A cooling circuit is provided disposed between a first wall portion and a second wall portion of a wall for use in a gas turbine engine, one or more inlet apertures, and one or more exit apertures. The inlet aperture(s) provides a cooling airflow path into the cooling circuit and the exit aperture(s) provides a cooling airflow path out of the cooling circuit. The cooling circuit includes a plurality of first pedestals extending between the first wall portion and the second wall portion. The first pedestals are arranged in one or more rows.

Method and apparatus for cooling a wall within a gas turbine engine

A turbine shroud apparatus is provided for a gas turbine engine having a central axis. The apparatus includes: (a) an annular support member (14); (b) a turbine shroud (12) disposed in the support member (14), the shroud (12) being a continuous ring comprising a low-ductility material and having opposed flowpath and back surfaces, and opposed forward and aft ends; and (c) a spring (32) mounted between the support member (14) and the shroud (12) and arranged to resiliently urge the shroud (12) to a concentric position within the support member (14).

Mounting apparatus for low-ductility turbine shroud

A cooling hole configuration for an air-cooled component, such as a gas turbine engine airfoil. The cooling hole is configured to have cross-sectional variations and a noncircular-shaped diffuser-type opening that significantly improve the cooling film distribution across the external surface of an airfoil, with the result that heat transfer from the surrounding environment to the airfoil is reduced. The cooling hole is configured to have its central axis at an acute angle to the exterior surface of the airfoil, and defines a noncircular-shaped opening at the airfoil surface. The cooling hole generally has a first region adjacent the airfoil surface and a second region interior to the airfoil. The cooling hole is configured such that the second region has an oblong or oval-shaped cross-section, with a major diameter approximately equal to the major diameter of the first region and a minor diameter less than the major diameter of the first region, with a smooth transition existing between the first and second regions. The oblong shape of the second region is preferably the result of a recess being present in the wall of the cooling hole opposite the direction in which the cooling hole extends toward the airfoil surface, causing the central axis of the cooling hole to have an arcuate shape in which the central axis is disposed at a lesser angle to the airfoil surface in the first region than the angle in the second region. The cooling hole can be generated using a water jet technique.

Shaped cooling hole for an airfoil

A gas turbine engine cooling system includes a heat exchanger in fluid communication with a source of cooling air, a first cooling circuit including a first heat exchanger circuit in the heat exchanger and a first bypass circuit with a first bypass valve for selectively bypassing at least a portion of first airflow around the first heat exchanger circuit. A second cooling circuit may be used having a second heat exchanger circuit in the heat exchanger and a shutoff control valve operably disposed in the second cooling circuit upstream of the second heat exchanger circuit and the heat exchanger. A circuit inlet of the first cooling circuit may be used to bleed a portion of compressor discharge bleed air for the first airflow to cool turbine blades mounted on a rotor disk using an annular flow inducer downstream of the first bypass valve and the heat exchanger.

Gas turbine engine temperature modulated cooling flow

A precooler for cooling compressor bleed air for an environmental control system includes a heat exchanger in fluid communication with a source of cooling air and operable for cooling the bleed air. A variable bypass valve between a bleed air source and environmental control system is operable for bypassing at least a portion of the compressor bleed air around the heat exchanger. The cooling air may be a portion of fan air modulated by a variable fan air valve. The bleed air source may be selectable between the low pressure bleed air source and a high pressure bleed air source. One method includes flowing the compressor bleed air from a single low pressure source only and increasing thrust sufficiently to meet a minimum level of pressure of the bleed air during one engine out aircraft operating condition during approach or loitering.

Environmental control system supply precooler bypass

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

A fluid cooled article having a protective coating on a surface, for example an air cooled gas turbine engine article having a Thermal Barrier Coating, includes through an article wall a fluid cooling passage, and typically a plurality of passages, having openings sized to maintain desired fluid flow, unobstructed by coating within the passage at an exit opening. The passage has a first or inlet opening, which establishes the amount of fluid flow through the passage, and a second opening through which the flow exits the passage through a wall surface on which the coating is deposited. A method for making the article to have such unobstructed flow includes selecting a cross sectional area for the first opening, selecting a coating method for deposit of the coating, in a selected thickness range, on the wall surface and partially within the passage at the second opening, and generating the passage to have a second opening with a cross sectional area greater than that of the first opening by an amount which is no less than a reduction in cross sectional area resulting from partial deposit of the coating through the second opening.

Fluid cooled article with protective coating

A combustor has a flow sleeve and a flow liner defining a generally axial flow direction of compressor discharge air toward combustor burners. A casing is secured to the forward end of the flow sleeve defining an annular plenum along the interior of the flow sleeve. Openings through the flow sleeve supply compressor discharge air into the plenum where the air changes direction for flow through apertures into and generally coaxially with the free air stream. The axial injection minimizes or eliminates energy losses due to cross flow injection within the axial air stream while continuing to cool the liner.

Axial flow sleeve for a turbine combustor and methods of introducing flow sleeve air

Method for making a fluid cooled article with protective coating

A row of turbine blades is surrounded by a turbine shroud. The shroud is supported by a turbine case having a pair of radially outwardly extending rails. The rails are perforate for channeling therethrough control air to control thermal radial movement of the case, and in turn radial position of the shroud supported therefrom.

Robert Paul Ziegler

Papers

Compound clearance control engine

Fluid cooled article with protective coating

Axial flow sleeve for a turbine combustor and methods of introducing flow sleeve air

Method for making a fluid cooled article with protective coating

Transpiration clearance control turbine