Patent ID: 12235186

DETAILED DESCRIPTION

FIG.1illustrates an aircraft engine depicted as a gas turbine engine10of a type preferably provided for use in subsonic flight, generally comprising in serial flow communication a fan12through which ambient air is propelled, a compressor section14for pressurizing the air, a combustor16in which the compressed air is mixed with fuel and ignited for generating an annular stream of hot combustion gases, and a turbine section18for extracting energy from the combustion gases. It will be appreciated that the principles of the present disclosure may apply to any suitable aircraft engines such as turbofan, turboshaft, turboprop, and so on.

The gas turbine engine10includes a plurality of parts that define apertures through them for allowing a flow of a fluid (e.g., air) for cooling components, pressurizing seals, bleeding air, providing air for combustion, or other purposes. For instance, the combustor16includes a combustor liner16A that may define a plurality of liner apertures16B (only one illustrated for clarity) therethrough. The liner apertures16B are used to allow air from the compressor section14to flow through the combustor liner16A into a combustion chamber of the combustor16. Heat shields16C, which may be located within the combustor16, may further have apertures16D therethrough for film cooling and/or impingement cooling and so on. The compressor section14has a diffuser14A, which may also define apertures14B (only one shown) therethrough. These apertures may be used to bleed air from the compressor section14. The turbine section18includes airfoils18A of either blades or vanes, the airfoils may define aperture18B for cooling of the airfoils18A.

These apertures are often manufactured with laser drilling and, in some cases, they are coated with a heat-resistant coating or any other appropriate coating. The laser drilling and/or the coating may change the shape of peripheral walls bounding the apertures. This change of shape may affect a mass flow rate flowing through the apertures. An exemplary shape of an aperture being coated and/or laser drilled is shown schematically with dashed lines inFIG.3. As one can appreciated by looking atFIG.3, the profile of the aperture may deviate from a purely cylindrical hole. Therefore, each of the above-mentioned aircraft engine parts are subjected to extensive experimental flow testing to establish whether or not they are compliant with flow requirements.

A typical experimental flow testing bench or rig200is shown schematically onFIG.2and includes a source of a fluid, which is typically air,201, a regulator202for controlling a mass flow rate supplied from the source of ambient air201, a valve203, a sonic nozzle204, and a plenum205. The aircraft engine component is pneumatically connected to the plenum205to undergo testing. A conduit206fluidly connects the source201to the plenum205. A first temperature sensor207A and a first pressure sensor208A may be operatively connected to the air conduit206upstream of the sonic nozzle204. A second temperature sensor207B and a second pressure sensor208B may be operatively connected to the air conduit206downstream of the sonic nozzle204and upstream of the plenum205relative to a flow of the fluid flowing from the source201to the plenum205. As shown inFIG.3, the conduit206that is pneumatically connected to the plenum205has a diameter D that is different than a diameter d of the apertures14A,16A,16D,18A defined through the aircraft engine component being tested. A third temperature sensor207C and a third pressure sensor208C may be operatively connected to the plenum205for measuring a pressure and temperature of the fluid (e.g., air) being injected into the apertures. It will be appreciated that, although only one aperture is shown inFIG.3, the plenum205is designed to simultaneously inject the flow through a plurality of the apertures of the aircraft engine component being tested. More detail about the diameter d are presented herein below.

Since each flow bench might be used for various types of parts, a diameter of the sonic nozzle204may be varied to have different mass flow rate going through the sonic nozzle204. The pressure and temperature of the flow coming out of the sonic nozzle204is being measured. The mass flow rate through the part is being measured. The third pressure sensor208C is used to measure the air pressure of the flow prior to pass through the part.

The rig200may use specific fixtures, nozzle with specific diameters (e.g., 0.085″, 0.125″, 0.250″, 0.350″, 0.500″ and so on) and an operator with sufficient knowledge on the flow bench and the flow characteristics operates the rig200. Each sonic nozzle204being used must be calibrated on a regular basis (usually every 6 months) by an expert. The pressure and temperature sensors must be also calibrated at the same frequency. Depending on the design requirements, the flow may be based on constant pressure or constant mass-flow rate to verify the conformity of the part. Leak tests may have to be performed to ensure that the rig200is operating as it should. A master part test may have to be performed to corroborate the results and confirm that the rig200is adequately calibrated. The production part (e.g., combustor liner, etc) may then undergo flow-tests and checked for design requirements and approval.

