Method and apparatus for real-time gap assessment using pressure measurements

Abstract

Methods and apparatus for real-time gap assessment using pressure measurements are disclosed. An exemplary method includes determining a first static pressure measurement and a second static pressure measurement at a first measurement location and a second measurement location, respectively, with respect to a blade tip gap, determining a normalized pressure measurement using the first static pressure measurement and the second static pressure measurement, generating a transition curve to correlate the normalized pressure measurement with a gap measurement, and adjusting an active clearance control of the blade tip clearance based on a comparison of the real-time in-flight pressure measurement to the transition curve.

Description

[0001] This application is a divisional application of the invention patent application with application number 202210005247.7 filed on January 4, 2022 and invention name “Method and device for real-time gap assessment using pressure measurement”. Technical Field

[0002] The present disclosure relates generally to turbine engines and, more particularly, to methods and apparatus for real-time clearance assessment using pressure measurements. Background Art

[0003] Turbine engines are some of the most widely used technologies for power generation. A gas turbine is an example of an internal combustion engine that uses a combusted air-fuel mixture to generate hot gases that spin a turbine, thereby generating power. Applications for gas turbines can be found in aircraft, trains, ships, generators, gas compressors, and pumps. For example, modern aircraft rely on a variety of gas turbine engines as part of a propulsion system to generate thrust, including turbojets, turbofans, turboprops, and afterburning turbojets. Such engines include a combustion section, a compressor section, a turbine section, and an inlet, providing high power output with high thermal efficiency.

[0004] Engine efficiency, stability, and operating temperature are significantly affected by blade tip clearance. For example, the turbine tip clearance represents the radial distance between the turbine blade tip and the turbine containment structure. Considering that the power provided by the turbine (or consumed by the compressor) depends on the airflow occurring through the blade location area, an increase in the tip clearance will result in a decrease in turbine efficiency. Therefore, the presence of the tip clearance will cause the airflow to change, affecting the expected flow path and affecting the turbine efficiency, including the potential increase in fuel consumption. Factors that cause the tip clearance to change include temperature and speed, among others. Active clearance control can be achieved using tip clearance optimization based on full authority digital engine control (FADEC). However, this optimization does not take into account the progression of blade tip losses, resulting in adjustments based on clearance measurements associated with new blade tip parameters. Therefore, real-time measurement of blade tip clearance taking into account blade tip losses would be welcome in this technology. Summary of the invention

[0005] A method and apparatus for real-time clearance assessment using pressure measurements are disclosed.

[0006] Certain examples include a method of assessing real-time blade tip clearance in a turbine engine, the method comprising determining first and second static pressure measurements at first and second measurement locations, respectively, relative to the blade tip clearance, and determining a normalized pressure measurement using the first and second static pressure measurements. The method also includes generating a conversion curve to relate the normalized pressure measurement to the clearance measurement, and adjusting active clearance control of the blade tip clearance based on a comparison of the real-time in-flight pressure measurement to the conversion curve.

[0007] Certain examples provide an apparatus for assessing real-time blade tip clearance in a turbine engine, the apparatus comprising: a pressure sensor that determines a first static pressure measurement and a second static pressure measurement at a first measurement location and a second measurement location, respectively, relative to the blade tip clearance; and a conversion curve generator that uses the first static pressure measurement and the second static pressure measurement to determine a normalized pressure measurement and generate a conversion curve to relate the normalized pressure measurement to the clearance measurement. The apparatus also includes an active clearance controller that adjusts active clearance control of the blade tip clearance based on a comparison of the real-time in-flight pressure measurement with the conversion curve.

[0008] Certain examples provide a non-transitory computer-readable medium including machine-readable instructions that, when executed, cause a processor to determine first and second static pressure measurements, respectively, at first and second measurement locations relative to blade tip clearance based on at least an input signal received as an input to the processor, determine a normalized pressure measurement using the first and second static pressure measurements, the instructions further cause the processor to generate a conversion curve to relate the normalized pressure measurement to the clearance measurement, and adjust active clearance control of the blade tip clearance based on a comparison of the real-time in-flight pressure measurement to the conversion curve. BRIEF DESCRIPTION OF THE DRAWINGS

[0009] Figure 1 is a cross-sectional schematic diagram of an exemplary high bypass turbofan gas turbine engine.

[0010] Figure 2A An exemplary single point pressure measurement at a first location is shown, illustrating airflow as the radial tip clearance increases.

[0011] Figure 2B An exemplary single point pressure measurement at a first location is shown, illustrating airflow as the radial tip clearance is reduced.

[0012] Figure 2C An exemplary single point pressure measurement at a second location is shown, illustrating airflow as the radial tip clearance increases.

[0013] Figure 2D An exemplary single point pressure measurement at a second location is shown, illustrating airflow as the radial tip clearance is reduced.

[0014] Figure 2E shows that during high power operation, based on Figures 2A-2D An exemplary conversion curve of a single point pressure measurement using gap and pressure efficiency determination.

[0015] Figure 2F shows that during low power operation, based on Figures 2A-2D An exemplary conversion curve of a single point pressure measurement using gap and pressure efficiency determination.

[0016] Figure 3A An exemplary two-point pressure measurement at a first location is shown, illustrating airflow as the radial tip clearance increases.

[0017] Figure 3B An exemplary two-point pressure measurement is shown, illustrating airflow as the radial tip clearance is reduced.

[0018] Figure 3C An exemplary two-point pressure measurement after blade tip loss has occurred is shown, illustrating the airflow as the radial tip clearance increases.

[0019] Figure 3D An exemplary two-point pressure measurement after blade tip loss has occurred is shown, illustrating the airflow as the radial tip clearance is reduced.

[0020] Figure 3E For the new blade, based on Figure 3A-3B Figure 2. Exemplary conversion curve of two-point pressure measurement using gap and pressure efficiency determination.

[0021] Figure 3F For a blade with tip losses, based on Figure 3C-3D Figure 2. Exemplary conversion curve of two-point pressure measurement using gap and pressure efficiency determination.

[0022] Figure 4A An exemplary three-point pressure measurement is shown, illustrating the airflow as the radial tip gap increases.

[0023] Figure 4B An exemplary three-point pressure measurement is shown, illustrating airflow as the radial tip clearance is reduced.

[0024] Figure 4C The results based on Figures 4A-4B An exemplary conversion curve of a three-point pressure measurement using gap and pressure efficiency determination.

[0025] Figure 4D The results show that the blades with tip losses and new blades have a good performance during low power operation. Figures 4A-4B An exemplary conversion curve of a three-point pressure measurement using gap and pressure efficiency determination.

[0026] Figure 5 An exemplary measurement of exhaust gas temperature (EGT) degradation over multiple flight cycles using a baseline measurement is shown compared to a real-time measurement achieved using the methods disclosed herein.

[0027] Fig. 6A An exemplary change in clearance with increased active clearance control is shown based on measurements of intermediate seal static pressures at the front and rear cavities.

[0028] Figure 6B An exemplary change in pressure efficiency measurement with increased active clearance control is shown based on measurements of intermediate seal static pressure at the front and rear cavities.

[0029] Figure 6C The figure shows the value obtained based on the flight data at the cruise point. Figure 6A-6B Measurements of an exemplary linear correlation between clearance and pressure efficiency.

[0030] Figure 7 is a block diagram of an exemplary implementation of a blade tip loss determiner in which examples disclosed herein may be implemented.

[0031] Figure 8 A flowchart representing exemplary machine readable instructions that may be executed to implement Figure 7 An example blade tip loss determiner.

[0032] Fig. 9 A flowchart representing exemplary machine readable instructions that may be executed to use Figure 7 An exemplary blade tip loss determiner of generates conversion curves for various power levels and / or altitudes.

[0033] Fig.10 A flowchart representing exemplary machine readable instructions that may be executed to use Figure 7 An exemplary blade tip loss determiner is provided to measure real-time blade tip losses.

[0034] Fig.11 is constructed to execute Figure 8-10 Instructions to achieve Figure 7 Block diagram of an exemplary processor platform of an exemplary blade tip loss determiner.

[0035] The drawings are not drawn to scale. On the contrary, the thickness of the layer or region can be magnified in the drawings. In general, the same reference numerals will be used throughout the drawings and the written descriptions of the combination to represent the same or similar parts. As used in this patent, it is stated that any component (e.g., layer, film, area, region, or plate, etc.) is in any way on (e.g., positioned on, located on, set on, or formed on, etc.) another component, indicating that the referenced component is in contact with another component, or indicating that the referenced component is above another component, wherein one or more intermediate components are located between them. Unless otherwise specified, connection references (e.g., attachment, connection, connection, engagement, etc.) will be interpreted widely and may include intermediate members between element sets and relative movement between elements. In this way, connection references do not necessarily infer that two elements are directly connected and are in a fixed relationship with each other. Stating that any component is "in contact" with another component means that there is no intermediate component between the two components. Although layers and regions with clear lines and boundaries are shown in the figure, some or all of these lines and / or boundaries may be idealized. In fact, boundaries and / or lines may be unobservable, mixed and / or irregular. DETAILED DESCRIPTION

[0036] During operation, turbine engines are exposed to high temperatures, high pressures, and high speeds. Active clearance control (ACC) can be used to improve engine performance, which manages the clearance between the gas turbine containment structure (e.g., casing) and the tip of the rotating blade (e.g., turbine tip clearance). For example, a turbine clearance control system uses a control valve to manage the thermal expansion of the turbine casing surrounding the engine stage, thereby controlling the tip clearance. The tip clearance is kept at a minimum to ensure maximum propulsion efficiency. For example, the temperature of the gas after combustion can exceed 1000 degrees Celsius, causing the expansion of the turbine blades and the expansion of the containment structure, increasing the tip clearance and reducing the overall turbine efficiency (e.g., increasing fuel combustion and fuel consumption). Controlling the thermal expansion and reduction of the containment structure allows the turbine tip clearance to be controlled. For example, circulating air can be used to cool and reduce the containment structure. The engine's full authority digital engine control (FADEC) engine parameters (e.g., using sensors or calculated air temperatures) use ACC via the incremental opening or closing of the control valve throughout the flight cycle, allowing the thermal expansion of the containment structure to be controlled to achieve optimal or otherwise improved blade tip clearance.

