Turbomachine stator assembly comprising rows of tandem blades

Abstract

The invention relates to a turbomachine stator assembly (32), characterized in that it comprises: inner and outer supports; two successive rows of stator blades (33, 35) defining a tandem configuration, comprising an annular row of upstream blades (33) extending substantially radially between the inner and outer supports, and an annular row of downstream blades (35), located downstream of the upstream blades (33), extending substantially radially between the inner and outer supports.

Description

[0001] DESCRIPTION

[0002] TITLE: Turbomachine stator assembly comprising tandem blade rows

[0003] TECHNICAL FIELD OF THE INVENTION

[0004]

[0001] The present invention relates to the general field of turbomachinery, and more specifically to the field of turbomachine rectifiers or stators, in particular turbomachine compressor rectifiers or stators.

[0005]

[0002] The invention applies to all types of aeronautical turbomachinery, and in particular to aircraft turbomachinery such as turbojets and turboprops. The invention can be applied to aircraft turbomachinery comprising at least one unfaired propeller, and also a pair of unfaired co-rotating or counter-rotating propellers, this type of turbomachine being also called "open rotor" or "propfan".

[0006]

[0003] The invention thus proposes a turbomachine stator assembly comprising successive rows of stator blades in a tandem configuration, a turbomachine compressor comprising such a stator assembly, as well as a turbomachine comprising such a stator assembly or such a compressor.

[0007] STATE OF THE ART

[0008]

[0004] Climate change is a major concern for many legislative and regulatory bodies worldwide. Indeed, various restrictions on carbon emissions have been, are being, or will be adopted by various states. In particular, an ambitious standard applies to both new types of aircraft and those already in service, requiring the implementation of technological solutions to bring them into compliance with current regulations. Civil aviation has been actively working for several years now to contribute to the fight against climate change.

[0009]

[0005] Technological research efforts have already led to very significant improvements in the environmental performance of aircraft. The Applicant takes into account the factors impacting all phases of design and development in order to obtain less energy-intensive and more environmentally friendly aeronautical components and products, the integration and use of which in civil aviation have moderate environmental consequences, with the aim of improving the energy efficiency of aircraft.

[0006] Consequently, the Applicant is constantly working to reduce its negative climate impact by using methods and operating virtuous development and manufacturing processes that minimize greenhouse gas emissions to the minimum possible in order to reduce the environmental footprint of its activity.

[0010]

[0007] This sustained research and development work focuses on new generations of aircraft engines, the lightening of aircraft, in particular through the materials used and lighter on-board equipment, the development of the use of electrical technologies to provide propulsion, and, as essential complements to technological progress, aviation biofuels.

[0011]

[0008] Thus, in order to reduce the size and weight of a turbomachine, it is possible to make the turbomachine modules more compact, and in particular the compressors, especially low-pressure compressors, by reducing the number of compression stages. However, in order to obtain an equivalent compression ratio with fewer stages, this requires driving the rotor at a higher rotational speed.

[0012]

[0009] It is possible to consider speeds close to the speed of sound. The compressor is then said to be transonic when at least a radially external part of the rotor moves at a speed greater than that of sound.

[0013]

[0010] Transonic compressors are characterized by a significant load on the rotor and stator blades because these must accelerate and straighten an airflow over a shorter axial distance. This load is even greater when the compressor includes variable-pitch blades, which, depending on the engine speed, impose a large deflection on the flow, and whose straightening therefore induces an even greater load. In this context, the blade camber must be significant, but excessive camber presents the risk of reaching boundary flow conditions, conditions in which the airflow separates from the blades under the effect of the adverse static pressure gradient. In the event of boundary layer separation, the acceleration and / or straightening of the flow is no longer guaranteed, which impairs the compressor's efficiency.

