EP2990660A1 - Verschleissmonitor für einen gasturbinenmotor - Google Patents

Verschleissmonitor für einen gasturbinenmotor Download PDF

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Publication number
EP2990660A1
EP2990660A1 EP15182102.2A EP15182102A EP2990660A1 EP 2990660 A1 EP2990660 A1 EP 2990660A1 EP 15182102 A EP15182102 A EP 15182102A EP 2990660 A1 EP2990660 A1 EP 2990660A1
Authority
EP
European Patent Office
Prior art keywords
depth
subgroup
liner
abradable
indicators
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Granted
Application number
EP15182102.2A
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English (en)
French (fr)
Other versions
EP2990660B1 (de
EP2990660B8 (de
Inventor
Michael Keenan
Nicholas Thomas
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Rolls Royce PLC
Original Assignee
Rolls Royce PLC
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Rolls Royce PLC filed Critical Rolls Royce PLC
Publication of EP2990660A1 publication Critical patent/EP2990660A1/de
Publication of EP2990660B1 publication Critical patent/EP2990660B1/de
Application granted granted Critical
Publication of EP2990660B8 publication Critical patent/EP2990660B8/de
Active legal-status Critical Current
Anticipated expiration legal-status Critical

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Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/08—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator
    • F01D11/12—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator using a rubstrip, e.g. erodible. deformable or resiliently-biased part
    • F01D11/122—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator using a rubstrip, e.g. erodible. deformable or resiliently-biased part with erodable or abradable material
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D27/00—Control, e.g. regulation, of pumps, pumping installations or pumping systems specially adapted for elastic fluids
    • F04D27/001—Testing thereof; Determination or simulation of flow characteristics; Stall or surge detection, e.g. condition monitoring
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00—Details, component parts, or accessories
    • F04D29/08—Sealings
    • F04D29/16—Sealings between pressure and suction sides
    • F04D29/161—Sealings between pressure and suction sides especially adapted for elastic fluid pumps
    • F04D29/164—Sealings between pressure and suction sides especially adapted for elastic fluid pumps of an axial flow wheel
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00—Details, component parts, or accessories
    • F04D29/40—Casings; Connections of working fluid
    • F04D29/52—Casings; Connections of working fluid for axial pumps
    • F04D29/522—Casings; Connections of working fluid for axial pumps especially adapted for elastic fluid pumps
    • F04D29/526—Details of the casing section radially opposing blade tips
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/08—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator
    • F01D11/14—Adjusting or regulating tip-clearance, i.e. distance between rotor-blade tips and stator casing
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D21/00—Shutting-down of machines or engines, e.g. in emergency; Regulating, controlling, or safety means not otherwise provided for
    • F01D21/003—Arrangements for testing or measuring
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/80—Diagnostics
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2270/00—Control
    • F05D2270/30—Control parameters, e.g. input parameters
    • F05D2270/305—Tolerances
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2270/00—Control
    • F05D2270/80—Devices generating input signals, e.g. transducers, sensors, cameras or strain gauges
    • F05D2270/821—Displacement measuring means, e.g. inductive

