Patent Publication Number: US-2017370220-A1

Title: Customized blend limit for gas turbine engine airfoils

Description:
BACKGROUND OF THE INVENTION 
     This application relates to a method and system for providing direction on blends in gas turbine engine airfoils. 
     Gas turbine engines are known and typically include a fan delivering air into a compressor where the air is compressed. The compressed air is delivered into a combustor where it is mixed with fuel and ignited. Products of this combustion pass downstream over turbine rotors, driving them to rotate. The compressor and turbine sections include a plurality of airfoils, including rotating blades and static vanes. 
     During operation, these airfoils can become damaged. There are two general ways that an airfoil can be repaired. In one, the turbine or compressor rotor is taken off the engine and repaired in a shop. This is undesirable, as there is down time and expense involved. A second way that an airfoil can be repaired is while still on the engine. For example in some engines, there are borescope openings which allow maintenance tools to enter through an engine housing. In other engines, large sections of the case may be removed to provide access to the damaged airfoil(s). If an airfoil can be repaired while still on the engine, the maintenance down time and cost is reduced. 
     One type of repair is a blend. A blend essentially removes material that encompasses the damage, and smoothes out the area of an airfoil around the location of where the damage occurred. Damage may consist of a dent, tear, crack or partially removed material. 
     Currently, there are limits on the blend sizes and locations which are set forth in guide manuals provided by the engine manufacturer to the engine maintainer. 
     If a required blend would exceed the limit, the airfoil in question must be removed from the engine. This requires removal of the engine from the aircraft, and may also require a disassembly and rebuild of the impacted module, and substantial down time and cost. 
     SUMMARY OF THE INVENTION 
     In a featured embodiment, a method of developing a suggested blend repair to an airfoil includes the steps of: (a) storing history with regard to a particular airfoil in a particular engine; (b) taking information with regard to new damage to the particular airfoil; (c) reaching an initial blend recommendation based upon step (b); (d) assessing whether the initial blend recommendation of step (c) would be appropriate to repair the new damage based upon a consideration of steps (a)-(c); and (e) reporting a final blend recommendation. 
     In another embodiment according to the previous embodiment, the history includes at least one of a number of flights of the engine, information with regard to an actual condition of the engine, and a surge history for the engine. 
     In another embodiment according to any of the previous embodiments, the final recommendation is that no blend is appropriate for the particular airfoil. 
     In another embodiment according to any of the previous embodiments, the initial blend recommendation is determined to not be suggested at step (d) if a module incorporating the component which is damaged would be subject to replacement or repair within a particular period of operation. 
     In another embodiment according to any of the previous embodiments, the final recommendation is also based upon stored history of components in the engine other than the particular airfoil. 
     In another embodiment according to any of the previous embodiments, the final blend recommendation is performed utilizing borescopic blending. 
     In another embodiment according to any of the previous embodiments, the history of prior blends in the particular airfoil is considered in the assessment of step (d). 
     In another embodiment according to any of the previous embodiments, the initial blend recommendation includes at least a recommended blend depth, flat length, and percent span. 
     In another embodiment according to any of the previous embodiments, at least one of pressure, temperature, and speed history for an engine incorporating the damaged airfoil is considered at step (d). 
     In another embodiment according to any of the previous embodiments, the final blend recommendation is transmitted to a maintenance operator. 
     In another embodiment according to any of the previous embodiments, the maintenance operator provides performed blend information back to a repair system with regard to the performance of the final blend recommendation. 
     In another embodiment according to any of the previous embodiments, the performed blend information is stored in the repair system, such that it can be considered if the particular airfoil, or a module incorporating the particular airfoil, is subject to further damage. 
     In another embodiment according to any of the previous embodiments, the final blend recommendation includes both a blend to be performed, but also instruction to use the particular airfoil in an engine subjecting the particular airfoil to less stress or to lower thrust. 
     In another featured embodiment, an airfoil repair recommendation system has a controller with a receiver, an evaluation module, and a repair module, the controller including at least a non-transitory memory and a processor. The receiver includes an input configured to receive information with regard to a particular airfoil, and an output communicatively coupled to the evaluation module. The evaluation module includes instructions configured to cause the controller to analyze the received information utilizing an algorithm and determine whether a blend may be recommended. The repair module includes instructions configured to cause the controller to develop a final recommendation with regard to the airfoil being considered and to cause the controller to transmit that final recommendation to a location of the airfoil. 
