is positioned in ...

FAN DRIVE GEAR SYSTEM

A gas turbine engine includes a fan section and a speed change mechanism for driving the fan section. A first fan section support bearing is mounted forward of the speed change mechanism and a second fan section bearing is mounted aft of the speed change mechanism. 1. A gas turbine engine comprising:a fan section;a speed change mechanism for driving the fan section; anda first fan section support bearing mounted forward of the speed change mechanism and a second fan section bearing mounted aft of the speed change mechanism.2. The gas turbine engine of claim 1 , wherein the speed change mechanism is a planetary gear system including a sun gear in communication with a fan drive turbine and a planet carrier in communication with the fan section.3. The gas turbine engine of claim 1 , wherein a torque frame surrounds the speed change mechanism.4. The gas turbine engine of claim 3 , wherein the torque frame includes a first end for engaging the fan section and second end supporting the second fan section bearing.5. The gas turbine engine of claim 4 , wherein the torque frame includes a plurality of fingers that surround a planet carrier of the speed change mechanism.6. The gas turbine engine of claim 5 , wherein the second fan section bearing is a fan thrust bearing supported on a bearing support attached to distal ends of the plurality of fingers on the torque frame.7. The gas turbine engine of claim 6 , wherein the bearing support includes at least one tang that engages a groove in at least one of the plurality of fingers.8. The gas turbine engine of claim 7 , wherein the groove is located on a radially inner side of at least one of the plurality of fingers.9. The gas turbine engine of claim 1 , wherein the speed change mechanism is at least partially axially aligned with a compressor section.10. The gas turbine engine of claim 1 , further comprising a high pressure compressor with a compression ratio of approximately 20:1 or greater.11. The gas turbine engine of claim ...

FAN DRIVE GEAR SYSTEM

A gas turbine engine includes a fan section and a speed change mechanism for driving the fan section. A first fan section support bearing is mounted forward of the speed change mechanism and a second fan section bearing is mounted aft of the speed change mechanism. 1. A gas turbine engine comprising:a fan section;a speed change mechanism for driving the fan section; anda first fan section support bearing mounted forward of the speed change mechanism and a second fan section bearing mounted aft of the speed change mechanism.2. The gas turbine engine of claim 1 , wherein the speed change mechanism is a planetary gear system including a sun gear in communication with a fan drive turbine and a planet carrier in communication with the fan section.3. The gas turbine engine of claim 1 , wherein a torque frame surrounds the speed change mechanism.4. The gas turbine engine of claim 3 , wherein the torque frame includes a first end for engaging the fan section and second end supporting the second fan section bearing.5. The gas turbine engine of claim 4 , wherein the torque frame includes a plurality of fingers that surround a planet carrier of the speed change mechanism.6. The gas turbine engine of claim 5 , wherein the second fan section bearing is a fan thrust bearing supported on a bearing support attached to distal ends of the plurality of fingers on the torque frame.7. The gas turbine engine of claim 6 , wherein the bearing support includes at least one tang that engages a groove in at least one of the plurality of fingers.8. The gas turbine engine of claim 7 , wherein the groove is located on a radially inner side of at least one of the plurality of fingers.9. The gas turbine engine of claim 1 , wherein the speed change mechanism is at least partially axially aligned with a compressor section.10. The gas turbine engine of claim 1 , further comprising a high pressure compressor with a compression ratio of approximately 20:1 or greater and a fan bypass ratio of ...