[0037] While the FADEC calculates tip clearance under operating conditions to control ACC and optimize / improve tip clearance, the FADEC does not compensate for blade tip loss progression (e.g., related to friction, oxidation, etc.). Therefore, blade tip clearance optimization associated with the FADEC is based on calculations determined using blade tip parameters associated with a newly installed blade, rather than calculations determined using real-time tip parameters that account for blade tip losses. Over time, actual tip clearance and corresponding engine efficiency calculations may not be representative of real-time parameters because these calculations are based on initial, inventory, or "ideal" measurements of design intent. The methods and apparatus disclosed herein for real-time clearance assessment using pressure measurements allow for accurate tip clearance control once blade tip losses occur.

[0038] In the following detailed description, reference is made to the accompanying drawings which form a part thereof, and in which specific examples that can be put into practice are shown by way of example. These examples are described in sufficient detail to enable those skilled in the art to practice the subject matter, and it should be understood that other examples may be utilized. Therefore, the following detailed description is provided to describe exemplary embodiments, rather than to limit the scope of the subject matter described in this disclosure. Certain features from the different aspects described below may be combined to form other new aspects of the subject matter discussed below.

[0039] "Include" and "comprising" (and all forms and states thereof) are used herein as open-ended terms. Thus, whenever a claim employs any form of "include" or "comprising" (e.g., includes, comprises, having, etc.) as a preamble or in any type of claim recitation, it should be understood that additional elements, terms, etc. may be present without exceeding the scope of the corresponding claim or recitation. As used herein, when the phrase "at least" is used as a transitional term, such as in the preamble of a claim, it is open-ended in the same manner as the terms "include" and "comprising" are open-ended. The term "and / or" when used, for example, in a form such as A, B, and / or C, refers to any combination or subset of A, B, and C, such as (1) only A, (2) only B, (3) only C, (4) A and B, (5) A and C, (6) B and C, and (7) A and B and C. As used herein in the context of describing structures, components, items, objects, and / or things, the phrase “at least one of A and B” is intended to mean embodiments that include any of the following: (1) at least one A, (2) at least one B, and (3) at least one A and at least one B. Similarly, as used herein in the context of describing structures, components, items, objects, and / or things, the phrase “at least one of A or B” is intended to mean embodiments that include any of the following: (1) at least one A, (2) at least one B, and (3) at least one A and at least one B. As used herein in the context of describing the performance or execution of processes, instructions, acts, activities, and / or steps, the phrase “at least one of A and B” is intended to mean embodiments that include any of the following: (1) at least one A, (2) at least one B, and (3) at least one A and at least one B. Similarly, as used herein in the context of describing the performance or execution of a process, instruction, action, activity, and / or step, the phrase "at least one of A or B" means an implementation that includes any of the following: (1) at least one A, (2) at least one B, and (3) at least one A and at least one B.

[0040] As used herein, singular references (e.g., "a", "an", "first", "second", etc.) do not exclude a plurality. The term "a" or "an" entity as used herein refers to one or more of that entity. The terms "a" (or "an"), "one or more" and "at least one" may be used interchangeably herein. Furthermore, although listed separately, multiple devices, elements or method actions may be implemented by, for example, a single unit or processor. In addition, although individual features may be included in different examples or claims, these features may be combined, and inclusion in different examples or claims does not mean that a combination of features is not feasible and / or advantageous.

[0041] As used herein, the terms "system," "unit," "module," "engine," "component," and the like may include hardware and / or software systems that operate to perform one or more functions. For example, a module, unit, or system may include a computer processor, controller, and / or other logic-based device that performs operations based on instructions stored on a tangible and non-transitory computer-readable storage medium (e.g., a computer memory). Alternatively, a module, unit, or system may include a hard-wired device that performs operations based on the hard-wired logic of the device. The various modules, units, engines, and / or systems shown in the figures may represent hardware that operates based on software or hard-wired instructions, software that directs hardware to perform operations, or a combination thereof.

[0042] A turbine engine, also known as a combustion turbine or gas turbine, is a type of internal combustion engine. Turbine engines are commonly used in aircraft and power generation applications. As used herein, the terms "asset", "aircraft turbine engine", "gas turbine", "land-based turbine engine" and "turbine engine" are used interchangeably. The basic operation of a turbine engine includes using a fan to draw a fresh atmospheric air flow through the front of the turbine engine. In some examples, the air flow passes through an intermediate pressure compressor or a booster compressor located between the fan and the high-pressure compressor. The booster compressor is used to supercharge or boost the air flow before the air flow enters the high-pressure compressor. The air flow can then pass through the high-pressure compressor, which further pressurizes the air flow. The high-pressure compressor includes a set of blades attached to a shaft. The blades rotate at high speed and then compress the air flow. The high-pressure compressor then feeds the compressed air flow into the combustion chamber. In some examples, the high-pressure compressor supplies the pressurized air flow at a speed of hundreds of miles per hour. In some cases, the combustion chamber includes one or more rings of fuel injectors, which inject a steady stream of fuel into the combustion chamber, where the fuel mixes with the pressurized air flow.

[0043] In the combustion chamber of the turbine engine, the fuel is ignited by an electric spark provided by an igniter, wherein, in some examples, the fuel burns at a temperature exceeding 1,000 degrees Celsius. The combustion performed generates a high-temperature, high-pressure gas flow (e.g., hot combustion gases) that passes through another set of blades of the turbine. The turbine includes an intricate array of alternating rotating and static airfoil section blades. When the hot combustion gases pass through the turbine, the hot combustion gases expand, causing the rotating blades to rotate. The rotating blades have at least two purposes. The first purpose of the rotating blades is to drive the boost compressor and / or the high-pressure compressor to draw more pressurized air into the combustion chamber. For example, in a direct drive configuration, the turbine is attached to the same shaft as the high-pressure compressor, so that the rotation of the turbine causes the high-pressure compressor to rotate. The second purpose of the rotating blades is to rotate a generator operably coupled to the turbine section to generate electricity, and / or drive a rotor, fan, or propeller. For example, the turbine can generate electricity for use in aircraft, power plants, and the like. In the example of an aircraft turbine engine, after passing through the turbine, the hot combustion gases leave the aircraft turbine engine through a nozzle located at the back of the aircraft turbine engine.

[0044] Referring now to the drawings, wherein like numerals represent like elements throughout the several views, Figure 1 is a schematic cross-sectional view of an exemplary high bypass turbofan gas turbine engine 110 (“turbofan 110”). Figure 1 As shown, turbofan 110 defines a longitudinal or axial centerline axis 112 extending therethrough for reference. Generally speaking, turbofan 110 includes a core turbine or gas turbine engine 114 disposed downstream of a fan section 116 .

[0045] The core turbine engine 114 generally includes a generally tubular casing 118 defining an annular inlet 120. The casing 118 may be formed of a single casing or multiple casings. The casing 118, in a series flow relationship, surrounds a compressor section having a supercharger or low-pressure compressor 122 ("LP compressor 122") and a high-pressure compressor 124 ("HP compressor 124"), a combustion section 126, a turbine section having a high-pressure turbine 128 ("HP turbine 128") and a low-pressure turbine 130 ("LP turbine 130"), and an exhaust section 132. A high-pressure shaft or spool 134 ("HP shaft 134") drivingly couples the HP turbine 128 and the HP compressor 124. A low-pressure shaft or spool 136 ("LP shaft 136") drivingly couples the LP turbine 130 and the LP compressor 122. The LP shaft 136 may also be coupled to a fan spool or shaft 138 of the fan section 116. In some examples, LP shaft 136 is directly coupled to fan shaft 138 (eg, a direct drive configuration). In an alternative configuration, LP shaft 136 may be coupled to fan shaft 138 via reduction gear 139 (eg, an indirect drive or geared configuration).

[0046] like Figure 1 As shown, fan section 116 includes a plurality of fan blades 140 that are coupled to and extend radially outward from fan shaft 138. An annular fan casing or nacelle 142 circumferentially surrounds fan section 116 and / or at least a portion of core turbine 114. Nacelle 142 may be supported relative to core turbine 114 by a plurality of circumferentially spaced outlet guide vanes 144. In addition, a downstream section 146 of nacelle 142 may surround an exterior of core turbine 114 to define a bypass airflow passage 148 therebetween.

[0047] like Figure 1 As shown, air 150 enters an inlet portion 152 of the turbofan 110 during operation thereof. A first portion 154 of the air 150 flows into the bypass airflow passage 148, while a second portion 156 of the air 150 flows into the inlet 120 of the LP compressor 122. One or more sequential stages of LP compressor stator blades 170 and LP compressor rotor blades 172 coupled to the LP shaft 136 progressively compress the second portion 156 of the air 150 flowing through the LP compressor 122 before flowing to the HP compressor 124. Next, one or more sequential stages of HP compressor stator blades 174 and HP compressor rotor blades 176 coupled to the HP shaft 134 further compress the second portion 156 of the air 150 flowing through the HP compressor 124. This feeds the compressed air 158 to the combustion section 126, where the remaining fuel is mixed and combusted to provide combustion gases 160.

[0048] The combustion gases 160 flow through the HP turbine 128, wherein one or more sequential stages of HP turbine stator blades 166 and HP turbine rotor blades 168 coupled to the HP shaft 134 extract a first portion of kinetic and / or thermal energy therefrom. This energy extraction supports the operation of the HP compressor 124. The combustion gases 160 then flow through the LP turbine 130, wherein one or more sequential stages of LP turbine stator blades 162 and LP turbine rotor blades 164 coupled to the LP shaft 136 extract a second portion of thermal and / or kinetic energy therefrom. This energy extraction causes the LP shaft 136 to rotate, thereby supporting the operation of the LP compressor 122 and / or rotation of the fan shaft 138. The combustion gases 160 then exit the core turbine 114 through its exhaust section 132. At Figure 1In the example of FIG. 1 , an exemplary turbine housing 157 surrounds LP turbine rotor blades 164 and / or HP turbine rotor blades 168. A turbine frame having a fairing assembly 161 is located between the HP turbine 128 and the LP turbine 130. The turbine frame 161 serves as a support structure, connects the rear bearing of the high pressure shaft to the turbine housing and forms an aerodynamic transition duct between the HP turbine 128 and the LP turbine 130. The fairing forms a flow path between the high pressure turbine and the low pressure turbine and can be formed using metal casting (e.g., nickel-based cast metal alloy, etc.).

