US3736751A - Gap control apparatus - Google Patents

Gap control apparatus Download PDF

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Publication number
US3736751A
US3736751A US00147265A US3736751DA US3736751A US 3736751 A US3736751 A US 3736751A US 00147265 A US00147265 A US 00147265A US 3736751D A US3736751D A US 3736751DA US 3736751 A US3736751 A US 3736751A
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United States
Prior art keywords
clearance
control
tube
fluid
hot
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.)
Expired - Lifetime
Application number
US00147265A
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English (en)
Inventor
J Rodney
D Fuller
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UK Secretary of State for Defence
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Secr Defence
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    • 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
    • F01D11/16—Adjusting or regulating tip-clearance, i.e. distance between rotor-blade tips and stator casing by self-adjusting means
    • F01D11/18—Adjusting or regulating tip-clearance, i.e. distance between rotor-blade tips and stator casing by self-adjusting means using stator or rotor components with predetermined thermal response, e.g. selective insulation, thermal inertia, differential expansion
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F16—ENGINEERING ELEMENTS AND UNITS; GENERAL MEASURES FOR PRODUCING AND MAINTAINING EFFECTIVE FUNCTIONING OF MACHINES OR INSTALLATIONS; THERMAL INSULATION IN GENERAL
    • F16J—PISTONS; CYLINDERS; SEALINGS
    • F16J15/00—Sealings
    • F16J15/44—Free-space packings
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y10—TECHNICAL SUBJECTS COVERED BY FORMER USPC
    • Y10S—TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y10S277/00—Seal for a joint or juncture
    • Y10S277/931—Seal including temperature responsive feature

