Rolls-royce Corporation

Leo Vivian Lewis

Marko Bacic

Timothy John Scanlon

Simon James Jones

Anthony G Razzell

Jonathan E Holt

Guy David Snowsill

Peter Beecroft

Colin Young

Lee Mansfield

Robert John Irving

Thomas Horace Hyde

Bengt E G Blidmark

Vinod Gopalkrishna

Barry J Marsden

Heinrich Fessler

Timothy Scanlon

Timothy John Scanlan

Bengt Eric Gustav Blidmark

Anthony Gordon Razzel

A clearance control arrangement () for a rotor () comprising a segment assembly () spaced radially inwards from a casing () and defining a radial clearance () between the rotor () and the segment assembly (). A front carrier () and a rear carrier () supporting the segment assembly (). A heat transfer cavity () formed between the carriers (), the segment assembly () and the casing (). A settling chamber () formed adjacent to the casing () and upstream of the heat transfer cavity (). The clearance control arrangement () is configured to receive air into the settling chamber () and thence to deliver it to the heat transfer cavity ().

Clearance control arrangement

An arrangement for a gas turbine engine that includes a turbine blade configured to rotate about an axis, a casing radially outside of the turbine blade, and a carrier segment mounted to the casing so as to define a first impingement space therebetween. The carrier segment is positioned radially outside the turbine blade and includes a first impingement carrier wall, a main carrier wall, and a cooling chamber. The first impingement carrier wall is adjacent to and radially inside of the first impingement space, and the first impingement carrier wall includes a first aperture. The main carrier wall is radially inside of the first impingement carrier wall. The cooling chamber is radially inside of the main carrier wall. Additionally, an intermediate chamber is disposed radially between the cooling chamber and the first impingement space. A second aperture is configured to allow ingress of air into the intermediate chamber.

Tip clearance control for turbine blades

A case cooling calibration system for a gas turbine engine includes a rotor casing, and a stage, rotatably mounted within the rotor casing. The rotor-casing includes in-series fluid communication, a duct, a flow control valve, and a manifold. The valve regulates an air flow through the duct to the manifold, which is arranged radially outwardly around the casing radially coplanar with the rotor stage. A temperature sensing apparatus is positioned on the rotor casing, radially outwardly of the rotor stage. A controller regulates the valve in a stepwise manner providing a variable cooling air-flow to the rotor-casing. A processor calculates a correction depending on a comparison between the rotor casing temperature and a predicted rotor-casing temperature for a given engine power condition. An engine controller controls the valve depending on the correction characteristic to maintain a predetermined operational radial clearance between the rotor-stage and the rotor-casing.

System and method for calibrating a case cooling system for a gas turbine engine

A carrier segment of a carrier section for circumscribing an array of circumferentially spaced turbine blades of a gas turbine engine, the blades being disposed radially inwardly of a turbine casing, the carrier segment including a carrier wall disposed radially inwardly of the casing and radially outwardly of the turbine blades, and the carrier wall including one or more portions facing the casing, wherein at least one of the one or more portions of the carrier wall is provided with one or more impingement apertures therein for passage therethrough of air of a predetermined temperature, heating air, from a feed source into impingement onto the turbine casing. The application of heating air onto the casing during transient stages of enhanced engine power, e.g. during step climbs, results in better matching of radial expansion of the casing relative to the turbine blades, thereby enabling improved blade tip clearance control.

A gas turbine engine comprising first and second axially spaced turbine rotor stages () and a turbine casing () radially outside the rotor stages. A first seal segment arrangement () forms a cavity () radially between the first turbine rotor stage () and the turbine casing (). A first air source () is coupled to the first seal segment arrangement (). A second seal segment arrangement () forms a cavity () radially between the second turbine rotor stage () and the turbine casing (). A heating chamber () is provided radially between the second seal segment arrangement () and the turbine casing (). A duct () is coupled between the first air source () and the heating chamber ().

Rotor blade tip clearance control

A method of characterizing mechanical and thermal growth of a rotor tip clearance arrangement and a method of controlling such an arrangement. The characterization method comprising constructing a finite element model and calibration cases; run the model for the calibration cases to obtain displacements; and calculating component growth therefrom. The growth is characterized and scaled for the operating conditions. The clearance is calculated in a time step from the growths and input to a control arrangement for subsequent time steps.

Rotor tip clearance

A gas turbine engine control apparatus comprises a controllera memoryassociated with the controllerand inputsfor measurement data from an engine. The controllerdetermines the start of a monitoring cycle atreceives measurement data at the inputsduring the monitoring cycle, manipulates the measurement data to provide an incremental deterioration value representing deterioration occurring within the engine and during the monitoring cycle, and uses the incremental deterioration value atto update a deterioration valuestored in the memoryand determines the start of a further monitoring cycle.