     In another embodiment according to any of the previous embodiments, each of the evaluation module and the repair module are software, and are stored in the non-transitory memory. 
     In another embodiment according to any of the previous embodiments, the receiver is configured for receiving information of a blend performed on the particular airfoil, and updating stored information with regard to the airfoil. 
     In another embodiment according to any of the previous embodiments, the algorithm is developed utilizing historic information. 
     In another embodiment according to any of the previous embodiments, the algorithm is developed utilizing analytic models. 
     In another embodiment according to any of the previous embodiments, the controller is configured to develop a recommendation at least in part upon the condition of other components in the engine. 
     In another embodiment according to any of the previous embodiments, the controller is configured to develop a final recommendation that a blend is not recommended for the airfoil. 
     These and other features may be best understood from the following drawings and specification. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1A  is a schematic view of a gas turbine engine. 
         FIG. 1B  shows an airfoil. 
         FIG. 1C  shows a repaired airfoil. 
         FIG. 2  shows the existing method of specifying blend repair allowables. 
         FIG. 3  is a schematic showing a series of information exchanges leading to a repair suggestion. 
         FIG. 4  schematically shows a repair operation. 
     
    
    
     DETAILED DESCRIPTION 
       FIG. 1A  schematically illustrates a gas turbine engine  20 . The gas turbine engine  20  is disclosed herein as a two-spool turbofan that generally incorporates a fan section  22 , a compressor section  24 , a combustor section  26  and a turbine section  28 . Alternative engines might include an augmentor section (not shown) among other systems or features. The fan section  22  drives air along a bypass flow path B in a bypass duct defined within a nacelle  15 , while the compressor section  24  drives air along a core flow path C for compression and communication into the combustor section  26  then expansion through the turbine section  28 . Although depicted as a two-spool turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to use with two-spool turbofans as the teachings may be applied to other types of turbine engines including three-spool architectures. 
     The exemplary engine  20  generally includes a low speed spool  30  and a high speed spool  32  mounted for rotation about an engine central longitudinal axis A relative to an engine static structure  36  via several bearing systems  38 . It should be understood that various bearing systems  38  at various locations may alternatively or additionally be provided, and the location of bearing systems  38  may be varied as appropriate to the application. 
     The low speed spool  30  generally includes an inner shaft  40  that interconnects a fan  42 , a first (or low) pressure compressor  44  and a first (or low) pressure turbine  46 . The inner shaft  40  is connected to the fan  42  through a speed change mechanism, which in exemplary gas turbine engine  20  is illustrated as a geared architecture  48  to drive the fan  42  at a lower speed than the low speed spool  30 . The high speed spool  32  includes an outer shaft  50  that interconnects a second (or high) pressure compressor  52  and a second (or high) pressure turbine  54 . A combustor  56  is arranged in exemplary gas turbine  20  between the high pressure compressor  52  and the high pressure turbine  54 . A mid-turbine frame  57  of the engine static structure  36  is arranged generally between the high pressure turbine  54  and the low pressure turbine  46 . The mid-turbine frame  57  further supports bearing systems  38  in the turbine section  28 . The inner shaft  40  and the outer shaft  50  are concentric and rotate via bearing systems  38  about the engine central longitudinal axis A which is collinear with their longitudinal axes. 
     The core airflow is compressed by the low pressure compressor  44  and high pressure compressor  52 , mixed and burned with fuel in the combustor  56 , then expanded over the high pressure turbine  54  and low pressure turbine  46 . The mid-turbine frame  57  includes airfoils  59  which are in the core airflow path C. The turbines  46 ,  54  rotationally drive the respective low speed spool  30  and high speed spool  32  in response to the expansion. It will be appreciated that each of the positions of the fan section  22 , compressor section  24 , combustor section  26 , turbine section  28 , and fan drive gear system  48  may be varied. For example, gear system  48  may be located aft of combustor section  26  or even aft of turbine section  28 , and fan section  22  may be positioned forward or aft of the location of gear system  48 . 
