US20050252000A1 - Method and system for improved blade tip clearance in a gas turbine jet engine - Google Patents

Method and system for improved blade tip clearance in a gas turbine jet engine Download PDF

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Publication number
US20050252000A1
US20050252000A1 US11/128,959 US12895905A US2005252000A1 US 20050252000 A1 US20050252000 A1 US 20050252000A1 US 12895905 A US12895905 A US 12895905A US 2005252000 A1 US2005252000 A1 US 2005252000A1
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Prior art keywords
stiffener ring
turbine case
notch
blade tip
jet engine
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Abandoned
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US11/128,959
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English (en)
Inventor
Louis Cardarella
John Usherwood
Andres Campo
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Carlton Forge Works
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Carlton Forge Works
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Priority to US11/128,959 priority Critical patent/US20050252000A1/en
Priority to PCT/US2005/017299 priority patent/WO2006043987A2/fr
Assigned to CARLTON FORGE WORKS reassignment CARLTON FORGE WORKS ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: CAMPO, ANDRES DEL, USHERWOOD, JOHN
Publication of US20050252000A1 publication Critical patent/US20050252000A1/en
Assigned to CARLTON FORGE WORKS reassignment CARLTON FORGE WORKS ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: CARDARELLA, L. JAMES
Abandoned legal-status Critical Current

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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
    • 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
    • F01D21/04—Shutting-down of machines or engines, e.g. in emergency; Regulating, controlling, or safety means not otherwise provided for responsive to undesired position of rotor relative to stator or to breaking-off of a part of the rotor, e.g. indicating such position
    • F01D21/045—Shutting-down of machines or engines, e.g. in emergency; Regulating, controlling, or safety means not otherwise provided for responsive to undesired position of rotor relative to stator or to breaking-off of a part of the rotor, e.g. indicating such position special arrangements in stators or in rotors dealing with breaking-off of part of rotor
    • 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
    • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F01D25/08—Cooling; Heating; Heat-insulation
    • F01D25/12—Cooling
    • 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
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/60—Efficient propulsion technologies, e.g. for aircraft
    • 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
    • Y10T—TECHNICAL SUBJECTS COVERED BY FORMER US CLASSIFICATION
    • Y10T29/00—Metal working
    • Y10T29/49—Method of mechanical manufacture
    • Y10T29/49316—Impeller making
    • Y10T29/49318—Repairing or disassembling
    • 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
    • Y10T—TECHNICAL SUBJECTS COVERED BY FORMER US CLASSIFICATION
    • Y10T29/00—Metal working
    • Y10T29/49—Method of mechanical manufacture
    • Y10T29/49316—Impeller making
    • Y10T29/4932—Turbomachine making
    • Y10T29/49321—Assembling individual fluid flow interacting members, e.g., blades, vanes, buckets, on rotary support member
    • 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
    • Y10T—TECHNICAL SUBJECTS COVERED BY FORMER US CLASSIFICATION
    • Y10T29/00—Metal working
    • Y10T29/49—Method of mechanical manufacture
    • Y10T29/49718—Repairing
    • Y10T29/49721—Repairing with disassembling
    • Y10T29/4973—Replacing of defective part
    • 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
    • Y10T—TECHNICAL SUBJECTS COVERED BY FORMER US CLASSIFICATION
    • Y10T29/00—Metal working
    • Y10T29/53—Means to assemble or disassemble