[0014]

[0011] Also, an excessive difference in angle of attack, also called "swing," between two extreme operating points of the turbomachine, combined with a large required deflection, can prove problematic for the compressor, and even more so if, in addition to the acceptable angle of attack range, there is a deceleration that is too difficult to manage. In particular, a point with a high negative angle of attack coupled with a high Mach number at the inlet can generate sonic blockages, just as a point with a high positive angle of attack coupled with a large deceleration to manage can present a high risk of separation. The operability of the compressor, as well as the rectification function for the correct supply of the downstream components, can then be jeopardized.

[0012] A particularly critical point in this regard is the final compression stage of the compressor, which must absolutely deliver a substantially axial flow.

[0015]

[0013] To straighten a flow in such a compressor while limiting the aerodynamic load experienced by the blades, a so-called "tandem" straightener can be used. This consists of a succession of two rows or grids of stator blades, each of which partially contributes to straightening the flow. Such examples are described in particular in patent applications BE 1 030421 A1, EP 2913 480 A1, EP 2409 002 A1, US 2020 / 0240283 A1 and DE 102018 108 940 A1.

[0016]

[0014] Thus, a first blade grid, whose purpose is to accommodate a flow with any angle of incidence, is followed by a second blade grid which is responsible for completing the remaining deviation with an operation always adapted thanks to the filtering work of the incidence of the first grid.

[0017]

[0015] However, the implementation of such a tandem configuration requires fine optimization of several geometric parameters which contribute to guaranteeing the operability required of the rectifier and its performance in terms of aerodynamic losses.

[0018]

[0016] In particular, since the tandem configuration consists of two blades, the positioning of the downstream blade relative to the upstream blade plays an important role in determining the flow state through this configuration. With a variable angle of attack, the behavior of the two blades is significantly affected, and therefore finding an optimal positioning of the downstream blade becomes essential for the operability of the tandem under different flight regimes. Furthermore, the positioning of the downstream blade is important because it typically conditions the behavior of the flow downstream of the low-pressure compressor, particularly in a gooseneck configuration. It is therefore necessary to control the parameters governing this positioning to ensure an efficient and operable tandem configuration.

[0019] DESCRIPTION OF THE INVENTION

[0020]

[0017] The invention aims to remedy at least partially the needs mentioned above and the disadvantages relating to the achievements of the prior art.

[0021]

[0018] To this end, the invention is the result of technological research aimed at significantly improving aircraft performance and, in this sense, contributes to reducing the environmental impact of aircraft.

[0022]

[0019] In particular, the invention aims to provide a turbomachine rectifier or stator configuration that enables the necessary flow deflection at all operating points of the turbomachine, while minimizing losses, and for any angle of attack, to axially straighten the upstream flow so as to properly supply a downstream area. It thus aims to provide a rectifier or stator design that improves the compactness of the turbomachine without compromising efficiency, across the entire operating range, including transonic, of the turbomachine.

[0023]

[0020] The invention thus relates, according to one of its aspects, to a turbomachine stator assembly, characterized in that it comprises:

[0024] - internal support and external support,

[0025] - two successive rows of stator blades defining a tandem configuration, comprising:

[0026] - an annular row of upstream blades extending substantially radially between the internal and external supports,

[0027] - an annular row of downstream blades, located downstream of the upstream blades, extending substantially radially between the internal and external supports,

[0028] an angular pitch being defined between two upstream blades, or two downstream blades, circumferentially adjacent, a circumferential spacing length between an upstream blade and a downstream blade circumferentially adjacent is defined between the trailing edge of one of the upstream blades and the downstream blade and the leading edge of the other of the upstream blade and the downstream blade, a total axial distance Cz of the tandem being defined between the leading edge of an upstream blade and the trailing edge of a downstream blade circumferentially adjacent and an axial offset AO between an upstream blade and a downstream blade circumferentially adjacent is defined between the trailing edge of the upstream blade and the leading edge of the downstream blade.

[0029]

[0021] Thanks to the invention, it is possible to design a rectifier or stator principle in tandem configuration capable of tolerating a large range of flow incidence and achieving a large flow deviation, through the optimization of the design parameters of the rows of blades constituting the stator assembly.

