EP2623724A1 - Conformal Liner for Gas Turbine Engine Fan Section - Google Patents

Conformal Liner for Gas Turbine Engine Fan Section Download PDF

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Publication number
EP2623724A1
EP2623724A1 EP13151360.8A EP13151360A EP2623724A1 EP 2623724 A1 EP2623724 A1 EP 2623724A1 EP 13151360 A EP13151360 A EP 13151360A EP 2623724 A1 EP2623724 A1 EP 2623724A1
Authority
EP
European Patent Office
Prior art keywords
fan
thermal expansion
coefficient
liner
fan case
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
EP13151360.8A
Other languages
German (de)
French (fr)
Other versions
EP2623724B1 (en
Inventor
Darin S. Lussier
Sreenivasa R. Voleti
Thomas J. Robertson
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.)
Raytheon Technologies Corp
Original Assignee
United Technologies Corp
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 United Technologies Corp filed Critical United Technologies Corp
Priority to EP14158310.4A priority Critical patent/EP2775104B1/en
Publication of EP2623724A1 publication Critical patent/EP2623724A1/en
Application granted granted Critical
Publication of EP2623724B1 publication Critical patent/EP2623724B1/en
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
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/02—Blade-carrying members, e.g. rotors
    • F01D5/08—Heating, heat-insulating or cooling means
    • 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
    • F05D2220/00—Application
    • F05D2220/30—Application in turbines
    • F05D2220/36—Application in turbines specially adapted for the fan of turbofan engines
    • 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
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/231—Preventing heat transfer

