Compressor blade

Abstract

The present invention relates to a compressor blade for a gas turbine engine. In particular, the blade of the compressor is modified according to predetermined requirements of both aerodynamic stability and fuel economy in a plurality of planes.

Description

Technical Field

[0001] This invention relates to an improved compressor blade arrangement. Specifically, but not exclusively, this invention relates to a compressor blade for use in low-pressure compressors (LPCs) and / or high-pressure compressors (HPCs). Background Technology

[0002] A typical gas turbine engine includes a pair of compressors: a first upstream low-pressure compressor and a second downstream high-pressure compressor. These compressors compress the air entering the engine in two stages, and the compressed gas is then delivered to the combustor, where fuel is introduced and the mixture is ignited. The operation of a gas turbine engine is well known to those skilled in the art.

[0003] Two key performance requirements for compressor blades are high efficiency at an operating point (typically the engine's cruise speed) and high stability across the engine's entire operating speed range. High efficiency is crucial for meeting engine specific fuel consumption targets, while high stability is necessary to ensure safe operation under a variety of flight conditions. These two requirements are often clearly conflicting, meaning designers must sacrifice efficiency to achieve sufficient internal stability within the compressor (stability implies airflow and vibration stability). In effect, a trade-off is made between ensuring safe engine operation at all engine speeds and achieving the best possible fuel economy.

[0004] One problem with the operation and design of gas turbine engines is that they operate at very different speeds. For example, during taxiing or waiting for takeoff, the engine will operate at relatively low speeds, with the compressor section operating at subsonic speeds (far less than 450 m / s). -1 The engine operates at significantly higher speeds during climbs or cruise, with the compressor section operating at near-supersonic speeds (at or close to 450 m / s). -1 It operates at a speed of ).

[0005] Achieving trade-offs in compressor blade design often requires other modifications to the engine. For example, blade shapes are designed to meet efficiency targets while achieving high stability through the introduction of technologies such as variable guide vanes and bleed-out mechanisms. These technologies prevent air from entering the compressor at low speeds to prevent, for example, engine stall. These technologies also allow for the construction of engines that maintain stability and achieve acceptable fuel consumption levels.

[0006] However, while both technologies effectively allow for engine design, they also increase engine weight and complexity, which in turn increases product cost and fuel consumption.

[0007] The inventors of this invention have developed a surprising alternative to compressor blade design that allows for simultaneous efficiency and stability, and further provides the possibility of eliminating variable guide vanes and bleed arrangement structures in gas turbine engines. This can significantly improve engine maintenance schedules. Summary of the Invention

[0008] The inventive aspects disclosed herein are described below.

[0009] From a first aspect of the invention described herein, a compressor blade is provided for a compressor comprising a plurality of blades, each blade extending radially from a central hub and including a pressure surface on one side of the blade and a suction surface on an opposite side of the blade, wherein the pressure surface and the suction surface extend from a common leading edge of the blade to a trailing edge of the suction surface and a trailing edge of the pressure surface, and wherein the trailing edges of the pressure surface and the suction surface of the blade are connected by a third trailing edge surface.

[0010] The invention described herein provides a novel approach to compressor blade shape. The strategy is to design the suction side of the blade at the design point based on achieving high efficiency, while simultaneously designing the pressure side of the blade based on achieving high-stability aerodynamics. These two parameters are implemented simultaneously in the blade design, which is counterintuitive in blade design.

[0011] In conventional blade design, designers must choose a compromise blade configuration that compromises either efficiency, blade stability, or both. Designers have achieved operable blades under these constraints, but a fully optimized blade design that is simultaneously and independently optimized was previously impossible.

[0012] The inventors of this invention have created a counterintuitive and inherently different blade configuration. According to this strategy of simultaneously and independently optimizing blade performance, the trailing edge of the blade—that is, the trailing edge surface connecting the trailing edges of the pressure and suction surfaces at the "downstream" end of the blade—becomes thicker than that of a conventional blade. This thicker blade would otherwise be undesirable, as it would introduce additional losses due to eddy current shedding from the trailing edge.