The process therefore comprises of a plethora of steps from multiple organization (for flow rig calibration), expert personnel and logistics around these processes. This process is highly resource intensive and expensive. The method described in the current disclosure aims to at least partially alleviate some of the aforementioned drawbacks.

Typically, a plurality of the aircraft engine components, such as the combustor liner16A, the heat shields,16C, the airfoils18A, the diffusers14A and so on are manufactured so that more than one gas turbine engine10may be produced. As explained above, all of these components may have to undergo extensive experimental flow testing to ensure compliance with flow requirements. For instance, it may be important to determine that the apertures16B through the combustor liner16A allow the prescribed mass flow rate therethrough for proper operation of the gas turbine engine10.

The method described herein proposes to perform the experimental flow testing for a first instance of a specific aircraft engine component (e.g., a first combustor liner of a series of combustor liners). This first instance may be referred to as a prototype of the specific aircraft engine component, however it is to be understood that this prototype could in certain embodiments in fact be a first production model. The second, third, and subsequent instances (e.g., second combustor liner, third combustor liner, etc.) of the same aircraft engine component, which are expected to be substantially identical to one another but for small deviations due to coatings and/or manufacturing tolerances as explained above, may be tested numerically to ensure compliance with the flow requirements of that specific aircraft engine component. These subsequent instances may be referred to as production models of the aircraft engine component, since they are manufactured with the intent of being installed on aircraft engines should they be compliant with the flow requirements. However, it is to be understood that in certain embodiments, the second, third and subsequent instances of the manufactured component may be early-stage test parts, with subsequent instances of the produced part being validated in a similar manner and implemented in full-scale production. Hence, these second, third, and subsequent instances may not undergo extensive experimental testing. The numerical testing process, which will be described below, may use some experimental data gathered from the experimental flow testing of the prototype to derive a computed mass flow rate through the one or more apertures defined through a subsequent instance of the component. This computed mass flow rate may then be used to determine whether or not the component meets its flow requirements. If so, the component may be installed on an aircraft engine. If not, the component may either be send back to undergo further manufacturing (e.g., milling, drilling, etc) or simply recycled or discarded.

Referring now toFIG.4, a method of evaluating compliance of a component of the engine10for compliance with flow requirements through apertures defined therethrough is shown at400. The method400includes obtaining experimental data from experimental testing on a prototype of the component at402, the experimental testing including flowing a flow of a fluid through apertures of the prototype. Then, a digitized model of a production model of the component is obtained at404. The digitized model includes digitized apertures having geometrical data corresponding to that of apertures defined in the production model. More detail about the geometrical data are presented below with reference toFIG.5. A nominal mass flow rate through the digitized apertures may be computed using the geometrical data and the fluid flow characteristics (e.g., pressure, temperature, density of the fluid, etc.) at406. Then, the nominal mass flow rate of the digitized model may be corrected using the experimental data to obtain a computed mass flow rate of the production model at408. And, at least one parameter may be assigned to the production model, the at least one parameter indicative of installation approval of the production model of the component for installation on the aircraft engine when the computed mass flow rate is determined to be within a prescribed range of the flow requirements at410. In some cases, the production model of the component may be marked as non-airworthy for installation on the aircraft engine when the computed mass flow rate is determined to be outside the prescribed range of the flow requirements. The method400may include conducting the experimental testing. The digitized model may be obtained by scanning the components with any suitable device. The digitized model may be fed to a metrology software able to extract data about the scanned component (e.g., diameters of apertures, detecting apertures, and so on).

If the component complies with the flow requirements, the component may be sent to be installed in an engine. If the component does not comply with the flow requirements, the component may be recycled, repaired, adjusted, or simply discarded.

Consequently, each of the subsequent instances of the engine component (e.g., second combustor liner, third combustor liner, etc) may be evaluated for compliance with flow requirements without going through extensive experimental testing. This may offer substantial cost and time savings during the manufacturing of the different aircraft engine components.

