US20130202424A1 - 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
US20130202424A1
US20130202424A1 US13/366,416 US201213366416A US2013202424A1 US 20130202424 A1 US20130202424 A1 US 20130202424A1 US 201213366416 A US201213366416 A US 201213366416A US 2013202424 A1 US2013202424 A1 US 2013202424A1
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United States
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.)
Abandoned
Application number
US13/366,416
Inventor
Darin S. Lussier
Sreenivasa R. VOLETI
Thomas J. Robertson
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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 US13/366,416 priority Critical patent/US20130202424A1/en
Assigned to UNITED TECHNOLOGIES CORPORATION reassignment UNITED TECHNOLOGIES CORPORATION ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: Lussier, Darin S., Robertson, Thomas J., VOLETI, SREENIVASA R.
Priority to EP14158310.4A priority patent/EP2775104B1/en
Priority to EP13151360.8A priority patent/EP2623724B1/en
Publication of US20130202424A1 publication Critical patent/US20130202424A1/en
Abandoned legal-status Critical Current

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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 ⁇ 10 ⁇ 6 /° F. (18 ⁇ 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 ⁇ 10 ⁇ 6 /° F. (1.8 ⁇ 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 ⁇ 10 ⁇ 6 /° F. (18 ⁇ 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 is a schematic, cross-sectional side view of an example gas turbine engine.
  • FIG. 2 is an enlarged, cross-sectional side view of a fan case structure in a fan section of the gas turbine engine shown in FIG. 1 .
  • FIG. 3 is a further enlarged view of the fan case structure shown in FIG. 2 .
  • FIG. 4 is a schematic, cross-sectional end view to the fan section.
  • FIG. 1 An example gas turbine engine 10 is schematically illustrated in FIG. 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 FIG. 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 ⁇ 10 ⁇ 6 /° F. (1.8 ⁇ 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 ⁇ 10 ⁇ 6 /° F. (18 ⁇ 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 44 is selected to accommodate changes in a diameter 50 (only radial lead line is shown in FIG. 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 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. 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 a desired radial tip clearance 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 10×10−6/° F. (18×10−6/° 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 1×10−6/° F. (1.8×10−6/° 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 10×10−6/° F. (18×10−6/° 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:
  • FIG. 1 is a schematic, cross-sectional side view of an example gas turbine engine.
  • FIG. 2 is an enlarged, cross-sectional side view of a fan case structure in a fan section of the gas turbine engine shown in FIG. 1.
  • FIG. 3 is a further enlarged view of the fan case structure shown in FIG. 2.
  • FIG. 4 is a schematic, cross-sectional end view to the fan section.
    •  DETAILED DESCRIPTION
  • An example gas turbine engine 10 is schematically illustrated in FIG. 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 FIG. 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 FIGS. 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 1×10−6/° F. (1.8×10−6/° 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 10×10−6/° F. (18×10−6/° 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 44 is selected to accommodate changes in a diameter 50 (only radial lead line is shown in FIG. 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)

What is claimed is:
1. A fan section of a gas turbine engine comprising:
a fan case structure having a first coefficient of thermal expansion;
a fan blade arranged within the fan case structure and having second coefficient of thermal expansion;
a continuous ring-shaped liner surrounding the fan blade 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 between the liner and the fan blade; and
an elastomeric adhesive operatively connecting the liner to the fan case structure, the adhesive configured to accommodate diametrical change in the liner and maintain the desired radial tip clearance throughout various fan section operating temperatures.
2. The fan section according to claim 1, wherein the adhesive has a 300% elongation or greater.
3. The fan section according to claim 2, wherein the adhesive is silicone rubber.
4. The fan section according to claim 1, wherein the second coefficient of thermal expansion is greater than the first coefficient of thermal expansion by at least 10×10−6/° F. (18×10−6/° C.).
5. The fan section according to claim 4, wherein the fan case structure includes a composite fan case.
6. The fan section according to claim 5, wherein the fan case structure includes a honeycomb structure operatively connected radially inward of and to the composite fan case.
7. The fan section according to claim 6, wherein the fan case structure includes a composite septum interconnecting the adhesive and the honeycomb.
8. The fan section according to claim 1, wherein the second and third coefficients of thermal expansion are within 1×10−6/° F. (1.8×10−6/° C.) of one another.
9. The fan section according to claim 8, wherein the fan blade and the liner are constructed from the same series aluminum alloy.
10. The fan section according to claim 8, wherein the desired radial tip clearance is about 0.030 inch at −65° F. (0.76 m at −54° C.) ambient.
11. The fan section according to claim 10, comprising a rub strip supported on and radially inward of the liner between the liner and the fan blade.
12. A fan case structure comprising:
a composite fan case structure having a first coefficient of thermal expansion;
an continuous ring-shaped liner 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 10×10−6/° F. (18×10−6/° C.) and
an elastomeric adhesive operatively connecting the liner to the fan case structure, wherein the adhesive has a 300% elongation or greater, the adhesive configured to accommodate diametrical change in the liner throughout various operating temperatures.
13. The fan case structure according to claim 12, wherein the composite fan case structure includes a structure constructed from resin and at least one of carbon fibers and fiberglass, and the liner is an aluminum alloy.
14. The fan case structure according to claim 13, wherein the adhesive is silicone rubber.
15. The fan case structure according to claim 13, wherein a rub strip is supported radially inward from and by the liner, and 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.
US13/366,416 2012-02-06 2012-02-06 Conformal liner for gas turbine engine fan section Abandoned US20130202424A1 (en)