[0013] However, according to the invention described herein, the inventors have maintained a thicker trailing edge surface and overcome the problem of eddy current generation. Specifically, the dual-optimized blade includes a modified trailing edge, which, in an arrangement, is shaped into a sinusoidal pattern that advantageously controls the shedding and thus minimizes losses.

[0014] In fact, the pressure surface is the surface that "first" faces the flow when the blades rotate, while the suction surface is the surface that "faces backward".

[0015] Therefore, it is possible to achieve blades that are fully optimized for both stability and efficiency under cruise and other operating conditions.

[0016] The third trailing edge surface is the surface that defines the trailing surface of the blade. In conventional blades, the trailing edges of the suction and pressure surfaces are irregular or tapered to define a single smooth surface where the two surfaces meet.

[0017] According to the invention described herein, the third trailing edge surface is not a single line or a single edge, but rather a surface that itself has a boundary defined between the blade hub and the blade tip and between the trailing edge of the pressure surface and the trailing edge of the suction surface.

[0018] Advantageously, the blade profiles of the pressure and suction surfaces can be optimized for the overall performance of the engine without compromise. Therefore, the surface profile of the pressure surface of the blade can be shaped according to a predetermined profile that provides predetermined aerodynamic stability during operation. Simultaneously, the surface profile of the suction surface of the blade can be shaped according to a predetermined profile that provides predetermined fuel efficiency during operation. Thus, both desired profiles can be accommodated without compromise.

[0019] For clarity, as mentioned above, the pressure surface always faces the flow surface "first" as the blades rotate, while the suction surface is the "backward" facing surface.

[0020] The length of the suction surface, measured from the leading edge to the trailing edge of the suction surface, can be greater than the length of the pressure surface, measured from the leading edge to the trailing edge of the pressure surface. This allows for a greater curvature of one surface relative to the other and allows for independent optimization of both surfaces.

[0021] The third trailing edge surface can have a uniform thickness measured in the circumferential direction between the trailing edge on the suction side and the trailing edge on the pressure side. In such an arrangement, the width of the third trailing surface (the tail surface of the blade), measured in the circumferential direction, is substantially constant along the radial extent of the blade.

[0022] The leading edge of the pressure surface can smoothly intersect the leading edge of the suction surface to define the smoothly curved leading edge surface of the blade. Therefore, the air impacting the blade is smoothly separated on both sides of the blade and directed toward the two trailing edges of the suction and pressure surfaces.

[0023] The pressure surface extending from the smooth leading edge can advantageously have curvature to allow air to exit the trailing edge of the pressure surface in a first direction; and the suction surface extending from the smooth leading edge can have curvature to allow air to exit the trailing edge of the suction surface in a second direction. In practice, the air exiting the rear of the blade can comprise two distinct airflows; each airflow has a slightly different direction due to the curvature of the pressure and suction surfaces.

[0024] As described above, the curvature of each surface is selected according to different and independent operating requirements, which results in two independent directions of airflow from the trailing or rear surface of the blade.

[0025] The difference in direction will depend on the difference in the necessary airflow curvature of the pressure surface and the suction surface, but can be separated at a predetermined angle according to the operating characteristics of the engine.

[0026] Advantageously, when measured along the blade radius, a portion of the third trailing edge surface of the blade alternates between a first circumferential direction and a second opposite circumferential direction. This effectively provides a serpentine profile for the third trailing surface. Specifically, unlike surfaces with two straight sides, each side of the third surface undulates left and right from the hub or near the hub to the blade tip or near the blade tip.

[0027] In one example, a portion of the third trailing edge surface may have a sinusoidal profile when measured along the blade radius. As described herein, the sinusoidal profile may have a predetermined amplitude and frequency according to the design requirements needed to shed eddies at a specific frequency.

[0028] The wavy or sinusoidal shape can be uniform along the blade from hub to tip. Alternatively, and advantageously, the third trailing edge profile may include:

[0029] (i) a first substantially straight portion extending radially from the hub; and

[0030] (ii) A second alternating portion that extends from the intersection of the first and second portions toward the tip of the blade.