[0030]

[0022] Furthermore, the optimization proposed by the invention makes it possible to obtain better performance in terms of aerodynamic losses. In addition, the capabilities of the tandem configuration in terms of resistance to angle of attack and aerodynamic load (high deflection) are improved.

[0031]

[0023] The stator assembly according to the invention may further comprise one or more of the following characteristics taken individually or according to any possible technical combinations.

[0032]

[0024] The relative azimuthal offset Da of a downstream blade with respect to an upstream blade can be defined as: Da = t / S, where:

[0033] t is the circumferential spacing length between the upstream and downstream blades,

[0034] S is the angular pitch defined between two upstream blades or two circumferentially adjacent downstream blades.

[0035]

[0025] An inlet throat section Å is defined as the distance between an upstream blade and the tangent to the camber line of a circumferentially adjacent downstream blade, at the leading edge of the circumferentially adjacent downstream blade. An outlet throat section Ås is defined as the distance between a downstream blade and the tangent to the camber line of a circumferentially adjacent upstream blade, at the trailing edge of the circumferentially adjacent upstream blade. Furthermore, a throat section ratio Å / Ås is defined as the ratio between the inlet throat section Å and the outlet throat section Ås.

[0036]

[0026] The azimuthal offset Da can be chosen such that 0.85 < Da < 0.95.

[0037]

[0027] Furthermore, the axial offset AO can be chosen such that 0 < AO / Cz < 7.5%, in particular to obtain a maximum angle of attack range. Specifically, with such values ​​of the azimuth offset Da and the axial offset AO, tandem configurations can be obtained that will maximize the angle of attack range (or "swing").

[0038]

[0028] The axial offset AO can also be chosen such that 0 < -AO < 8%, particularly to obtain minimal aerodynamic losses. In particular, with such values ​​of the azimuthal offset Da and the axial offset AO, tandem configurations can be obtained that will exhibit minimal aerodynamic losses.

[0039]

[0029] The throat area ratio Ae / As, defined as the ratio between the inlet throat area Ae and the outlet throat area As, can be chosen such that 1.05 < Ae / As < 1.20, preferably such that 1.05 < Ae / As < 1.18, in order to provide the best tandem performance by maximizing the angle of attack range (or "swing") and minimizing aerodynamic losses. Such a range of values ​​can allow for re-energization of the boundary layer of the lower upper surface of the blade, improving its maximum deflection capacity without lift-off.

[0040]

[0030] The number of upstream blades and the number of downstream blades can be identical. The internal support can be an internal ferrule, and the external support can be an external ferrule or a housing.

[0041]

[0031] The upstream paddle rope may be different from the downstream paddle rope.

[0042] In other words, the chord ratio can be chosen such that Rc

[0043]

[0044] 1.

[0045]

[0032] In addition, the invention also relates, according to another of its aspects, to a turbomachine compressor, in particular a high pressure compressor or a low pressure compressor, characterized in that it comprises a stator assembly as defined above.

[0046]

[0033] Preferably, the compressor is a low pressure compressor.

[0047]

[0034] Preferably still, the stator assembly belongs to the last compression stage of the compressor.

[0048]

[0035] In addition, the invention also relates, according to another of its aspects, to a turbomachine, characterized in that it comprises a stator assembly as defined above or a compressor as defined above.

[0049]

[0036] The turbomachine may include a moving wheel and a separating nozzle, arranged downstream of the moving wheel and separating an annular airflow into a primary flow through an internal channel and a secondary flow through an external channel, the stator assembly being arranged in the internal channel upstream of a gooseneck shape.

[0050]

[0037] Furthermore, the turbomachine may include a high-pressure compressor arranged downstream of the stator assembly, the gooseneck shape being in particular located between the stator assembly and the high-pressure compressor.

