Patent Publication Number: US-2006013681-A1

Title: Turbine case reinforcement in a gas turbine jet engine

Description:
CROSS REFERENCE TO RELATED APPLICATIONS  
      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.” 
      A nonprovisional U.S. application entitled “METHOD AND SYSTEM FOR IMPROVED BLADE TIP CLEARANCE IN A GAS TURBINE JET ENGINE” is being filed concurrently by L. James Cardarella, John Usherwood and Andres Del Campo, wherein the contributions by John Usherwood and Andres Del Campo have been assigned to Carlton Forge Works, a California corporation. 
    
    
     BACKGROUND  
      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&#39;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&#39;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&#39;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 present description.  
       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 present description.  
       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 present description.  
       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 present description.  
       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 present description.  
       FIG. 8  shows an improvement in clearance under load in an embodiment of the present description.  
       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 present description.  
       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 present description.  
       FIG. 11  shows a schematic cross-sectional diagram of a low pressure turbine case having stiffener rings. 
    
    
     DETAILED DESCRIPTION OF THE DRAWINGS  
      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 and rotates on an axis of rotation along centerline  202 . 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 retard air from escaping around the tips of Blades  208 . It is appreciated that seals performing similar functions are often referred to by other names. 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 can 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 can be tremendous benefit in narrowing Blade Tip Clearance  212  during the cruise condition. This is often where the greatest reduction in SFC can be gained (longest part of the flight profile). On the other hand, rubbing is generally to be avoided. Minimal clearance typically is 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 generally play a large 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  can reduce air leakage over the tips of Blades  208 . This can increase turbine efficiency and permit the engine to meet performance and thrust goals with less fuel bum and lower rotor inlet temperatures. Because the turbine runs at lower temperatures, while producing the same work, hot section components can have increased cycle life. The increased cycle life of hot section components can increase engine service life (TOW) by increasing the time between overhauls.  
      Engine SFC and EGT are generally 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, it is believed that a 0.010 inch HPTC blade tip clearance decrease may 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 may produce large savings in annual fuel and engine maintenance costs amounting to over hundreds of millions of dollars per year.  
      Reducing fuel consumption may also reduce 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 (CO 2 ). Modem aero engine emissions are made up of over seventy-one % CO 2  with about twenty-eight % water (H 2 O) and 0.3% nitrogen oxide (NO 2 ) along with trace amounts of carbon monoxide (CO), sulfur dioxide (SO 2 ), etc. Air transport accounts for 2.5% (600 million tons) of the world&#39;s CO 2  Production. Emissions from land-based engines, primarily for power generation, contributes amounts in addition to these totals. Clearly a reduction in fuel bum can 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 generally works harder (hotter) to produce the same work and is therefore less efficient. This increase in operating temperature, particularly takeoff EGT, can further promote the degradation of hot section components due to thermal fatigue. It is believed that 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 often 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&#39;s and early 1980&#39;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.  
      It is believed that 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.  
       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 present description.  FIG. 11  shows a cross-sectional schematic diagram of the low pressure turbine case of  FIG. 2  fitted with stiffener rings in an embodiment of the present description. Referring now to  FIGS. 3, 11 , one or more features of the present description may be applied to existing gas turbine jet engines, or may be incorporated into the design and build of new gas turbine jet engines, for a variety of applications including aviation, marine and land-based engines. Features of the present description are applicable to the HPTC as well as the LPTC, and the description and figures in relation to the LPTC also apply equally to the HPTC and are 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. It is appreciated that the stiffener rings may be located at other positions of a turbine case, depending upon the particular application It is further appreciated that sizes, dimensions, shapes, materials and clearances may vary, depending upon the particular application.  
      In one embodiment, Stiffener Rings  304  (shown in cross section in  FIG. 3 ) are shrink interference fit into each Notch  302  so that the Stiffener ring  304  encircles the circumferential Notch  302  as shown in  FIG. 11 . Since Low Pressure Turbine Case  204  is conical in shape, each Stiffener Ring  304  may have a different diameter. In each case, the inside diameter of each Stiffener Ring  304  may 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 present description. Referring now to  FIG. 4 , Notch  302  is manufactured circumferentially with a reverse taper relative to the taper of the low pressure turbine case in one embodiment. 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  may be 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 can add extra security so that Stiffener Ring  304  is inhibited from slipping axially on Low Pressure Turbine Case  204 . If Notch  302  was manufactured flat without the taper, there may be an increased possibility of slippage in some applications. 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 diameter of the inner surface of the Stiffener Ring  304  to the diameter of the outer surface of the Notch  302 , a shrink with an interference fit results, with radially 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 embodiment, the radially compressive forces may be centered on the axis of rotation defined by center line  202  as schematically shown by arrows in  FIG. 11 . In one embodiment, the radially compressive forces are applied continuously around the entire circumference of the Notch  302  and the Turbine Case  204  without interruption.  