Definitions

  • This invention relates to gas turbine jet engines, and more particularly to the high pressure turbine case and low pressure turbine case of gas turbine jet engines, and even more particularly to improving the clearance of the blade tips within the interior of the high and low pressure turbine cases, to stiffening the high and low pressure turbine cases, and to cooling the high and low pressure turbine cases.
  • Reduced clearance in both the HPTC and the LPTC can provide dramatic reductions in specific fuel consumption (“SFC”), compressor stall margin and engine efficiency, as well as increased payload and mission range capabilities for aero engines. Improved clearance management can dramatically improve engine service life for land-based engines and time-on-wing (“TOW”) for aero engines.
  • Deterioration of exhaust gas temperature (“EGT”) margin is the primary reason for aircraft engine removal from service.
  • the Federal Aviation Administration (“FAA”) certifies every aircraft engine with a certain EGT limit. EGT is used to indicate how well the HPTC is performing. Specifically, EGT is used to estimate the disk temperature within the HPTC.
  • FIG. 1 shows a schematic diagram of the overall structure of a typical gas turbine jet engine.
  • FIG. 2 shows a sectional schematic diagram of a low pressure turbine case of a typical gas turbine jet engine.
  • FIG. 3 shows a sectional schematic diagram of the low pressure turbine case of FIG. 2 fitted with stiffener rings in an embodiment of the method and system for improved blade tip clearance, case deformation, and cooling of the present invention.
  • FIG. 4 shows a sectional schematic diagram of Section A of the low pressure turbine case of FIG. 3 , showing the stiffener ring about to be seated in an embodiment of the method and system for improved blade tip clearance, case deformation, and cooling of the present invention.
  • FIG. 5 shows a sectional schematic diagram of a section of a low pressure turbine case showing the stiffener ring about to be seated in another embodiment of the method and system for improved blade tip clearance, case deformation, and cooling of the present invention.
  • FIG. 6 shows a sectional schematic diagram of a section of a low pressure turbine case showing the stiffener ring seated in another embodiment of the method and system for improved blade tip clearance, case deformation, and cooling of the present invention.
  • FIG. 7 shows a sectional schematic diagram of a section of a low pressure turbine case showing the stiffener ring seated in another embodiment of the method and system for improved blade tip clearance, case deformation, and cooling of the present invention.
  • FIG. 8 shows the improvement in clearance under load in an embodiment of the method and system for improved clearance of the present invention.
  • FIGS. 9A, 9B , and 9 C show sectional schematic diagrams of a section of a low pressure turbine case having the stiffener ring positioned on the low pressure turbine case with a hydraulic nut and secured with a locking nut in another embodiment of the method and system for improved blade tip clearance, case deformation, and cooling of the present invention.
  • FIG. 10 shows a schematic diagram of a low pressure turbine case having stiffener rings actuated by hydraulic, electric, or other means in another embodiment of the method and system for improved clearance of the present invention.
  • FIG. 1 shows a schematic diagram of the overall structure of a typical gas turbine jet engine.
  • Gas Turbine Jet Engine 100 has Fan 102 for air intake within Fan Frame 104 .
  • High Pressure Compressor Rotor 106 and its attached blades and stators force air into Combustor 108 , increasing the pressure and temperature of the inlet air.
  • High Pressure Turbine Rotor 110 and its accompanying blades and stators are housed within High Pressure Turbine Case 112 .
  • Low Pressure Turbine Rotor 114 and its accompanying blades and stators are housed within Low Pressure Turbine Case 116 .
  • the turbine extracts the energy from the high-pressure, high-velocity gas flowing from Combustor 108 and is transferred to Low Pressure Turbine Shaft 118 .
  • FIG. 2 shows a sectional schematic diagram of a low pressure turbine case of a typical gas turbine jet engine.
  • Centerline 202 runs through the center of Low Pressure Turbine Case 204 (shown in cross-section).
  • Rotor 206 (shown in cross-section) has Blade 208 attached thereto.
  • Blade 208 is shown for simplicity.
  • Labyrinth seal designs vary by application. Sometimes the labyrinth seals are located on the blade tips, and sometimes they are located on the inside diameter of the cases as shown in FIG. 2 .
  • Labyrinth Seals 210 (shown in cross-section) line the inside diameter of Low Pressure Turbine Case 204 forming a shroud around each rotating Blade 208 , limiting the air that spills over the tips of Blades 208 .