     The engine  20  in one example is a high-bypass geared aircraft engine. In a further example, the engine  20  bypass ratio is greater than about six (6), with an example embodiment being greater than about ten (10), the geared architecture  48  is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3 and the low pressure turbine  46  has a pressure ratio that is greater than about five. In one disclosed embodiment, the engine  20  bypass ratio is greater than about ten (10:1), the fan diameter is significantly larger than that of the low pressure compressor  44 , and the low pressure turbine  46  has a pressure ratio that is greater than about five 5:1. Low pressure turbine  46  pressure ratio is pressure measured prior to inlet of low pressure turbine  46  as related to the pressure at the outlet of the low pressure turbine  46  prior to an exhaust nozzle. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present disclosure is applicable to other gas turbine engines including direct drive turbofans. 
     A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The fan section  22  of the engine  20  is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet (10,668 meters). The flight condition of 0.8 Mach and 35,000 ft (10,668 meters), with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption (‘TSFCT’)”—is the industry standard parameter of lbm of fuel being burned divided by lbf of thrust the engine produces at that minimum point. “Low fan pressure ratio” is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.45. “Low corrected fan tip speed” is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram ° R)/(518.7° R)] 0.5 . The “Low corrected fan tip speed” as disclosed herein according to one non-limiting embodiment is less than about 1150 ft/second (350.5 meters/second). 
       FIG. 1B  shows a turbine or compressor blade  80 . As shown, a crack  82  has formed in an airfoil  84 . While the crack is shown in a blade, it should be understood that static vanes also have airfoils and they may also receive a crack. 
     One current method of repairing the crack is to form a blend, or to remove a portion of material that includes the crack. Thus, as shown in  FIG. 1C , there is a blend formed to remove the crack  82 . There are three parameters to the blend including a length L which is a flat length between two curved portions C at each radial end of the flat portion  86 . There is also a blend depth D and a percentage span S which is the location along the span of the airfoil measured from some reference distance that is receiving the blend. In addition, an aspect ratio (its length-to-depth ratio) of the blend is also considered. 
     As can be appreciated, at some point a component may have received too many blends and, thus, further blends would not be advised. In addition, in some cases, the damage, such as a crack, may be so extreme that a blend would not be indicated. 
     Thus,  FIG. 2  shows a known blade manual  90  which sets forth the limits of a particular blend and, in particular, areas on the blade. Note, in practice, the limits are likely not all the same. As can be appreciated, these limits are selected to be conservative and to err on the side of removing a part rather than performing a blend. As can be appreciated, there may thus be circumstances in the current method where a part could receive an additional blend, yet such an operation would not take place. 
     The present disclosure replaces the manual  90 . Stated simply, a method and system under this disclosure would initially develop a required blend. Then, the system will perform the step of evaluating whether the initial recommended blend would actually be advisable taking into account the history of not only the component to receive the blend, but perhaps other components on the associated engine. A final recommendation is then reached and transmitted to the location of the component. 
     As part of this, statistical surrogate models are developed and utilized to determine optimal blend geometries on a case by case basis. The statistical surrogate models are developed from analytical engineering analyses and/or a library of acceptable conditions to repair and center on the parameters D, S and L as shown in  FIG. 1C . The use of statistical surrogate models enables a system to rapidly customize a blend solution for the engine that allows the engine to be returned to service quickly and safely. In fact, some engine have ports that allow access to the compressor such that the blending can be done without removing the engine from the aircraft. The acceptable parameters that can be accommodated on a given blade and a given compressor are calculated given the history of the compressor to be repaired. Additionally, multiple blend types can be considered including spherical, flat bottom, and tip blends. Note while a compressor is mentioned, this disclosure also provides benefits to turbine sections. 
     The surrogate models assess the blended airfoils high cycle fatigue capabilities and an operability model will determine the blends impact to a compressor&#39;s stall margin. The software may then output optimized geometric parameters, blend depth, aspect ratio, flat length, and percent span to produce an acceptable blend. 
     By utilizing this disclosed method, an aircraft operator will be allowed to provide an increased number of blends and continue engine operation. Alternatively, in an overhaul shop environment, a blend could be used to eliminate further work such as replacement of an integrally bladed rotor. This will reduce the cost of further work that may be unnecessary. Alternatively, some stages may be provided with a blend and a more heavily damaged stage could be replaced, thereby salvaging at least part of the compressor section. 
     A service operator in the field will measure the damage, such as its depth, its span-wise location, and damage type, as examples. The operator inputs these parameters to be sent to the system to develop a recommended blend geometry utilizing these statistical surrogate models. 