[0049] Together with the turbofan 110, the core turbine 114 is used for similar purposes and is exposed to similar environments in land-based gas turbines, turbojet engines, where the ratio of the first portion 154 of air 150 to the second portion 156 of air 150 is less than that of turbofans and non-ducted fan engines (where the fan section 116 does not have a nacelle 142). In each of the turbofan engine, the turbojet engine, and the non-ducted engine, a speed reduction device (e.g., a speed reduction gearbox 139) may be included between any shaft and the spool. For example, the speed reduction gearbox 139 is disposed between the LP shaft 136 and the fan shaft 138 of the fan section 116.

[0050] Figure 2A An example single point pressure measurement 202 at an example first location 205 is shown, illustrating an example air flow path 215 as an example radial tip gap 225 increases. Figure 2B An exemplary single point pressure measurement 238 at an exemplary first position 205 is shown, showing an exemplary airflow path 240 when an exemplary radial tip clearance 245 is reduced. The radial distance of the tips of the fan blades 140 and / or rotor blades 164, 168 (e.g., LP turbine rotor blades 164, HP turbine rotor blades 168) from the casing 142, 157 (e.g., fan casing, turbine casing, etc.) defines the blade tip clearance 225, 245. Figure 2A The example shows that the radial tip gap 225 increases, while Figure 2B The example shows that the radial tip gap 245 is reduced, which is an active clearance control (ACC) (e.g., using the exemplary active clearance controller 705, such as in combination with Figure 7, 245). For example, ACC maintains tight clearances for engine performance but attempts to avoid the risk of rubbing between the tips of the rotor blades 164, 168 and the turbine casing 157. For example, the ACC system includes a butterfly valve with an angle change that varies the amount of cooling airflow to achieve the desired tip clearance to maintain engine efficiency. In some examples, the fan blades 140 and / or rotor blades 164, 168 grow relative to the casing 142, 157 (e.g., during aircraft takeoff, etc.), which is caused by an increase in shaft speed. By controlling the clearances 225, 245, ACC improves the overall efficiency of the engine 110 by reducing operating temperatures, allowing the engine 110 to operate with less fuel combustion. Lowering operating temperatures reduces engine degradation and increases time-on-wing while reducing maintenance costs.

[0051] The turbine housing 157 may include a containment structure (e.g., a shroud made of a superalloy-based material, etc.). In some examples, the exterior of the housing 157 containment structure (e.g., a shroud) may be constructed using a Figure 1 The bypass flow of the high pressure compressor 124 is used for cooling. Figure 2B In the example of FIG. 1 , the airflow path 215 represents the flow of combustion gases (eg, represented using the exemplary flow profiles 220 , 240 ) that causes the rotor blades 164 , 168 to expand and the casing 157 to expand. Figure 2A In the example of FIG. 2 , the combustion gas flow path 215 originates from the example leading edge 230 and flows in the direction of the example trailing edge 235 . Figure 2A In the example of FIG. 1 , the airflow based on the ACC causes the containment structure (e.g., the retaining shield) to expand, while in Figure 2B In some examples, the ACC-based airflow causes the structure to shrink. Therefore, the radial tip gaps 225, 245 between the blades 164, 168 and the containment structure 157 can be adjusted using the ACC-based airflow. For example, the ACC can allow a minimum gap to maintain thrust generation, such as during aircraft takeoff. The gap setting is critical to ensure that rubbing does not occur. For example, flight conditions that cause the blades 164, 168 to heat up and close the generated gap will cause the rotor blades 164, 168 to rub against the structure 157. Similarly, in some examples, Figure 2B The reduction in radial tip clearance 245 may be due to a reduction (eg, contraction) of housing 157. For example, when the aircraft is in cruise mode, the ACC system adjusts the closing of the clearance because the greatest reduction in specific fuel consumption (SFC) may be achieved during the longest portion of the overall flight profile.

[0052] To facilitate real-time assessment of blade tip clearance, pressure measurements may be obtained at at least one location (e.g., mid, front, rear, etc.) relative to blade tip clearance 225. Figure 2A-2B In the example of , a sample measurement is obtained at a first position 205, which is located in the middle of the tip gap 225, 245. Figure 2C-2D In the example of , a sample measurement is obtained at a second position 250, which is behind the tip gap 225, 245. In some examples, pressure measurements at multiple positions (e.g., two-point pressure measurements, three-point pressure measurements, etc.) can be made, such as in combination with Figures 3A-3D and Figures 4A-4B For example, the front, middle, and / or rear positions relative to the tip gap allow for identification of gap changes based on pressure changes. Figure 2E-2F As described, a new engine test at different power levels (e.g., high, low, etc.) may be used to determine a conversion curve, wherein any blade tip losses (e.g., fan blades 140, LP turbine rotor blades 164, HP turbine rotor blades 168, etc.) are evaluated by deviation from the determined curve. Thus, real-time signal measurement and use of the conversion method described herein allows for immediate adjustment of ACC modulation and maintenance of tight clearances even if engine degradation has begun (e.g., blade tip losses due to oxidation, friction, thermal fatigue, etc.).

[0053] exist Figure 2A and 2B In the example, at any location (for example, Figure 2A and 2B The intermediate points shown in the examples and / or Figure 2C and 2D The static pressure measurement (P S ) and total pressure measurement (P T ). For example, a single point location (e.g., middle point 205, rear point 250, etc.) can be used to simultaneously measure total (e.g., reference) pressure and static pressure. For example, total pressure (P T ) can be defined as static pressure (P S ) and dynamic pressure (P D ), where the dynamic pressure can be defined as 1 / 2*ρ*V 2 , ρ represents the gas density (e.g., kg / m 3 ), V represents the gas velocity (e.g., m / s). In some examples, the gas velocity V affects the tip clearance reading (e.g., due to changes in radial clearances 225, 245). Tip clearance changes may occur directly due to modulation based on the ACC system and / or due to degradation of the rotor blades 164, 168 (e.g., reduction in blade tips). Figure 2A and Figure 2B In the example, V is based on opening ( Figure 2A ) or close ( Figure 2B) gaps 225, 245 resulting in changes in the airflow distribution 220, 240. Therefore, the total pressure measurement (P T ) and static pressure measurement (P S ) is obtained at a first position (e.g., a reference point) 205 (e.g., a midpoint position). Figure 2A and 2B shows a pressure measurement based on a reference point at a first position 205, and Figure 2C-2D 2 represents a pressure measurement based on a reference point at a second location 250, which is located downstream of the first location 205, where the gaps 225, 245 of the ACC modulation are opened and closed respectively. Therefore, the location of the pressure measurement based on a reference point can vary and is not limited to Figures 2A-2D Therefore, when determining the relationship between pressure efficiency (Pη) and blade tip clearance, the first position 205 or the second position 250 can be used to generate Figure 2E-2F The conversion curve.

[0054] Figure 2E The results for new blades and blades with tip losses are shown based on Figure 2A-2B and / or Figure 2C-2D An exemplary conversion curve 264 is determined at an exemplary high power level 275 using an exemplary clearance 265 and an exemplary pressure efficiency 270 measurement of a single point pressure measurement of a new blade and a blade with tip losses based on Figure 2A-2B and / or Figure 2C-2D A single point pressure measurement, determined at an exemplary low power level 290 Figure 2F Example conversion curve 288. Thus, conversion curves may be determined for different power levels (e.g., high, low, etc.) and / or different altitudes (e.g., low, medium, high, etc.). Figure 2A-2B and / or Figure 2C-2D Single point pressure measurement (e.g., P T , P S The conversion curve can be determined by calculating the static pressure (P S ) and total pressure (P T ) to determine the normalized equation for pressure efficiency (Pη), where P T Used as a reference point. Using normalized pressure efficiency allows Pη output values ​​to range from 0-1 (unitless).

[0055] Based on pressure efficiency calculation, Figure 2E-2FThe conversion curves 264, 288 can be determined using specific gap 265 measurements identified at a given value of Pη. For example, under a given set of conditions, Pη can range from 0.80-0.90 (normalized values), while the gap 265 measurements can range from 0-40 mils (e.g., 1 mil corresponds to one thousandth of an inch). Over time, as the range of Pη values ​​increases, the gap 265 range can show a corresponding increase after multiple flights. Thus, conversion curves can be developed under various conditions (e.g., altitude, power level, number of flights, etc.). Figure 2F As shown in the example, Figure 2E The Pη to clearance conversion curve 288 shows a lower slope of the curve 280 for the new engine at low power 290, as compared to the conversion curve 264 for the new engine, corresponding to the lower overall temperature to which the system is exposed at low power 290 (compared to the system at high power 275), which translates into an overall lower clearance 265 value. In addition to the conversion curves obtained during new engine testing, such curves may also be obtained for degraded engine conditions (e.g., an engine with rotor blade 164, 168 tip losses). Figure 2E-2F As shown in the example of , the conversion curve of the degraded engine 285 is lower than the conversion curve of the new engine 280. Figure 3C-3D As described in more detail, the transfer curve for the degraded engine 285 may be determined based on blade tip losses measured using the methods described herein.