  • the shape of Labyrinth Seals 210 is designed to create air turbulence between the tips of each Blade 208 and the corresponding Labyrinth Seal 210 . The air turbulence acts as a barrier to prevent air from escaping around the tips of Blades 208 .
  • Blade Tip Clearance 212 defined as the distance between the tip of Blade 208 and Labyrinth Seal 210 , will vary over the operating points of the engine.
  • the mechanisms behind Blade Tip Clearance 212 variations come from the displacement or distortion of both static and rotating components of the engine due to a number of loads on these components and expansion due to heat.
  • Axis-symmetric clearance changes are due to uniform loading (centrifugal, thermal, internal pressure) on the stationary or rotating structures that create uniform radial displacement. Centrifugal and thermal loads are responsible for the largest radial variations in Blade Tip Clearance 212 .
  • Labyrinth Seal 210 can be generally categorized into three major categories: rubbing (blade incursion), thermal fatigue, and erosion.
  • Engine build clearances in both high pressure and low pressure turbine cases are chosen to limit the amount of blade rubbing.
  • LCC life cycle cost
  • Blade Tip Clearance 212 As a cold engine is started, a certain amount of Blade Tip Clearance 212 exists between each Labyrinth Seal 210 and the tip of Blades 208 . Blade Tip Clearance 212 is rapidly diminished as the engine speed is increased for takeoff due to the centrifugal load on Rotor 206 as well as the rapid heating of Blades 208 , causing the rotating components to grow radially outward. Meanwhile, Low Pressure Turbine Case 204 expands due to heating but at a slower rate. This phenomenon can produce a minimum Blade Tip Clearance 212 “pinch point.” As Low Pressure Turbine Case 204 expands due to heating after the pinch point, Blade Tip Clearance 212 increases.
  • Rotor 206 Shortly after Low Pressure Turbine Case 204 expansion, Rotor 206 begins to heat up (at a slower rate than Low Pressure Turbine Case 204 due to its mass) and Blade Tip Clearance 212 narrows. As the engine approaches the cruise condition, Low Pressure Turbine Case 204 and Rotor 206 reach thermal equilibrium and Blade Tip Clearance 212 remains relatively constant.
  • Gas turbine performance, efficiency, and life are directly influenced by Blade Tip Clearances 212 .
  • TOW engine service life
  • Previous attempts at blade tip clearance management can generally be categorized by two control schemes, active clearance control (“ACC”) and passive clearance control (“PCC”).
  • PCC is defined as any system that sets the desired clearance at one operating point, namely the most severe transient condition (e.g., takeoff, re-burst, maneuver, etc.).
  • ACC is defined as any system that allows independent setting of a desired blade tip clearance at more than one operating point.
  • the problem with PCC systems is that the minimum clearance, the pinch point, that the system must accommodate leaves an undesired larger clearance during the much longer, steady state portion of the flight (i.e., cruise).
  • Typical PCC systems include better matching of rotor and stator growth throughout the flight profile, the use of abradables to limit blade tip wear, the use of stiffer materials and machining techniques to limit or create distortion of static components to maintain or improve shroud roundness at extreme conditions, and the like.
  • Engine manufacturers began using thermal ACC systems in the late 1970's and early 1980's. These systems utilized fan air to cool the support flanges of the HPTC, reducing the case and shroud diameters, and hence blade tip clearance, during cruise conditions.
  • FIG. 3 shows a sectional schematic diagram of the low pressure turbine case of FIG. 2 fitted with stiffener rings in an embodiment of the method and system for improved blade tip clearance, case deformation, and cooling of the present invention.
  • the method and system of the present invention may be applied to existing gas turbine jet engines, or may be incorporated into the design and build of new gas turbine jet engines.
  • the method and system of the present invention is applicable to the HPTC as well as the LPTC, and the description of the invention and figures in relation to the LPTC also apply equally to the HPTC and is not limited to the LPTC.
  • Notches 302 which may be of several different geometries as described in detail below, are manufactured circumferentially, typically through machining, into the outside diameter of Low Pressure Turbine Case 204 to coincide with one or more locations of the Labyrinth Seals 210 .
  • notches may be machined circumferentially in locations corresponding to “hot spots” that have been identified in Low Pressure Turbine Case 204 through computer modeling, through monitoring surface temperatures, or through visual inspections for cracks when the engine is overhauled.
