Abstract
A method for producing a structural component including the following steps: a) forming of the structural component; b) removing material from the structural component; c) thermal treatment of the structural component following step b; d) removing further material from the structural component following step c.
Claims
1. A method for producing a structural component with the following steps: a) forming the structural component; b) removing material from the structural component; c) thermal treatment of the structural component following step b; d) removing further material from the structural component following step c.
2. The method according to claim 1, further comprising the following step: e) non-destructive inspection, in particular ultrasonic inspection, of the structural component following step d.
3. The method according to claim 1, wherein further a calculation or simulation of an internal stress field of the structural component is performed.
4. The method according to claim 3, wherein the removing of material in step b is performed according to a result of the calculation or simulation.
5. The method according to claim 1, wherein material is removed at the same or at neighboring areas of the structural component in steps b and d.
6. The method according to claim 1, wherein material is removed at a locally limited area of the structural component in step b.
7. The method according to claim 1, wherein the forming in step a comprises forging.
8. The method according to claim 1, wherein material is removed by cutting in at least one of step b and step d.
9. The method according to claim 1, wherein, following step a and prior to step b, the structural component has an outer shape that has a total dimension (G) in one direction, wherein in step b material is removed in this direction over a length that is a seventh of the total dimension (G).
10. The method according to claim 1, wherein, following step a and prior to step b, the structural component has an outer shape that has a local dimension (L) in one direction, wherein in step b material is removed in this direction over a length that is a third of the local dimension (L).
11. The method according to claim 1, wherein the finished structural component is a disc for a gas turbine engine.
12. The method according to claim 11, wherein in step b material is removed at a radially inwardly positioned area of the structural component.
13. A disc for a gas turbine engine, wherein the disc is embodied for the purpose of supporting multiple airfoils, wherein the disc is produced or can be produced according to a method according to claim 1.
14. A gas turbine engine for an aircraft, comprising: a core engine that comprises a turbine, a compressor and a core shaft that connects the turbine to the compressor; a fan that is positioned upstream of the core engine, wherein the fan comprises multiple fan blades; a gearbox that can be driven by the core shaft, wherein the fan can be driven by means of the gearbox with a lower speed than the core shaft, and at least one disc according to claim 13.
15. The gas turbine engine according to claim 14, wherein: the turbine is a first turbine, the compressor is a first compressor and the core shaft is a first core shaft; the core engine further comprises a second turbine, a second compressor and a second core shaft that connects the second turbine to the second compressor; and the second turbine, the second compressor and the second core shaft are arranged in such a manner that they rotate with a higher speed than the first core shaft.
Description
[0050] Embodiments will now be described by way of example only, with reference to the Figures, in which:
[0051] FIG. 1 shows a sectional side view of a gas turbine engine;
[0052] FIG. 2 an enlarged sectional side view of an upstream section of a gas turbine engine;
[0053] FIG. 3 shows a partial cut-away view of a gearbox for a gas turbine engine;
[0054] FIG. 4 shows a method for producing a structural component;
[0055] FIG. 5A to 5G show a structural component during multiple processing steps in the method according to FIG. 4;
[0056] FIG. 6A shows a cross section through a comparative structural component in multiple phases of manufacturing in a comparative method;
[0057] FIG. 6B shows a cross section through a structural component in multiple phases of manufacturing according to a method according to FIG. 4;
[0058] FIG. 7A shows a cross section through the finished comparative structural component according to FIG. 6A with indicated isobars;
[0059] FIG. 7B shows a cross section through the finished structural component according to FIG. 6B with indicated isobars;
[0060] FIG. 8 shows a cross section through a structural component in multiple phases of manufacturing according to a method according to FIG. 4; and
[0061] FIG. 9 shows a cross section through a structural component in multiple phases of manufacturing according to the method according to FIG. 4.