[0031] Therefore, the third rear surface can be divided into multiple zones with different surface profiles or geometries that vary from the portion of the blade closer to the hub to the portion of the blade closer to the tip.

[0032] The location of the contour change on the third surface can be selected according to design requirements. For example, the intersection of the first and second parts can be at a predetermined radius measured from the hub of the blade.

[0033] Specifically, the predetermined radius can correspond to a radius at which the airflow velocity at the leading edge of the blade exceeds approximately Mach 1 when the compressor operates under its normal cruise conditions. The term "cruise" refers to the engine's operating state when the aircraft is at normal cruise speed and altitude. The reason for applying the invention above this radius is that the design challenge of balancing cruise and partial speed operating points is greatest when the inlet flow is supersonic. In fact, it is beneficial to begin the undulation of the blade at a location where the inlet airflow on the blade exceeds Mach 1 (during cruise).

[0034] As described above, the undulating profile can be determined based on the engine's expected operating conditions and using industry-standard fluid dynamics and finite element analysis equipment. This allows for the determination of the amplitude and frequency of the sinusoidal or undulating profile to optimize eddy current damping effects.

[0035] In one example, the sinusoidal undulating profile, measured from a baseline extending radially outward along the blade, can be uniform in amplitude. In another example, the amplitude of the undulating profile, measured from a baseline extending radially outward along the blade, can be non-uniform.

[0036] For example, when measured radially outward from the blade, the frequency of the undulations can be uniform, or when measured radially outward from the blade, the frequency of the undulations can be non-uniform.

[0037] In one example, the surface of the third trailing edge can be substantially flat or planar. In another arrangement, the surface can be wholly or partially convex and / or concave to induce three-dimensional eddy shedding and damping from that surface.

[0038] On the other hand, a method for manufacturing compressor blades as described herein is provided.

[0039] From another aspect of the invention described herein, a compressor blade is provided for a compressor comprising a plurality of blades, each blade extending radially from a central hub and including a pressure surface on one side of the blade and a suction surface on the opposite side of the blade, wherein the pressure surface and the suction surface extend from a common leading edge of the blade to trailing edges of the suction surface and the pressure surface, and wherein the trailing edges of the pressure surface and the suction surface intersect to form a single trailing edge, and wherein a portion of the intersecting trailing edge has a circumferentially oscillating profile extending radially along the blade.

[0040] In another aspect, a gas turbine engine including a compressor is provided, the compressor comprising compressor blades according to the disclosure herein. In yet another aspect, an aircraft including one or more engines incorporating the blades described herein is provided.

[0041] In another view, a method for manufacturing a blade for use in a compressor is provided, the compressor having a plurality of blades, wherein each blade is arranged to extend radially from a central hub and includes a pressure surface on one side of the blade and a suction surface on the opposite side of the blade, wherein the pressure surface and the suction surface extend from a common leading edge of the blade to a trailing edge of the suction surface and a trailing edge of the pressure surface, and wherein the trailing edges of the pressure surface and the suction surface of the blade are connected by a third trailing edge surface.

[0042] This manufacturing method can be implemented using conventional machining techniques, such as computer numerical control machining centers, which can process complex three-dimensional shapes, such as blade profiles.

[0043] Alternatively, the blades can be manufactured using additive manufacturing technology.

[0044] Various additive manufacturing techniques can be used to generate contours and apply the surface modifications of this invention. In fact, these geometries make additive manufacturing particularly suitable because complex internal geometries and surface finishes can be produced without the need for grinding or polishing tools.

[0045] The term additive manufacturing is intended to refer to a technology in which blades are produced layer by layer until a complete blade or hub and blade structure are formed.

[0046] Examples of readily applicable additive manufacturing technologies include powder bed techniques such as electron beam welding, selective laser melting, selective laser sintering, or direct metal laser sintering. Alternative technologies may include wire feeding processes, such as electron beam forming.

[0047] The invention extends to methods of forming the structures described herein using additive manufacturing. Attached Figure Description

[0048] aspects of the invention will now be described by way of example only with reference to the accompanying drawings, in which:

[0049] Figure 1 A cross-sectional view of a gas turbine engine incorporating a compressor according to the present invention is shown;

[0050] Figure 2 A schematic diagram of a single compressor blade is shown;

[0051] Figure 3 Two configurations of compressor blades viewed radially inward are shown;

[0052] Figure 4 The tip portion of a compressor blade, which combines optimized pressure and suction surfaces, is shown.