[0051] BRIEF DESCRIPTION OF THE FIGURES

[0052]

[0038] Other advantages, purposes and special features of the invention will become apparent from the following non-limiting description of at least one embodiment of the present invention, with reference to the accompanying figures, in which:

[0053] Figure 1 schematically represents, in axial cross-section, an example of a turbomachine comprising a stator assembly according to the invention; Figure 2 schematically illustrates, in a transverse view with respect to the radial direction, an example of a stator assembly according to the invention comprising two rows of stator blades, only one blade being represented per row;

[0054] Figure 3 is a cross-sectional view along AA' of Figure 2, illustrating in a very schematic way, perpendicular to the radial direction, the example of a stator assembly comprising two rows of stator blades, with only two blades shown per row; and

[0055] Figure 4 schematically represents, according to a partial axial cross-sectional view, an example of the implementation of a stator assembly according to the invention in a turbomachine similar to that of Figure 1.

[0056]

[0039] Throughout these figures, identical references may designate identical or analogous elements.

[0057]

[0040] Furthermore, the different parts shown in the figures are not necessarily shown to a uniform scale, in order to make the figures more legible. DETAILED DESCRIPTION OF THE INVENTION

[0058]

[0041] Throughout the description, given by way of non-limiting example, it is noted that the terms upstream and downstream are to be considered with respect to a principal direction F of normal gas flow (from upstream to downstream) for a turbomachine 1. Furthermore, the axis X of the turbomachine 1 is called the radial axis of symmetry of the turbomachine 1. The axial direction of the turbomachine 1 corresponds to the axis of rotation X of the turbomachine 1. A radial direction of the turbomachine 1 is a direction perpendicular to the axis X of the turbomachine 1.

[0059]

[0042] Furthermore, unless otherwise specified, the adjectives and adverbs axial, radial, axially, and radially are used with reference to the aforementioned axial and radial directions. In addition, unless otherwise specified, the terms inside and outside are used with reference to a radial direction such that the inside portion of an element is closer to the X-axis of the turbomachine 1 than the outside portion of the same element.

[0060]

[0043] Furthermore, the term height refers to a dimension measured along the longest dimension of the blades, which may be substantially radial. The chord is the straight line segment connecting the leading edge to the trailing edge in a plane perpendicular to a radius. The camber refers to the median curve connecting the leading edge to the trailing edge equidistant from the lower and upper surfaces.

[0061]

[0044] Figure 1 shows, in an axial cross-sectional view, an example of a turbomachine 1 according to the invention.

[0062]

[0045] An inner casing 2 guides a primary flow F1 passing successively through a low-pressure compressor 4, a high-pressure compressor 4', a combustion chamber 6 and high- and low-pressure turbines 8 before exiting through a nozzle 10. The energy of the combustion drives the turbines 8 in rotation about the longitudinal axis X of the turbomachine 1. The turbines 8 drive the compressors 4, 4', directly via transmission shafts or indirectly by means of a reduction gear 23. The turbines 8 also drive in rotation a rotor 12 with fan blades 1 which set in motion a secondary flow F2.

[0063]

[0046] A fairing 16 and a nacelle 18 define a passage 19 through which the secondary flow F2 flows. Structural arms 20 carry the forces between the nacelle 18 and the inner casing 2.

[0064]

[0047] An annular row of stator blades 22, also called OGVs for "outlet guide vanes," is arranged downstream of the rotor 12 to straighten the secondary flow F2.

[0048] The turbomachine 1 has a separation nozzle 48 to separate the annular flow F into two primary flows F1 and secondary flows F2. The invention can be applied downstream of any type of flow separation nozzle and is not limited to the separation nozzle 48.

[0065]

[0049] Each compressor 4, 4' is formed of a succession of compression stages. Each stage includes rotating or rotor blades providing momentum to the flow and stationary or stator blades straightening the flow direction of the flow.

[0066]

[0050] In this example, and in no way limiting, the low pressure compression 4 of the turbomachine 1 includes a final compression stage, as detailed later with reference to Figure 4, in which is located a stator assembly 32 according to the invention comprising rows of blades in tandem.