      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. In one embodiment, the Stiffener Ring  304  may be fabricated as a solid, unitary or one-piece, continuous or seamless member forged or machined in a closed loop shape. In another embodiment, the Stiffener Ring  304  may be fabricated using an open loop-shaped member and bonding the ends together by welding, for example, to form a closed loop shape. 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. It is appreciated that the stiffener ring and the turbine case may be made of a variety of materials, depending upon the particular application. 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. Alternatively, 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 radially 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 movement of Stiffener Ring  304  within Notch  302 , including axial or rotational movement. The plurality of grooves on Stiffener Ring  304 , which may be made of a nickel-base super alloy for example, may be 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 can dent into or form an indentation in the softer titanium, steel, or aluminum grooves. Alternatively, Stiffener Ring  304  may 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 or otherwise moving in relation to Notch  302 . Machining in cross directions may not be needed in this case.  
      By thus positioning Stiffener Rings  304  in the manner described, Blade Tip Clearance  212  may be improved in some applications, especially during cruise operation of the engine in some applications. An engine designer may as a result, design the engine to have a reduced blade tip clearance than may otherwise be appropriate for a given engine design absent such stiffener rings. It is also appreciated that other or different benefits, advantages, improvements or other features may be utilized alone or in combination, depending upon the particular application. In one application, the radially compressive circumferential force (represented by arrows in  FIG. 11 ) applied by the Stiffener Rings  304  can prevent Low Pressure Turbine Case  204  from expanding due to heat as much as it would otherwise expand. In one aspect, the Stiffener Rings  304  function as a girdle for the Turbine Case  204 , to inhibit expansion or going out of round and otherwise reinforce the Turbine Case  204 . 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 thermal expansion, which would increase the radially compressive circumferential force applied over that of a stiffener ring of the same material as the case as the temperature rises. The compressive forces may be sufficient to form an indentation in the turbine case such as in the Notch  302 .  
      In many engine designs, heat is mainly dissipated from the outside surface area of Low Pressure Turbine Case  204  by convection. Another benefit which may be achieved by adding Stiffener Rings  304  to Low Pressure Turbine Case  204  is that heat may be dissipated at a greater rate because Stiffener Rings  304  can act as cooling fins, which can result in cooler operating temperatures within Low Pressure Turbine Case  204 . This cooling may also contribute to less expansion and smaller Blade Tip Clearance  212 . Also, Stiffener Rings  304  can help to maintain roundness of Low Pressure Turbine Case  204 . Again, it is appreciated that other or different benefits, advantages, improvements or other features may be utilized alone or in combination, depending upon the particular application.  
       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 present description. Referring now to  FIG. 5 , Notch  502  is machined circumferentially with a chevron shape in one embodiment. 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 can add extra security to inhibit the Stiffener Ring  304  from slipping off of Low Pressure Turbine Case  204 . 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 radially 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 present description. 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 radially 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 present description. 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 radially compressive and tensile stress requirements. In addition to aero or aviation applications, it is appreciated that flush embodiments as well as other embodiments may be utilized in land-based and marine applications as well.  
      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 or similar or different 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. Alternatively, the notch and stiffener ring may have an inverted chevron shape. In other embodiments, a notch may not be utilized. Many other such shapes may be envisioned without departing from the scope of the present description.  
       FIG. 8  shows the improvement in blade tip clearance under load in an embodiment of the present description. 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 radially 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, in one aspect of the present description, the amount of expansion that would normally occur due to heating in the LPTC and the HPTC, is reduced, and consequently blade tip clearance may be improved. As stated above, increased blade tip clearance can accelerate the effects of low cycle fatigue and erosion due to increased temperatures in the HPTC and LPTC, and degrade EGT margin and engine life. In general, for large gas turbine engines, it is believed that blade tip clearance reductions on the order of 0.010 inch can produce decreases in SFC of one % and EGT of ten ° C. It is believed that improved blade tip clearance of this magnitude can produce fuel and maintenance savings of over hundreds of millions of dollars per year. Reduced fuel bum can also reduce aircraft emissions, which currently account for thirteen % of the total U.S. transportation sector emissions of CO 2 . In another aspect, blade tip clearances can be reduced at cruise condition to make a significant impact on SFC and EGT margin and improve turbine efficiency. Moreover, the increased outer surface area of the HPTC and LPTC due to the stiffener rings can, in certain embodiments, increase cooling and result in lower internal temperatures which can lengthen the cycle life of the engine. In yet another aspect, an increase in payload per engine may be achieved due to the improvement in blade tip clearance. Additional pounds of freight may be transported per takeoff and landing. It is further appreciated that features of the present description could readily replace expensive passive clearance control options. It is appreciated that reductions in one or more of out-of-roundness, blade tip clearance, SFC, EGT or polluting emissions may be achieved utilizing one or more features herein described. For example, fabricating a stiffener ring from a material having a lower coefficient of thermal expansion than that of the turbine case material, may facilitate achieving one or more of these or other reductions. Similarly, it is appreciated that one or more of these reductions or other benefits may be achieved fabricating a turbine case and stiffener ring of the same material.  
       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. 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 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 present description. 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, Stiffener C-Rings  1004  are not shrink interference fit onto Low Pressure Turbine Case  1000 . A notch for each Stiffener C-Ring  1004  may still be 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 including radially 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 provide a suitable benefit such as maintaining an optimum blade tip clearance or counterbalancing a “hot spot”, for example.  
      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, 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 various features, it will be understood by those skilled in the art that many and widely differing embodiments and applications will suggest themselves without departing from the scope of the present description.