[0053] Figure 5 This explains the generation and shedding of eddies in the modified compressor blades as described in this article;

[0054] Figures 6A to 6C Two example arrangements of conventional compressor blades and modified compressor blades as described herein are shown;

[0055] Figure 7 An alternative trailing edge profile of a compressor blade according to the invention described herein is shown as viewed in the axial direction of the compressor.

[0056] Figure 8 An alternative trailing edge profile of a compressor blade according to the invention described herein is shown as viewed in the circumferential direction of the compressor.

[0057] Figure 9A and Figure 9B A perspective view showing the outline of the pipe illustrates its geometry; and

[0058] Figure 10A and Figure 10B Example hybrid functions are shown, which can be used to combine a highly efficient blade with a highly stable blade.

[0059] While this teaching is readily available in various modifications and alternative forms, specific embodiments thereof have been illustrated by way of example in the accompanying drawings and will be described in detail herein. However, it should be understood that the drawings and their detailed description are not intended to limit the scope to the particular forms disclosed, but rather to encompass all modifications, equivalents, and alternatives falling within the spirit and scope defined by the appended claims.

[0060] As used in this specification, the words “including,” “comprising,” and similar terms should not be interpreted in an exclusive or exhaustive sense. In other words, they are intended to mean “including but not limited to.”

[0061] It will be appreciated that the features of the various aspects of the invention described herein can be readily and interchangeably used in any suitable combination. It will also be appreciated that the invention covers not only the individual embodiments, but also combinations of the embodiments discussed herein. Detailed Implementation

[0062] Figure 1 A cross-sectional view of a gas turbine engine 1 is shown, which may be incorporated with compressor blades according to the invention as described in detail below.

[0063] Technicians will understand the main components of the gas turbine engine and its operation. In summary, engine 1 includes an air intake 2 that allows air to flow into the engine and onto a fan 3 located at the upstream end of the engine. All components are housed within the engine compartment 4.

[0064] The engine includes a bypass passage downstream of the fan and a central engine core containing a compressor, combustor, and turbine. The engine core consists of a first low-pressure compressor (LPC) 5 and a second high-pressure compressor (HPC) 6. This multi-stage compressor arrangement draws air from the ambient pressure and temperature and compresses it to high temperature and pressure. The compressed air is then delivered to a combustion chamber 7, where fuel is injected and combustion occurs.

[0065] Combustion gases are exhausted from the rear of combustion chamber 7 and first impact the high-pressure turbine 9, then the second low-pressure turbine 10, before exiting the rear of the engine through the core nozzle 11. Thrust from the engine is generated by two airflows: the first airflow comes from the fan nozzle 8 (receiving thrust from the fan), and the second airflow comes from the exhaust gases at the core nozzle 11.

[0066] The present invention relates to blades that exist in both low-pressure compressor 5 and high-pressure compressor 6.

[0067] Each compressor consists of a series of compressor blade rows, each series connected to the central shaft via a hub. By adjusting the spacing between the successive blades, the compression ratio along the compressor can be increased, as understood by those skilled in compressor design.

[0068] Figure 2 This is an exaggerated schematic diagram of a single compressor blade 12 located on hub 13. Each blade includes a leading edge 14 and a trailing edge 15. Only a single blade is shown, but it should be understood that each hub 13 includes multiple circumferentially positioned and radially extending blades.

[0069] The hub rotates in the direction indicated by arrow 16, and air impacts the leading edge 14 and is guided along the pressure surface 17 toward the trailing edge 15 (the opposite side of the blade is called the suction surface of the blade). The airflow is as shown by arrow series 18.

[0070] As the hub and blades rotate, the distal end 19 of each blade also rotates in the circumferential path around the hub (and within the compressor housing, not shown). The blade tip can move between subsonic to supersonic or near-supersonic speeds, a phenomenon known as transonic speeds. It can be appreciated that while the revolutions per minute (RPM) of the blade is uniform along its radial length, the instantaneous tangential rotational speed increases with increasing distance from the tip radius of the blade operating at its highest rotational speed.