Definitions

  • This disclosure relates to a fan section for a gas turbine engine, and, in particular, a conformal liner for the fan section.
  • One type of gas turbine engine includes a core engine having compressor and turbine sections that drive a fan section.
  • the fan section includes circumferentially arranged fan blades disposed within a fan case.
  • the fan section is subject to large temperature fluctuations throughout engine operation. A minimized clearance tight seal is desired between the tips of the fan blades and the fan case throughout engine operation at the various operating temperatures.
  • One system has been proposed to accommodate thermal expansion and contraction in a fan section having composite fan blades.
  • the composite fan blades are arranged within a composite liner of generally the same material.
  • Several pins at discrete circumferential locations along the liner are used to support the liner relative to a metallic fan case and permit the fan case to expand and contract relative to the composite liner.
  • a fan section of a gas turbine engine includes a fan case structure having a first coefficient of thermal expansion.
  • a fan blade is arranged within the fan case structure and has a second coefficient thermal expansion.
  • a continuous, ring-shaped liner surrounds the fan blade and includes a third coefficient of thermal expansion that is substantially similar to the second coefficient of thermal expansion and substantially different than the first coefficient of thermal expansion.
  • a desired radial tip clearance is provided between the liner and the fan blade.
  • An elastomeric adhesive operatively connects the liner to the fan case structure. The adhesive is configured to accommodate diametrical change in the liner and maintain the desired radial tip clearance throughout various fan section operating temperatures.
  • the adhesive has a 300% elongation or greater.
  • the adhesive is silicone rubber.
  • the second coefficient of thermal expansion is greater than the first coefficient of thermal expansion by at least 10 x 10 -6 /°F (18 x 10 -6 /°C).
  • the fan structure includes a composite fan case.
  • the fan case structure includes a honeycomb structure operatively connected radially inward of and to the composite fan case.
  • the fan case structure includes a composite septum interconnecting the adhesive and the honeycomb.
  • the second and third coefficients of thermal expansion are within 1 x 10 -6 /°F (1.8 x 10 -6 /°C) of one another.
  • the fan blade and the liner are constructed from the same series aluminum alloy.
  • the desired radial tip clearance is about 0.030 inch at -65°F (0.76 m at -54°C) ambient.
  • a rub strip is supported on and radially inward of the liner between the liner and the fan blade.
  • a fan case structure includes a composite fan case structure having a first coefficient of thermal expansion.
  • a continuous, ring-shaped liner has a second coefficient of thermal expansion that is substantially different than the first coefficient of thermal expansion.
  • the second coefficient of thermal expansion is greater than the first coefficient of thermal expansion by at least 10 x 10 -6 /°F (18 x 10 -6 /°C).
  • An elastomeric adhesive operatively connects the liner to the fan case structure.
  • the adhesive has a 300% elongation or greater. The adhesive is configured to accommodate diametrical change in the liner through various operating temperatures.
  • the composite fan case structure includes a structure constructed from resin and at least one of carbon fibers and fiberglass.
  • the liner is an aluminum alloy.
  • the adhesive is silicone rubber.
  • a rub strip is supported radially inward from and by the liner.
  • the composite fan case structure includes a composite septum interconnecting the adhesive to a honeycomb structure that is supported by and radially inward from a composite fan case.
  • FIG. 1 An example gas turbine engine 10 is schematically illustrated in Figure 1 .
  • the gas turbine engine 10 includes a compressor section 12, a combustor section 14 and a turbine section 16, which are arranged within a core housing 24.
  • high pressure stages of the compressor section 12 and the turbine section 16 are mounted on a first shaft 20, which is rotatable about an axis A.
  • Low pressure stages of the compressor section 12 and turbine section 16 are mounted on a second shaft 22 which is coaxial with the first shaft 20 and rotatable about the axis A.
  • the first and second shafts 20, 22 are supported for rotation within the core housing 24.
  • a fan section 18 is arranged within a fan case structure 30, which provides a bypass flow path 28 between the fan case structure 30 and the core housing 24.
  • the first shaft 20 rotationally drives circumferentially arranged fan blades 26 that provide flow through the bypass flow path 28.
  • the fan blades 26 are constructed from an aluminum alloy. It should be understood that the configuration illustrated in Figure 1 is exemplary only, and the disclosure may be used in other configurations. Although a high bypass engine is illustrated, it should be understood that the disclosure also relates to other types of gas turbine engines, such as turbo jets.
  • the fan section 18 includes a fan case structure 30 comprising multiple components in one example.
  • a honeycomb structure 40 which may be constructed from aluminum, is supported radially inward from and on the fan case 32.
  • a septum 42 is arranged radially inward from and supported by the honeycomb structure 40.
  • the fan case structure 30 includes a composite fan case 32, which is constructed from carbon fiber and resin in one example.
  • the septum 42 is a composite structure constructed from fiberglass and resin. As can be appreciated, composite structures have relatively low coefficients of thermal expansion and are dimensionally stable throughout the various operating temperatures.
  • a continuous, ring-shaped liner 44 which is an aluminum alloy, for example, is supported by the fan case structure 30, and in the example shown, by the septum 42, using an elastomeric adhesive 46.
  • the adhesive 44 has a room temperature radial thickness 48 of 0.100 in. (2.54 mm) and greater than 300% elongation, which may be provided by a silicone rubber.
  • the liner 44 has a coefficient of thermal expansion that is substantially the same as the coefficient of thermal expansion of the fan blades 26 and substantially different than the fan case structure 30.
  • the fan blades 26 and liner 44 have coefficients of thermal expansion that are within 1 x 10 -6 /°F (1.8 x 10 -6 /°C) of one another and are constructed from the same series aluminum alloy, which may be AM54027 in one example.
  • the liner/fan blade coefficient of thermal expansion is greater than the fan case structure thermal expansion by at least 10 x 10 -6 /°F(18 x 10 -6 /°C)
  • the liner 44 includes a rub strip 36 that provides an abradable material immediately adjacent to tips 34 of the fan blades 26, providing a blade tip clearance 38. It is desirable to maintain a desired radial blade tip clearance throughout various fan section operating temperatures. In one example, a desired radial tip clearance is about 0.030 in. at -65°F (0.76 mm at -54°C) ambient, which is typically encountered during cruise altitude.
  • the elastomeric adhesive 46 is selected to accommodate changes in a diameter 50 (only radial lead line is shown in Figure 3 ) of the liner 44 as the liner 44 expand and contract during operation.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)

Abstract

A fan section (18) of a gas turbine engine (10) includes a fan case structure (30) having a first coefficient of thermal expansion. A fan blade (26) is arranged within the fan case structure (30) and has a second coefficient thermal expansion. A continuous ring-shaped liner (44) surrounds the fan blade (26) and includes a third coefficient of thermal expansion that is substantially similar to the second coefficient of thermal expansion and substantially different than the first coefficient of thermal expansion. An elastomeric adhesive (46) operatively connects the liner (44) to the fan case structure (30). The adhesive (46) is configured to accommodate diametrical change in the liner (44) and maintain a desired radial tip clearance (38) throughout various fan section operating temperatures.