     Each component may have separate models, but there is one integrated process for various disciplines. As an example, the advisability of an additional blend may be evaluated based upon stability, aerodynamic performance, vibratory stress and mistuning assessments. Further, this determination can be based on knowledge of conditions of other stages within the particular compressor section which is to receive the blend, and including blends that have been perhaps performed on components in that compressor section in the past. 
     As shown in  FIG. 3 , a maintenance individual at the engine will then get an answer in a very short period of time utilizing the statistical surrogates. 
     A maintenance operator, such as an airline, submits information at  100  to the repair system. The information  100  may include an engine serial number, a module serial number, the number of flights to this point for the engine, and any particular components to be considered. In addition, various pressure and temperature information recorded in the engine&#39;s electronic control (or other recording device, or streamed off of the aircraft in some other manner) may be provided. Also, total engine cycles may be provided which can be used in an assessment of the engine&#39;s overall health, time to next overhaul, and compressor efficiency and stall margin in the engine&#39;s current state where a blend is being considered. 
     This pressure and temperature data can be broadly considered “conditions” for purposes of this disclosure. As an example, if the temperature downstream of the last compressor stage is different than expected, such as the temperature being higher than expected with the pressure being lower than expected, this may imply that a component within the compressor has significant wear. Further, if the engine has a lower than expected exhaust gas temperature margin, or a turbine temperature margin is not as expected, then this might point to a decision not to suggest further blends. 
     In an alternative method, an assessment of the physical erosion of a particular component may be made utilizing representative photos. In the repair system, this information is compared to historic experience in the surrogate model. 
     From this initial communication, an engine history can be developed at the repair system. If the number of existing blends are already very high, it may be desirable not to repair the blend on the engine. Further, if other components or modules of the engine are reflecting a major repair being required soon, a blend may not be recommended. 
     The repair system then develops preliminary stability assessment based upon hours and cycles in flight and also experience. As an example, the amount of surge margin in the particular engine&#39;s series and the amount of erosion for a particular airline operator may be considered. Also, the historic number of blends in the component to be repaired, as well as the overall compressor may be considered. 
     A preliminary stability assessment is then made which may indicate that a blend may not be recommended and that an overhaul or redo may be suggested. This may be communicated back to the maintenance operator at line  102 . 
     The maintenance operator also provides information at  104  to the repair system including damage measurements of any airfoil to be repaired. Pictures of the new damage or scalable pictures of the location of previous blend repairs may be provided. This is input into an emulator algorithm  105  to assess the stability margin if the new damage is repaired by a blend. 
     The analysis is based upon an emulator algorithm and the emulator algorithm may be a regression of at least two analysis points of different blend geometries, locations or sizes. In one embodiment, the analysis is complemented by at least one of a stability and a performance assessment. 
     Then, a blend recommendation is sent back at  106 . The emulator will produce the suggested blends quickly and will also check the aggregate, for example, the impact of proposed blends on compressor exit temperature for turbine cooling, speed margins versus redline speeds for expected ambient and expected thrust requirements, and for other useful EGT targets. 
     As mentioned elsewhere, the final recommendation may be that a blend should not be performed. Also, the final recommendation may be that a blend could be performed if the component, or perhaps its engine, is utilized in a less stressful application. As one example, the component could be moved to an engine operating at a lower thrust level. 
     Again, the algorithms may be developed utilizing analytical models, electronic engine models populated with discrete sensor data, and/or historic data from real world components which have been evaluated and stored in the emulator  105 . The analytic models may be physics based, empirical, or semi-empirical. 
     The recommended blend repair may be developed based upon the measured damage to at least one airfoil using a structural and vibratory algorithm that includes a structural assessment of the blade geometry of the blade post blending based upon the dimensions of the damage to the airfoil and material remaining after repair. 
     The stability or performance may be assessed for at least one of a compressor section stability margin, performance change based upon a suggested blend, performance level, exhaust gas temperature margin relative to EGT limits, or a rotor speed margin to its redline limit. These assessments are based on at least one measured engine parameter, such as exhaust gas temperature margin relative to a maximum exhaust gas temperature. The limits will reflect the effect that a proposed blend repair will have on the engine or compressor. 