Definitions

  • FIG L FIG 5 GAP CONTROL APPARATUS The present invention relates to a clearance control mechanism.
  • the pneumatic sensing system and the servo mechanism have a number of moving parts each of which added to the complexity of the system and was subject to possible failure.
  • the invention proposes the use of a temperature responsive mechanism which relies on thermal expansion and contraction of a control member to maintain the desired clearance.
  • the invention is particularly applicable to maintaining substantially constant clearance between two sealing members in the hot part of a gas turbine engine where minimum weight, reliability and simplicity are important design criteria.
  • a self-adjusting clearance control mechanism for maintaining a clearance substantially constant at a desired value, comprises two members spaced apart to define the clearance, a temperature responsive control member which is arranged to expand or contract so as to act to produce relative movement between the two members in response to variations in its temperature, and means for varying the temperature of the control member when the size of the clearance changes from said desired value so that the control member acts to oppose any variation in clearance.
  • the control member may be a solid metal rod or strip, or may alternatively be in the form of a tube. Faster response times may be achieved by making the control member thin. In a preferred embodiment therefore, the member comprises a thin-walled tube.
  • the two sealing members may be parts of a seal between rotating and static parts of a gas turbine engine.
  • one member is formed on a rotor of a gas turbine engine while the other is associated with adjacent static structure.
  • Said other member may thus be an annular sealing ring, and a plurality of control members are then preferably used, said control members being equally angularly spaced apart around the sealing ring.
  • the means for heating the control member is conveniently a supply of hot gas which can be supplied to heat the control member in a direct relationship to the size of the sealing clearance.
  • tubular control member is supplied at one end with hot gas from the vicinity of the sealing members, and is also supplied with a colder gas from the other end.
  • the tube may be vented through a restriction at some region along its length to a pressure which is lower than the hot gas supply pressure and the cold gas pressure.
  • the tube may be supplied with hot gas on the inside, the outside, or both, and the cold gas may similarly be supplied to the inside, the
  • apertures may be provided which communicate between the inside and outside of the tube.
  • the temperature of the tube wall may be affected by either hot gas alone, or a mixture of hot and cold gases and hence the response time of the tube in varying its temperature can be arranged to be any desired value.
  • FIG. 1 illustrates a gas turbine engine in which the invention has been applied to a turbine seal.
  • FIG. 2 shows in more detail the turbine seal and a control member of the engine of FIG. 1.
  • FIG. 3 is a section on the line 3-3 of FIG. 2.
  • FIG. 4 illustrates the simplest form of control tube which may be used.
  • FIG. 5 illustrates an alternative form of control tube.
  • FIG. 6 illustrates a further alternative form of control tube.
  • FIG. 1 a gas turbine engine is shown in FIG. 1 having compressor means 1, combustion equipment 2, turbine means 3 and a propulsion nozzle 4 in flow series. A portion of the engine casing has been cut away in order to illustrate a turbine rotor of the engine to which the invention has been applied.
  • Part of the air compressed by the compressor means passes into a by-pass duct 6 and by-passes the combustion equipment and the turbine means.
  • FIG. 2 illustrates the turbine of the engine in greater detail.
  • the clearance which it is desired to control is that between the first rotor stage 10 of the turbine means and the adjacent upstream static structure. At this position the pressure of the hot gases emerging from the combustion chamber'12 through nozzle guide vanes 14, is very nearly the highest in the engine, so that leakage at this point has the maximum effect on the efficiency of the engine cycle.
  • Sealing ribs 16 are provided on shroud 18 of the turbine rotor blades 20.
  • Each of the ribs 16 constitute a first sealing member and is to be maintained at a substantially constant small running clearance from a second sealing member 22.
  • the clearance should be the minimum possible and in practical terms this will be of the order of 0.010 ins.
  • the member 22 forms part of a sealing ring 24 which is attached to a control tube 26, so as to be movable with the tube as the'tube expands and contracts.
  • the tube 26 is anchored at a point 28 well upstream of the turbine, so as to provide the required length of tube to give the necessary expansion and contraction to control the seal clearance over the whole range of engine running conditions.
  • the sealing ring 24 is provided with dogs 30 for centralizing it while allowing axial movement and radial thermal expansion.
  • a piston ring type of seal 32 prevents leakage of hot gas around the sealing members 16 and 22.
  • a plurality of such tubes for example six, are equiangularly spaced around the ring 24 and the following description applies to each tube.
  • a substantially radial bore 34 extending through the sealing ring 24 communicates at one end with an orifice 36, and at the other end with the inside of the control tube 26.
  • the orifice 36 is placed close to a radially extending flange 38 on the radially outer surface of the turbine blade shroud.
  • the flange 38 restricts the area of the orifice 36 so that the flow of gas into the interior of the tube is controlled by the distance between the orifice and the flange.
  • This distance bears a direct relation to the size of the clearance between the sealing members 116 and 22.
  • control tube in this example is surrounded by an intermediate tube 40 and an outer tube 42 to define flow passages 41 and 82 respectively.
  • An aperture 44 communicates between the control tube and the intermediate tube, and the intermediate tube is vented to a lower pressure via a control orifice 46 (See FIG. 3).
  • the intermediate tube is mounted from the point 28 and is slidingly supported at its other end on the center tube at 47.
  • the outer tube communicates at one end with the region surrounding the turbine rotor blades so that it is fed with hot gas which varies in pressure in the same manner as the hot gas fed to the control tube. This eliminates variation of the sealing clearance by the control system, due purely to changes in engine operating conditions.
  • a first pair of control orifices 48 control the flow of hot gas down the passage 43 and a second pair of control orifices 49 vent the outer tube to the lower pressure.
  • the outer tube communicates via radial drillings 50 with the interior of the control tube.
  • the nest of tubes is disposed outside the inner casing of the engine and in this particular example is exposed to the relatively cold bypass air flow in the bypass duct 6.
  • Hot gas from the region surrounding the turbine rotor enters the control tube 26 through the orifice 36 and the radial drilling 34. This gas is vented to the bypass duct through apertures 44, passage d] and control orifices 46. Due to the flow passing through two orifices in series the intermediate pressure, i.e. that in the control tube, can be pre-selected for a given set of pressures in the hot gas, and in the bypass duct.
  • the passage 43 between the intermediate tube and the outer tube, is fed with hot gas from the region surrounding the turbine rotor, and this gas flows through the control orifices 48 and 49 into the bypass duct.