Definitions

  • This invention relates to an abradable liner for a compressor of a gas turbine engine.
  • the invention relates to a wear indication system incorporated into the abradable liner.
  • Figure 1 shows a typical three shaft gas turbine engine 10.
  • the gas turbine engine 10 includes an air intake 1, a fan 2 having rotating blades, a bypass duct 18 and an engine core 12.
  • the engine core 12 includes an intermediate pressure compressor 3, a high pressure compressor 4, a combustor 5, a turbine arrangement comprising a high pressure turbine 6 an intermediate pressure turbine 7, a low pressure turbine 8 and an exhaust nozzle 9. Air entering the intake 1 is accelerated by the fan 2 and directed into two air flows. The first air flow passes into the engine core 12, and the second air flows along the bypass 18 to provide propulsive thrust.
  • the engine core air flow travels through the intermediate 3 and high 4 pressure compressors in turn.
  • the compressed air exhausted from the high pressure compressor 4 is mixed with fuel from an injector 14 and burnt in the combustor 5.
  • the hot gas expands through and drives the high 6, intermediate 7 and low 8 pressure turbines before being exhausted through the nozzle 9 and adding to the propulsive thrust created by the first air flow.
  • the high 6, intermediate 7 and low 8 pressure turbines respectively drive the high 4 and intermediate 3 pressure compressors and the fan 2 via respective shafts.
  • the gap between compressor blade tips and the engine casing is closely controlled to minimise the leakage of compressed air over the blade tips and back upstream.
  • the engine casings often include an attrition or abradable liner which provides a close fitting seal with the blade tips.
  • the abradable liner is initially installed so as to be in contact with the compressor blade tips such that, the liner is scored by the rotating compressor (or fan as the case may be) during the first few rotations which removes enough material to allow a close fitting free rotation of the blades.
  • the compressor casing may vary in its symmetry during operation from circular to elliptical or other non-round shape and it is difficult to accurately predict the actual axisymmetric and asymmetric rubs in a compressor.
  • the rubs in an abradable liner are typically measured after running and striping a compressor. This information is used to adjust the geometry of the blades and/or casing liner to ensure the optimum sealing.
  • the present invention seeks to provide a solution to help monitor and control abradable liner damage.
  • an abradable liner for a gas turbine engine compressor stage, the liner extending about an axis and comprising: an abradable body having a radially inner surface over which an aerofoil tip passes in use; a plurality depth indicators extending into the body from the surface; wherein, the depth indicators are arranged in two or more groups each group being circumferentially spaced on the surface from another group; wherein each group has a first subgroup of depth indicators that are circumferentially offset from other depth indicators in the first subgroup, the depth indicators also being axially offset from other depth indicators in the first subgroup; each group further having a second subgroup of depth indicators circumferentially offset from the first subgroup, the depth indicators in the second subgroup being circumferentially aligned with other depth indicators in the second subgroup.
  • the depth indicators may be visible using a borescope or other inspection equipment allowing an easy determination of whether the depth indicator is still present in the abradable body or whether material has been removed from the abradable body that is at least equal to the depth of the depth indicator.
  • At least one depth indicator extends into the body a greater distance than at least one other depth indicator. This may be used to provide a scale from which the amount of material removed can be more easily calculated
  • Each depth indicator may be a blind aperture and may present a circular or elliptical cross-section to the surface.
  • the maximum length of the longest axis of the cross-section may be 1cm but more preferably may be less than 0.5cm.
  • the relatively small size of the apertures presented to the surface ensures that aerodynamic effects on the compressor are limited so as to be almost negligible. This allows for depth indicators to be provided on the first engines entering development or service to confirm and check operation but not provided on subsequent engines and performance of the engines to be the same regardless of whether the holes are present or not.
  • Each depth indicator may be filled with a visual indicator.
  • the visual indicator further reduces the aerodynamic effects of the depth indicators on the engine performance.
  • the visual indicator may be coloured or fluorescent. The colouring may be used to help the operator of the inspection device confirm the portion of the liner that he is viewing.
  • the depth indicators may be arranged in two or more groups with the depth indicators of each group being filled with coloured material of a different colour to at least one of the other groups.
  • the different colours can allow easier inspection as it can easily be determined which group is being inspected.
  • the liner may extends about an axis with the surface defining an inner wall of a tube, wherein the groups extend in at least one array extending circumferentially along the inner wall.
  • Two or more arrays may be provided with each array being axially spaced from another of the arrays.
  • the liner may extend about an axis with the surface defining an inner wall of a tube, wherein the depth indicators are arranged in two or more groups each group being circumferentially spaced on the surface.
  • the liner may be used in a gas turbine engine compressor unit having a row of blades which rotate about an axis, and a liner extending about the axis; the liner may have an abradable body with a radially inner surface over which an blade tip passes in use; a plurality of abradable depth indicators extending into the body from the surface; wherein, the depth indicators are arranged in two or more groups each group being circumferentially spaced on the surface from another group; each group having a plurality of axially spaced depth indicators aligned to respectively correspond with a leading edge, a trailing edge and a mid-chord of the blade tip.
  • the axially spaced depth indicators may be circumferentially offset from each other.
  • Each group may have a first subgroup of depth indicators that are circumferentially offset from each depth indicator in the first subgroup and one depth indicator is axially aligned to correspond with the leading edge, a second depth indicator is axially aligned to correspond with a trailing edge and third depth indicator is axially aligned to correspond with a mid-chord of the aerofoil tip; each group further having a second subgroup of depth indicators circumferentially offset from the first subgroup, the depth indicators in the second subgroup being circumferentially aligned with each other.