[0067]

[0051] An example of a stator assembly 32 according to the invention is described below with reference to figures 2 and 3. The stator assembly 32 thus has a tandem configuration with two rows or grids of stator blades forming a double grid.

[0068]

[0052] Specifically, a first row or grid of upstream stator blades 33 is provided, moderately deviating and intended to accommodate the flow of the primary flow F1 with a wide range of angle of incidence and to provide the second row or grid of downstream stator blades 35 with a supply that is always uniform at every point of operation of the turbomachine 1. The second row or grid of stator blades 35 operates in quasi-single incidence, and completes the deviation without having to manage the variations in incidence.

[0069]

[0053] Figures 2 and 3 allow for the definition of design parameters that help define the geometry of the stator assembly 32, which here, and without limitation, forms the stator of the last compression stage of the low-pressure compressor 4, also comprising a rotor upstream of the stator. The rotor has rotor blades that accelerate the fluid flow thanks to the energy transmitted by the drive shaft, while the stator transforms the kinetic energy into pressure thanks to the shape of the stator blades 33, 35.

[0070]

[0054] The stator assembly 32 comprises an upstream row of blades 33 and a downstream row of blades 35, extending from a radially internal support 36 to a radially external support 38, as shown in Figure 2. The internal supports 36 and external supports 38 may be mounting platforms for attachment to a supporting structure. The internal supports 36 and external supports 38 may also be ferrules describing 360° around the X-axis or angular sectors of ferrules describing an angle of a few degrees or a few tens of degrees around the X-axis. The upstream blades 33 and downstream blades 35 are supported by a common internal support and a common external support.

[0055] The camber of the upstream stator blades 33 and downstream blades 35 is denoted A1 and A2 respectively and is represented by dashed lines. The angular pitch between two circumferentially adjacent upstream 33 or downstream 35 blades is noted S as seen in Figure 3, and can be identical for both rows of blades 33, 35.The radial height, along a radial axis perpendicular to the X axis, of each upstream blade 33 and downstream blade 35 is respectively noted H33 and H35 in Figure 2.

[0071]

[0056] The chord of the upstream blades 33 and downstream blades 35, which connects the leading edge to the trailing edge, is respectively denoted C33 and C35. The chord C33 of the upstream blades 33 may be different from the chord C35 of the downstream blades 35.

[0072]

[0057] Furthermore, the axial length AO formed between the two rows of blades, upstream 33 and downstream 35, is also shown in Figures 2 and 3. This length AO can be considered an axial overlap length when the position of the blades 33, 35 is such that they at least partially overlap, thus exhibiting axial overlap over at least part of their radial height. In the example shown in Figure 3, the rows of blades 33, 35 overlap.

[0073]

[0058] This length AO defines the axial offset (or “Axial Overlap” in English) between an upstream blade 33 and an adjacent downstream blade 35, which corresponds to the axial distance between the trailing edge of an upstream blade 33 and the leading edge of an adjacent downstream blade 35, the length AO being positive in the case of overlap.

[0074]

[0059] In addition, the spacing between the trailing edge of an upstream blade 33 and the leading edge of an adjacent downstream blade 35 is quantified by a circumferential spacing length between blades of the tandem, denoted t in Figure 3, measured perpendicular to the X axis in a similar way to the measurement of the pitch S between two adjacent blades of the same row.

[0075]

[0060] Furthermore, the total axial distance of the tandem, corresponding to the axial distance between the leading edge of an upstream blade 33 and the trailing edge of an adjacent downstream blade 35, is noted Cz in Figure 3.

[0076]

[0061] The geometry of the upstream stator blades 33 and downstream stator blades 35 can be further described by the inlet throat section Ae and the outlet throat section As. Thus, as shown in Figure 3, the inlet throat section Ae is the distance between an upstream blade 33 and the tangent to the camber line of an adjacent downstream blade 35, at the leading edge of the adjacent downstream blade 35. The outlet throat section As is the distance between a downstream blade 35 and the tangent to the camber line of an adjacent upstream blade 33, at the trailing edge of the adjacent upstream blade 33.