The step406of the computing of the nominal mass flow rate through the apertures of the component may include calculating a coefficient of discharge Cdfrom the experimental data gathered from the experimental testing of the prototype, and a total flow passage area of the flow through the apertures of the prototype from a digitized model of the prototype. At which point, a reference nominal mass flow rate through the apertures of the prototype may be derived with the following equation:
qm=CdAtotal√{square root over (2ΔPρ)}

where qmis the reference nominal mass flow rate, Cdis a coefficient of discharge obtained from the experimental data, Atotalis the total flow passage area obtained from the geometrical data of a digitized model of the prototype, ρ is the density of the fluid flown through the apertures of the prototype during the experimental testing and taken upstream of the apertures of the prototype, and Δp is a pressure differential between pressures respectively upstream and downstream of the apertures of the prototype.

In the embodiment shown, the step402of obtaining the experimental data may include obtaining the coefficient of discharge Cdfrom the experimental data and an experimental mass flow rate through the apertures of the prototype from the experimental data. The computing of the coefficient of discharge Cdmay be done using the following equation:

Cd=4⁢qrig⁢D4-d4ϵπ⁢D2⁢d2⁢2⁢Δ⁢pp

where qrigis the experimental mass flow rate measured from the experimental data, D is the diameter of the conduit206(FIG.3) feeding the flow to the apertures of the prototype during the experimental testing, d is the mean diameter of maximum diameters of cylinders able to fit in the apertures of the prototype as detailed above, and is an expansion coefficient, which accounts for gas compressibility effects and corrects the fluid density, calculated as follows:

ϵ=1-[(0.351+0.256β4+0.93β8)⁢(1-γ⁢p-Δ⁢pp)]
where p is a pressure of the flow upstream of the apertures obtained from the experimental testing, γ is a specific heat ratio of the fluid (e.g., 1.4 for air), and β is a ratio of the mean diameter d to the diameter D of the conduit.

Referring back toFIG.3, the determining of the total flow passage area includes multiplying a number of the apertures (e.g., 40) by a mean flow passage area of the apertures. That is, the apertures defined through the aircraft engine components may present some variations and irregularities as explained above. An exemplary shape of a peripheral wall of one of the apertures is shown with a dashed line inFIG.3. These variations are amplified for better understanding. These variations may be caused by the manufacturing process of these components. For instance, the apertures may be drilled with a laser and/or a coating may be applied on the components. The laser drilling and/or coating may create these deviations from a purely cylindrical aperture as exhibited inFIG.3. These deviations may also have an impact on the flow rate flowing through the apertures as explained above.

To obtain the mean diameter d, the method400includes the calculation of a maximum diameter d of a cylinder14B′,16B′,16D′,18B′ (FIG.3) that is able to fit inside each of the apertures14B,16B,16D,18B. This maximum diameter may vary from aperture to aperture to aperture. If a cylinder having a diameter greater than the maximum diameter, it would not be able to extend throughout the aperture; it would be blocked by crests or other features defined by the peripheral wall bounding the aperture. The maximum diameter may correspond to a throat of the aperture, although the aperture are not necessarily shaped with convergent-divergent shape. Nevertheless, the smallest cross-sectional area of the aperture may be the limiting factor to the flow flowing through the aperture.

Referring also toFIG.5, the digitized model of any of the aircraft engine component may include a plurality of data points, which are represented by the markers500onFIG.5. The determining of the maximum diameter of a cylinder that fit insides each of the apertures may include generating a first cylinder representation502from the digitized model. The cylinder representation502may be created by fitting part of a cylindrical surface through some of the data points500. The part of the cylindrical surface fitted through the data points500may then be used to derive a radius (or diameter), an orientation (e.g., i, j, k) and a location of the cylindrical representation502. Cylinder representations502may have a substantially circular cross-sectional profile. The method of computing the cylindrical representation and the associated diameter is presented in U.S. patent application Ser. No. 16/920,868 filed on Jul. 6, 2020, the entire contents of which are incorporated herein by reference.

In some embodiments, a plurality of planes504,506,508may be obtained from the data points500. The apertures may be re-created from the data points500and for each of the planes504,506,508. Once the apertures are re-created, algorithm may be used to determine the maximum diameter of a cylinder502that may fit through them.

Once the maximum diameters of each of the apertures is determined, an average of the maximum diameters may be computed. This average diameter may be used to calculate the mean flow passage area by computing the area of a circle from the average diameter d. And, the total flow passage area may be computed by multiplying the number of apertures by the mean flow passage area.