  • Low Pressure Turbine Case 204 is typically removed in order to repair cracks resulting from the these “hot spots”. After such repairs, groves may then be applied through a weld repair through machining. The external rings would then be shrink interference fit in the grooves.
  • Stiffener Rings 304 are then shrink interference fit into each Notch 302 . Since Low Pressure Turbine Case 204 is conical in shape, each Stiffener Ring 304 will have a different diameter. In each case, the inside diameter of each Stiffener Ring 304 will be slightly less than the outside diameter of its corresponding Notch 302 . Each Stiffener Ring 304 is heated, starting with the largest diameter Stiffener Ring 304 . Heating causes each Stiffener Ring 304 to expand, increasing the inside diameter to a diameter that is greater than the outside diameter of its corresponding Notch 302 . Once positioned in Notch 302 , Stiffener Ring 304 is allowed to cool, which shrinks with an interference fit into its corresponding Notch 302 .
  • FIG. 4 shows a sectional schematic diagram of Section A of the low pressure turbine case of FIG. 3 , showing the stiffener ring about to be seated in an embodiment of the method and system for improved blade tip clearance, case deformation, and cooling of the present invention.
  • Notch 302 is manufactured circumferentially with a reverse taper in one embodiment of the invention. Angle 402 for the taper will vary from case to case, ranging from just greater than 0° for a cylindrical case to an appropriate degree that would depend upon the specific geometry of a conical case.
  • Stiffener Ring 304 is machined circumferentially on its inside diameter to match this same taper.
  • Stiffener Ring 304 is shrink interference fit onto Low Pressure Turbine Case 204 , the taper adds extra security so that Stiffener Ring 304 will not slip axially on Low Pressure Turbine Case 204 , which could possibly happen if Notch 302 was manufactured flat without the taper.
  • Stiffener Ring 304 When Stiffener Ring 304 has been heated it expands, giving rise to Ring Clearance 404 , enabling Stiffener Ring 304 to be positioned as shown against Heel 406 of Notch 302 . As Stiffener Ring 304 cools, it shrinks in diameter and seats itself circumferentially into Notch 302 .
  • Low Pressure Turbine Case 204 may be fifty inches in outside diameter at the portion where Blade 208 and Labyrinth Seal 210 are located.
  • Low Pressure Turbine Case 204 is made of nickel-based super alloy, such as Inconel 718 , as is Stiffener Ring 304 through a forging process.
  • Super alloy Inconel 718 is a high-strength, complex alloy that resists high temperatures and severe mechanical stress while exhibiting high surface stability, and is often used in gas turbine jet engines. Heating Stiffener Ring 304 to a calculated temperature will cause Stiffener Ring 304 to expand, yielding an appropriate Ring Clearance 404 when Low Pressure Turbine Case 204 is at ambient air temperature of approximately seventy ° F.
  • Low Pressure Turbine Case 204 may be cooled with liquid nitrogen or other means to a calculated temperature to cause Low Pressure Turbine Case 204 to shrink in diameter, yielding an appropriate Ring Clearance 404 when Stiffener Ring 304 is at ambient air temperature of approximately seventy ° F.
  • an appropriate Ring Clearance 404 may be achieved through a combination of cooling Low Pressure Turbine Case 204 and heating Stiffener Ring 304 , each to various calculated temperatures. Increasing or decreasing the inside diameter of Stiffener Ring 304 will result in more or less compressive circumferential force and tensile stress as required for a particular application, and within the stress limits of the material that Stiffener Ring 304 is made from.
  • the machining for Low Pressure Turbine Case 204 may be done in a first direction, such as radially, and the machining for Stiffener Ring 304 may be done in a second direction, such as axially, which is more or less perpendicular to the first direction. Since machining leaves a spiral, or record, continuous groove on the machined surfaces, the grooves on each surface will align in a cross-hatch manner to each other, increasing the frictional forces between the two surfaces and reducing the potential for spinning of Stiffener Ring 304 within Notch 302 .
  • the plurality of grooves on Stiffener Ring 304 which is typically made of a nickel-base super alloy, are harder than the plurality of grooves on Notch 302 of Low Pressure Turbine Case 204 , which is typically made of titanium, or in other low pressure turbine casings, possibly steel or aluminum.
  • the nickel-base super alloy grooves will dent into the softer titanium, steel, or aluminum grooves.
  • Stiffener Ring 304 could simply be spot welded in one or more locations to Notch 302 , or bolted to one or more flanges secured to Notch 302 , to keep Stiffener Ring 304 from spinning in relation to Notch 302 . Machining in cross directions would not be needed in this case.
