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

Abstract

A low or high pressure turbine case is machined on its outside surface to form circumferential notches. The notches coincide with the internal locations of labyrinth seals for the blades, or with “hot spots” that have been identified. A stiffener ring is shrunk with an interference fit into each notch through inducing temperature differentials between the ring and the case. The compressive circumferential force exerted by each ring prevents the low or high pressure turbine case from expanding as much as it would otherwise, thus improving blade tip clearance or counterbalancing the “hot spots”, stiffening the case, and improving case cooling. In an alternate embodiment a hydraulic nut may be used to push the ring in place and held by a locking nut. Alternatively, C-rings, or multiple segmented rings, may be coupled together by hydraulic, electrical, or other means and actuated by a controller to exert adjustable compressive circumferential force.

Description

CROSS REFERENCE TO RELATED APPLICATION

This application claims the benefit of U.S. Provisional Application No. 60/571,701, filed on May 17, 2004, titled “METHOD AND SYSTEM FOR IMPROVED BLADE TIP CLEARANCE IN A GAS TURBINE JET ENGINE.”

FIELD OF THE INVENTION

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.

BACKGROUND OF THE INVENTION

Since the development of the gas turbine jet engine, blade tip clearance within the interior of the casing has been a challenging problem. Blade tip and inter-stage sealing have taken on a prominent role in engine design since the late 1960's. This is because the clearance between the blade tips and surrounding casing tends to vary due primarily to changes in thermal and mechanical loads on the rotating and stationary structures. On today's largest land-based and aero turbine engines, the high pressure turbine case (“HPTC”) and low pressure turbine case (“LPTC”) have such large diameters that they are more susceptible to expanding excessively and becoming out-of-round, exacerbating the blade tip clearance problem.

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. As components degrade and clearance between the blade tips and the seal on the interior of the casing increase, the engine has to work harder (and therefore runs hotter) to develop the same thrust. Once an engine reaches its EGT limit, which is an indication that the high pressure turbine disk is reaching its upper temperature limit, the engine must be taken down for maintenance. Maintenance costs for major overhauls of today's large commercial gas turbine jet engines can easily exceed one million dollars.

BRIEF DESCRIPTION OF THE DRAWINGS

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 9C 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.

DETAILED DESCRIPTION OF THE INVENTION

Referring now to the Figures, in which like reference numerals and names refer to structurally and/or functionally similar elements thereof, FIG. 1 shows a schematic diagram of the overall structure of a typical gas turbine jet engine. Referring now to FIG. 1, 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. Referring now to FIG. 2, 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. One skilled in the art will recognize that many more blades and stators would normally be present within Low Pressure Turbine Case 204. Only one 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.

Wear mechanisms for 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. Studies have shown that improved blade tip clearances in the high pressure and low pressure turbine cases result in significant life cycle cost (“LCC”) reductions.

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. 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.

There is tremendous benefit in narrowing Blade Tip Clearance 212 during the cruise condition. This is where the greatest reduction in SFC can be gained (longest part of the flight profile). On the other hand, it is greatly desirable to avoid rubbing. Minimal clearance must be maintained at takeoff to ensure thrust generation as well as keeping EGT below its established limit. Hence, it has been the goal of many control systems to attempt to maintain a minimal Blade Tip Clearance 212 while avoiding rubbing over the entire flight profile.

Engine temperatures play a huge role in determining the operational Blade Tip Clearances 212. Gas turbine performance, efficiency, and life are directly influenced by Blade Tip Clearances 212. Tighter Blade Tip Clearances 212 reduce air leakage over the tips of Blades 208. This increases turbine efficiency and permits the engine to meet performance and thrust goals with less fuel burn and lower rotor inlet temperatures. Because the turbine runs at lower temperatures, while producing the same work, hot section components would have increased cycle life. The increased cycle life of hot section components increases engine service life (TOW) by increasing the time between overhauls.