[0062] FIG. 1 shows a gas turbine engine 10 having a principal rotational axis 9. The gas turbine engine 10 comprises an air intake 12 and a fan 23 that generates two airflows: a core airflow A and a bypass airflow B. The core engine 11 comprises, as viewed in the axial flow direction, a low-pressure compressor 14, a high-pressure compressor 15, combustion device 16, a high-pressure turbine 17, a low-pressure turbine 19 and a core thrust nozzle 20. An engine nacelle 21 surrounds the gas turbine engine 10 and defines the bypass channel 22 and a bypass thrust nozzle 18. The bypass airflow B flows through the bypass channel 22. The fan 23 is attached at the low-pressure turbine 19 via a shaft 26 and an epicyclic planetary gearbox 30, and is driven by the low-pressure turbine 19.
[0063] During operation, the core airflow A is accelerated and compressed by the low-pressure compressor 14 and guided into the high-pressure compressor 15, where further compression takes place. The air that is discharged from the high-pressure compressor 15 in a compressed state is directed into the combustion device 16 where it is mixed with fuel and combusted. The resulting hot combustion products are then propagated through the high-pressure and the low-pressure turbine 17, 19, thus driving it before being discharged through the nozzle 20 for providing a certain thrust. The high-pressure turbine 17 drives the high-pressure compressor 15 via a suitable interconnecting shaft 27. The fan 23 usually provides the greatest portion of the propulsive thrust. The epicyclic planetary gearbox 30 is a reduction gear.
[0064] An exemplary arrangement for a geared fan gas turbine engine 10 is shown in FIG. 2. The low pressure turbine 19 (see FIG. 1) drives the shaft 26, which is coupled to a sun gear 28 of the epicyclic planetary gearbox 30. Located radially outwardly of the sun gear 28 and intermeshing therewith are multiple planetary gears 32 that are coupled with each other by a planet carrier 34. The planet carrier 34 guides the planetary gears 32 in such a manner that they rotate synchronously around the sun gear 28 whilst enabling each planet gear 32 to rotate about its own axis. Via linkages 36, the planet carrier 34 is coupled to the fan 23 in such a manner that it causes it to rotate about the rotational axis 9. An external gear or ring gear 38 that is coupled via linkages 40 to a stationary supporting structure 24 is located radially outside of the planetary gears 32 and intermeshes therewith.
[0065] Note that the terms low-pressure turbine and low-pressure compressor as used herein may be taken to refer to the turbine stage with the lowest pressure or the compressor stage with the lowest pressure (i.e., not including the fan 23) and/or refer to the turbine and compressor stage that are connected by the interconnecting shaft 26 with the lowest rotational speed in the engine (i.e., not including the gearbox output shaft that drives the fan 23). In some documents, a low-pressure turbine and a low-pressure compressor referred to herein may alternatively also be known as an intermediate-pressure turbine and an intermediate-pressure compressor. Where such alternative nomenclature is used, the fan 23 may be referred to as a first or lowest pressure stage.
[0066] The epicyclic planetary gearbox 30 is shown by way of example in greater detail in FIG. 3. The sun gear 28, planetary gears 32 and the ring gear 38 respectively have teeth at their circumference to facilitate intermeshing with the other gears. However, for reasons of clarity, only exemplary portions of the teeth are illustrated in FIG. 3. Here, four planetary gears 32 are illustrated, although it will be apparent to the person skilled in the art that more or fewer planetary gears 32 may be provided within the scope of the claimed invention. Practical applications of a planetary epicyclic gearbox 30 generally comprise at least three planetary gears 32.
[0067] The epicyclic planetary gearbox 30 illustrated by way of example in FIGS. 2 and 3 is a planetary gearbox in which the planetary carrier 34 is connected to an output shaft via linkages 36. However, it is also possible to use any other suitable type of a planetary gearbox 30. By way of further example, the planetary gearbox 30 may comprise a star arrangement, in which the planet carrier 34 is supported in a fixed manner, with the ring gear (or external gear) 38 being allowed to rotate. In such an arrangement, the fan 23 is driven by the ring gear 38. By way of further alternative example, the gearbox 30 may be a differential gearbox in which the ring gear 38 as well as the planet carrier 34 are allowed to rotate.