The method400may then include computing a correction factor corresponding to a ratio of the reference nominal mass flow rate calculated above for the prototype to the experimental mass flow rate through the apertures of the prototype and measured from the experimental data. The correction factor is therefore obtained by dividing the reference nominal mass flow rate of the flow through the apertures of the prototype by the experimental mass flow rate through the same apertures of the prototypes, and which has been measured during the experimental testing. This correction factor therefore takes into account differences between simulations and actual experimental testing. These differences may be explained by complicated flow patterns of the flow flowing through the apertures, such as turbulence, secondary flows, friction, roughness of walls bounding the apertures, and so on. At step408, the correcting of the nominal mass flow rate therefore include multiplying the correction factor by the nominal mass flow rate through the apertures of the component.

To determine the robustness of this method, a plurality of experimental tests were conducted. Namely, the pressure differential across the apertures was varied to create a variation in the measured mass flow rate through the apertures. For each respective pressure differentials, a nominal mass flow rate was computed and a correction factor was derived. It was observed that, for a given aircraft engine component (e.g., combustor liner), the correction factor remains substantially unchanged regardless of the mass flow rate flowing through the apertures. Moreover, the pressure differential across the apertures of the prototype was varied and the coefficient of discharge was measured for each of the pressure differentials. This showed that the equations for the nominal mass flow rate and for the coefficient of discharged are suitable.

The nominal mass flow rate of the component (e.g., second combustor liner of the series of combustor liners) may then be obtained by calculating the total flow passage area of the flow through the apertures of the component from the digitized model of the component as explained above. Then, the equation below may be used to compute the nominal mass flow rate qmthrough the apertures of the component:
qm=CdAtotal√{square root over (2ΔPρ)}

where Cdis the coefficient of discharge obtained from the experimental data of the prototype as detailed above, Atotalis the total flow passage area of the apertures of the component, which is calculated using the method described above, ρ is the density of the fluid flown through the apertures of the prototype and taken upstream of the apertures of the prototype, and Δp is a pressure differential between pressures respectively upstream and downstream of the apertures of the prototype.

At which point, the step406of computing the mass flow rate through the apertures of the component includes multiplying the nominal mass flow rate qmof the component by the correction factor obtained from the experimental flow testing. Then, the computed mass flow rate may be compared with the range of the flow requirements to determine whether the component meets its flow requirements at the steps408A,408B.

The present disclosure illustrates a method to calculate the mass-flow through perforated aircraft engine components using the digitized part model. The method400also establish a correlation or correction factor between the actual airflow vs. expected or calculated airflow obtained by simulated model for each part number. Presently, each and every manufactured parts are undergoes airflow test in order to validate design conformity. The process is time consuming and required customized fixture and specialized personnel who experience on airflow test bench. This method400may allow the use of the digitized part model, which is readily available after certain operation such as coating or laser drilling operation, to establish an airflow mass discharge coefficient. This discharge coefficient and the digitized part is used to compute a correction factor. This correction factor is applied to calculate the mass flow rate of any give part of that specific part number. Henceforth, this method400may facilitate the avoidance of the logistics involves in the process of airflow evaluation. This method400may reduce the process cost, facilitate automation and improve stability since there are no human intervention may be required. The method400assumes that, for each component of the series of components, the flow requirements are consistent and that pressure and temperature differences and air density are being constant during the experimental flow testing.

The proposed method400may eliminate the use of airflow test bench and all the relevant processes, manpower and cost related to the flow test. The method400may not only facilitate the automation capability of a manufacturing cell but may also reduce the logistics behind it by optimizing the process and may eliminate human errors. The method may comprise the digitized part model obtain after the last operation before airflow (such as laser drilling, plasma coating, etc.). As a part of the standard inspection process, the manufactured part undergoes3D scanning operation that collect the digitized model and processed in a metrology software. This method is use the readily available3D scanned data of the part as an input hence no additional setup, fixture setting, equipment may be required. The digitized part model is processed in a metrology software using certain inspection algorithm and strategies.

With reference toFIG.6, an example of a computing device600is illustrated. For simplicity only one computing device600is shown but the system may include more computing devices600operable to exchange data. The computing devices600may be the same or different types of devices. The computing device600comprises a processing unit602and a memory604which has stored therein computer-executable instructions606. The processing unit602may comprise any suitable devices configured to implement the method400such that instructions606, when executed by the computing device600or other programmable apparatus, may cause the functions/acts/steps performed as part of the method400as described herein to be executed. The processing unit602may comprise, for example, any type of general-purpose microprocessor or microcontroller, a digital signal processing (DSP) processor, a central processing unit (CPU), an integrated circuit, a field programmable gate array (FPGA), a reconfigurable processor, other suitably programmed or programmable logic circuits, or any combination thereof.