  • Each depth indicator in the second subgroup may extend into the body the same distance as the one depth indicators in the second subgroup.
  • Each depth indicator in the first subgroup may extend into the body to a different distance as the other depth indicators in the first subgroup.
  • Each depth indicator may be a hole.
  • Each depth indicator may taper as it extends into the abradable liner. This allows the cross-sectional size of the hole to be measured and, provided the taper angle of the hole is known and the original cross-sectional area presented to the surface it is possible to calculate the amount of the liner removed.
  • coloured material it may be provided in two or more layers of different colour. There may be a distinct boundary between the layers.
  • Figure 2 shows a cross section of a compressor casing 110 having an attrition liner 112 which incorporates an annulus of abradable material.
  • the liner 112 is located around the compressor blade 2 so as to provide a sealing function.
  • the casing of the embodiment shown is a ring compressor casing although this application is equally applicable to split casings, or casings used in turbine or fan applications.
  • Split casings are usually formed as two semicircles joined together at an attachment flange whilst the ring casings are provided by a single hoop of material.
  • a separate case may be provided for one or more stages of the compressor and are joined together to form a casing that extends the length of the compressor.
  • the attrition liner is sprayed onto the compressor casing or rotor path
  • the liner includes an abradable portion 116 which is designed to be contacted and abraded by the compressor blade in use.
  • the abradable portion may be any suitable type known in the industry such as polyester based material.
  • one or more depth indicators that provide a visual indication of the wear of the abradable liner around the circumference of the liner.
  • the relative axial and radial position of the blade tips and casing varies throughout the flight cycle and understanding the relative movement around the full circumference assists in providing more efficient components by helping to ensure that the optimum blade length can be provided to deliver the smallest running clearances between the blade tip and liner and also that the liner shape could be adjusted to correct for flight asymmetries.
  • Each depth indicator is provided by a hole or that is provided within the liner at a known circumferential and axial position and extending from the surface of the liner into the liner body.
  • the holes and grooves can be arranged in arrays extending about the circumference with each array having depth indicators of identical, or different depths, as required.
  • each depth indicator 55, 56, 57 are provided with each indicator being of a different depth.
  • the depth indicators extend into the liner from the abradable surface 50.
  • each depth indicator may be observed visually from a borescope inserted into the engine at a routine, or scheduled, inspection. The operator need not conduct time consuming measurement activities to determine that each depth indicator is observed.
  • the liner may be eroded to a new liner surface 50 depicted in Figure 4 .
  • the erosion may be to an extent greater than the depth of one or more of the depth indicators which means that one or more of the depth indicators is not visible.
  • depth indicator 55 is not visible but indicators 56 and 57 can still be observed. Without extensive measurement and calculation it is possible for the inspector to determine that the liner has been worn to a distance between the depth of 55' and the depth of 56'.
  • each depth indicator may be filled with a material that is easily identifiable against the background of the liner.
  • the material may be identifiable by colour or by the presence of a marker e.g. a fluorescent compound. Where colour is used the depth indicators of a particular depth may each have the same colour which is different to the depth indicators of another depth.
  • the colours or pattern of colours indicate the depth to which the liner has been removed.
  • the filled depth indicators may be circumferential slots and / or drilled holes arranged at a plurality of different circumferential locations. Where holes are used they may be 1cm in diameter, or less. A diameter of 4-5mm has been found to be a suitable size.
  • depth indicators are provided at four circumferential locations around the liner and they may be offset from each other to limit the danger of a weakened liner caused by too many depth indicators in a given locale.
  • the depth indicators may be arranged in a group of five as shown in Figure 5 with two of the holes 58a, 58b aligned circumferentially to help to understand if there is any asymmetry in the angle of the blade tip 60 relative to the liner as it passes over the liner. Greater wear in one or more of the depth indicators 58a, 58b is indicative that the blade tip is not parallel to the liner surface 50.
  • the depth indicators 62 have a conical profile such that the depth that the liner is abraded can be determined by measuring the diameter of the hole at the surface.
  • Each hole or groove can be filled with a coloured material in order to provide a substantially smooth surface free of the substantial discontinuities that would be exhibited if the holes or grooves were not filled.
  • the coloured material may be provided as one colour per hole, or arranged in layers such that there are multiple colours per hole. Where layers are provided the boundaries may be distinct or blended.
  • all the holes of a given depth may be filled with the same colour material to help identify which hole is part of which group.
  • the invention provides a simple mechanical visual indication which can be used to provide quick and reliable information as to the extent and pattern of wear in an attrition liner. This can be used by maintenance staff to determine when an engine requires an overhaul and allows for efficient scheduling.
  • the arrangements allow an early indication of rubs around the circumference of the compressor casing. This allows the geometry of the compressor to be modified as soon as an engine is run - rather than waiting for a full strip and measurement.
  • depth indicators conical, cylindrical or grooved may be used independently or together as deemed appropriate to achieve the best depth determination. Where grooves are used they should be kept small to keep the aerodynamic effects to a minimum.
  • the embodiments may find application in determining rub depth and location on any appropriate liner in pump, fan or turbine applications.
EP15182102.2A 2014-08-28 2015-08-24 Verschleissmonitor für einen gasturbinenmotor Active EP2990660B8 (de)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
GBGB1415201.1A GB201415201D0 (en) 2014-08-28 2014-08-28 A wear monitor for a gas turbine engine fan