  • the pressure between these orifices can also be initially determined for a given set of hot gas and bypass duct pressures.
  • the system is initially set up so that when the seal clearance is at the desired value the pressures between the two sets of control orifices are equal, and there is no flow in the control tube forward of orifice 444.
  • the area of the orifice 36 will decrease and flow of .hot gas into the control tube will be further restricted by the flange 38.
  • the pressure in the control tube will drop and gas from the passage 43 will pass into the control tube from the upstream end thereof. Since the outer tube 42 has been exposed to the relatively cold air in the bypass duct 6, the gas in the passage 43 will have been cooled so that the control tube will be cooled by this flow and will contract, thus opening the sealing clearance and the area of the orifice 36 to restore the desired clearance.
  • the above described system has several advantages over possible alternative arrangements. For example, by selecting the source of the flow which opposes the control tube flow to be at the same pressure as the control tube flow, variation of sealing clearance due to pressure fluctuations only can be avoided. Additionally, when either relatively cold flow or hot flow is taking place in the tube system, the flow washes both the internal and external surfaces of the control tube to heat the tube quickly and give a fast response time.
  • FIGS. 4, 5, and 6 show some of the alternative arrangements of tubes and flow paths which may be used depending on the application of the invention and the speed of response required. Each of these figures is described with reference to a gas turbine application and common .components will be given the same reference numerals.
  • FIG. 4 the simplest form of the invention is shown, which consists of a control tube 60 supplied at one end, referred to as the hot end, with hot high pressure gas which has leaked through a sealing clearance between sealing members 61 and 62.
  • the control tube is exposed to a supply of low pressure cooling air.
  • the tube 60 is anchored at the cold end and carries sealing member 61 at the hot end so that any thermal expansion or contraction of the'tube moves the sealing member 61 towards or away from the sealing member 62.
  • Sealing member 62 is a rib which is formed on a shroud 63 of a turbine blade 64.
  • An orifice 65 is provided at the hot end of the tube and its flow area is variable due to the restricting effect of an adjacent flange 66 on the shroud 63.
  • FIG. 6 shows a further variation in which the tube is washed on both its internal and external surfaces by either a hot, a cold or a mixed hot and cold flow, so that its response time is decreased.
  • the vent 70 to low pressure is controlled by a plurality of orifices 75 in the tube 60 itself, and the tube 60 is surrounded by an outer tube 76 which provides a passage 77 for flow of gas from the orifices 75 to the vent 70.
  • This type of expansion compensated sealing clearance control has a great advantage in that there are virtually no moving parts.
  • the number of tubes and vents can be arranged to provide differing degrees of hot, cold, or mixed hot and cold flows on either the inside or outside or on both sides of the tube, for varying response times.
  • the tube may be replaced by a solid rod or strip or even a bimetallic strip.
  • the leakage it is desired to prevent may be a relatively cold fluid, and a relatively hotter or colder fluid may be used to increase or decrease the temperature of the control member.
  • leakage fluid may be used to perform the heating and cooling of the control member provided their operation is dependent on the size of the sealing clearance.
  • control member may open and close ports to supply a combustible material to burners which play hot burned gases onto the control member.
  • a self-adjusting clearance control mechanism for maintaining a clearance substantially constant at a desired value, comprising two members spaced apart to define the clearance, a temperature responsive control member which is arranged to expand or contract so as to act to produce relative movement between the two members in response to variations in its temperature, and means controlled by the size of the clearance for varying the temperature of the control member when the size of the clearance changes from said desired value so that the control member acts to oppose any variation in clearance.
  • a clearance control mechanism in which the temperature responsive control member is a metal tube and means are provided for connecting the tube at one end to one of the members defining the clearance and at the other end to a static structure.
  • a clearance control mechanism in which the two members defining the clearance form parts of a seal between rotating and static parts of a gas turbine engine.
  • a clearance control mechanism in which the seal comprises a non-rotating sealing ring, and a plurality of provided control tubes are provided connected to the ring at equal angularly spaced apart intervals.
  • a clearance control mechanism in which the means for varying the temperature of the control member comprises a supply of hot control fluid from the engine, said supply varying directly with the size of the sealing clearance.
  • a clearance control mechanism according to claim 5 in which the hot control fluid is fluid which has leaked through the sealing clearance.
  • control tube is supplied with hot control fluid at one end and with a relatively colder control fluid at the other end, and a vent being provided between the ends of the tube, said vent including a restriction to flow of fluid therethrough.
  • a clearance control mechanism in which the supply of said relatively colder control fluid originates from an environment in the engine which varies in pressure in proportion to the pressure conditions of the hot control fluid.
  • a clearance control mechanism in which the engine is a by-pass gas turbine engine, air from a by-passduct being used for the supply of said relatively colder air, and the control tube is vented to the by-pass duct.
  • a clearance control mechanism in which the engine is'a by-pass engine, the control tube is disposed in the by-pass passage of the engine and is surrounded by two further concentric tubes to define two annular flow passages around the tube, the hot control fluid being supplied to the control tube, at one end, the relatively colder control fluid being obtained from a source of hot fluid adjacent the seal and being passed down the outer of said two flow passages in contact with the outer tube so as to be cooled before being fed to the other end of the control tube, and vent means including flow restrictors being provided for venting both the control tube and the outer flow passages to the by-pass passage to maintain intermediate pressures therein.
  • a clearance control mechanism in which the control tube is vented to the by pass passage via the inner of the flow passages surrounding the control tube so that at least part of the outer surface of the control tube is contacted by the high temperature control fluid.
  • a self adjusting clearance control mechanism for maintaining a clearance substantially constant at a desired value comprising two members spaced apart to define the clearance, a temperature responsive control member connected at one end to one of the members defining the clearance, and at the other end to a static structure, said control member being arranged to expand or contract so as to act to produce relative movement between the two members in response to variations in its temperature, and means controlled by the size of the clearance for varying the temperature of the control member when the size of the clearance changes from said desired value so that the control member acts to oppose any variation in clearance.