The memory604may comprise any suitable known or other machine-readable storage medium. The memory604may comprise non-transitory computer readable storage medium, for example, but not limited to, an electronic, magnetic, optical, electromagnetic, infrared, or semiconductor system, apparatus, or device, or any suitable combination of the foregoing. The memory604may include a suitable combination of any type of computer memory that is located either internally or externally to device, for example random-access memory (RAM), read-only memory (ROM), compact disc read-only memory (CDROM), electro-optical memory, magneto-optical memory, erasable programmable read-only memory (EPROM), and electrically-erasable programmable read-only memory (EEPROM), Ferroelectric RAM (FRAM) or the like. Memory604may comprise any storage means (e.g., devices) suitable for retrievably storing machine-readable instructions506executable by processing unit502.

The methods and systems for evaluating an aircraft engine component for compliance with flow requirements described herein may be implemented in a high level procedural or object oriented programming or scripting language, or a combination thereof, to communicate with or assist in the operation of a computer system, for example the computing device600. Alternatively, the methods and systems for evaluating an aircraft engine component for compliance with flow requirements may be implemented in assembly or machine language. The language may be a compiled or interpreted language. Program code for implementing the methods and systems for evaluating an aircraft engine component for compliance with flow requirements may be stored on a storage media or a device, for example a ROM, a magnetic disk, an optical disc, a flash drive, or any other suitable storage media or device. The program code may be readable by a general or special-purpose programmable computer for configuring and operating the computer when the storage media or device is read by the computer to perform the procedures described herein. Embodiments of the methods and systems for evaluating an aircraft engine component for compliance with flow requirements may also be considered to be implemented by way of a non-transitory computer-readable storage medium having a computer program stored thereon. The computer program may comprise computer-readable instructions which cause a computer, or more specifically the processing unit602of the computing device600, to operate in a specific and predefined manner to perform the functions described herein, for example those described in the method400.

Computer-executable instructions may be in many forms, including program modules, executed by one or more computers or other devices. Generally, program modules include routines, programs, objects, components, data structures, etc., that perform particular tasks or implement particular abstract data types. Typically the functionality of the program modules may be combined or distributed as desired in various embodiments.

The embodiments described herein are implemented by physical computer hardware, including computing devices, servers, receivers, transmitters, processors, memory, displays, and networks. The embodiments described herein provide useful physical machines and particularly configured computer hardware arrangements. The embodiments described herein are directed to electronic machines and methods implemented by electronic machines adapted for processing and transforming electromagnetic signals which represent various types of information. The embodiments described herein pervasively and integrally relate to machines, and their uses; and the embodiments described herein have no meaning or practical applicability outside their use with computer hardware, machines, and various hardware components. Substituting the physical hardware particularly configured to implement various acts for non-physical hardware, using mental steps for example, may substantially affect the way the embodiments work. Such computer hardware limitations are clearly essential elements of the embodiments described herein, and they cannot be omitted or substituted for mental means without having a material effect on the operation and structure of the embodiments described herein. The computer hardware is essential to implement the various embodiments described herein and is not merely used to perform steps expeditiously and in an efficient manner.

The term “connected” or “coupled to” may include both direct coupling (in which two elements that are coupled to each other contact each other) and indirect coupling (in which at least one additional element is located between the two elements).

The technical solution of embodiments may be in the form of a software product. The software product may be stored in a non-volatile or non-transitory storage medium, which can be a compact disk read-only memory (CD-ROM), a USB flash disk, or a removable hard disk. The software product includes a number of instructions that enable a computer device (personal computer, server, or network device) to execute the methods provided by the embodiments.

The embodiments described in this document provide non-limiting examples of possible implementations of the present technology. Upon review of the present disclosure, a person of ordinary skill in the art will recognize that changes may be made to the embodiments described herein without departing from the scope of the present technology. Yet further modifications could be implemented by a person of ordinary skill in the art in view of the present disclosure, which modifications would be within the scope of the present technology.