Gas turbine engine control

A rotor blade tip clearance control apparatus for a gas turbine engine includes a plurality of circumferentially distributed segments which form an annular shroud surrounding the outer tips of a row of rotor blades. A mechanical arrangement is operatively connected to the segments. Actuation of the arrangement causes the segments to move in a radial direction thereby controlling a clearance between the segments and the outer tips. A case cooling system supplies cooling air to an engine case to which the segments are mounted. The cooling air regulates thermal expansion of the case and thereby also controls the clearance between the segments and the outer tips.

Combination of mechanical actuator and case cooling apparatus

A gas turbine engine has a row of circumferentially spaced rotor blades and a plurality of seal segments circumscribing the rotor blade tips and attached to a radially inward side of a casing of the engine. The seal segments are spaced from the casing by a spacing cavity. A flow of relatively hot cooling air is routed to the spacing cavity to cool the seal segments. The engine has an external cooling arrangement for impinging relatively cold cooling air on a radially outward side of the casing. The engine has a wall containing a plurality of through-holes which is attached to a radially inward side of the casing adjacent the seal segments. The wall is spaced from the casing to define a heating control chamber between the wall and the casing. The engine has one or more closable ducts which allow air to be exhausted from the heating control chamber.

A method of controlling a system that has a finite execution time in the optimization of a cost function: where N is selected such that J is convergent; k+i is the time in the future after the passage of i time increments; u(k+i|k) and y(k+i|k) are restricted to values of the respective control variables u and performance variables y that are compatible with predetermined system operational constraints; and for each time increment i=1 . . . N, values for F(u(k+i|k), y(k+i|k)) are calculated using a system model M to determine y(k+i|k) for a plurality of respective discrete values of u(k+i|k) that are limited to and representatively sampling the range of possible values of u that are compatible with the predetermined system operational constraints, J being optimised by identifying the combination {u(k+1|k),u(k+2|k), . . . u(k+N|k)} of respective discrete values u(k+i|k) over N time increments that combination yields an optimised value for J.

System control

A gas turbine engine has in flow series a compressor section, a combustor, and a turbine section. The engine includes a turbine section rotor disc, and a stationary wall forward of a front face or rearward of a rear face of the rotor disc. The wall defines a cavity between the stationary wall and the rotor disc, and has a plurality of air entry nozzles through which cooling air can be delivered into the cavity at an inlet swirl angle. The engine further includes a cooling air supply arrangement which accepts a flow of compressed air and supplies the compressed air to the nozzles for delivery into the cavity. The cooling air supply arrangement and the nozzles are configured such that the inlet swirl angle of the air delivered into the cavity can be varied between a first inlet swirl angle and a second inlet swirl angle.

Gas turbine engine swirled cooling air

Some embodiments control rotor blade over-tip leakage of a working fluid in a rotating machine, which includes a circumferential row of rotor blades. A circumferential row of seal segments seal with the radially outer tips of the rotor blades to reduce over-tip leakage of the working fluid. The seal segments have an inboard abradable layer adapted to be abraded by the blade tips to form a wear track for the blade tips. Tip wearing bodies are adapted to wear down the blade tips. The abradable layer thickness and cold build radial positions of the blade tips, seal segments, and tip wearing bodies are arranged such that, during running-in of the machine, a wear track having a uniform radius is formed in the abradable layer by the blade tips, and the tip wearing bodies wear down the blade tips to provide a uniform blade tip radius.

Rotor blade over-tip leakage control

Sensors () for determining a gap between a conductive member () such as a blade in a gas turbine engine and a seal segment () are known to use capacitive variants in order to create an electrical signal indicative of the gap width. Thermal disparities can create problems with regard to sensor aging and accuracy. By creating a sensor incorporating a metal rod () typically integrally formed or associated with the seal segment () and coupled through inductive coupling loops () it is possible to create a tuned circuit with a Q value which is more stable and therefore acceptable with regard to producing more accurate results at elevated temperatures with less problems with regard to thermal disparities.

Sensor

A rotor assembly for a gas turbine engine, the rotor assembly comprises at least two rotors defining a cavity therebetween. A first rotor defines a cooling air inlet in its radially inward portion. A second rotor defines a cooling air outlet in its radially outward portion, such that the cooling air passes radially outwardly through the cavity.

Gas turbine engine rotor ventilation arrangement

A spacing arrangement for a gas turbine engineThe spacing arrangement being arranged such that as a rotor bladeof a compressorrotates, the rotor discwill move outwards and forwards to maintain a substantially constant gapbetween the rotor bladeand an inclined casing

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Kenilworth, United Kingdom

Leo Vivian Lewis