     In another embodiment, the stability or performance assessment may propose a blend based on at least one of the number of flights on a particular module to be prepared, a physical examination of hardware, including the shape of an outer air seal rub, or the leading edge condition of a blade relative to an uneroded portion. 
     In another embodiment, the assessment may be based on at least one of a normalized and measured compressor parameter, such as compressor exit temperature at a flight condition normalized to an ambient temperature, or a corrected speed, or a ratio of temperatures and pressure at a constant speed. In a further alternative, a simple comparison of fleet data for these parameters at a particular flight condition may be assessed to support an approximate assessment that judges a compressor as being suitable or unsuitable for additional blending based upon fleet surge, history, and experience. 
     In another embodiment, the assessment may be based on at least one normalized compressor parameter within fleet data and a quantitative assessment of the compressor as being suitable or unsuitable relative to a history of similar engines to support a decision to incorporate a proposed blend. Blending of the airfoil changes the overall incidence of air. At high incidence, the flow begins to separate. The foil no longer is able to sustain the pressure rise required of it, and the foil stalls. In a turbine engine, since the foils can no longer support the pressure rise, the higher pressure downstream of the stalled foil or row rapidly propagates forward in the compressor and the complete compressor no longer functions. The engine fails to operate, as a stalled compressor results in an engine that generates insufficient thrust. Thus, a blend should not be implemented if there is a risk of this occurring. 
     Further, past repaired damage is considered when considering a feature future recommended blend. 
     Further, the algorithm may include an assessment of erosion of the blade, stage or module to assess the stability characteristics of the blended component. 
     As can be appreciated, more than one blend may be suggested, or not suggested, and provided to a maintenance location as a combined recommendation. 
     The assessment may also include assessment of margins relative to at least one operational limit after the repairs are completed. Further, the assessment includes structural, stability and performance assessments after the blends are completed. 
     This disclosure could be summarized as a method of suggesting appropriate blends to an airfoil, wherein the suggested blend is considered for structural adequacy based on at least one of system frequencies, mode shapes, mistuning characteristics, vibratory stresses, or material capability from a variety of engine conditions. Further, this analysis of a blend recommendation may be customized for a specific blade whereby the customer receives an acceptable blend in minutes. 
     The maintenance location may then make the blend repairs. As shown for example, in  FIG. 4 , an integrally bladed rotor  121  is shown with a blade  122  having a crack  124 . An engine housing  120  surrounds the rotor  121 , which is part of a compressor section. A blending tool  126  is shown schematically entering into a bore in the housing  120  such that a blend may be provided, as suggested by the above method, while the rotor is still on the engine and aircraft. 
     Returning to  FIG. 3 , at  108 , the maintenance system then reports what was actually done with regard to the recommended blends back to the repair system. The repair system now updates and stores at storage  110  a summary of the actual blends as performed. 
     The repair may actually be made through a borescope port in an engine housing. Alternatively, in some conditions, the repair may occur after removing a portion of a compressor casing to permit access to the component to be repaired. 
     The system as suggested provides a powerful maintenance tool and also facilitates further optional steps. 
     The communication between the maintenance operator and the repair system may be computer-based, wireless, or any other method of communication. 
     A method of developing a suggested blend repair to an airfoil could be summarized as including the steps of (a) storing history with regard to a particular airfoil in a particular engine; (b) taking information with regard to new damage to the particular airfoil; (c) reaching an initial blend recommendation based upon step (b); (d) assessing whether the initial blend recommendation of step (c) would be appropriate to repair the new damage based upon a consideration of steps (a)-(c); and (e) reporting a final recommended blend. 
     A worker of ordinary skill in the art would recognize the appropriate computer technology and transmission technologies to provide repair system algorithms as shown in  FIG. 3 . 
     As an example, an airfoil repair recommendation system may include a controller having a receiver, an evaluation module, and a repair module. The controller includes at least a non-transitory memory and a processor. The receiver may include an input configured to receive information with regard to a particular airfoil, and an output communicatively coupled to the evaluation module. The evaluation module may include instructions configured to cause the controller to analyze the received information utilizing algorithms and determine whether a blend may be recommended. The repair module may include instructions configured to cause the controller to develop a recommendation with regard to the airfoil being considered and to cause the controller to transmit that information to a location of the airfoil. 
     Although an embodiment of this invention has been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of this invention. For that reason, the following claims should be studied to determine the true scope and content of this invention.