[0071] As discussed above, Figure 2 The diagram shown greatly exaggerates the blade curvature. In fact, the blade appears almost straight.

[0072] The reason blades are conventionally straight is that this shape provides maximum efficiency at high speeds, and therefore maximum fuel economy during cruise. This leads to the preference for straight blades. Conversely, the optimal design for aerodynamic stability is a more curved shape, particularly a more curved shape towards the trailing edge. This leads to the preference for curved blades.

[0073] Designers typically have to compromise between straight and curved shapes and choose the right location to optimize engine operation during flight.

[0074] However, as described herein, the inventors of this invention have created a configuration that benefits from both of these arrangements while continuing to operate efficiently.

[0075] This reference Figure 3 illustrate.

[0076] Figure 3 This shows that when radially inward, that is, in Figure 2 The outlines of the two compressors when viewed from the direction shown by reference A.

[0077] The dashed lines indicate high-efficiency blade profiles, while solid lines indicate blade profiles for high aerodynamic stability. As shown, the optimal design for cruise is a generally straight profile extending outwards from the hub. As indicated by arrow B, the optimal design for stability is more curved. In the stability-optimized profile, the section of blade C extending from the leading edge 14 to the trailing edge 15 is substantially straight, and the trailing edge profile is curved, as shown... Figure 3 The straight yet curved profile of the blade is shown in the solid line.

[0078] As discussed above, the inventors have determined that these two profiles can be combined to simultaneously achieve the advantages of both high stability and high efficiency. This is in Figure 4 As shown in the image.

[0079] Figure 4 A hybrid compressor blade design with modified tips or trailing edges is shown. Specifically, both the pressure side 17 and the suction side 20 are modified according to the optimal profile for stability and efficiency.

[0080] This is achieved by allowing the suction side profile to follow the... Figure 3 The optimal shape of the profile shown by the dashed line is achieved. Simultaneously, the relative pressure side of the blade has a profile curved according to the optimal stability profile (corresponding to...). Figure 3 (The solid line shown).

[0081] This configuration has been found to provide an optimized blade design that can operate over a wide range of engine speeds, from coasting to cruise, and offers improved fuel efficiency.

[0082] Furthermore, because aerodynamic constraints can be taken into account without compromising efficiency requirements, the compressor can be adapted to operate at lower speeds without stalling, which subsequently eliminates the need for a bleed valve in the transition pipe or compressor inlet. This significantly reduces the complexity and weight of the engine core and offers numerous associated advantages.

[0083] However, as well as Figure 4As shown, the profile of the trailing edge of the blade cannot have the conventional sharp or converging profiles of the pressure and suction sides. Instead, due to the different profiles of the pressure and suction sides, the thicker portion 21 is used as the distal trailing edge of the blade.

[0084] Figure 4 This illustrates how the mixed blade surface terminates at the trailing edge of the blade. Specifically, a surface extending between the edge defining the trailing edge of the suction surface and the edge defining the trailing edge of the pressure surface defines a third trailing edge surface, or tip 21. As shown, as a result of these two profiles, a thicker tip 21 is formed. This tip forms the (third) trailing edge of the blade. The trailing edge surface extends radially from the hub to the farthest portion of each blade and extends circumferentially between the trailing edges of the pressure and suction surfaces. This trailing edge profile is further described below.

[0085] Providing each compressor blade with a pressure surface optimized for aerodynamic stability and a suction surface optimized for efficiency offers numerous technical advantages, including, but not limited to:

[0086] - The possibility that a venting system is not required;

[0087] - The possibility of not requiring complex variable still blades;

[0088] - Improved fuel efficiency

[0089] - Improved reliability and simpler maintenance; and

[0090] -Simplify manufacturing.

[0091] However, the inventors also determined that while hybrid blades can offer many technical advantages, such as Figure 4 The resulting trailing edge configuration shown may produce harmful aerodynamic effects in the form of vortices.