  • Stiffener Rings 304 By thus positioning Stiffener Rings 304 in the manner described, Blade Tip Clearance 212 is improved, especially during cruise operation of the engine.
  • the compressive circumferential force applied by the Stiffener Rings 304 prevent Low Pressure Turbine Case 204 from expanding due to heat as much as it would otherwise expand.
  • Stiffener Rings 304 may be made of the same material as Low Pressure Turbine Case 204 , or may be made of a different material with a lower coefficient of expansion, which would increase the compressive circumferential force applied over that of a stiffener ring of the same material as the case as the temperature rises.
  • Heat is mainly dissipated from the outside surface area of Low Pressure Turbine Case 204 by convection.
  • Another benefit to adding Stiffener Rings 304 to Low Pressure Turbine Case 204 is that heat is dissipated at a greater rate because Stiffener Rings 304 act as cooling fins, which results in cooler operating temperatures within Low Pressure Turbine Case 204 , also contributing to less expansion and smaller Blade Tip Clearance 212 .
  • Stiffener Rings 304 help to maintain roundness of Low Pressure Turbine Case 204 .
  • FIG. 5 shows a sectional schematic diagram of a section of a low pressure turbine case showing the stiffener ring about to be seated in another embodiment of the method and system for improved blade tip clearance, case deformation, and cooling of the present invention.
  • Notch 502 is machined circumferentially with a chevron shape in one embodiment of the invention. Angle 508 may vary by application.
  • Stiffener Ring 504 is machined circumferentially on its inside diameter to match this same chevron shape.
  • Stiffener Ring 504 is shrink interference fit onto Low Pressure Turbine Case 204 , the chevron shape adds extra security so that Stiffener Ring 304 will not slip off of Low Pressure Turbine Case 204 , which could possibly happen if Notch 502 was manufactured flat without the chevron shape.
  • Stiffener Ring 504 has been heated it expands, giving rise to Ring Clearance 404 , enabling Stiffener Ring 504 to be positioned as shown against Heel 506 of Notch 502 . As Stiffener Ring 504 cools, it shrinks in diameter and seats itself circumferentially into Notch 502 .
  • FIG. 6 shows a sectional schematic diagram of a section of a low pressure turbine case showing the stiffener ring seated in another embodiment of the method and system for improved blade tip clearance, case deformation, and cooling of the present invention.
  • Stiffener Ring 604 is manufactured to have a profile that, when seated as shown in FIG. 6 , is substantially flush with the outer surface of Low Pressure Turbine Case 204 .
  • Notch 302 with a reverse taper as shown in FIG. 4 is machined into Low Pressure Turbine Case 204 .
  • Notch 302 may be machined deeper, and/or wider, and Stiffener Ring 604 given added depth, and/or width, in order to meet the compressive and tensile circumferential stress requirements.
  • FIG. 7 shows a sectional schematic diagram of a section of a low pressure turbine case showing the stiffener ring seated in another embodiment of the method and system for improved blade tip clearance, case deformation, and cooling of the present invention.
  • Stiffener Ring 704 is manufactured to have a profile that, when seated as shown in FIG. 6 , is substantially flush with the outer surface of Low Pressure Turbine Case 204 .
  • Notch 502 with a chevron shape as shown in FIG. 5 is machined into Low Pressure Turbine Case 204 .
  • Notch 502 may be machined deeper and/or wider, and Stiffener Ring 704 given added depth, and/or width, in order to meet the compressive and tensile stress requirements.
  • the notch may have one or more ridges and channels, angular or undulating, that will match up with one or more channels and ridges, angular or undulating, on the inside surface of the stiffener ring.
  • the notch and stiffener ring may have an inverted chevron shape. Many other such shapes may be envisioned without departing from the scope of the present invention.
  • FIG. 8 shows the improvement in blade tip clearance under load in an embodiment of the method and system for improved clearance of the present invention.
  • Stiffener Ring 304 as shown in FIG. 4 has been shrink interference fit onto Low Pressure Turbine Case 204 , and the engine is now under load, such as during cruise operation.
  • Labyrinth Seal 210 and Low Pressure Turbine Case 204 with Inner Surface 802 and Outer Surface 804 are depicted with solid lines in the positions they would be in without Stiffener Ring 304 .
  • Low Pressure Turbine Case 204 would have expanded in diameter, and Labyrinth Seal 210 would have moved away from Blade 208 , giving rise to a wider Blade Tip Clearance 212 .
  • Labyrinth Seal 210 is in the position indicated in phantom as 210 ′, and Ring 304 , Inner Surface 802 and Outer Surface 804 of Low Pressure Turbine Case 204 are in the positions indicated in phantom as 304 ′, 802 ′, and 804 ′, thus reducing Blade Tip Clearance 212 ′.