[0077]

[0062] Design parameters are advantageously provided for the stator assembly 32 according to the invention. These parameters, in particular the axial offset AO and azimuthal offset Da, are advantageously optimized to minimize losses. Similarly, the throat generated between the two adjacent upstream 33 and downstream 35 blades constitutes an aerodynamic parameter that can also be optimized to achieve the same objectives. The convergence of this throat ensures the minimization of losses, and an optimization of the ratio between Àe and Às ensures optimal operability by maximizing the angle of attack range.

[0078]

[0063] The relative azimuthal offset Da (also called "clocking" in English) of a downstream blade 35 with respect to an upstream blade 33 is defined as: Da = t / S.

[0079]

[0064] In particular, the azimuthal offset Da can be chosen such that 0.85 < Da < 0.95, and the axial offset AO is chosen such that 0 < AO / Cz < 7.5%. Advantageously, these ranges of values ​​for these parameters make it possible to obtain tandem configurations that will maximize the range of incidence (or "swing").

[0080]

[0065] The azimuthal offset Da can further be chosen such that 0.85 < Da < 0.95, and the axial offset AO is chosen such that 0 < -AO / Cz < 8%. Advantageously, these ranges of values ​​for these parameters make it possible to obtain tandem configurations that will exhibit minimal aerodynamic losses.

[0081]

[0066] Furthermore, as can be seen in Figure 3, the tandem configuration is characterized by the presence of a channel delimited by the tangent to the leading edge of the blade 35 downstream and the tangent to the trailing edge of the blade 33 upstream. A converging channel with a throat area ratio Ae / As, defined as the ratio between the inlet throat area Ae and the outlet throat area As, such that 1.05 < Ae / As < 1.20, can advantageously offer the best tandem performance by maximizing the angle of attack range ('swing') and minimizing aerodynamic losses.

[0082]

[0067] Advantageously, the choice of values ​​for the parameters mentioned above can be made according to the goal(s) sought, in particular in terms of maximizing the range of incidence and minimizing losses.

[0083]

[0068] Figure 4 illustrates an example of a possible implantation of the stator assembly 32 according to the invention in a turbomachine 1 such as that shown in Figure 1, preferably downstream of a flow separator.

[0084]

[0069] The annular flow F is split into two flows F1 and F2. The annular flow F flows into an annular vein 42, and the flows F1 and F2 flow respectively into a primary annular vein 44 and a secondary annular vein 46. The separation of the flows is carried out by the separating nozzle 48.

[0085]

[0070] Directly or not upstream of the separation nozzle 48 is a rotating or rotor assembly in the form of a movable wheel 50, in particular a blower, whose blades 52 extend radially upstream of the primary annular veins 44 and secondary 46.

[0086]

[0071] The stator assembly 32, consisting of an annular row of stator blades 33 and a row of stator blades 35 forming the tandem, is preferably arranged in the primary annular flow 44, in the low-pressure compressor 4 which also includes rotor blades 30, and precedes a portion of the flow shaped like a gooseneck 54 which is arranged upstream of the high-pressure compressor 4'. Thus, the stator assembly 32 constitutes the last blades 33, 35 of the low-pressure compressor 4 and allows the primary flow F1, coming from the upstream stages and entering the flow 44 by first encountering fixed blades 56 or straightening arms, in order to properly supply the gooseneck 54 located downstream.

[0087]

[0072] The low-pressure compressor 4 may include variable-pitch stator vanes, or VSV for "Variable Stator Vanes," and the stator assembly 32 may include only the stator vanes of the low-pressure compressor 4 that are not variable-pitch. The low-pressure compressor 4 may have between 1 and 4 compression stages, each formed of at least one row or annular grid of rotor vanes directly followed by at least one row or grid of stator vanes.

[0088]

[0073] Of course, the invention is not limited to the embodiments just described. Various modifications can be made to them by a person skilled in the art.

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