[0092] Figure 5 A first embodiment of the hybrid blade of the present invention described herein is shown.

[0093] refer to Figure 5 The airflow over the blades can be observed. As described above, the trailing edge S... t The surface must extend between the end of the suction surface A and the end of the pressure surface B. This results in an unconventional, thicker trailing edge surface S. t .

[0094] The inventors have determined that the high-speed air leaving the trailing edge of the blade generates vortices 22, such as Figure 5 As shown.

[0095] The generation of eddies at the trailing edge, or more specifically, shedding, leads to pressure loss in the compressor, which is undesirable because it impairs the efficiency of the compressor (and therefore the engine).

[0096] However, the inventors have determined a solution to the problem caused by the trailing edge profile S. t The methods addressing the resulting technical problems. Solving the problems of eddy current generation and shedding further improves the efficiency of the blades and compressor described in this paper. Therefore, the combination of hybrid blade surfaces and modified trailing edges offers greater technical advantages than existing compressor technologies.

[0097] Now from the reference Figure 6A and Figure 6B Begin describing the modified trailing edge.

[0098] Figure 6A and Figure 6B The illustrations show the trailing edge arrangement of a conventional compressor blade and an arrangement of the trailing edge according to the invention described herein. Figure 6A and Figure 6B The view shown corresponds to an axial view of the compressor from rear to tail, that is, a view of the downstream end of the compressor viewed forward toward the air inlet.

[0099] like Figure 6A As shown, conventional compressor blades have a generally uniform shape. That is, the trailing edge extends radially outward from the hub 13 toward the housing 23. The tip 21 shown has a radial space between the inner surface of the outer compressor housing and the tip of the blade.

[0100] In conventional blades, the trailing edges of the pressure surface P and the suction surface S are at the centerline PS. C Gathering places, such as Figure 6A As shown. This is the abrupt termination of the two surfaces midway between the suction surface and the pressure surface.

[0101] Figure 6B This illustrates a modified trailing edge configuration of the hybrid blade, as determined by the inventors. As shown, due to the aforementioned reasons, the trailing edge is much thicker and does not... Figure 6A The convergence line PS shown C Conversely, the trailing edge surface S t It is confined between the ends of the pressure surface P and the suction surface S, as described above.

[0102] To solve the surface S t The problems that arise (such as) Figure 5 As shown, the trailing edge of the blade has an alternating profile that alternates in direction between the pressure side and the suction side of the blade. In practice, the trailing edge of the blade can be provided with a wavy or serpentine shape. In one example, the profile can be sinusoidal.

[0103] exist Figure 6B In the middle, the alternative profile extends from the hub all the way to the tip. However, the wavy undulations may be more pronounced or concentrated at the most radial part of each blade.

[0104] like Figure 6B As shown, providing a wavy profile to the trailing edge creates corresponding wavy undulations on the trailing edges of both the pressure and suction surfaces. This, in turn, creates different initiation points for vortex shedding on each surface measured in the airflow direction. By dispersing or distributing the initiation points of each vortex, the vortices interact and mix as they leave the blade trailing edge. This advantageously results in the vortices mixing in a destructive manner, thereby reducing the impact of vortex shedding on pressure loss. In effect, the vortices are mixed to reduce their effects and return them to a more normal linear airflow. In effect, energy is redistributed in the airflow. Specifically, vortex shedding occurs at a specific frequency, but the current arrangement is designed to suppress this effect, and energy is redistributed to other frequencies. The specific frequency will vary depending on the speed, and the sinusoidal shape described herein provides optimal performance under aircraft cruise conditions.

[0105] Depending on the specific compressor and engine, the alternating profile can be any suitable shape.

[0106] In a certain arrangement structure, such as Figure 6B As shown, the wavy profile can extend from the hub 13 to the tip 21 along the entire radial length of the blade. Thus, eddy shedding can be suppressed along the entire radial length of the blade.

[0107] However, it has been determined that vortex generation is related to airspeed, therefore greater vortex generation occurs towards the blade tip. More specifically, the difference between the pressure and suction sides is greatest at the tip, which also means that the trailing edge thickness is thickest at the tip—therefore, vortex shedding should be highest there compared to the hub.