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US13/366,416 US20130202424A1 (en) 2012-02-06 2012-02-06 Conformal liner for gas turbine engine fan section
EP14158310.4A EP2775104B1 (en) 2012-02-06 2013-01-15 Conformal liner for gas turbine engine fan section
EP13151360.8A EP2623724B1 (en) 2012-02-06 2013-01-15 Conformal liner for gas turbine engine fan section

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Cited By (14)

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US20130302154A1 (en) * 2012-05-11 2013-11-14 Rolls-Royce Plc Casing
DE102014215693A1 (en) 2014-08-07 2016-02-11 Technische Universität Dresden Strain-adapted engine inter-housing in composite construction and modular system for an engine intermediate housing
US20160312641A1 (en) * 2015-04-23 2016-10-27 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
US9828876B2 (en) 2014-04-28 2017-11-28 Rolls-Royce Corporation Fan containment case
US9957807B2 (en) 2015-04-23 2018-05-01 Pratt & Whitney Canada Corp. Rotor assembly with scoop
US20180238346A1 (en) * 2017-02-21 2018-08-23 General Electric Company Turbine engine and method of manufacturing
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
US10480530B2 (en) 2017-08-25 2019-11-19 United Technologies Corporation Fan Containment case for gas turbine engines
US10690146B2 (en) 2017-01-05 2020-06-23 Pratt & Whitney Canada Corp. Turbofan nacelle assembly with flow disruptor
US10724540B2 (en) 2016-12-06 2020-07-28 Pratt & Whitney Canada Corp. Stator for a gas turbine engine fan
US11313325B2 (en) * 2016-03-15 2022-04-26 Safran Aircraft Engines Gas turbine engine with minimal tolerance between the fan and the fan casing
US11939871B1 (en) 2022-10-28 2024-03-26 Rtx Corporation Abradable material and design for jet engine applications

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Cited By (22)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20130302154A1 (en) * 2012-05-11 2013-11-14 Rolls-Royce Plc Casing
US9429039B2 (en) * 2012-05-11 2016-08-30 Rolls-Royce Plc Casing
US9828876B2 (en) 2014-04-28 2017-11-28 Rolls-Royce Corporation Fan containment case
DE102014215693A1 (en) 2014-08-07 2016-02-11 Technische Universität Dresden Strain-adapted engine inter-housing in composite construction and modular system for an engine intermediate housing
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
US10837361B2 (en) 2014-09-23 2020-11-17 Pratt & Whitney Canada Corp. Gas turbine engine inlet
US11118601B2 (en) 2014-09-23 2021-09-14 Pratt & Whitney Canada Corp. Gas turbine engine with partial inlet vane
US10378554B2 (en) 2014-09-23 2019-08-13 Pratt & Whitney Canada Corp. Gas turbine engine with partial inlet vane
US10145301B2 (en) 2014-09-23 2018-12-04 Pratt & Whitney Canada Corp. Gas turbine engine inlet
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
US20160312641A1 (en) * 2015-04-23 2016-10-27 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
US11313325B2 (en) * 2016-03-15 2022-04-26 Safran Aircraft Engines Gas turbine engine with minimal tolerance between the fan and the fan casing
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
WO2018156265A1 (en) * 2017-02-21 2018-08-30 General Electric Company Turbine engine and method of manufacturing
US10677260B2 (en) * 2017-02-21 2020-06-09 General Electric Company Turbine engine and method of manufacturing
CN110506150A (en) * 2017-02-21 2019-11-26 通用电气公司 Turbogenerator and manufacturing method
US20180238346A1 (en) * 2017-02-21 2018-08-23 General Electric Company Turbine engine and method of manufacturing
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US11939871B1 (en) 2022-10-28 2024-03-26 Rtx Corporation Abradable material and design for jet engine applications

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EP2623724A1 (en) 2013-08-07
EP2775104A1 (en) 2014-09-10
EP2775104B1 (en) 2017-03-29

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