[0068] It will be obvious that the arrangement shown in FIGS. 2 and 3 serves merely as an example, and the scope of the present disclosure also comprises various alternatives. Purely by way of example, any suitable arrangement may be used for arranging the gearbox 30 in the gas turbine engine 10 and/or for connecting the gearbox 30 to the gas turbine engine 10. By way of further example, the connections (such as the linkages 36, 40 in the example according to FIG. 2) between the gearbox 30 and other parts of the gas turbine engine 10 (such as the input shaft 26, the output shaft and the stationary support structure 24) may have any certain degree of stiffness or flexibility. By way of further example, any suitable arrangement of the bearings between rotating and stationary parts of the gas turbine engine 10 (for example between the input and output shafts of the gearbox 30 and the fixed structures, such as for example the gearbox casing) may be used, and the disclosure is not limited to the exemplary arrangement of FIG. 2. For example, where the gearbox 30 has a star arrangement (described above), the person skilled in the art would readily understand that the arrangement of output and support linkages and bearing locations would typically be different from that shown in FIG. 2.
[0069] Accordingly, the present disclosure extends to a gas turbine engine 10 having any arrangement of gearbox styles (for example star arrangement or planetary arrangements), support structures, input and output shaft arrangement, and bearing locations.
[0070] Optionally, the gearbox may drive additional and/or alternative components (e.g. the intermediate-pressure compressor and/or a booster compressor).
[0071] Other gas turbine engine 10 to which the present disclosure may be applied may have alternative configurations. For example, such engines may have an alternative number of compressors and/or turbines and/or an alternative number of interconnecting shafts. By way of further example, the gas turbine engine 10 shown in FIG. 1 has a split flow nozzle 20, 22, meaning that the flow through the bypass channel 22 has its own nozzle that is separate from and arranged radially outside of the engine core nozzle 20. However, this is not to be taken in a limiting manner, and any aspect of the present disclosure may also apply to engines in which the flow through the bypass channel 22 and the flow through the core 11 is intermixed or combined in front of (or upstream of) a single nozzle which may be referred to as a mixed flow nozzle. One or both nozzles may have a fixed or variable cross section (independently of whether a mixed or a partial flow is present). Whilst the example described herein relates to a turbofan engine, the disclosure may apply, for example, to any type of gas turbine engine, such as an engine with an open rotor (in which the fan stage is not surrounded by an engine nacelle) or to a turboprop engine, for example.
[0072] The geometry of the gas turbine engine 10, and components thereof, is defined by a conventional axis system, comprising an axial direction (which is aligned with the rotational axis 9), a radial direction (in the bottom-to-top direction in FIG. 1), and a circumferential direction (perpendicular to the page in the FIG. 1 view). The axial, radial and circumferential directions are mutually perpendicular.
[0073] FIG. 4 illustrates the individual steps of a method for producing a structural component 100A. In the subsequent description of this method, the FIGS. 5A to 5G are additionally referred to, showing the structural component 100A in multiple steps of manufacturing.
[0074] In a first step S1, the structural component 100A is formed. Here, different methods are possible, in particular primary forming and/or forming. In the example shown in FIG. 5A, a forging material 300 is formed in a forging device 400. For easier reference, the product that is produced in this process is referred to as the structural component 100A from this step on, even if this final shape will only be reached after multiple further method steps. Following forming, internal stresses (inner stresses, residual stresses) are present inside the structural component 100A. The manifestations of the internal stresses have consequences for the mechanical properties and the stability of the structural component 100A.