[0108] In another arrangement, only a portion of the radial length of the blade may have a wavy trailing edge, such as... Figure 6C As shown. Here, only the most radially shaded portion of the trailing edge of the blade has a wavy, undulating structure. The root portion (towards hub 13) has a straight profile.

[0109] The inventors have determined the point at which the undulations should begin, i.e., the blade radius at which they should begin, which is the specific location r for optimal performance. b This should correspond to the radial distance at which the airflow becomes transonic, that is, the radius at which the air on the blades reaches Mach 1.

[0110] When the eddy current shedding profile should be introduced, i.e., the radius r at which the wavy undulations should begin. vs It is V c The radius r at which the speed is close to or equal to Mach 1 b This is in cruise mode during aircraft flight. V c Points that are close to or equal to Mach 1 can be determined through aerodynamic modeling.

[0111] By mixing vortices, accumulated turbulence at the trailing edge is reduced, thus lowering the overall pressure loss. Therefore, combined with the modified mixing pressure and intake surface profile of the compressor blades, highly advantageous blades can be produced for a given engine application.

[0112] Figure 7 An example profile of the trailing edge of the hybrid blade described in this paper is illustrated.

[0113] like Figure 8 As shown, further modifications can be applied to the trailing edge in the longitudinal direction. Figure 8 The view shown illustrates the side projection of the blades, with the leading edge 14 on the left and the trailing edge 15 on the right. As illustrated, the trailing edge profile can also be modified from the root (near the hub) to the radially distal tip of each blade, and simultaneously modified in the airflow or chordal direction of that blade. Therefore, complex profiles incorporating wavy undulations can be generated, such as sinusoidal profiles in the circumferential direction across the trailing edge and in the chordal direction of the trailing edge. The profile can be selected based on the specific engine and compressor characteristics. In the example shown, the radial distance required for vortex shedding is represented by the radius r. b express.

[0114] The modified trailing edge, combined in two planes, produces a complex trailing edge surface, leading to vortex shedding at many different locations in both the radial and circumferential directions. As can be recognized from the teachings of this paper, a precise shape can be optimized using modeling of a given engine and its desired operating characteristics.

[0115] Each blade may have the same modified profile in both the radial and circumferential directions, but it should also be recognized that adjacent blades or groups of blades may have different and non-uniform trailing edge profiles. In practice, it can be advantageous to generate almost randomly mixed vortices at the trailing edge of the blade. This can advantageously minimize pressure loss. In effect, the vortices are mixed or blended to generate a more uniform pressure behind the trailing edge.

[0116] Figure 9A and Figure 9B An alternative embodiment of the hybrid blade according to the invention described herein is shown. In this embodiment, the thicker trailing edge described above with reference to the first embodiment can be avoided.

[0117] Specifically, in this embodiment, the surface S extending from the trailing edge between the pressure surface P and the suction surface S... t Merging. More specifically, it provides a normal, abrupt convergence of the trailing edges of the pressure surface and the suction surface, which itself exhibits wavy undulations between line A, indicating the end of the suction surface optimized for efficiency, and line B, optimized for stability (see...). Figure 5 ). Figure 9A This is illustrated using dashed lines A and B.

[0118] Similarly, the radial distance at which the wavy undulations are introduced in the trailing edge design is r. b .

[0119] Figure 9B yes Figure 9A The image shows an end view of the trailing edge of the blade. The wavy surface is radii r from the hub. b It begins at a point and has a pressure surface optimized for stability and a suction side optimized for efficiency. The trailing edges of the pressure and suction surfaces converge along the profile shown.

[0120] Figure 10A and 10B The illustration shows an example mathematically accurate description of how to program the deformation between the ideal shape of the "optimal efficiency" blade and the "optimal partial speed" blade. The advantages of this invention can be achieved by combining these two mathematical profiles and initiating a wavy undulation at the aforementioned location.

[0121] Any invention described herein can be used for both high-pressure compressor blades and low-pressure compressor blades.