Publications (3)

Publication Number Publication Date
EP2990660A1 true EP2990660A1 (de) 2016-03-02
EP2990660B1 EP2990660B1 (de) 2020-02-05
EP2990660B8 EP2990660B8 (de) 2020-03-11

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EP15182102.2A Active EP2990660B8 (de) 2014-08-28 2015-08-24 Verschleissmonitor für einen gasturbinenmotor

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US (1) US20160061050A1 (de)
EP (1) EP2990660B8 (de)
GB (1) GB201415201D0 (de)

Cited By (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US10428674B2 (en) * 2017-01-31 2019-10-01 Rolls-Royce North American Technologies Inc. Gas turbine engine features for tip clearance inspection
WO2020117882A1 (en) * 2018-12-06 2020-06-11 Gulfstream Aerospace Corporation Visual detection of fan case liner damage for turbine engine
FR3091548A1 (fr) * 2019-01-09 2020-07-10 Safran Aircraft Engines Elément abradable de turbomachine pourvu de témoins d’usure visuels
WO2020208316A1 (fr) 2019-04-12 2020-10-15 Safran Aircraft Engines Procedure de detection d'une asperite sur une couche abradable dans un carter de soufflante
FR3125316A1 (fr) * 2021-07-16 2023-01-20 Safran Aircraft Engines Element abradable comportant un temoin d'usure