[0075] In a further step S2, which can also be performed prior to the previously described step S1, a calculation or simulation of an internal stress field of the structural component 100A is performed (e.g. based on the finite element method). The calculation/simulation can comprise an optimization algorithm. The simulation can comprise an iterative process simulation. The calculation or simulation can be based on a model, in particular a numerical model. The calculation or simulation can be performed by means of a computer 402 (see in particular FIG. 5B). Here, that internal stress field may for example be calculated or simulated that the structural component 100A would have if a certain amount of material was removed at a certain position of the structural component. Here, multiple variants can be calculated/simulated, and respectively that variant or those variants can be selected in multiple iterations that has/have the best internal stress field. The best internal stress field is e.g. an internal stress field that is as homogenous as possible, and/or an internal stress field that [has] a compression stress at areas of the structural component 100A that are subjected to tension during the later use of the structural component 100A. Alternatively or additionally, it is also possible to calculate/simulate variants that have a volume addition with respect to an initial shape (in particular locally). In this case, the structural component is formed in step S1 with the volume addition (and the step S2 is performed prior to step S1, as is illustrated by the dashed lines in FIG. 4).
[0076] An input parameter of the simulation (apart from a geometry and a material of the structural component, e.g. a steel, nickel or titanium alloy) can be the indication of an area that is supposed to have increased residual compressive stresses (or alternatively tensile residual stresses, depending on the requirement). The simulation can further comprise a service life calculation. The result of the simulation may e.g. be an indication as to how much material has to be removed at what position of the structural component 100A, and/or how large the initial shape of the structural component 100A has to be planned (e.g. 100A according to FIG. 5A).
[0077] To influence the internal stresses (more precisely, the internal stress field) in the structural component 100A, material is removed from the structural component 100A in one subsequent step S3. The material can be removed at a locally limited area of the structural component 100A, e.g. only at one side or only a portion of a side of the structural component 100A. According to FIG. 5B, the material is removed by means of lathe 401 (alternatively, other types of machining are conceivable). The structural component 100A is attached at a drive and rotated by means of the drive. A blade or the like is moved, in a manner controlled by a computer 402, to the rotating structural component 100A, so that material of the structural component 100A is removed. The computer 402 can select the lathe 401 depending on the result of the calculation or simulation.
[0078] In a basic subsequent step S4, the structural component 100A is thermally treated. This may for e.g. be performed in a furnace 403 (see FIG. 5C). In an exemplary structural component made of steel, a thermal treatment can be performed as follows. In a warm-up period of e.g. 4 hours, the structural component is heated up to e.g. 600 C. This temperature may e.g. be maintained for 6 hours. Subsequently, cooling e.g. with a rate of 35 C. per hour is performed. Other materials may e.g. be heated up to such a temperature that they begin to plastically flow. Through the thermal treatment, the internal stresses of the structural component 100A can be reduced.
[0079] Following the thermal treatment in step S4, material is again removed in a subsequent step S5. According to FIG. 5D, this can again be carried out with a lathe 401. Here, too, the removing of material can occur depending on a result of the simulation, e.g. through selection by a computer 402. The material is optionally removed at a location of the structural component 100A that adjoins the area at which material had already been removed in the previous step S3.
[0080] Optionally, the structural component 100A is ultrasonically inspected in a step S6 following step S5. According to FIG. 5E, this can be performed with an ultrasound testing device 404. The ultrasound testing device can be coupled to the computer 402 to analyze the measuring results of the ultrasound testing device 404. In an ultrasonic inspection e.g. material defects can be detected. Structural components 100A in which a material defect has been detected, can be taken out or reworked, e.g. by once more performing one or multiple of the steps S3 to S5 (in that case, inspection optionally takes place in a renewed step S6).
[0081] In an optional subsequent step S7, material is again removed from the structural component 100A (see FIG. 5F), optionally at a location at the structural component 100A that adjoins an area at which material has already been removed in a previous step S3, S5. In this step, the final geometry of the structural component 100A can be created, if this has not already happened before.
[0082] FIG. 5G shows the finished structural component 100A, which in the present case is the ring gear 38 of the gearbox 30 according to FIG. 3, by way of example.