Engine SFC and EGT are directly related to HPTC blade tip clearance. One study has shown that for every 0.001 inch increase in HPTC blade tip clearance, SFC increases approximately 0.1%, while EGT increases one ° C. Therefore, a 0.010 inch HPTC blade tip clearance decrease would roughly produce a one % decrease in SFC and a ten ° C. decrease in EGT. Military engines generally show slightly greater HPTC blade tip clearance influence on SFC and EGT due to their higher operating speeds and temperatures over large commercial engines. Improvements of this magnitude would produce huge savings in annual fuel and engine maintenance costs amounting to over hundreds of millions of dollars per year.

Reducing fuel consumption also reduces aero engine total emissions. Recent estimates indicate that Americans alone now fly 764 million trips per year (2.85 airline trips per person). The energy used by commercial aircraft has nearly doubled over the last three decades. The increased fuel consumption accounts for thirteen % of the total transportation sector emissions of carbon dioxide (CO₂). Modern aero engine emissions are made up of over seventy-one % CO₂ with about twenty-eight % water (H₂O) and 0.3% nitrogen oxide (NO₂) along with trace amounts of carbon monoxide (CO), sulfur dioxide (SO₂), etc. Air transport accounts for 2.5% (600 million tons) of the world's CO₂ Production. Emissions from land-based engines, primarily for power generation, contributes amounts in addition to these totals. Clearly a reduction in fuel burn will significantly reduce aero and land-based engine emissions.

Current large commercial engines have cycle lives (defined as the time between overhauls) that vary significantly, ranging typically between 3,000 to 10,000 cycles. The cycle life is primarily determined by how long the engine retains a positive EGT margin. New engines or newly overhauled engines are shipped with a certain cold build blade tip clearance which increases with time. As the engine operating clearances increase, the engine must work harder (hotter) to produce the same work and is therefore less efficient. This increase in operating temperature, particularly takeoff EGT, further promotes the degradation of hot section components due to thermal fatigue. Retaining engine takeoff EGT margin by maintaining tight blade tip clearances can dramatically increase engine cycle life. This could also lead to huge savings in engine maintenance over a period of years due to the large overhaul costs.

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, on the other hand, 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.

All of the approaches described above have significant problems associated with them. Some are quite expensive, others achieve little results, especially during cruise where the greatest advantages are gained, or require actuation through the case due to the lack of current high temperature actuator capabilities, which raise secondary sealing issues and added weight and mechanical complexity. However, none of the approaches heretofore attempted matches the effectiveness of the present invention.

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. Referring now to FIG. 3, 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. In addition to locations corresponding to one or more of the 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. For existing engines, 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 (shown in cross section) 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. Referring now to FIG. 4, 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. Even though 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. 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. At ambient temperature, due to the smaller inside diameter of Stiffener Ring 304 to the outside diameter of Notch 302, a shrink with an interference fit results, with compressive circumferential force being applied to Low Pressure Turbine Case 204 by Stiffener Ring 304, and tensile circumferential force is applied to Stiffener Ring 304 by Low Pressure Turbine Case 204.

In one example, 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. Alternatively, 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. Or, 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.

In addition, 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. Alternatively, 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.

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. Also, 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. Referring now to FIG. 5, 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. Even though 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. When 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. At ambient temperature, due to the smaller inside diameter of Stiffener Ring 504 to the outside diameter of Notch 502, a shrink with an interference fit results, with compressive circumferential force being applied to Low Pressure Turbine Case 204 by Stiffener Ring 504, and tensile circumferential force is applied to Stiffener Ring 504 by Low Pressure Turbine Case 204.

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. Referring now to FIG. 6, for aero applications, where added weight to the engine is a concern, 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. In addition, based on the engine to be designed or to be retrofitted, 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. Referring now to FIG. 7, for aero applications, where added weight to the engine is a concern, 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. In addition, based on the engine to be designed or to be retrofitted, 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.

One skilled in the art will recognize that, in addition to the reverse taper and chevron designs for the notch and stiffener ring as shown in FIGS. 4-7, various other designs may be utilized to accomplish the same goals. For example, 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. Or, 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. Referring now to FIG. 8, 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. However, due to the compressive force exerted by Stiffener Ring 304 on Low Pressure Turbine Case 204, 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′.