Description

    BACKGROUND
  • This disclosure relates to a fan section for a gas turbine engine, and, in particular, a conformal liner for the fan section.
  • One type of gas turbine engine includes a core engine having compressor and turbine sections that drive a fan section. The fan section includes circumferentially arranged fan blades disposed within a fan case. The fan section is subject to large temperature fluctuations throughout engine operation. A minimized clearance tight seal is desired between the tips of the fan blades and the fan case throughout engine operation at the various operating temperatures.
  • One system has been proposed to accommodate thermal expansion and contraction in a fan section having composite fan blades. The composite fan blades are arranged within a composite liner of generally the same material. Several pins at discrete circumferential locations along the liner are used to support the liner relative to a metallic fan case and permit the fan case to expand and contract relative to the composite liner.
    • SUMMARY
  • A fan section of a gas turbine engine includes a fan case structure having a first coefficient of thermal expansion. A fan blade is arranged within the fan case structure and has a second coefficient thermal expansion. A continuous, ring-shaped liner surrounds the fan blade and includes a third coefficient of thermal expansion that is substantially similar to the second coefficient of thermal expansion and substantially different than the first coefficient of thermal expansion. A desired radial tip clearance is provided between the liner and the fan blade. An elastomeric adhesive operatively connects the liner to the fan case structure. The adhesive is configured to accommodate diametrical change in the liner and maintain the desired radial tip clearance throughout various fan section operating temperatures.
  • In a further embodiment of any of the above, the adhesive has a 300% elongation or greater.
  • In a further embodiment of any of the above, the adhesive is silicone rubber.
  • In a further embodiment of any of the above, the second coefficient of thermal expansion is greater than the first coefficient of thermal expansion by at least 10 x 10-6/°F (18 x 10-6/°C).
  • In a further embodiment of any of the above, the fan structure includes a composite fan case.
  • In a further embodiment of any of the above, the fan case structure includes a honeycomb structure operatively connected radially inward of and to the composite fan case.
  • In a further embodiment of any of the above, the fan case structure includes a composite septum interconnecting the adhesive and the honeycomb.
  • In a further embodiment of any of the above, the second and third coefficients of thermal expansion are within 1 x 10-6/°F (1.8 x 10-6/°C) of one another.
  • In a further embodiment of any of the above, the fan blade and the liner are constructed from the same series aluminum alloy.
  • In a further embodiment of any of the above, the desired radial tip clearance is about 0.030 inch at -65°F (0.76 m at -54°C) ambient.
  • In a further embodiment of any of the above, a rub strip is supported on and radially inward of the liner between the liner and the fan blade.
  • A fan case structure includes a composite fan case structure having a first coefficient of thermal expansion. A continuous, ring-shaped liner has a second coefficient of thermal expansion that is substantially different than the first coefficient of thermal expansion. The second coefficient of thermal expansion is greater than the first coefficient of thermal expansion by at least 10 x 10-6/°F (18 x 10-6/°C). An elastomeric adhesive operatively connects the liner to the fan case structure. The adhesive has a 300% elongation or greater. The adhesive is configured to accommodate diametrical change in the liner through various operating temperatures.
  • In a further embodiment of any of the above, the composite fan case structure includes a structure constructed from resin and at least one of carbon fibers and fiberglass. The liner is an aluminum alloy.
  • In a further embodiment of any of the above, the adhesive is silicone rubber.
  • In a further embodiment of any of the above, a rub strip is supported radially inward from and by the liner. The composite fan case structure includes a composite septum interconnecting the adhesive to a honeycomb structure that is supported by and radially inward from a composite fan case.
    • BRIEF DESCRIPTION OF THE DRAWINGS
  • The disclosure can be further understood by reference to the following detailed description when considered in connection with the accompanying drawings wherein:
    • Figure 1 is a schematic, cross-sectional side view of an example gas turbine engine.
    • Figure 2 is an enlarged, cross-sectional side view of a fan case structure in a fan section of the gas turbine engine shown in Figure 1.
    • Figure 3 is a further enlarged view of the fan case structure shown in Figure 2.
    • Figure 4 is a schematic, cross-sectional end view to the fan section.
    DETAILED DESCRIPTION
  • An example gas turbine engine 10 is schematically illustrated in Figure 1. The gas turbine engine 10 includes a compressor section 12, a combustor section 14 and a turbine section 16, which are arranged within a core housing 24. In the example illustrated, high pressure stages of the compressor section 12 and the turbine section 16 are mounted on a first shaft 20, which is rotatable about an axis A. Low pressure stages of the compressor section 12 and turbine section 16 are mounted on a second shaft 22 which is coaxial with the first shaft 20 and rotatable about the axis A. The first and second shafts 20, 22 are supported for rotation within the core housing 24.
  • A fan section 18 is arranged within a fan case structure 30, which provides a bypass flow path 28 between the fan case structure 30 and the core housing 24. In the example illustrated, the first shaft 20 rotationally drives circumferentially arranged fan blades 26 that provide flow through the bypass flow path 28. In one example, the fan blades 26 are constructed from an aluminum alloy. It should be understood that the configuration illustrated in Figure 1 is exemplary only, and the disclosure may be used in other configurations. Although a high bypass engine is illustrated, it should be understood that the disclosure also relates to other types of gas turbine engines, such as turbo jets.
  • Referring to Figures 2-4, the fan section 18 includes a fan case structure 30 comprising multiple components in one example. A honeycomb structure 40, which may be constructed from aluminum, is supported radially inward from and on the fan case 32. A septum 42 is arranged radially inward from and supported by the honeycomb structure 40.
  • In one example, the fan case structure 30 includes a composite fan case 32, which is constructed from carbon fiber and resin in one example. In one example, the septum 42 is a composite structure constructed from fiberglass and resin. As can be appreciated, composite structures have relatively low coefficients of thermal expansion and are dimensionally stable throughout the various operating temperatures.
  • A continuous, ring-shaped liner 44, which is an aluminum alloy, for example, is supported by the fan case structure 30, and in the example shown, by the septum 42, using an elastomeric adhesive 46. In one example, the adhesive 44 has a room temperature radial thickness 48 of 0.100 in. (2.54 mm) and greater than 300% elongation, which may be provided by a silicone rubber.
  • The liner 44 has a coefficient of thermal expansion that is substantially the same as the coefficient of thermal expansion of the fan blades 26 and substantially different than the fan case structure 30. In one example, the fan blades 26 and liner 44 have coefficients of thermal expansion that are within 1 x 10-6/°F (1.8 x 10-6/°C) of one another and are constructed from the same series aluminum alloy, which may be AM54027 in one example. In one example, the liner/fan blade coefficient of thermal expansion is greater than the fan case structure thermal expansion by at least 10 x 10-6/°F(18 x 10-6/°C)
  • The liner 44 includes a rub strip 36 that provides an abradable material immediately adjacent to tips 34 of the fan blades 26, providing a blade tip clearance 38. It is desirable to maintain a desired radial blade tip clearance throughout various fan section operating temperatures. In one example, a desired radial tip clearance is about 0.030 in. at -65°F (0.76 mm at -54°C) ambient, which is typically encountered during cruise altitude. Thus, the elastomeric adhesive 46 is selected to accommodate changes in a diameter 50 (only radial lead line is shown in Figure 3) of the liner 44 as the liner 44 expand and contract during operation.
  • Although an example embodiment has been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of the claims. For that reason, the following claims should be studied to determine their true scope and content.