  • the present invention reduces the amount of expansion that would normally occur due to heating in the LPTC and the HPTC, and consequently improving blade tip clearance.
  • increased blade tip clearance accelerates the effects of low cycle fatigue and erosion due to increased temperatures in the HPTC and LPTC, and degrades EGT margin and engine life.
  • blade tip clearance reductions on the order of 0.010 inch can produce decreases in SFC of one % and EGT of ten ° C.
  • Improved blade tip clearance of this magnitude can produce fuel and maintenance savings of over hundreds of millions of dollars per year.
  • Reduced fuel burn will also reduce aircraft emissions, which currently account for thirteen % of the total U.S. transportation sector emissions of CO 2 .
  • the present invention reduces blade tip clearances at cruise condition to make a significant impact on SFC and EGT margin and improving turbine efficiency.
  • the increased outer surface area of the HPTC and LPTC due to the stiffener rings in certain embodiments will increase cooling and result in lower internal temperatures which will lengthen the cycle life of the engine.
  • Another result of the present invention is an increase in payload per engine due to the improvement in blade tip clearance. Additional pounds of freight may be transported per takeoff and landing.
  • the present invention could easily replace more expensive passive clearance control options.
  • FIGS. 9A, 9B , and 9 C show sectional schematic diagrams of a section of a low pressure turbine case having the stiffener ring positioned on the low pressure turbine case with a hydraulic nut and secured with a locking nut in another embodiment of the method and system for improved blade tip clearance, case deformation, and cooling of the present invention.
  • Stiffener Ring 904 is sized to fit without pressure in a location near an internal Blade 208 and Labyrinth Seal 210 , or previously identified “hot spot”, and placed in position there.
  • a Hydraulic Nut 902 is threadably mounted to Low Pressure Turbine Case 204 . Hydraulic Nut 902 has Piston 906 which engages with Stiffener Ring 904 .
  • Piston 906 has extended from Hydraulic Nut 902 , pushing Stiffener Ring 904 toward the larger diameter end of Low Pressure Turbine Case 204 , thus positioning Stiffener Ring 904 in the optimum location in relation to the internal Blade 208 and Labyrinth Seal 210 and resulting in an interference fit.
  • the amount that Piston 906 is extended by Hydraulic Nut 902 is calculated to produce a desired compressive circumferential force by Stiffener Ring 904 .
  • FIG. 10 shows a schematic diagram of a low pressure turbine case having stiffener rings actuated by hydraulic, electric, or other means in another embodiment of the method and system for improved clearance of the present invention.
  • Low Pressure Turbine Case 1000 has Stiffener C-Rings 1004 positioned at predetermined locations to coincide with blade/labyrinth seals and/or “hot spots”.
  • Stiffener C-Rings 1004 are not shrink interference fit onto Low Pressure Turbine Case 1000 .
  • a notch for each Stiffener C-Ring 1004 is still machined into Low Pressure Turbine Case 1000 , but the stiffener rings are c-rings rather than continuous rings.
  • Each end of Stiffener C-Ring 1004 is linked to an Actuator Means 1002 , which when actuated, pulls each end of Stiffener C-Ring 1004 together, exerting compressive force on Low Pressure Turbine Case 1000 .
  • the inside surface of each Stiffener C-Ring 1004 , or the notch surface, or both, may be coated with Teflon® or some other lubricating substance to facilitate slippage when tightened.
  • Each Actuator Means 1002 is connected to Controller 1008 through Electrical/Electronic Connections 1006 .
  • Controller 1008 receives temperature readings from multiple temperature sensors located near each Stiffener C-Ring 1004 (not shown). It is also possible to derive the LPTC temperature from EGT temperature readings and use these readings for feedback to Controllers 1008 . As the temperatures being monitored throughout Low Pressure Turbine Case 1000 rise, Controller 1008 processes the temperature data and determines how much each of the ends of each Stiffener C-Ring 1004 need to be pulled together by each Actuator Means 1002 in order to exert the proper compressive circumferential force on Low Pressure Turbine Case 1000 to either maintain an optimum blade tip clearance or counterbalance the “hot spot”.
  • a chain-like multiple segmented ring may be coupled together by Actuator Means 1002 .
  • the stiffener rings may be made of a strip of non-metallic material, such as Kevlar®.
  • the inside surface of the Kevlar®, or the notch surface, or both may also be coated with Teflon® or some other lubricating substance to facilitate slippage when tightened.