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  • Engineering & Computer Science (AREA)
  • General Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
US00147265A 1970-05-30 1971-05-26 Gap control apparatus Expired - Lifetime US3736751A (en)

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Application Number Priority Date Filing Date Title
GB2618670 1970-05-30

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US00147265A Expired - Lifetime US3736751A (en) 1970-05-30 1971-05-26 Gap control apparatus

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DE (1) DE2126270C3 (enExample)
FR (1) FR2093885A5 (enExample)
GB (1) GB1308963A (enExample)

Cited By (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE2532415A1 (de) * 1974-07-31 1976-02-19 Snecma Vorrichtung zum selbsttaetigen regeln des spiels zwischen den spitzen von rotorschaufeln einer turbine und einer gegenueberliegenden wand
US3982850A (en) * 1974-06-29 1976-09-28 Rolls-Royce (1971) Limited Matching differential thermal expansions of components in heat engines
US4069662A (en) * 1975-12-05 1978-01-24 United Technologies Corporation Clearance control for gas turbine engine
US4338061A (en) * 1980-06-26 1982-07-06 The United States Of America As Represented By The Administrator Of The National Aeronautics And Space Administration Control means for a gas turbine engine
US6146090A (en) * 1998-12-22 2000-11-14 General Electric Co. Cooling/heating augmentation during turbine startup/shutdown using a seal positioned by thermal response of turbine parts and consequent relative movement thereof
US20060159547A1 (en) * 2005-01-15 2006-07-20 Siemens Westinghouse Power Corporation Warning system for turbine component contact

Families Citing this family (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB1605255A (en) * 1975-12-02 1986-08-13 Rolls Royce Clearance control apparatus for bladed fluid flow machine
GB1581566A (en) * 1976-08-02 1980-12-17 Gen Electric Minimum clearance turbomachine shroud apparatus
GB9808656D0 (en) * 1998-04-23 1998-06-24 Rolls Royce Plc Fluid seal
GB201113165D0 (en) 2011-08-01 2011-09-14 Rolls Royce Plc A tip clearance control device

Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US1857961A (en) * 1927-12-15 1932-05-10 Westinghouse Electric & Mfg Co Bi-metal packing
US3029064A (en) * 1958-07-11 1962-04-10 Napier & Son Ltd Temperature control apparatus for turbine cases
US3146992A (en) * 1962-12-10 1964-09-01 Gen Electric Turbine shroud support structure
US3375015A (en) * 1966-09-30 1968-03-26 Judson S. Swearingen Shaft seal employing seal gas with means for indicating proper flow thereof
US3391904A (en) * 1966-11-02 1968-07-09 United Aircraft Corp Optimum response tip seal
US3623736A (en) * 1968-09-26 1971-11-30 Rolls Royce Sealing device

Patent Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US1857961A (en) * 1927-12-15 1932-05-10 Westinghouse Electric & Mfg Co Bi-metal packing
US3029064A (en) * 1958-07-11 1962-04-10 Napier & Son Ltd Temperature control apparatus for turbine cases
US3146992A (en) * 1962-12-10 1964-09-01 Gen Electric Turbine shroud support structure
US3375015A (en) * 1966-09-30 1968-03-26 Judson S. Swearingen Shaft seal employing seal gas with means for indicating proper flow thereof
US3391904A (en) * 1966-11-02 1968-07-09 United Aircraft Corp Optimum response tip seal
US3623736A (en) * 1968-09-26 1971-11-30 Rolls Royce Sealing device

Cited By (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3982850A (en) * 1974-06-29 1976-09-28 Rolls-Royce (1971) Limited Matching differential thermal expansions of components in heat engines
DE2532415A1 (de) * 1974-07-31 1976-02-19 Snecma Vorrichtung zum selbsttaetigen regeln des spiels zwischen den spitzen von rotorschaufeln einer turbine und einer gegenueberliegenden wand
US4069662A (en) * 1975-12-05 1978-01-24 United Technologies Corporation Clearance control for gas turbine engine
US4338061A (en) * 1980-06-26 1982-07-06 The United States Of America As Represented By The Administrator Of The National Aeronautics And Space Administration Control means for a gas turbine engine
US6146090A (en) * 1998-12-22 2000-11-14 General Electric Co. Cooling/heating augmentation during turbine startup/shutdown using a seal positioned by thermal response of turbine parts and consequent relative movement thereof
EP1013892A3 (en) * 1998-12-22 2002-05-08 General Electric Company Cooling/heating during turbine startup/shutdown
KR100471958B1 (ko) * 1998-12-22 2005-03-07 제너럴 일렉트릭 캄파니 터빈
US20060159547A1 (en) * 2005-01-15 2006-07-20 Siemens Westinghouse Power Corporation Warning system for turbine component contact
US7207768B2 (en) 2005-01-15 2007-04-24 Siemens Power Generation, Inc. Warning system for turbine component contact

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Publication number Publication date
DE2126270B2 (de) 1979-02-15
FR2093885A5 (enExample) 1972-01-28
GB1308963A (en) 1973-03-07
DE2126270C3 (de) 1979-10-04
DE2126270A1 (de) 1971-12-23

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