Claims (15)

  1. A fan section (18) of a gas turbine engine (10) comprising:
    a fan case structure (20) having a first coefficient of thermal expansion;
    a fan blade (26) arranged within the fan case structure (30) and having second coefficient of thermal expansion;
    a continuous ring-shaped liner (44) surrounding the fan blade (26) and having a third coefficient of thermal expansion that is substantially similar to the second coefficient of thermal expansion and substantially different than the first coefficient of thermal expansion, and a desired radial tip clearance (38) between the liner (44) and the fan blade (27); and
    an elastomeric adhesive (46) operatively connecting the liner (44) to the fan case structure (30), the adhesive (46) configured to accommodate diametrical change in the liner (44) and maintain the desired radial tip clearance (38) throughout various fan section operating temperatures.
  2. The fan section (18) according to claim 1, wherein the adhesive (46) has a 300% elongation or greater.
  3. The fan section (18) according to claim 1 or 2, wherein the adhesive (46) is silicone rubber.
  4. The fan section (18) according to any of claims 1 to 3, wherein the second coefficient of thermal expansion is greater than the first coefficient of thermal expansion by at least 10 x 10-6/°F (18 x 10-6/°C).
  5. The fan section (18) according to any preceding claim, wherein the fan case structure (30) includes a composite fan case.
  6. The fan section (18) according to claim 5, wherein the fan case structure (30) includes a honeycomb structure (40) operatively connected radially inward of and to the composite fan case (32).
  7. The fan section (18) according to claim 6, wherein the fan case structure (30) includes a composite septum (42) interconnecting the adhesive (46) and the honeycomb structure (40).
  8. The fan section (18) according to any preceding claim, wherein the second and third coefficients of thermal expansion are within 1 x 10-6/°F (1.8 x 10-6/°C) of one another.
  9. The fan section (18) according to any preceding claim, wherein the fan blade (26) and the liner (44) are constructed from the same series aluminum alloy.
  10. The fan section (18) according to any preceding claim, wherein the desired radial tip clearance (38) is about 0.030 inch at -65°F (0.76 m at -54°C) ambient.
  11. The fan section (18) according to any preceding claim, comprising a rub strip (36) supported on and radially inward of the liner (44) between the liner (44) and the fan blade (27).
  12. A fan case structure comprising:
    a composite fan case structure (30) having a first coefficient of thermal expansion;
    an continuous ring-shaped liner (44) having a second coefficient of thermal expansion that is substantially different than the first coefficient of thermal expansion, wherein the second coefficient of thermal expansion is greater than the first coefficient of thermal expansion by at least 10 x 10-6/°F (18 x 10-6/°C) and
    an elastomeric adhesive (46) operatively connecting the liner (44) to the fan case structure (30), wherein the adhesive (46) has a 300% elongation or greater, the adhesive (46) configured to accommodate diametrical change in the liner (44) throughout various operating temperatures.
  13. The fan case structure according to claim 12, wherein the composite fan case structure (30) includes a structure constructed from resin and at least one of carbon fibers and fiberglass, and the liner (44) is an aluminum alloy.
  14. The fan case structure according to claim 12 or 13, wherein the adhesive (46) is silicone rubber.
  15. The fan case structure according to any of claims 12 to 14, wherein a rub strip (36) is supported radially inward from and by the liner (44), and the composite fan case structure (30) includes: a composite septum (42) interconnecting the adhesive (46) to a honeycomb structure (40) that is supported by and radially inward from a composite fan case (32).
EP13151360.8A 2012-02-06 2013-01-15 Conformal liner for gas turbine engine fan section Active EP2623724B1 (en)