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WO2014158236A1 (en) * 2013-03-12 2014-10-02 United Technologies Corporation Cantilever stator with vortex initiation feature
WO2015065712A1 (en) * 2013-10-30 2015-05-07 United Technologies Corporation Nosecap
US10612413B2 (en) * 2017-03-06 2020-04-07 United Technologies Corporation Wear indicator for determining wear on a component of a gas turbine engine
US11326469B2 (en) * 2020-05-29 2022-05-10 Rolls-Royce Corporation CMCs with luminescence environmental barrier coatings

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US4329308A (en) * 1976-01-30 1982-05-11 Societe Nationale D'etude Et De Construction De Moteurs D'aviation Method of making an abradable stator joint for an axial turbomachine
EP2065566A1 (de) * 2007-11-28 2009-06-03 United Technologies Corporation Segmentierte keramische Schicht für ein Element eines Gasturbinenantriebs
FR2929349A1 (fr) * 2008-03-28 2009-10-02 Snecma Sa Carter pour roue a aubes mobiles de turbomachine
WO2013050688A1 (fr) * 2011-10-07 2013-04-11 Turbomeca Compresseur centrifuge equipe d'un marqueur de mesure d'usure et procede de suivi d'usure utilisant ce marqueur

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US4329308A (en) * 1976-01-30 1982-05-11 Societe Nationale D'etude Et De Construction De Moteurs D'aviation Method of making an abradable stator joint for an axial turbomachine
EP2065566A1 (de) * 2007-11-28 2009-06-03 United Technologies Corporation Segmentierte keramische Schicht für ein Element eines Gasturbinenantriebs
FR2929349A1 (fr) * 2008-03-28 2009-10-02 Snecma Sa Carter pour roue a aubes mobiles de turbomachine
WO2013050688A1 (fr) * 2011-10-07 2013-04-11 Turbomeca Compresseur centrifuge equipe d'un marqueur de mesure d'usure et procede de suivi d'usure utilisant ce marqueur

Cited By (11)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US10428674B2 (en) * 2017-01-31 2019-10-01 Rolls-Royce North American Technologies Inc. Gas turbine engine features for tip clearance inspection
WO2020117882A1 (en) * 2018-12-06 2020-06-11 Gulfstream Aerospace Corporation Visual detection of fan case liner damage for turbine engine
CN113227541A (zh) * 2018-12-06 2021-08-06 湾流航空公司 涡轮发动机的风扇壳体衬里损坏的视觉检测
FR3091548A1 (fr) * 2019-01-09 2020-07-10 Safran Aircraft Engines Elément abradable de turbomachine pourvu de témoins d’usure visuels
GB2580789A (en) * 2019-01-09 2020-07-29 Safran Aircraft Engines Abradable turbomachine element provided with visual wear indicators
US11225879B2 (en) 2019-01-09 2022-01-18 Safran Aircraft Engines Abradable turbomachine element provided with visual wear indicators
GB2580789B (en) * 2019-01-09 2022-12-14 Safran Aircraft Engines Abradable turbomachine element provided with visual wear indicators
WO2020208316A1 (fr) 2019-04-12 2020-10-15 Safran Aircraft Engines Procedure de detection d'une asperite sur une couche abradable dans un carter de soufflante
FR3095045A1 (fr) * 2019-04-12 2020-10-16 Safran Aircraft Engines Procede de detection d’une aspérité sur une couche abradable dans un carter de soufflante
US11753957B2 (en) 2019-04-12 2023-09-12 Safran Aircraft Engines Method for detecting a roughness in an abradable layer in a fan casing
FR3125316A1 (fr) * 2021-07-16 2023-01-20 Safran Aircraft Engines Element abradable comportant un temoin d'usure

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Publication number Publication date
US20160061050A1 (en) 2016-03-03
EP2990660B1 (de) 2020-02-05
EP2990660B8 (de) 2020-03-11
GB201415201D0 (en) 2014-10-15

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