[0083] FIG. 6A shows a comparative structural component V that will be explained in more detail in the following.
[0084] FIG. 6B shows multiple contours of a structural component 100B during manufacturing in the method according to FIG. 4. This structural component 100B is a disc of the gas turbine engine 10 according to FIGS. 1 and 2, concretely a disc of the high-pressure turbine 17 (see FIG. 1). The structural component 100B is configured for supporting vanes or blades at an outer circumference. The structural component 100B has an inner passage opening (for receiving the shaft 26). FIG. 6B shows the structural component 100B in the cross section, wherein only a half is shown. The edge of the structural component 100B that is located at the left in FIG. 6B delimits the passage opening.
[0085] Following the step S1 of forming the structural component 100B (in the present example by means of forging), the structural component 100B has a shape that is referred to as the black forging shape 200 here. After removing material in step S3, the structural component 100B has a shape that is referred to as the thermal treatment shape 201 here. As can be seen in FIG. 6B, material has been locally removed here in the area of the passage opening, thus forming an (inner circumferential) groove, for example. The thermal treatment shape 201 differs from the black forging shape 200. The thermal treatment shape 201 has at least one local deviation from the black forging shape 200. In the present example, the thermal treatment shape 201 has an indentation that had not been formed in the black forging shape 200. By removing material in step S3, a curvature is for example introduced or reinforced in a surface of the structural component 100B. Alternatively or additionally, a curvature is removed from the surface of the structural component 100B, or is reduced.
[0086] The black forging shape 200 of the structural component 100B has a total dimension G (with respect to the sections of the structural component 100B that are connected to each other in the cross section along the direction of this total dimension G). As shown in FIG. 6B, the total dimension G can be measured the in radial direction of the structural component 100B that is embodied as a disc.
[0087] At the position at which material is removed in step S3 (prior to thermal treatment in step S4), the black forging shape 200 of the structural component 100B has a local dimension L that is parallel to the total dimension G. In step S3, a material volume up to a depth T (in parallel to the total dimension G and to the local dimension L) is removed from the structural component 100B at this position. The depth T can e.g. be 1/10 to of the total dimension G, in particular 1/9 to , in particular 1/7. Optionally, the depth T is to of the local dimension L, in particular .
[0088] In step S3 (prior to thermal treatment in step S4), material can e.g. be removed from the structural component 100B at such a position at which the black forging shape 200 and a testing shape 202 and/or a final shape 203 differ from each other most strongly. In other words, material can be removed prior to thermal treatment at a location where particularly much, in particular the largest amount of material has to be removed until the finished structural component is formed.
[0089] The thermal treatment shape 201 has a smaller volume than the black forging shape 200. Material is again removed after the thermal treatment in the step S4. In that case, the structural component has the testing shape 202. The testing shape 202 has a smaller volume than the thermal treatment shape 201.
[0090] Following ultrasonic inspection, material is again removed, so that the structural component 100B has the final shape 203.
[0091] FIG. 7A shows the finished comparative structural component V.
[0092] FIG. 7B shows the structural component 100B with the final shape 203. The finished structural component 100B has a flange 102 and an outer circumference 103 regarding the passage opening 101 auf. Further, isobars with respectively constant internal stress are indicated in FIG. 7.
[0093] FIG. 6A shows the contours of the comparative structural component V which is thermally treated directly after forming. In this structural component, the black forging shape 200 and the thermal treatment shape 201 are thus identical. FIG. 7A shows the final shape of this comparative structural component V, with isobars are being indicated.
[0094] As a comparison of FIGS. 7A and 7B shows, the structural component 100B shown in FIG. 7B which has been manufactured according to the method according to FIG. 4, has particularly evenly distributed internal stresses in the area of the passage opening 101, more evenly than in the comparative structural component V. In addition, it has been shown that the internal stresses in the area of the passage opening 101 and of the flange 102 are stronger than compression stresses. These compression stresses can be compensated by tensile stresses performed externally during use of the structural component 100B in the gas turbine engine 10. In this manner, the structural component 100B can have a longer service life.
