Patent Application
20250179921
This claims priority to German Patent Application DE 102024101944.1 filed on Jan. 24, 2024 which is hereby incorporated by reference herein.
The present invention relates to a turbofan engine having a low-pressure turbine.
The present subject matter relates to a turbofan engine having a low-pressure turbine, whose blade rings are typically coupled to a fan via a gearbox. An engine of this type is also referred to as “geared turbofan (GTF) engine.” During operation, the fan rotates at a lower speed than the low-pressure turbine (sometimes also referred to as “low-pressure turbine module”) due to the gear reduction. The term “low-pressure turbine” as used herein refers to that section of the aircraft gas turbine of the turbofan engine which is downstream of (i.e., comes after) an upstream-most turbine module immediately downstream of the combustor in the direction of flow. The low-pressure turbine may drive a middle shaft and/or an innermost shaft of the turbofan engine.
The present invention addresses the technical problem of providing an advantageous turbofan engine having a low-pressure turbine.
The present invention provides a turbofan engine. The low-pressure turbine thereof has at least four stages. At least one of the four stages has a transonic airfoil cascade, i.e., a transonic blade cascade and/or a transonic vane cascade. As will be discussed in detail below, a transonic airfoil cascade can be understood to be an airfoil cascade or airfoil ring in which, at least in some areas of an airfoil passage defined along the transonic airfoil cascade, the flow reaches velocities greater than or equal to a sonic velocity of a fluid flowing through the transonic airfoil cascade during operation.
The transonic design of the low-pressure turbine may, for example, be an expression of a modified or optimized work distribution in the engine, namely a work output that is proportionally shifted axially further toward the rear (i.e., in a downstream direction in relation to the working gas). All things considered, this can increase the efficiency of the turbofan engine for the same total work output, for example. In simplified terms, a modified work split within the turbofan engine from the high-pressure turbine into the low-pressure turbine is achieved while at the same time increasing the overall efficiency of the turbofan engine. Consequently, however, the flow velocities in the low-pressure turbine increase, which is why it is provided with one or more transonic airfoil cascade(s).
As an alternative to the low-pressure turbine with at least four stages, it may also have exactly three stages. At least two of these three stages each have a transonic airfoil cascade, i.e., at least one of a transonic blade cascade and a transonic vane cascade. In this case, too, the work output is shifted axially toward the rear (as explained above), thereby increasing the overall efficiency of the turbofan engine, the at least two transonic airfoil cascades accounting for the resulting higher flow velocities. In general, the transonic design can reduce any influence on the flow, for example, prevent significant additional acceleration or deceleration of the fluid flow near or in the supersonic range (both of which could be detrimental).
Preferred embodiments will be apparent from the dependent claims and the entire disclosure. In the description of the features, a distinction is not always specifically made between the different claim categories. For example, when a turbofan engine designed for a specific operation is described, such description should be read to refer also to a corresponding use or a corresponding method for operating the turbofan engine. Furthermore, information regarding the low-pressure turbine always also refers to the turbofan engine having such a low-pressure turbine.
When reference is made to a flow “during operation,” this may refer, for example, to a design under take-off or, in particular, cruise conditions, i.e., to the Aerodynamic Design Point (ADP). Unless expressly indicated otherwise, the terms “axial,” “radial,” and “circumferential” as well as the corresponding directions (axial direction, etc.) are taken with respect to a centerline of the turbofan engine, which may coincide with an axis of rotation of the blade cascades. The terms “upstream” and “downstream” are taken with respect to a direction of flow of the working gas, i.e., of the hot gas flowing through the turbine.
In a preferred embodiment, a transonic airfoil of the transonic airfoil cascade has a suction side surface including an at least approximately linear portion that has an at least approximately linear profile when viewed in a sectional plane of the airfoil. The sectional plane of the airfoil is orthogonal to a stacking axis, which is obtained as a line connecting the centroids of area of the cross-sectional profiles, a respective centroid of area being taken in a respective sectional plane of the airfoil at the respective cross-sectional profile. With the at least approximately linear profile, any influence on the fluid flow at the suction side surface can be reduced (no additional acceleration and also no additional deceleration).
In a preferred embodiment, the at least approximately linear profile in the sectional plane of the airfoil extends over at least 10%, preferably 20%, of an axial width of the transonic airfoil (possible upper limits may be 60%, 50%, or 40%, or 30%). The axial width is taken as the axial chord length of the cross-sectional profile in the respective sectional plane of the airfoil.
