A turbine includes a plurality of inner shell sections mounting first and second stage nozzle and shroud portions. The inner shell sections are pinned to an outer containment shell formed of sections to preclude circumferential movement of the inner shell relative to the outer shell and enable thermal expansion and contraction of the inner shell relative to the outer shell. Positive bucket tip clearance control is afforded by passing a thermal medium about the inner shell in heat transfer relation with the shrouds about the first and second stage bucket tips, the thermal medium being provided from a source of heating/cooling fluid independent of the turbine. Access is provided to the rotor and turbine buckets by removing the outer and inner shell sections.

Removable inner turbine shell with bucket tip clearance control

A gas turbine engine (10) includes a compressor (18), combustor (20), and turbine (22) having a row of blades (40) mounted inside a surrounding turbine shroud (44). A heat exchanger (56) is used for cooling pressurized air bled from the compressor (18). A distribution network (60-66) joins the heat exchanger (56) to the turbine (22) for selectively channeling air from the heat exchanger (56) below the blades (40) and above the shroud (44) for controlling blade tip clearance.

Compound clearance control engine

An apparatus is provided to thermally control a section of engine casing by flowing heat transfer fluid within an axially disposed set of circular flowpaths within a section of engine casing or flange such as a thermal control ring. Flowpaths are preferably arranged in alternating counterflowing directions in the ring such that there is substantially no circumferential gradient in the mass flowrate weighted average temperature of the heat transfer fluid supplied by any two adjacent counterflowing fluid flowpaths. Each flowpath may have two or more circuits having inlets and outlets for flowing the heat transfer fluid.

Gas turbine engine case thermal control flange

A casing for the fan section of an aircraft gas turbine engine of the turbofan type. The casing includes an inner filler material and is defined by an outer shell and an inner liner that is spaced radially inwardly from the outer shell. The inner liner is held in position relative to the outer shell by a plurality of radially-inwardly-extending support pins. The pins are carried by the outer shell and are slidably received in and extend at least partially into the inner liner. The support pins retain the inner liner against axial and rotational movement relative to the outer shell, while allowing relative radial movement therebetween in response to thermal expansion and contraction, as well as in response to aircraft maneuver loads imposed on the engine.

Fan casing liner support

A method of operating a gas turbine engine is provided. The gas turbine engine includes at least one engine casing and at least one rotor assembly. The method includes directing airflow through a supply pipe and into a heat exchanger, lowering the temperature of the airflow in the heat exchanger, and directing the cooled airflow into the engine casing to cool the casing.

Methods and apparatus for maintaining rotor assembly tip clearances

Improved operation can be achieved from an enhanced gas turbine engine having a segmented annular shroud which radially expands and contracts to match the expansion and contraction of engine rotor blades. The segmented annular shroud is supported by a structure which includes an annular ring having two radially outwardly extending flanges, and forward and aft annular segmented brackets which attach the segmented shroud to the forward and aft side of the ring respectively. In a preferred embodiment of the invention, two circumferentially extending separate and distinct air impingement manifolds surrounds each of the outwardly extending flanges. Each manifold is provided with a valve for controlling the amount and temperature of the airflow entering each manifold. The air from each manifold then impinges upon each of the outwardly extending flanges, thereby controlling the radial movement of the corresponding shroud segments and the associated clearances with the rotor blade tips. The use of separate and distinct manifolds and the corresponding values which regulate the amount and temperature of airflow to each manifold allows individual shroud portions to be separately expanded or contracted.

Clearance control system for separately expanding or contracting individual portions of an annular shroud

Gas turbine engine shroud clearance control

Tip clearance control system having a plurality of impingement manifolds (40) to unevenly heat or cool a turbine shroud (28) such creating a non-uniform circumferential temperature distribution to produce ovalization of said turbine shroud (28) to conform to high load induced non-concentricities of the turbine rotor (52).

Clearance control system

L'invention concerne un dispositif qui controle la temperature, et par consequent les dimensions, d'une enveloppe 39 entourant une turbine dans un moteur d'avion a turbine a gaz Une ou plusieurs conduites de decharge 120 A C sont annulaires autour de l'enveloppe, et lui fournissent de l'air dans le but d'en modifier la temperature Une chambre 117, qui est egalement annulaire autour de l'enveloppe, fournit de l'air aux conduites de decharge Le dispositif sert a maintenir l'air decharge a une temperature uniforme pour eviter que l'enveloppe ne prenne une forme non-circulaire a cause des variations de sa temperature De plus, l'air consomme, apres fourniture a l'enveloppe, est maintenu en contact avec la chambre afin de l'isoler et reduire encore les variations de temperature de l'air delivre Application aux moteurs a turbine a gaz

Shotts L D

Papers

Clearance control system for separately expanding or contracting individual portions of an annular shroud

Gas turbine engine shroud clearance control

Clearance control system

Dispositif de controle de jeu dans une turbine a gaz