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US11/128,959 2004-05-17 2005-05-13 Method and system for improved blade tip clearance in a gas turbine jet engine Abandoned US20050252000A1 (en)

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PCT/US2005/017299 WO2006043987A2 (fr) 2004-05-17 2005-05-17 Procede et systeme permettant d'ameliorer le jeu a l'extremite des aubes dans un moteur a reaction de turbine a gaz

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Cited By (4)

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US20060013681A1 (en) * 2004-05-17 2006-01-19 Cardarella L J Jr Turbine case reinforcement in a gas turbine jet engine
US20080199301A1 (en) * 2004-09-23 2008-08-21 Cardarella Jr L James Fan Case Reinforcement in a Gas Turbine Jet Engine
US20180142620A1 (en) * 2012-09-20 2018-05-24 United Technologies Corporation Fan drive gear system module and inlet guide vane coupling mechanism
US10358933B2 (en) 2016-09-15 2019-07-23 Rolls-Royce Plc Turbine tip clearance control method and system

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Cited By (8)

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Publication number Priority date Publication date Assignee Title
US20060013681A1 (en) * 2004-05-17 2006-01-19 Cardarella L J Jr Turbine case reinforcement in a gas turbine jet engine
US20080199301A1 (en) * 2004-09-23 2008-08-21 Cardarella Jr L James Fan Case Reinforcement in a Gas Turbine Jet Engine
US8191254B2 (en) 2004-09-23 2012-06-05 Carlton Forge Works Method and apparatus for improving fan case containment and heat resistance in a gas turbine jet engine
US8317456B2 (en) 2004-09-23 2012-11-27 Carlton Forge Works Fan case reinforcement in a gas turbine jet engine
US8454298B2 (en) 2004-09-23 2013-06-04 Carlton Forge Works Fan case reinforcement in a gas turbine jet engine
US20180142620A1 (en) * 2012-09-20 2018-05-24 United Technologies Corporation Fan drive gear system module and inlet guide vane coupling mechanism
US10767555B2 (en) * 2012-09-20 2020-09-08 Raytheon Technologies Corporation Fan drive gear system module and inlet guide vane coupling mechanism
US10358933B2 (en) 2016-09-15 2019-07-23 Rolls-Royce Plc Turbine tip clearance control method and system

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WO2006043987A3 (fr) 2006-09-14

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