Thus, 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. As stated above, 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. In general, for large gas turbine engines, 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 CO₂. 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 9C 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. Referring now to FIG. 9A, 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. Next, 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.

In FIG. 9B, 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.

In FIG. 9C, Hydraulic Nut 902 has been removed, and Locking Nut 908 has been threadably attached in its place onto Low Pressure Turbine Case 204. Retainer 910 of Locking Nut 908 engages with Stiffener Ring 904, thus securing Stiffener Ring 904 in place. This process is repeated for as many stages as required based upon turbine design. This embodiment of the invention may add excessive weight and would most likely be best suited for land based applications where weight is not of such concern.

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. Referring now to FIG. 10, Low Pressure Turbine Case 1000 has Stiffener C-Rings 1004 positioned at predetermined locations to coincide with blade/labyrinth seals and/or “hot spots”. In this embodiment of the invention, 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”.

In an alternate embodiment, instead of a c-ring, a chain-like multiple segmented ring may be coupled together by Actuator Means 1002. In another embodiment of the invention, 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.

Having described the present invention, it will be understood by those skilled in the art that many and widely differing embodiments and applications of the invention will suggest themselves without departing from the scope of the present invention.

Claims

1. A method for improved blade tip clearance in a gas turbine jet engine, the method comprising the steps of:

(a) machining at least one stiffener ring to fit without pressure near a predetermined location on an outer surface of a turbine case of the gas turbine jet engine;

(b) mounting a hydraulic nut on said turbine case adjacent to said at least one stiffener ring;

(c) engaging a piston of said hydraulic nut with said at least one stiffener ring;

(d) pushing said at least one stiffener ring by extending said piston of said hydraulic nut in a direction toward a larger diameter end of said turbine case resulting in an interference fit of said at least one stiffener ring into said predetermined location;

(e) replacing said hydraulic nut with a locking nut, said locking nut having a retainer; and

(f) engaging said retainer of said locking nut with said at least one stiffener ring, wherein said at least one stiffener ring is secured in said predetermined location in said interference fit;

wherein said at least one stiffener ring dissipates heat, applies compressive circumferential force to said turbine case, and prevents said turbine case form going out-of-round during cruise operation of the gas turbine jet engine, improving blade tip clearance.

2. The method according to claim 1 wherein pushing step (d) further comprises the step of:

extending said piston of said hydraulic nut by an amount calculated to produce a desired said compressive circumferential force on said turbine case by said at least one stiffener ring at said predetermined location.

3. The method according to claim 1 wherein machining step (a) further comprises the step of:

machining said at least one stiffener ring from a nickel-base super alloy.

4. The method according to claim 1 wherein machining step (a) further comprises the step of:

machining said at least one stiffener ring from a material that is different from a material of said turbine case, said material having a lower coefficient of expansion than said material of said turbine case.

5. The method according to claim 1 wherein machining step (a) further comprises the step of:

machining said at least one stiffener ring to fit at an outer surface of said turbine case at a location coinciding with a labyrinth seal on an inner surface of said turbine case.

6. The method according to claim 1 wherein machining step (a) further comprises the step of:

machining said at least one stiffener ring to fit at an outer surface of said turbine case at a location coinciding with a hot spot of said turbine case.

7. An apparatus for improving blade tip clearance in a gas turbine jet engine, the apparatus comprising:

at least one stiffener ring machined to fit without pressure near a predetermined location on an outer surface of a turbine case of the gas turbine jet engine;

a hydraulic nut mounted on said turbine case adjacent to said at least one stiffener ring;

a piston of said hydraulic nut which engages with said at least one stiffener ring and pushes said at least one stiffener ring by extending from said hydraulic nut in a direction toward a larger diameter end of said turbine case resulting in an interference fit of said at least one stiffener ring into said predetermined location;

a locking nut for replacing said hydraulic nut; and

a retainer of said locking nut which engages with said at least one stiffener ring and secures said at least one stiffener ring in said predetermined location in said interference fit;

wherein said at least one stiffener ring dissipates heat, applies compressive circumferential force to said turbine case, and prevents said turbine case form going out-of-round during cruise operation of the gas turbine jet engine, improving blade tip clearance.

