Rolls-royce Corporation

Michiel Kopmels

Keith C Sadler

Peter Jeffrey Goodman

Ian Tibbott

Andrew Hewitt

Edwin Dane

Mark Timothy Mitchell

Michael J Jago

Simon John Crowther

Darren L Howard

Philip Ridland

Timothy M Mitchell

Simon Crowther

Darren Howard

Keith Christopher Sadler

Michael John Jago

A method of selectively sealing at least one cooling hole provided on an aerofoil component (). The method comprises applying a thermoplastic material to one or more selected cooling holes () in a molten phase, and cooling the thermoplastic material to set the thermoplastic material.

Method of sealing cooling holes

A master component is provided for use in checking for any variation in the performance of a gas turbine test facility. The master component replaces a standard component, which is typically a turbine blade, vane or segment in the test facility. The master component is attached to the test facility in the same manner as the standard component. The flow rate through the master component is substantially the same as the flow rate through the standard component for a given inlet and outlet temperature and pressure. The flow through the master component is simpler than the flow through the standard component. This means that the flow through the master component is less likely to vary over time. As such, the master component is more suitable for use in checking for any variation in the performance of a test facility.

Master component for flow calibration

An aerofoil comprising a pressure-side wall, a suction-side wall and an intermediate wall extending from a free end of the pressure-side wall at an acute angle relative thereto towards the suction-side wall. A cooling fluid passageway extends through a region where the intermediate wall meets the pressure-side wall at an apex. The fluid passageway has an opening, at least in part, in the face of the pressure-side wall.

Aerofoil

A component, such as a turbine blade of a gas turbine engine, has an internal cooling system which includes a passage () having a passage inlet () and a passage exit (), which may be in the form of passageways at or adjacent to the trailing edge of the component. The passage () is divided into chambers () by partitions () which extend from one wall () of the component and terminate short of the opposite wall () to provide gaps () to permit chord-wise cooling air flow from the passage inlet () to the passage exit (). The gaps () force the cooling air to pass adjacent the hot walls (), so increasing the heat transfer coefficient between the cooling air and the material of the component.

Cooled component

A shroudless, air cooled gas turbine blade has an internal structure including load-carrying members that support the blade outer wall and divide the blade interior into cooling passages. The dimensions of the structure in particular the load-carrying members are chosen with respect to inward heat flow during operation to tend to reduce heat transfer from the outer wall to the load-carrying members and with respect to mass distribution in the structure to reduce centrifugal forces during operation.

Internally cooled airfoils with load carrying members

A component such as a turbine blade of a gas turbine engine has a cooling arrangement comprising a cascade impingement array in which cooling air flows from a supply chamber through impingement passages in webs to an internal passageway comprising first and second limbs. The cooling air flow through the limbs, provides improved heat transfer compared with continued impingement cooling in the chordwise direction of the blade.

Air-cooled component

A cooling arrangement is provided in which an incident coolant flow is presented to a passage such that a proportion of that flow is diverted in order to create a standing relative overpressure in a chamber compared to the actual static pressure of the flow in the passage. In such circumstances through flow transfer apertures in the walls of the chamber forced coolant flows are presented for impingement upon surfaces and therefore greater heat transfer. Generally, coolant flow is presented to both ends of the chamber in order to create the relative standing overpressure in the chamber relative to the current static pressure presented in the passage through the incident flow. In such circumstances through utilizing the dynamic rotation or other motion of a component incorporating the arrangement a higher relative pressure is achieved for greater force coolant flow projection compared to the incident or actual static pressure of the incident coolant flow in the.

Cooling arrangement

An internal fluid cooling system for a gas turbine aerofoil comprises a plurality of multi-pass cooling arrangements each of which consists of a serpentine passage in the interior of the aerofoil. The cooling fluid&#x2014;in particular air&#x2014;is supplied to an inlet end of each passage and exhausted through a multiplicity of discharge holes to provide tip, leading edge, trailing edge and surface film cooling. The inlet end of a first serpentine passage is positioned close to the leading edge and flows rearwards while the inlet end of the second serpentine passage is positioned close to the trailing edge and flows forwards. These serpentine passages are disposed side-by-side, one adjacent the pressure surface and the other adjacent the suction surface on opposite sides of a main load carrying member which comprises a major part of the internal structure of the aerofoil.

Cooled gas turbine aerofoil

A component such as a turbine blade for a gas turbine engine has a cooling arrangement comprising a supply passagewithin the blade from which extend cooling passagesof a first array which open at discharge openings, for example at a trailing edge of the blade. Cooling passagesof a second array extend into the blade from discharge openings, and intersect the passagesThe passagesterminate short of the supply passageAs a result of this arrangement, the limiting flow area for cooling air is defined by the cooling passagesof the first array, and is not affected by the accuracy with which the cooling passagesof the second array intersect the cooling passagesof the first array.

Component having a cooling arrangement

A component such as a turbine blade of a gas turbine engine includes a film cooling arrangement which is optimized by a process in which a component is manufactured having a film cooling arrangement of an initial design and evaluation of the performance of the film cooling arrangement is conducted, for example on the basis of the blowing rate of cooling holes present in the initial design. The configuration of a cooling hole of the initial design is subsequently modified in order to improve the performance of the film cooling arrangement and a component is manufactured in accordance with the modified design. By way of example, the initial design may provide a single cooling holeand the modification may comprise the provision of additional cooling holeseach intersecting the initial cooling holeadjacent its inlet end which opens into a source of cooling fluid

Component having a film cooling arrangement

A heat transfer arrangement comprises a target member (or), and an impingement member (). The impingement member () defines a fluid path (or) there through to direct fluid onto the target member. The arrangement includes a fluid directing formation () on the impingement member (). The fluid path (or) extends through the fluid directory formation () such that the fluid path directs fluid to exit there from at an exit angle that is substantially orthogonal to the fluid directing formation.

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Bristol, United Kingdom

Michiel Kopmels