[0056] Figure 3A Shows Figures 2A-2D The exemplary two-point pressure measurements 305 at the first location 205 and the second location 308 illustrate the airflow 215 as the radial tip gap 225 increases. Figure 3B An exemplary two-point pressure measurement 310 is shown, showing the airflow 215 as the radial tip gap 245 is reduced. Figure 2A-2B In contrast, the two-point pressure measurements 205, 308 allow for more than one measurement to be used in the determination of the pressure efficiency (Pη) calculation. While the single-point pressure measurement of FIG. 2 relies on a static pressure measurement and a total pressure measurement, the multi-point pressure measurement (e.g., two-point and / or three-point pressure measurement) relies on a static pressure measurement (e.g., using a static pressure sensor at one or more locations). Thus, the normalized pressure efficiency (Pη) can be calculated according to Equation 1:

[0057] Pη=(P high –P S _ local ) / (P high –P low )(Equation 1)

[0058] In the example of Equation 1, Phigh represents the maximum pressure obtained in the system (e.g., the burner pressure upstream), P low represents the lowest pressure obtained in the system during the measurement (e.g., rear pressure measurement), while P S _ local Indicates the local static pressure measurement (P S ). In some examples, Equation 2 can be used to determine the normalized pressure efficiency (Pη) based on the positioning of the static pressure measurement sensor:

[0059] Pη=(P high –P S_forward ) / (P high –P S_aft )(Equation 2).

[0060] In the example of Equation 2, P high represents the maximum pressure obtained in the system (e.g., the burner pressure upstream), P S_aft represents the static pressure measured downstream of the airflow 215 represented by the flow profile 220 (eg, a rear pressure measurement), and P S_forward Represents the static pressure measured upstream of the airflow 215 represented by the flow distribution 220. Since a single-point pressure measurement requires measuring both the total pressure and the static pressure, the pressure sensor used for such a measurement may need to be able to withstand the harsh environment in a turbine engine (e.g., a high-pressure turbine, etc.). Therefore, as described in conjunction with the methods disclosed herein, a single-point pressure measurement system may require a pressure sensor with a higher tolerance level, which is different from multiple pressure measurements (e.g., two-point and / or three-point measurements of Figures 3 and 4) using traditional static pressure sensors. The methods and apparatus disclosed herein allow real-time clearance assessment using such traditional pressure sensors without the need for more advanced sensor designs. In some examples, in connection with single-point pressure measurements as shown in Figure 2, optical sensors can generate highly accurate measurements and have a long operating life, but may require cooling flows or material improvements to withstand the 1,000 degree Celsius temperature limit based on the turbine. For example, an optical sensor (e.g., an intensity-based optical pressure sensor, etc.) can have a long service life (e.g., more than 20,000 flight hours), can capture pressures of 150 pounds per square inch (psi)-1,000 psi, can withstand temperatures of approximately 1,000 degrees Celsius, and does not require active cooling via gas path components. In some examples, sensor-based measurement points and / or locations (e.g., single-point measurement, two-point measurement, three-point measurement, etc.) can be based on an air-filled conduit (e.g., a sensing line), which can be used to indicate pressure changes at selected measurement points.

[0061] As combined Figures 2A-2FAs described, based on the measured total pressure and static pressure obtained during the single-point pressure measurement, a conversion curve can be generated for new engines and / or engines that have some degradation due to long-term use and exposure to high combustible gas temperatures (e.g., reduced blade tips, etc.). For example, an engine may be out of service due to a failure of the exhaust gas temperature (EGT) margin caused by degradation of high-pressure turbine components, where increased blade tip clearance is a major factor in degradation of hot section engine components. For example, as engine components degrade and clearances increase, the internal temperature of the engine will increase as it becomes hotter to achieve the same thrust level. The engine has reached the EGT limit of the engine, indicating that the disk of the high-pressure turbine has reached its upper temperature limit, resulting in the engine being off the wing for expensive maintenance. Therefore, blade tip clearance management is critical to ensuring improved engine efficiency, stability, and overall service life. The described real-time clearance assessment based on pressure measurements and using conversion curves using the method disclosed herein allows the engine's active clearance control (ACC) system to receive real-time input about clearances, not only for new engines, but also for engines that have begun to show signs of degradation (e.g., reduced tip blades). This allows the ACC system to adjust the lash appropriately (e.g., via opening and / or closing the lash), thereby improving engine efficiency, allowing for longer service life, such as in conjunction with Figure 5 Described below. Figure 3C-3D The blade tip reduction and resulting pressure measurements which can be obtained to allow a conversion curve to be developed for a degraded engine are described in more detail.

[0062] Figure 3C An exemplary two-point pressure measurement 315 is shown after fan bucket 140 and / or rotor blade 164 , 168 tip loss has occurred, showing an exemplary airflow profile 330 as an exemplary radial tip clearance 325 increases. Figure 3D An exemplary two-point pressure measurement 340 is shown after blade 140, 164, 168 tip loss has occurred, showing an exemplary airflow distribution 345 as an exemplary radial tip clearance 360 ​​is reduced. Figure 3CIn the example of , the original length 318 of the tip of the blade 140, 164, 168 shows the total reduction 320 due to the loss of the blade 140, 164, 168 tip. Such tip loss may occur due to friction, oxidation, erosion, corrosion and / or coating fatigue. Therefore, the reduction of the blade tip can cause clearance and overall clearance changes, which change the airflow in the engine - affecting operating behavior, fuel consumption and / or performance. Too large a clearance can cause increased internal leakage, resulting in thrust loss, increased fuel consumption and / or increased hot air flow temperature, while insufficient clearance can cause the blade 140, 164, 168 to rub against the casing 142, 157 (e.g., shroud). This friction can cause rotor blade failure (e.g., dynamic fatigue), overheating and / or damage to surfaces exposed to friction. As previously described, the engine's FADEC system can digitally control engine performance, including calculating tip clearance under operating conditions to allow ACC-based tip clearance optimization. However, the FADEC system calculations rely on "new" tip blade clearances associated with tip blades that do not have any tip losses due to operating conditions (e.g., friction, oxidation, etc.). Figures 2A-2D and / or Figure 3A-3B This new blade based pressure measurement is obtained as described to obtain Figure 3E However, if combined with Figure 5 As described, to allow ACC-based tip clearance to be effectively optimized and / or otherwise improved, real-time clearance measurements may be obtained. To obtain such real-time measurements, a transfer curve may also be obtained for blades with sustained degradation (e.g., blade tip loss) that occurs gradually during engine cycles and throughout engine life. Figure 3C-3D Two point pressure measurements of a blade with tip losses are shown, allowing a transfer curve to be drawn taking into account the reduction in blade tip over time.

[0063] exist Figure 3C In the example of , two-point pressure measurements may be obtained at locations 250, 308. However, any other pressure-based measurement locations (e.g., 205, 250, and / or 308) may be used. Figure 3A The opening of the gap (e.g., via active gap control) introduces a larger gap 325 (e.g., as shown by the exemplary blade tip reduction 320 relative to the original blade length 318) than the gap 225 in the absence of blade 140, 164, 168 losses. This causes a change in the airflow path because the airflow of the combustible gas shown by the flow distribution 330 is increased compared to the original flow distribution 220 in the absence of blade tip losses. Therefore, the tip gap not only affects the generated flow field, but also affects the heat transfer performance. Similarly, in Figure 3D In the example, Figure 3BThe airflow distribution 345 for closing the gap in the presence of blade 140, 164, 168 tip losses (e.g., blade tip reduction 355 relative to the original blade length 350 as shown) is increased compared to the airflow distribution 240 in the absence of blade 140, 164, 168 tip losses in the embodiment of the present invention. Therefore, when blade tip losses are not considered when determining the optimal gap, the real-time gap 360 in the presence of blade 140, 164, 168 tip losses (e.g., tip reduction 320, 355) can be larger than the desired gap 245 achieved using ACC. By developing a calibration curve that takes into account blade tip reduction 320, 355 based on pressure measurements at one or more locations (e.g., locations 205, 250 and / or 308), the ACC-based gap modulation can be adjusted to reflect the real-time blade tip conditions to achieve the desired gap (e.g., gap 225, 245) instead of a larger than expected gap (e.g., gap 325, 360).

[0064] Figure 3E For new (non-degraded) blades 140, 164, 168, based on Figure 3A-3B An exemplary conversion curve 370 is determined using the gap 265 and the pressure efficiency 270 for a two-point pressure measurement. Figure 3F For blades 140, 164, 168 with tip loss (deterioration), based on Figure 3C-3D An exemplary conversion curve 380 is determined using the gap 265 and the pressure efficiency 270. Figure 3A As described, Equations 1 and 2 can be used based on the location of the static pressure sensor (e.g., Figure 3A-3B Position 308, 205 and / or Figure 3C-3D The positions 308, 250 of the tip are used to determine a two-point pressure measurement (e.g., using two local static pressures). As the tip is reduced 320, Figure 3A The air flow velocity of the flow distribution 225 will be different from Figure 3C The air flow velocity of the flow distribution 330. Therefore, Figure 3E Conversion curve 370 of pressure efficiency (Pη) (eg, pressure η 270) versus clearance (eg, clearance 265) for a new engine 375 (eg, an undegraded engine without blade tip losses) and Figure 3D The conversion curves are 380 different, where Figure 3D A transfer curve is included for a degraded engine 385 (e.g., an engine with blade tip loss). The availability of two transfer curves (e.g., for a new engine 375 and a degraded engine 385) allows clearance 265 to be determined for engines at different life cycle stages (which results in changes in pressure 270).