Claims

1. A compressor blade for a compressor, the compressor comprising a plurality of blades, each blade extending radially from a central hub and including a pressure surface on one side of the blade and a suction surface on an opposite side of the blade, wherein the pressure surface and the suction surface extend from a common leading edge of the blade to a trailing edge of the suction surface and a trailing edge of the pressure surface, and wherein the trailing edges of the pressure surface and the suction surface of the blade are connected by a third trailing edge surface, wherein, When measured along the radius of the blade, a portion of the third trailing edge surface of the blade alternates between a first circumferential direction and a second opposite circumferential direction, thereby forming a wavy profile, wherein the amplitude of the wavy profile is non-uniform when measured from a reference line passing radially along the blade.

2. The compressor blade according to claim 1, wherein, The third trailing edge surface has a boundary defined between the central hub of the blade and the tip of the blade, and between the trailing edge of the pressure surface and the trailing edge of the suction surface.

3. The compressor blade according to claim 1 or 2, wherein The surface profile of the pressure surface of the blade is shaped according to a predetermined profile that provides predetermined aerodynamic stability during use; and The surface profile of the suction surface of the blade is shaped according to a predetermined profile that provides a predetermined fuel efficiency in use.

4. The compressor blade according to claim 1 or 2, wherein, The length of the suction surface, measured from the common leading edge to the trailing edge of the suction surface, is greater than the length of the pressure surface, measured from the common leading edge to the trailing edge of the pressure surface.

5. The compressor blade according to claim 1 or 2, wherein, The third trailing edge surface has a uniform thickness measured in the circumferential direction between the trailing edge of the suction surface and the trailing edge of the pressure surface.

6. The compressor blade according to claim 1 or 2, wherein, The leading edge of the pressure surface smoothly intersects the leading edge of the suction surface to define a smoothly curved leading edge surface of the blade.

7. The compressor blade according to claim 6, wherein, The pressure surface extending from the smooth common leading edge has curvature to allow air to exit the trailing edge of the pressure surface in a first direction; and the suction surface extending from the smooth common leading edge has curvature to allow air to exit the trailing edge of the suction surface in a second direction.

8. The compressor blade according to claim 1 or 2, wherein, When measured along the radius of the blade, a portion of the third trailing edge surface has a sinusoidal profile.

9. The compressor blade according to claim 1 or 2, wherein, The third trailing edge profile includes: The first straight portion extends radially from the central hub; and The second alternating portion extends from the intersection of the first straight portion and the second alternating portion toward the tip of the blade.

10. The compressor blade according to claim 9, wherein, The intersection of the first straight portion and the second alternating portion is located at a predetermined radius measured from the center hub of the blade.

11. The compressor blade according to claim 10, wherein, The predetermined radius corresponds to a radius at which the airflow across the pressure surface of the blades reaches Mach 1 when operating the compressor in high-power mode.

12. The compressor blade according to claim 1 or 2, wherein, The frequency of the wavy undulations is uniform along the blade.

13. The compressor blade according to claim 1 or 2, wherein, The frequency of the wavy undulations is not uniform along the blade.

14. The compressor blade according to claim 1 or 2, wherein, The third trailing edge surface is flat.

15. A gas turbine engine including a compressor, said compressor comprising compressor blades according to any of the preceding claims.

16. An aircraft comprising one or more engines, said engines being coupled with compressor blades according to any one of claims 1-14 or a compressor of a gas turbine engine according to claim 15.

17. A method of manufacturing blades for a compressor, the compressor having a plurality of blades, wherein each blade is arranged to extend radially from a central hub and includes a pressure surface on one side of the blade and a suction surface on an opposite side of the blade, wherein the pressure surface and the suction surface extend from a common leading edge of the blade to a trailing edge of the suction surface and a trailing edge of the pressure surface, and wherein the trailing edge of the pressure surface of the blade and the trailing edge of the suction surface of the blade are connected by a third trailing edge surface, wherein... When the radius of the blade is measured, a portion of the third trailing edge surface of the blade alternates between a first circumferential direction and a second opposite circumferential direction, thereby forming a wavy profile, wherein the amplitude of the wavy profile is non-uniform when measured from a reference line passing radially along the blade.

Images