In a preferred embodiment, the at least approximately linear portion extends radially over at least 10%, preferably at least 20%, particularly preferably at least 30% or 40%, of a radial length of the transonic airfoil. In the case of a linear portion with a radial extent that varies over its axial width, its maximum radial extent is taken as a basis, which maximum radial extent may be, for example, at the downstream end of the linear portion. In more detail, the radial extent and the radial length may be taken at the same axial position, for example in a section perpendicular to the centerline. If the suction side surface in this section also has a proportional extent in the circumferential direction (the airfoil is, for example, leaned), only the respective radial portion is taken as a basis.
In a preferred embodiment, the at least approximately linear profile has a curvature when viewed in the sectional plane of the airfoil, the curvature deviating from a straight line by at most 5%, preferably by at most 3% or at most 2%, particularly preferably by at most 1%. For illustration and further details, reference is made to the exemplary embodiment (see
In a preferred embodiment, the linear portion in the sectional plane of the airfoil is located immediately adjacent a trailing edge radius or a trailing edge curvature of the transonic airfoil. In a simplified summary, the linear portion merges directly into the trailing edge portion; i.e., it extends, for example, tangentially to or into the trailing edge radius.
In an embodiment, a total area of the linear portion increases in the downstream direction, i.e., becomes larger, for example, from cascade to cascade or from stage to stage.
In preferred embodiments, at least or exactly two, at least or exactly three, or at least or exactly four stages of the low-pressure turbine have a transonic airfoil cascade. Moreover, in one embodiment, the low-pressure turbine has exactly four stages. Alternatively, exactly three stages may be provided, and each of the three stages may have a transonic airfoil cascade.
In a preferred embodiment, the stator vanes and/or the rotor blades are uncooled airfoils. In other words, the stator vanes and/or the rotor blades do not have any cooling channels, cooling openings, or the like, to reduce the temperature of the airfoil during operation by means of a fluid (e.g., air, water, or other fluids).
In a preferred embodiment, the “aspect ratio” of the stator vane or rotor blade, which defines the ratio of the height to the axial width of the stator vane or rotor blade, is greater than or equal to 5, preferably greater than or equal to 4, particularly preferably greater than or equal to 3.5. The height of the airfoil is the distance in the circumferential direction (i.e., about the axis of rotation of the airfoil cascade) between the “uppermost” and “lowermost” ends of the airfoil, and the axial width is the distance in the axial direction between the “forwardmost” and the “rearwardmost” ends of the airfoil. In contrast to strongly curved airfoils, airfoils which are rather flat in profile correspond with these aspect ratio values, which allows for improved control of the flow situation (acceleration, velocity).
In a preferred embodiment, the transonic design of the turbofan engine is configured such that, as a start, only the last, or, in other words, the downstream-most airfoil cascade (blade cascade) is a transonic airfoil cascade. Further, the immediately adjacent upstream airfoil cascades (vane and blade cascades) may then, by-and-by, also be transonic. Thus, blade cascades and vane cascades could then be transonic, alternately from the flow exit end toward the flow inlet. Such extension may be carried out airfoil cascade by airfoil cascade, or may directly involve several immediately adjacent airfoil cascades. Alternatively, non-adjacent airfoil cascades (vane and/or blade cascades) could have a transonic design.
The invention further relates to a use of low-pressure turbine in a herein-described turbofan engine, where the flow through the at last one transonic airfoil cascade is transonic.
The invention will now be described in more detail with reference to an exemplary embodiment. The individual features may also be essential to the invention in other combinations within the scope of the other independent claims, and, as above, no distinction is specifically made between different claim categories.
During operation, transonic airfoil 50 is exposed to a fluid flow and has a suction side (the region above transonic airfoil 50 in
In the context of the present disclosure, the terms “approximately linear portion” and “approximately linear profile” are understood to refer to a shape/profile where a quotient of distance 82 to axial width 75 (distance/axial width) is at most 5%, in the order given at least 3%, 2% or 1%, or differs by no more than this value from the respective quotient of 0% of a linear shape. This substantially straight extent of suction side surface 55 is not necessarily limited to the 25% of the axial width, but may, for example, also extend axially further forward (the 25% range is, however, taken as a basis for the consideration of the linearity and should be met at least there).
Preferably, the maximum deviation of 5% (of at most 3%, 2%, or 1%) is present at least at a radial airfoil height between 50% and 90%. However, the maximum deviation criterion may additionally also be met radially inwardly and/or outwardly of this range (50%-90%).
| Number | Date | Country | Kind |
|---|---|---|---|
| 102024101944.1 | Jan 2024 | DE | national |