[0065] Figure 4A An exemplary three-point pressure measurement 400 is shown, illustrating the airflow profile 220 as the radial tip gap 225 increases. Figure 4B An exemplary three-point pressure measurement 425 is shown, showing the airflow distribution 245 as the radial tip clearance 245 is reduced. Figures 4A-4B , an example of three-point pressure measurements (e.g., front (P1), middle (P2), and rear (P3)) is shown to illustrate that any of the previous measurement locations (e.g., locations 308, 205, 250) can be used for multi-point pressure measurements that track changes in static pressure (e.g., based on the airflow 215 velocity shown using flow profiles 220, 240) during ACC-based gap opening (e.g., gap 225) and ACC-based gap closing (e.g., gap 245). For example, P2 based on the middle measurement will generate different outputs based on pressure changes and / or airflow profile changes. Therefore, multiple pressure point measurements can allow for the development of more accurate conversion curves (e.g., various altitudes, power levels, etc.) under various operating conditions (e.g., various altitudes, power levels, etc.). Figure 4C-4D The conversion curves 450, 475) are used, and the blade tip losses can be evaluated based on the deviation from the conversion curves developed for the new engine during ground and / or flight testing. For example, pressure measurements can be obtained using one or more sensors. In some examples, multi-point pressure measurements can rely on a simple traditional static pressure sensor. In some examples, any other type of pressure sensor (e.g., optical, laser, capacitive, eddy current, microwave, etc.) can be used. For example, the use of pressure sensors in harsh environments (such as high pressure turbines) requires that the pressure sensor design can withstand the harsh environment of the turbine engine and include a rugged design, a long operating life, high vibration and shock tolerance, easy maintenance, no cooling flow required during operation, improved signal-to-noise ratio and / or low cost suitable for mass-produced engines.

[0066] Figure 4C Shown based on Figures 4A-4B An exemplary transfer curve 450 is determined using the three-point pressure measurements 400 , 425 at high power 275 operation using the clearance 265 and pressure efficiency 270 , including transfer curves for a new engine 280 and a degraded engine 285 with tip loss. Figure 4D The graph based on Figures 4A-4BThe three-point pressure measurements 400, 425 of the engine are used to determine an exemplary transfer curve 475 under low power 290 operation using the clearance 265 and the pressure efficiency 270. For example, during operation, the FADEC system and associated ACC system of the engine can determine blade tip loss based on the offset of the transfer curve of the degraded engine 285 from the transfer curve of the new engine 280, which is developed when the new engine is tested on the ground and / or in flight (e.g., high power 275 and / or low power 290). Figure 5 As described, this allows the ACC system to correct the gaps 225 , 245 in the event of blade 140 , 164 , 168 tip loss.

[0067] Figure 5 An exemplary measurement 502 of exhaust gas temperature (EGT) degradation 504 over a plurality of flight cycles 506 using an exemplary baseline measurement 508 is shown compared to a real-time clearance adjustment 512 implemented using the methods disclosed herein. As previously described, EGT can be used as an indicator of whether an engine requires off-wing maintenance due to degradation and / or has reached its maximum service capability. Thus, EGT allows for management and diagnosis of the engine and provides protection for engine components that are sensitive to thermal loads. Figure 5 In the example of , EGT refers to the turbine exhaust temperature during exit from the turbine unit, which is measured using thermocouples mounted in the exhaust stream. Active clearance control maintains optimal clearances to, in part, ensure that EGT remains below its limits, thereby improving engine efficiency and time on board. Similarly, tighter blade tip clearances are maintained to reduce air leakage over the tips of the blades 140, 164, 168, otherwise the rotor inlet temperature would increase to achieve the same performance level and the hot section components would experience a shortened life cycle due to increased temperatures to produce the same amount of work (e.g., thermal fatigue). In addition, tip clearances can be optimized by ACC to ensure engine efficiency and reduce maintenance costs. Figure 5 In the example of , an increasing number of flight cycles 506 results in higher EGT degradation 504, which includes tip loss. Using baseline measurements 508, it can be estimated that a new engine has a certain level of EGT degradation 504 over a given number of flight cycles 506. However, such baseline measurements 508 are not necessarily representative of the actual EGT degradation 504 of a given engine over time.

[0068] like Figure 5As shown in the example of , the EGT degradation rate is highest during initial operation and then stabilizes to reach a steady-state level (e.g., baseline measurement 508). For example, baseline measurement 508 may indicate an installed loss of EGT degradation of 25 degrees Celsius for the first 2,000 flight cycles, compared to a steady-state loss of approximately 5 degrees Celsius every 1,000 flight cycles after initial operation. Unlike mature engines, engines operating for the first time (e.g., new engines) have higher EGT margins and lower EGT degradation rates. A borescope inspection (BSI) may provide a manual method to determine engine degradation. Such an inspection may be used to reset the EGT degradation 504 measurement, such as Figure 5 (e.g., through appropriate maintenance and / or parts replacement, etc.). However, BSI requires a lot of manual effort for tip notch detection, may not guarantee high quality data, and requires manual tracking rather than an automated solution that can be implemented in real time. In contrast, real-time clearance adjustment 512 can reduce EGT degradation by allowing optimized ACC-based clearance modulation based on the real-time state of the engine (e.g., blade tip loss progression). As previously described, a pressure to clearance conversion curve can be developed to allow identification of offsets relative to the conversion curve corresponding to engine degradation. Figures 6A-6C Determination of a clearance to pressure conversion curve during engine-based testing (eg, ground testing and / or flight testing) is further described.

[0069] Fig. 6A An exemplary graph 602 is shown showing the change in clearance with the addition of active clearance control based on the measurement of the intermediate seal static pressure at the front and rear cavities. Fig. 6A In the example of FIG. 6A , gap 606 decreases as ACC 604 is used (e.g., resulting in Figure 2A For example, the ACC system includes a butterfly valve that moves at various angles to change the cooling flow of the containment structure 142 (e.g., shroud), thereby controlling the expansion and / or contraction of the structure and maintaining a precise gap between the containment structure 142 and the blade tip. Thus, the ACC 604 valve can be fully closed (0%), partially open, or fully open (100%). Fig. 6A In the example of FIG. 6 , increased ACC 604 causes gap 606 to decrease, as shown using an average measurement 608 obtained from multiple tests. Figure 6B An exemplary graph 612 is shown showing the change in pressure efficiency 614 as the ACC 604 increases based on two test locations 610, 616 (e.g., static pressure measurements based on multi-point pressure measurements of FIGS. 3 and 4). Figure 6BIn the example of FIG. 6 , for two test positions 610, 616, the pressure efficiency (Pη) decreases as the ACC 604 increases. In some examples, the test positions may correspond to the front position 308, the middle position 205, and / or the rear position 250 of FIG. 2, 3, and / or 4. Based on Fig. 6A and 6B , a conversion curve can be prepared (for example, based on flight data obtained at the cruise point), such as Figure 6C As shown, Figure 6C An exemplary linear correlation 620 is shown between the gap 606 and the pressure efficiency 614. The linear correlation 620 indicates that an increase in the pressure efficiency 614 results in a corresponding increase in the gap 606, as previously shown using the conversion curves 264, 288, 370, 380, 450, 475 of FIGS. Figure 6C The exemplary transfer curve 620 may be developed during engine ground testing and / or flight testing, including at different altitudes and / or power levels (eg, at different crankshaft rotations, such as 16,400 revolutions per minute (rpm)).

[0070] Figure 7 7 is a block diagram 700 of an exemplary implementation of a blade tip loss determiner 710 by which the examples disclosed herein may be implemented. Figure 7 In the example of , active clearance controller 705 communicates with blade tip loss determiner 710. Exemplary blade tip loss determiner 710 includes measurement initiator 715, reference point selector 720, pressure sensor 725, transfer curve generator 730, test result analyzer 735, and data storage 740.

[0071] Active clearance controller 705 is part of a full authority digital engine control (FADEC) system that is used to maintain tight blade tip clearances to reduce hot gas leakage and improve engine performance (e.g., fuel burn, life cycle, service life, etc.). Controller 705 allows for real-time modulation of turbine clearances. For example, controller 705 can actuate a butterfly valve (e.g., via the FADEC system) to adjust the blade tip clearances around the engine (e.g., Figure 1 The controller 705 controls the blade tip clearance (e.g., blade tip clearance 225, 245 of FIG. 2 ) by distributing cooling air to the housing and / or containment structure 142, 157, thereby causing the structure to narrow. In some examples, the controller 705 maintains a circumferentially uniform clearance to take into account engine-to-engine manufacturing variability and real-time load effects on engine structural components. In some examples, the actuation mechanism of the controller 705 moves the housing 142, 157 components (e.g., shrouds) against large pressure differentials (e.g., 100-200 pounds of force per square inch (PSI)) to allow the clearance to open (e.g., as Figure 2A , 2C, 3A, 3C and / or 4A) and / or gap closure (e.g., as shown in Figure 2B , 2D , 3B, 3D and / or 4B). Figure 7 In the example of FIG. 7 , the controller 705 receives input from the blade tip loss determiner 710 to achieve tight clearance based on pressure measurements obtained in real time. This allows the controller 705 to modify the clearance accordingly even in the presence of blade tip loss, where the blade tip loss may result in a larger than expected clearance (e.g., as shown in FIG. 7 ). Figure 3C-3D gaps 325, 360 in the example).

[0072] The blade tip loss determiner 710 can be used during initial testing of the engine to generate conversion curves 264, 288, 370, 450, 475, and / or 620 at different power levels and / or altitudes, and during in-flight monitoring of clearance by the controller 705 to make real-time clearance adjustments that reflect engine conditions (e.g., progressive blade tip loss). The blade tip loss determiner 710 includes a measurement initiator 715 to determine when a pressure-based measurement (e.g., a single-point, two-point, and / or three-point pressure measurement) is needed (e.g., during testing and / or in-flight data collection). In some examples, the measurement initiator 715 initiates pressure measurements using one or more sensors (e.g., conventional static pressure sensors, optical sensors, laser-based sensors, capacitive sensors, eddy current sensors, microwave sensors). In some examples, the measurement initiator 715 initiates measurements at a rear, intermediate, and / or forward position relative to a given clearance, the given clearance being determined based on the direction of the combustible gas airflow (e.g., airflow 215 of FIG. 2). In some examples, during initial testing to develop a conversion curve that relates pressure to clearance measurements, the measurement initiator 715 can determine when to initiate pressure measurements based on a given power level (e.g., low power, high power), a particular altitude (e.g., at 35 kilometers, etc.), and / or a particular flight period.