[0095] FIG. 8 shows a structural component 100C which is formed with a material addition in step S1. In this case, the black forging shape 200 has a reinforced local dimension L as compared to the local dimension L of an adjacent area. The area with the reinforced local dimension L adjoins the area at which material is removed in step S3 (before thermal treatment). The difference between the reinforced local dimension L and the local dimension L of the adjacent area can e.g. be 1/10 to of the total dimension G, in particular 1/9 to , in particular 1/7. Alternatively, the local dimension L can be larger by 1/7 than a local dimension at the same position of an initial shape (prior to a corresponding calculation or simulation).
[0096] FIG. 9 shows a structural component 100D which is formed with a material addition in step S1_similar to the structural component 100C according to FIG. 8. Compared to an initial shape (e.g. the black forging shape 200 according to FIG. 6A or 6B), the black forging shape 200 of the structural component 100D has a (local) volume addition. Here, the volume addition is provided in the area of the outer circumference 103 of the structural component 100D (at which blades can later be attached). The size of this volume addition represents the result of a calculation and/or simulation in which in particular the internal stress field of the structural component 100D has been optimized. Thus, the total dimension G is enlarged by for example to , e.g. by of the original total dimension with respect to the original total dimension. Here, the thermal treatment shape 201 is also locally enlarged with respect to the initial shape (e.g. the thermal treatment shape of the comparative structural component V). Optionally, the testing shape 202 is unchanged with respect to the testing shape of the comparative structural component V according to FIG. 6A. The thermal treatment shape 201 has a smaller volume than the black forging shape 200. The testing shape 202 has a smaller volume than the thermal treatment shape 201. The final shape 203 has a smaller volume than the testing shape 202.
[0097] Alternatively to the described simulation, empirical values can be included. Further, test measurements can be taken at structural components in which material is to be removed in different amounts or at different locations in step S3. The material can then be removed in steps S3, S5 and/or S7 according to one of these structural components. Optionally, simulation or test measurements is/are performed with a structural component size, and the results of the simulation or test measurements are extrapolated to another structural component size. The method according to FIG. 4 is then performed with these results (optionally without step S2).
[0098] It is to be understood that the invention is not limited to the embodiments described herein and that various modifications and improvements can be realized without departing from the concepts described herein. Except where they are mutually exclusive, any of the features can be used separately or in combination with any other features, and the disclosure extends to all combinations and sub-combinations of one or multiple features described herein, and includes the same.
PARTS LIST
[0099] 9 main rotational axis [0100] 10 gas turbine engine [0101] 11 core engine [0102] 12 air intake [0103] 14 low-pressure compressor [0104] 15 high-pressure compressor [0105] 16 combustion device [0106] 17 high-pressure turbine [0107] 18 bypass thrust nozzle [0108] 19 low-pressure turbine [0109] 20 core thrust nozzle [0110] 21 engine nacelle [0111] 22 bypass channel [0112] 23 fan [0113] 24 stationary support structure [0114] 26 shaft [0115] 27 connecting shaft [0116] 28 sun gear [0117] 30 gearbox [0118] 32 planetary gears [0119] 34 planetary carrier [0120] 36 linkage [0121] 38 ring gear [0122] 40 linkage [0123] 100A-100D structural component [0124] 101 passage opening [0125] 102 flange [0126] 103 outer circumference [0127] 200 black forging shape [0128] 201 thermal treatment shape [0129] 202 testing shape [0130] 203 final shape [0131] 300 forging material [0132] 400 forging device [0133] 401 lathe [0134] 402 computer [0135] 403 furnace [0136] 404 ultrasound testing device [0137] A core airflow [0138] B bypass airflow [0139] G total dimension [0140] L local dimension [0141] L reinforced local dimension [0142] T depth [0143] V comparative structural component