Priority Applications (1)

Application Number Priority Date Filing Date Title
EP14158310.4A EP2775104B1 (en) 2012-02-06 2013-01-15 Conformal liner for gas turbine engine fan section

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US13/366,416 US20130202424A1 (en) 2012-02-06 2012-02-06 Conformal liner for gas turbine engine fan section

Related Child Applications (2)

Application Number Title Priority Date Filing Date
EP14158310.4A Division EP2775104B1 (en) 2012-02-06 2013-01-15 Conformal liner for gas turbine engine fan section
EP14158310.4A Division-Into EP2775104B1 (en) 2012-02-06 2013-01-15 Conformal liner for gas turbine engine fan section

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EP2623724A1 true EP2623724A1 (en) 2013-08-07
EP2623724B1 EP2623724B1 (en) 2015-05-27

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EP14158310.4A Active EP2775104B1 (en) 2012-02-06 2013-01-15 Conformal liner for gas turbine engine fan section
EP13151360.8A Active EP2623724B1 (en) 2012-02-06 2013-01-15 Conformal liner for gas turbine engine fan section

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EP (2) EP2775104B1 (en)

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DE102014215693B4 (en) 2014-08-07 2017-11-16 Technische Universität Dresden Strain-adapted engine inter-housing in composite construction and modular system for an engine intermediate housing
US10145301B2 (en) 2014-09-23 2018-12-04 Pratt & Whitney Canada Corp. Gas turbine engine inlet
US10378554B2 (en) 2014-09-23 2019-08-13 Pratt & Whitney Canada Corp. Gas turbine engine with partial inlet vane
US9957807B2 (en) 2015-04-23 2018-05-01 Pratt & Whitney Canada Corp. Rotor assembly with scoop
US9938848B2 (en) * 2015-04-23 2018-04-10 Pratt & Whitney Canada Corp. Rotor assembly with wear member
US20170234160A1 (en) * 2016-02-11 2017-08-17 General Electric Company Aircraft engine with an impact panel
FR3048999B1 (en) * 2016-03-15 2018-03-02 Safran Aircraft Engines TURBOREACTOR LOW GAME BETWEEN BLOWER AND BLOWER HOUSING
US10724540B2 (en) 2016-12-06 2020-07-28 Pratt & Whitney Canada Corp. Stator for a gas turbine engine fan
US10690146B2 (en) 2017-01-05 2020-06-23 Pratt & Whitney Canada Corp. Turbofan nacelle assembly with flow disruptor
US10677260B2 (en) * 2017-02-21 2020-06-09 General Electric Company Turbine engine and method of manufacturing
US10480530B2 (en) 2017-08-25 2019-11-19 United Technologies Corporation Fan Containment case for gas turbine engines
US11939871B1 (en) 2022-10-28 2024-03-26 Rtx Corporation Abradable material and design for jet engine applications

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EP1510657A1 (en) * 2003-08-18 2005-03-02 Snecma Moteurs Device adradable for the fan casing of a gas turbine

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Publication number Publication date
US20130202424A1 (en) 2013-08-08
EP2775104B1 (en) 2017-03-29
EP2775104A1 (en) 2014-09-10
EP2623724B1 (en) 2015-05-27

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