[0073] Reference point selector 720 determines whether to perform a single point measurement (e.g., as shown in FIG. 2 ), a two point measurement (e.g., as shown in FIG. 3 ), and / or a three point measurement (e.g., as shown in FIG. 4 ). In some examples, the total number of reference points used for pressure measurement can be determined based on engine type and / or other parameters, such as altitude, flight cycle, and / or power level. In some examples, multiple point measurements (e.g., three point measurements) can introduce greater accuracy to the final conversion curve prepared based on the acquired data. In some examples, reference point selector 720 determines the locations of reference points that will be used to obtain pressure measurements during testing and / or in flight. For example, Figures 2A-2DAs shown, a single point pressure measurement can be based on a front (e.g., upstream), middle and / or rear (e.g., downstream) position (e.g., positions 205, 250), and / or a single point pressure measurement can require both static pressure and total pressure measurements. Thus, the selection of a single point or multiple point pressure measurement can be based on whether a total pressure measurement is available (e.g., depending on the type of pressure sensor used). For example, a measurement of total pressure may require a more robust sensor (e.g., an optical sensor with cooling flow), while multiple point pressure measurements can be based on local static pressures, which can be obtained using a conventional static pressure sensor. As shown in FIG. Figures 3A-3D As shown, two-point measurements can similarly be based on specified locations and / or reference points (eg, locations 308 , 205 , and / or 250 ). Thus, reference point selector 720 can identify a location for use with pressure-based measurements, which can be based on the positioning of pressure sensor 725 .

[0074] The pressure sensor 725 can be designed to withstand the harsh environment of the turbine engine, with a long operating life, high vibration and shock tolerance, easy maintenance, no cooling flow required during operation, improved signal-to-noise ratio and / or low cost suitable for mass-produced engines (e.g., including cooling technology for increasing the life span of the sensor). As described in conjunction with FIG. 2, such advanced pressure sensors can be used for pressure measurements based on a single point. However, local static pressure measurements (e.g., using traditional pressure sensors) can be used to obtain multi-point pressure measurements. In some examples, the pressure sensor 725 (e.g., a static pressure sensor) can be installed on any area of ​​the engine, allowing clearance pressure measurements to be obtained through reliable data collection. In some examples, the pressure sensor 725 may include a transducer for converting the pressure measurement into an electrical signal that is transmitted to the controller 705. In some examples, the pressure sensor 725 can be positioned at a cooler location on the engine casing (e.g., the casing 157 surrounding the LP turbine rotor blades 164 and / or the HP turbine rotor blades 168) to avoid damage caused by high temperatures at the location where the pressure measurement is collected. For example, pressure sensor 725 can sense engine pressure via an air charging line (e.g., a sensing line), which can indicate a pressure change at a point of interest. In some examples, pressure sensor 725 can be mounted directly to an area of ​​interest to collect the desired data without the need for a sensing line. In some examples, pressure sensor 725 can be used to obtain various pressure measurements, including static pressure, peak pressure (P high ) (e.g., the burner pressure at the upstream) and / or the minimum pressure (P low )(e.g. rear pressure measurement).

[0075] The conversion curve generator 730 generates a conversion curve (e.g., conversion curves 264, 288, 370, 450, and / or 475 of Figures 2, 3, and / or 4) to determine the relationship between pressure efficiency (Pη) and clearance. For example, the conversion curve generator 730 receives input from the pressure sensor 725 and uses equations 1-2 (e.g., for multi-point pressure measurements) as described in conjunction with Figures 2 and 3. Thus, a specific clearance (mils) can be determined based on the pressure measurements obtained, allowing a conversion curve to be developed that can be used by the blade tip loss determiner 710 to identify deviations from the curve (e.g., due to blade tip losses) and thereby communicate the deviations to the controller 705, thereby achieving more accurate clearance adjustments based on real-time pressure data, such as in conjunction with Fig.10 As described.

[0076] The test results analyzer 735 determines changes in pressure measurements obtained using the pressure sensor 725 and / or identifies deviations from transfer curves generated using the transfer curve generator 730. For example, as the engine degrades and blade tip losses occur, any deviations from transfer curves developed for a new engine (e.g., identified in the sample transfer curves 2E-2F of FIG. 2 ) can be determined, allowing for real-time blade 140 , 164 , 168 tip loss assessments. In some examples, the test results analyzer 735 provides the controller 705 with real-time clearance measurements that account for blade tip loss progression, allowing the controller 705 to adjust clearances accordingly based on observed blade tip losses, avoiding the presence of larger clearances (e.g., as blade tip losses are observed). Figure 3C-3D ), rather than using gaps 325 , 360 in the example shown, a target and / or optimized gap is achieved to ensure engine efficiency.

[0077] The data storage 740 may be used to store any information related to the blade loss determiner 710. For example, the database 740 may store pressure measurements obtained using the one or more pressure sensors 725, transfer curves generated using the transfer curve generator 730, and / or the test results analyzer 735 output used by the controller 705 to make clearance adjustments based on real-time data. Figure 7 The exemplary data storage device 740 of the illustrated example is implemented by any memory, storage device, and / or storage disk for storing data, such as flash memory, magnetic media, optical media, etc. In addition, the data stored in the exemplary data storage device 740 can be in any data format, such as binary data, comma-delimited data, tab-delimited data, structured query language (SQL) structures, image data, etc.

[0078] Although Figure 7 An exemplary embodiment of a blade tip loss determiner 710 is shown, but in Figure 7One or more of the elements, processes and / or devices shown in the examples may be combined, divided, rearranged, omitted, eliminated and / or implemented in any other manner. In addition, the exemplary measurement initiator 715, the exemplary reference point selector 720, the exemplary pressure sensor 725, the exemplary conversion curve generator 730, the exemplary test result analyzer 735 and / or more generally, Figure 7 The exemplary blade tip loss determiner 710 may be implemented by hardware, software, firmware and / or any combination of hardware, software and / or firmware. Thus, the exemplary measurement initiator 715, the exemplary reference point selector 720, the exemplary pressure sensor 725, the exemplary conversion curve generator 730, the exemplary test result analyzer 735 and / or more generally, Figure 7 Any of the exemplary blade tip loss determiners 710 of the present invention may be implemented by one or more analog or digital circuits, logic circuits, programmable processors, programmable controllers, graphics processing units (GPUs), digital signal processors (DSPs), application specific integrated circuits (ASICs), programmable logic devices (PLDs), and / or field programmable logic devices (FPLDs). When any device or system claim of this patent is read to cover pure software and / or firmware implementations, the exemplary measurement initiator 715, the exemplary reference point selector 720, the exemplary pressure sensor 725, the exemplary conversion curve generator 730, the exemplary test result analyzer 735, and / or more generally, Figure 7 At least one of the exemplary blade tip loss determiners 710 is expressly defined herein as comprising a non-transitory computer readable storage device or storage disk, such as a memory, a digital versatile disk (DVD), a compact disk (CD), a Blu-ray disk, etc., including software and / or firmware. Further, in addition to or in lieu of Figure 7 Those shown in Figure 7 The exemplary blade tip loss determiner 710 may include one or more elements, processes and / or devices, and / or may include more than one of any or all of the illustrated elements, processes and devices. As used herein, the expression "communication" includes variations thereof, including direct communication and / or indirect communication through one or more intermediate components, and does not require direct physical (e.g., wired) communication and / or continuous communication, but also includes selective communication at periodic intervals, predetermined intervals, non-periodic intervals and / or one-time events.

[0079] Figure 8-10 shows a representative for implementing Figure 7 Flowchart of exemplary hardware logic, machine readable instructions, hardware implemented state machine and / or any combination thereof of blade tip loss determiner 710. Machine readable instructions may be executed by a computer processor (e.g., as described below in conjunction with Fig.11The exemplary processor platform 1100 discussed herein may be a processor 1112 or a portion of an executable program executed by the processor 1112. The program may be embodied in software stored on a non-transitory computer-readable storage medium, such as a CD-ROM, floppy disk, hard drive, DVD, Blu-ray disk, or memory associated with the processor 1112, but the entire program and / or portions thereof may alternatively be executed by a device other than the processor 1112 and / or embodied in firmware or dedicated hardware. In addition, although reference is made to Figure 8-10 The flowchart shown in describes an exemplary procedure, but many other methods of implementing the exemplary blade tip loss determiner 710 may alternatively be used. For example, the order of execution of the blocks may be changed, and / or some of the blocks described may be changed, eliminated, or combined. Additionally or alternatively, any or all of the blocks may be implemented by one or more hardware circuits (e.g., discrete and / or integrated analog and / or digital circuits, FPGAs, ASICs, comparators, operational amplifiers (op-amps), logic circuits, etc.) configured to perform the corresponding operations without executing software or firmware.

[0080] The machine-readable instructions described herein may be stored in one or more of a compressed format, an encrypted format, a segmented format, a compiled format, an executable format, a packaged format, etc. The machine-readable instructions as described herein may be stored as data (e.g., portions of instructions, codes, representations of codes, etc.) that may be used to create, manufacture, and / or generate machine-executable instructions. For example, the machine-readable instructions may be segmented and stored on one or more storage devices and / or computing devices (such as servers). The machine-readable instructions may require one or more of installation, modification, adaptation, update, combination, supplementation, configuration, decryption, decompression, unpacking, distribution, redistribution, compilation, etc., so that they can be directly read, interpreted, and / or executed by a computing device and / or other machine. For example, the machine-readable instructions may be stored in multiple parts that are individually compressed, encrypted, and stored on separate computing devices, where the parts, when decrypted, decompressed, and combined, form a set of executable instructions that implement a program such as described herein.

[0081] In another example, the machine-readable instructions may be stored in a state that can be read by a computer, but require the addition of a library (e.g., a dynamic link library (DLL)), a software development kit (SDK), an application programming interface (API), etc., to execute the instructions on a particular computing device or other device. In another example, the machine-readable instructions and / or corresponding programs may need to be configured in whole or in part (e.g., stored settings, data inputs, recorded network addresses, etc.) before the machine-readable instructions and / or corresponding programs can be executed in part or in whole. Therefore, the disclosed machine-readable instructions and / or corresponding programs are intended to cover such machine-readable instructions and / or programs, regardless of the specific format or state of the machine-readable instructions and / or programs when stored or otherwise stationary or transported.