8. The apparatus according to claim 7 wherein said piston of said hydraulic nut is extended by an amount calculated to produce a desired said compressive circumferential force on said turbine case by said at least one stiffener ring at said predetermined location.

9. The apparatus according to claim 7 wherein said at least one stiffener ring is machined from a nickel-base super alloy.

10. The apparatus according to claim 7 wherein said at least one stiffener ring is machined from a material that is different from a material of said turbine case, said material having a lower coefficient of expansion than said material of said turbine case.

11. The apparatus according to claim 7 wherein said predetermined location for machining said at least one notch circumferentially into said outer surface of said turbine case is at a location coinciding with a labyrinth seal on an inner surface of said turbine case.

12. The apparatus according to claim 7 wherein said predetermined location for machining said at least one notch circumferentially into said outer surface of said turbine case is at a location coinciding with a hot spot of said turbine case.

13. A method for improved blade tip clearance in a gas turbine jet engine, the method comprising the steps of:

(a) machining at least one notch circumferentially at a predetermined location into an outer surface of a turbine case of the gas turbine jet engine;

(b) coating an inside surface of a stiffener ring with a lubricating substance, said stiffener ring having a first end and a second end;

(c) seating said stiffener ring in each said at least one notch;

(d) linking said first end and said second end of said stiffener ring to an actuator; and

(e) actuating said actuator to pull said first and second ends of said stiffener ring together, wherein said coating facilitates slippage of said stiffener ring in said at least one notch when said actuator is actuated;

wherein said stiffener ring dissipates heat, applies compressive circumferential force to said turbine case, and prevents said turbine case form going out-of-round during cruise operation of the gas turbine jet engine, improving blade tip clearance.

14. A method for improved blade tip clearance in a gas turbine jet engine, the method comprising the steps of:

(a) machining at least one notch circumferentially at a predetermined location into an outer surface of a turbine case of the gas turbine jet engine;

(b) coating a surface of said at least one notch with a lubricating substance;

(c) seating a stiffener ring in each said at least one notch, said stiffener ring having a first end and a second end;

(d) linking said first end and said second end of said stiffener ring to an actuator; and

(e) actuating said actuator to pull said first and second ends of said stiffener ring together, wherein said coating facilitates slippage of said stiffener ring in said at least one notch when said actuator is actuated;

wherein said stiffener ring dissipates heat, applies compressive circumferential force to said turbine case, and prevents said turbine case form going out-of-round during cruise operation of the gas turbine jet engine, improving blade tip clearance.

15. An apparatus for improving blade tip clearance in a gas turbine jet engine, the apparatus comprising:

at least one notch machined circumferentially into an outer surface of a turbine case of the gas turbine jet engine at a predetermined location;

a stiffener ring seated in each said at least one notch, said stiffener ring having a first end and a second end;

a lubricating substance coated on an inside surface of said stiffener ring; and

an actuator, wherein said first and second ends are linked to said actuator and said actuator when actuated pulls said first and second ends together, said lubricating substance facilitating slippage of said stiffener ring in said notch when said actuator is actuated;

wherein said stiffener ring dissipates heat, applies compressive circumferential force to said turbine case, and prevents said turbine case form going out-of-round during cruise operation of the gas turbine jet engine, improving blade tip clearance.

16. An apparatus for improving blade tip clearance in a gas turbine jet engine, the apparatus comprising:

at least one notch machined circumferentially into an outer surface of a turbine case of the gas turbine jet engine at a predetermined location;

a stiffener ring seated in each said at least one notch, said stiffener ring having a first end and a second end;

a lubricating substance coated on a surface of said notch; and

an actuator, wherein said first and second ends are linked to said actuator and said actuator when actuated pulls said first and second ends together, said lubricating substance facilitating slippage of said stiffener ring in said notch when said actuator is actuated;

wherein said stiffener ring dissipates heat, applies compressive circumferential force to said turbine case, and prevents said turbine case form going out-of-round during cruise operation of the gas turbine jet engine, improving blade tip clearance.