[0082] The machine-readable instructions described herein may be represented by any past, present, or future instruction language, scripting language, programming language, etc. For example, the machine-readable instructions may be represented using any of the following languages: C, C++, Java, C#, Perl, Python, JavaScript, Hypertext Markup Language (HTML), Structured Query Language (SQL), Swift, etc.

[0083] As mentioned above, Figure 8-10 The exemplary processes of are implemented using executable instructions (e.g., computer and / or machine readable instructions) stored on a non-transitory computer and / or machine readable medium, such as a hard drive, flash memory, read-only memory, compact disk, digital versatile disk, cache, random access memory, and / or any other storage device or storage disk where information is stored for any duration (e.g., for an extended period of time, permanently, for a brief instance, temporarily buffered, and / or for caching information). As used herein, the term non-transitory computer readable medium is expressly defined to include any type of computer readable storage device and / or storage disk and does not include propagating signals and does not include transmission media.

[0084] Figure 8 A flowchart representing exemplary machine readable instructions 800 is shown that may be executed to implement Figure 7 Example blade tip loss determiner 710 of FIG. Figure 8In the example of , reference point selector 720 identifies reference points for obtaining pressure measurements using one or more pressure sensors 725 (block 805). In some examples, the reference points are determined based on whether a single-point, two-point, or three-point pressure measurement will be used to obtain the expected pressure measurement. For example, the determination of the reference point may depend on whether a conversion curve is being generated and / or whether data is collected during a subsequent flight cycle. Thus, the reference point may be determined based on the positioning and / or availability of pressure sensor 725. In some examples, pressure sensor 725 measures static pressure (P) at the inlet and / or outlet. s ), an identified reference point for a multi-point pressure measurement (e.g., as described in conjunction with FIGS. 3-4 ). In some examples, the pressure sensor 725 measures a total pressure (reference pressure) and a static pressure (local pressure) for a single-point pressure measurement (e.g., as described in conjunction with FIG. 2 ) (block 810). Once the pressure measurements are obtained, the conversion curve generator 730 generates a conversion curve for the engine being tested (e.g., during ground-based testing and / or flight testing) (block 815). For example, the conversion curve generator 730 determines a normalized equation for the pressure efficiency (Pη) to generate a linear relationship between the measured pressure efficiency and the corresponding clearance. As described in conjunction with Figures 6A-6C As described, the generation of the transfer curve may involve the measurement of the gap 606 at a given active gap control percentage 604 (e.g., as Fig. 6A ) and a measurement of pressure efficiency 614 at the same active clearance control percentage 604 (e.g., as Figure 6B The conversion curve generator 730 generates a conversion curve based on these measurements to obtain a linear relationship between the pressure efficiency 614 and the gap 606 (e.g., as shown in FIG. 1 ). Figure 6C shown).

[0085] Once the conversion curves are generated (e.g., for different engine power levels, altitudes, etc.), the blade tip loss determiner 710 measures real-time blade tip losses during actual engine flight cycles (block 820), such as in conjunction with Fig.10More specifically described. The blade tip loss determiner 710 identifies any measurable blade tip losses using the test result analyzer 735 (block 825). For example, the test result analyzer 735 compares the obtained pressure measurement data to the generated conversion curve so that any deviation of the curve indicates engine-based degradation (e.g., blade tip losses due to oxidation, thermal combustion, etc.). If the blade tip loss determiner 710 does not identify any blade tip losses based on the measurement data and / or the test result analyzer 735 output, the blade tip loss determiner 710 continues to monitor and / or measure any real-time blade tip losses (block 820). If the blade tip loss determiner 710 identifies a blade tip loss, the data is provided to the controller 705 (block 830). For example, the controller 705 uses the input received from the blade tip loss determiner 710 to adjust and / or optimize the tip clearance (block 835). In some examples, controller 705 may adjust cooling airflow, causing contraction and / or expansion of housing 142 , 157 to achieve tighter clearances and / or avoid any risk of friction between blades 140 , 164 , 168 and housing 142 , 157 .

[0086] Fig. 9 A flowchart representing exemplary machine readable instructions 815 is shown that may be executed to use Figure 7 The exemplary blade tip loss determiner 710 generates conversion curves for various power levels and / or altitudes. Fig. 9 In the example of , reference point selector 720 may identify one or more reference points at which pressure measurements may be made. For example, reference point selector 720 may identify a front (e.g., upstream) and / or a rear (e.g., downstream) reference point for obtaining one or more pressure measurements (e.g., Figure 3A-3BP1 and / or P2). The pressure sensor 725 mounted on the engine obtains front and / or rear pressure measurement data (box 905). The conversion curve generator 730 receives sensor-based input data and determines the correlation between the normalized pressure efficiency (Pη) and the blade clearance (box 910). For example, based on the input data provided via the pressure sensor 725, the conversion curve generator 730 can determine the normalized pressure efficiency (Pη). In some examples, the conversion curve generator 730 can generate such a conversion curve for a range of test flights, not only for new engines, but also for engines under different flight cycles (box 915). This allows verification of the conversion curve and allows observation and / or testing of engines with gradual blade loss to investigate the impact of blade length changes on pressure efficiency measurements. In some examples, testing can be performed at different power levels (e.g., low power, high power, etc.) and altitude ranges (box 920). Thorough testing and the generation of conversion curves allow the use of the blade loss determiner 710 during actual in-flight monitoring of the clearance and facilitate more accurate adjustment of the clearance by the active clearance controller 705.

[0087] Fig.10 A flowchart representing exemplary machine readable instructions 820 is shown that may be executed to use Figure 7 The exemplary blade tip loss determiner 710 measures real-time blade tip loss. Figure 8-9 The blade tip loss determiner 710 may be used to evaluate real-time blade tip loss in flight by using the transfer curves described above. For example, the reference point selector 720 identifies the pressure measurement location based on the pressure sensor location and / or program instructions (block 1005). As previously described, pressure measurements may be obtained using single-point, two-point, and / or three-point measurements, depending on factors such as the test site location, the type of sensor used, etc. The pressure sensor 725 measures the static pressure (P) at the identified measurement location. S ) and / or total pressure (P T), depending on whether the measurement is a single-point pressure measurement (e.g., including a total pressure measurement) or a multi-point pressure measurement (e.g., including a local static pressure measurement). For example, the blade tip loss determiner 710 identifies whether to perform a single-point pressure measurement (block 1010) or a multi-point pressure measurement (block 1015). For example, a pressure sensor 725 capable of measuring total pressure can be used as part of a single-point pressure measurement. Thus, the pressure sensor 725 can be used to measure total pressure and static pressure (block 1020). In some examples, if the pressure sensor 725 is configured and / or capable of measuring static pressure rather than total pressure, then the pressure sensor 725 measures the local static pressure for the multi-point pressure measurement, according to equations 1-2 (block 1025). In some examples, the measurements can be repeated and / or obtained at set time intervals, as required by the controller 705. For example, the measurements can be made more and / or less frequently depending on the flight conditions (e.g., takeoff, landing, cruising, etc.), or the measurements can be obtained continuously throughout the flight duration. The blade tip loss determiner 710 determines blade tip losses based on the shift in the transfer curve identified using the test results analyzer (block 1030). For example, the test results analyzer 735 compares the pressure efficiency data obtained during testing for generating the transfer curve for the new engine with the pressure efficiency measurements obtained during real-time, in-flight data collection. The shift from the expected pressure measurement may indicate blade tip losses, resulting in pressure changes and resulting in larger clearances than would actually be expected by the controller 705 in the absence of real-time pressure measurement data. Therefore, the controller 705 determines the corrective clearances based on the real-time pressure measurements, accounting for any clearance changes introduced due to gradual blade tip losses.

[0088] Fig.11 is constructed to execute Figure 8-10 The instructions to achieve Figure 7 1. Block diagram of an exemplary processor platform of an exemplary blade tip loss determiner of FIG. The processor platform 1100 may be a server, a personal computer, a workstation, a self-learning machine (eg, a neural network), or any other type of computing device.

[0089] The processor platform 1100 of the illustrated example includes a processor 1112. The processor 1112 of the illustrated example is hardware. For example, the processor 1112 can be implemented by one or more integrated circuits, logic circuits, microprocessors, GPUs, DSPs, or controllers from any desired family or manufacturer. The hardware processor can be a semiconductor-based (e.g., silicon-based) device. In this example, the processor 1112 implements the exemplary blade tip loss determiner 710, including the exemplary measurement initiator 715, the exemplary reference point selector 720, the exemplary pressure sensor 725, the exemplary conversion curve generator 730, and / or the exemplary test result analyzer 735.

[0090] The processor 1112 of the illustrated example includes a local memory 1113 (e.g., a cache). The processor 1112 of the illustrated example communicates with a main memory via a bus 1118, and the main memory includes a volatile memory 1114 and a non-volatile memory 1116. The volatile memory 1114 may be comprised of a synchronous dynamic random access memory (SDRAM), a dynamic random access memory (DRAM), Dynamic Random Access Memory The main memory 1114 and 1116 may be implemented by a memory controller.

[0091] The processor platform 1100 of the illustrated example also includes an interface circuit 1120. The interface circuit 1120 may be implemented by any type of interface standard, such as an Ethernet interface, a universal serial bus (USB), interface, a near field communication (NFC) interface, and / or a PCI Express interface.

[0092] In the example shown, one or more input devices 1122 are coupled to the interface circuit 1120. The input devices 1122 allow a user to input data and / or commands into the processor 1112. The input devices 1122 may be implemented by, for example, an audio sensor, a microphone, a camera (still or video), a keyboard, buttons, a mouse, a touch screen, a trackpad, a trackball, an isopoint, and / or a voice recognition system.

[0093] One or more output devices 1124 are also connected to the interface circuit 1120 of the illustrated example. The output device 1124 can be implemented, for example, by a display device (e.g., a light emitting diode (LED), an organic light emitting diode (OLED), a liquid crystal display (LCD), a cathode ray tube display (CRT), an in-place switching (IPS) display, a touch screen, etc.), a tactile output device, a printer, and / or a speaker. Therefore, the interface circuit 1120 of the illustrated example typically includes a graphics driver card, a graphics driver chip, and / or a graphics driver processor.

[0094] The interface circuitry 1120 of the illustrated example also includes communication devices such as transmitters, receivers, transceivers, modems, residential gateways, wireless access points, and / or network interfaces to facilitate data exchange with external machines (e.g., any type of computing device) via the network 1126. Communications may be via, for example, an Ethernet connection, a digital subscriber line (DSL) connection, a telephone line connection, a coaxial cable system, a satellite system, a line-of-site wireless system, a cellular telephone system, etc.

[0095] The processor platform 1100 of the illustrated example also includes one or more further mass storage devices 1128 for storing software and / or data. Examples of such mass storage devices 1128 include floppy disk drives, hard drive disks, compact disk drives, Blu-ray disk drives, redundant arrays of independent disks (RAID) systems, and digital versatile disk (DVD) drives.

[0096] Figure 8-10 The machine executable instructions 1132 may be stored in the mass storage device 1128, the volatile memory 1114, the non-volatile memory 1116, and / or a removable non-transitory computer-readable storage medium (e.g., a CD or DVD). For example, one or more of the volatile memory 1114, the non-volatile memory 1116, the mass storage device 1128, etc. may also be used to implement the exemplary data storage device 740.

[0097] From the above, it should be understood that the disclosed method and apparatus allow for real-time measurement of blade tip clearance taking into account blade tip loss. Given that the power provided by the turbine (or consumed by the compressor) depends on the airflow generated by the blade position area, an increase in the tip clearance will result in a decrease in turbine efficiency. Therefore, the presence of the tip clearance will result in a change in airflow, affecting the expected flow path and affecting turbine efficiency, including a potential increase in fuel consumption. The method and apparatus disclosed herein allow for the production of a conversion curve that can be used to determine blade tip loss based on an offset identified from the conversion curve. Therefore, active clearance control can be used to more accurately calculate and adjust the clearance based on real-time data input by taking into account blade tip loss, wherein blade tip loss can result in greater clearance and reduced engine efficiency, resulting in shorter engine life and on-board time. Although the examples disclosed herein describe real-time clearance assessment in an example aircraft engine, the methods and apparatus disclosed herein can be used for any turbine engine system. In addition, although the examples disclosed herein describe real-time clearance assessment based on low-pressure turbine rotor blades and / or high-pressure turbine rotor blades, the clearance modulation using the methods and apparatus disclosed herein can be applied to any other blades used in aircraft engines and / or any turbine engine system.

[0098] Although certain example methods, apparatus, and articles of manufacture have been disclosed herein, the scope of coverage of this patent is not limited thereto. On the contrary, this patent covers all methods, apparatus, and articles of manufacture that fully fall within the scope of the claims of this patent.

[0099] The following claims are incorporated into this Detailed Description by reference, with each claim standing on its own as a separate embodiment of the present disclosure.

[0100] Further aspects of the invention are provided by the subject matter of the following clauses:

[0101] A method for assessing real-time blade tip clearance in a turbine engine, the method comprising: determining a first static pressure measurement and a second static pressure measurement at a first measurement location and a second measurement location, respectively, relative to the blade tip clearance; determining a normalized pressure measurement using the first static pressure measurement and the second static pressure measurement; generating a conversion curve to relate the normalized pressure measurement to the clearance measurement; and adjusting active clearance control of the blade tip clearance based on a comparison of the real-time in-flight pressure measurement to the conversion curve.

[0102] A method according to any preceding clause, wherein the first pressure measurement or the second pressure measurement is obtained using a static pressure sensor.

[0103] A method according to any preceding clause, wherein the first static pressure measurement or the second static pressure measurement is obtained at a rearward position, an intermediate position or a forward position relative to the blade and the casing.

[0104] A method according to any preceding clause, wherein the transfer curve is developed for the turbine engine during testing at multiple altitudes.

[0105] A method according to any preceding clause, wherein the transfer curve is developed for the turbine engine during testing at a plurality of power levels, the plurality of power levels comprising at least one of low power or high power.

[0106] A method according to any preceding clause, wherein the conversion curve is determined based on clearance measurements and normalized pressure measurements obtained at different percentages of active clearance control, and the clearance measurements and normalized pressure measurements are correlated based on the percentage of active clearance control with which the two measurements are related.

[0107] A method according to any preceding clause, wherein the blade tip clearance is based on a distance between the blade and the casing, the blade comprising a fan blade, a high pressure rotor blade, or a low pressure rotor blade.

[0108] A method according to any preceding clause, wherein the housing is a fan housing or a turbine housing.

[0109] An apparatus for assessing real-time blade tip clearance in a turbine engine, the apparatus comprising: a pressure sensor that determines a first static pressure measurement and a second static pressure measurement at a first measurement location and a second measurement location, respectively, relative to the blade tip clearance; a transfer curve generator that uses the first static pressure measurement and the second static pressure measurement to determine a normalized pressure measurement and to generate a transfer curve to relate the normalized pressure to the clearance measurement; and an active clearance controller that adjusts active clearance control of the blade tip clearance based on a comparison of the real-time in-flight pressure measurement to the transfer curve.

[0110] Apparatus according to any preceding clause, further comprising a reference point selector, the reference point selector obtaining the first pressure measurement or the second pressure measurement at a rear position, an intermediate position or a forward position relative to the blade and the housing.

[0111] Apparatus according to any preceding clause, wherein the transfer curve generator generates transfer curves for a plurality of heights.

[0112] Apparatus according to any preceding clause, wherein the transfer curve generator generates the transfer curves for a plurality of power levels, the plurality of power levels comprising at least one of low power or high power.

[0113] An apparatus according to any preceding clause, wherein the conversion curve generator determines the conversion curve based on gap measurements and normalized pressure measurements obtained at different percentages of active gap control, and the gap measurement and the normalized pressure measurement are associated based on the percentage of active gap control corresponding to the two measurements.

[0114] An apparatus according to any preceding clause, further comprising a test results analyzer that compares the in-flight pressure measurement data with a generated transfer curve for the new engine.

[0115] A non-transitory computer-readable medium comprising machine-readable instructions that, when executed, cause a processor to at least: determine first and second static pressure measurements at first and second measurement locations, respectively, relative to blade tip clearance based on a signal received as an input to the processor; determine a normalized pressure measurement using the first and second static pressure measurements; generate a conversion curve to relate the normalized pressure measurement to the clearance measurement; and adjust active clearance control of the blade tip clearance based on a comparison of the real-time in-flight pressure measurement to the conversion curve.

[0116] The non-transitory computer readable medium of any preceding clause, wherein the location of the static pressure measurement is at least one of a rearward position, an intermediate position, or a forward position relative to the blade and the casing.

[0117] The non-transitory computer readable medium of any preceding clause, wherein the instructions cause the processor to generate a transfer curve for the turbine engine at a plurality of altitudes.

[0118] The non-transitory computer readable medium of any preceding clause, wherein the instructions cause the processor to generate a conversion curve for the turbine engine at a plurality of power levels, the plurality of power levels comprising at least one of low power or high power.

[0119] A non-transitory computer readable medium according to any preceding clause, wherein the instructions cause a processor to generate a conversion curve based on clearance measurements and normalized pressure measurements obtained under different percentages of active clearance control, and to associate the clearance measurements and the normalized pressure measurements based on the percentage of active clearance control corresponding to the two measurements.

[0120] The non-transitory computer readable medium of any preceding clause, wherein the instructions cause the processor to adjust the blade tip clearance based on a distance between the blade and the housing, the blade comprising a fan blade, a high pressure rotor blade, or a low pressure rotor blade.

Claims

1. A device for evaluating blade tip clearance in a turbine engine, characterized in that The device comprises: a pressure sensor that determines a first static pressure measurement at a first measurement location relative to a blade tip clearance and a second static pressure measurement at a second measurement location relative to the blade tip clearance; and an active clearance controller that adjusts active clearance control of the blade tip clearance based on a transfer curve that relates a normalized pressure measurement to a clearance measurement, the normalized pressure measurement being determined using the first static pressure measurement and the second static pressure measurement, the transfer curve being developed for a turbine engine during testing of a plurality of operating conditions.

2. The device according to claim 1, characterized in that The transfer curve is determined based on the gap measurements and the normalized pressure measurements obtained at different percentages of active gap control.

3. The device according to claim 2, characterized in that The clearance measurement and the normalized pressure measurement are correlated based on a percentage of active clearance control corresponding to both measurements.

4. The device according to claim 1, characterized in that wherein the active clearance controller adjusts active clearance control of the blade tip clearance based on comparison of real-time in-flight pressure measurements with the transfer curve.

5. The device according to claim 1, characterized in that Further included is a reference point selector that obtains the first static pressure measurement or the second static pressure measurement at a rear position, an intermediate position, or a forward position relative to the blade and the housing.

6. The device according to claim 1, characterized in that Further included is a test results analyzer that compares in-flight pressure measurement data to the conversion curve generated for the new engine.

7. The device according to claim 1, characterized in that Wherein the transition curves are identified for a plurality of power levels, the plurality of power levels comprising at least one of low power or high power.

8. A method for evaluating blade tip clearance in a turbine engine, characterized in that The method comprises: determining a first static pressure measurement at a first measurement location relative to a blade tip clearance and determining a second static pressure measurement at a second measurement location relative to the blade tip clearance; and Active clearance control that adjusts the blade tip clearance based on a transfer curve that relates a normalized pressure measurement to a clearance measurement, the normalized pressure measurement being determined using the first static pressure measurement and the second static pressure measurement, the transfer curve being developed for the turbine engine during testing of a plurality of operating conditions.

9. The method according to claim 8, characterized in that Further comprising determining the transfer curve based on the clearance measurements and the normalized pressure measurements obtained at different percentages of active clearance control.

10. The method according to claim 9, characterized in that Further comprising correlating the clearance measurement and the normalized pressure measurement based on a percentage of active clearance control corresponding to both measurements.

Images