An inlet guide vane for an aero-engine and an aero-engine

Abstract

This invention provides an inlet guide vane for an aero-engine's duct and an aero-engine in the field of aero-engine technology. The inlet guide vane has a blade body. The blade body has a root section and a tip section at both ends along the blade height direction. Simultaneously, the blade body also has an intermediate section located between the root section and the tip section along the blade height direction. The portion of the blade body between the tip section and the intermediate section forms a thickened section, and the ratio of the maximum thickness at the tip section to the maximum thickness at the intermediate section is 2-3. By increasing the maximum thickness from the intermediate section to the tip section, the thickness is increased over a wide range of blade height conditions, thereby improving the adaptability of the inlet guide vane to incoming flow conditions, and thus contributing to improved aerodynamic performance and stability.

Description

Technical Field

[0001] This invention relates to the field of aero-engine technology, and more specifically, to an inlet guide vane for an aero-engine and an aero-engine. Background Technology

[0002] The fan is one of the core components of an aircraft engine, its function being to compress air to increase the total temperature and pressure of the gas. After the air flows through the fan blades, it is split by the splitter ring; the passage through which the air passes in the core part is called the inner duct, and the passage through which the air passes in the outer part is called the outer bypass duct. The fan and low-pressure compressor sections of the inner duct are often referred to as the fan booster stage.

[0003] The splitter ring is a distribution structure that allows air to enter the inner duct and outer bypass duct separately. Under different engine operating conditions, the flow state of the airflow near the splitter ring changes accordingly. The airflow state near the splitter ring directly affects the airflow quality at the turbocharger stage inlet, that is, it affects the incoming flow conditions of the inner duct inlet guide vanes (also known as turbocharger stage inlet guide vanes), and thus affects the aerodynamic performance and stability of the turbocharger stage. Therefore, how to improve the aerodynamic performance and stability of the inner duct inlet guide vanes has become a pressing technical problem to be solved in this field. Summary of the Invention

[0004] The following provides a brief overview of one or more aspects to offer a basic understanding of them. This overview is not an exhaustive summary of all conceived aspects, nor is it intended to identify key or decisive elements of all aspects, nor to define the scope of any or all aspects. Its sole purpose is to present some concepts of one or more aspects in a simplified form to prepare for the more detailed descriptions that follow.

[0005] The purpose of this invention is to provide an inlet guide vane for an aero-engine duct that can improve the aerodynamic performance and stability of the inlet guide vane.

[0006] Another objective of this invention is to provide an aero-engine that can improve the aerodynamic performance and stability of the inlet guide vanes of the inner duct.

[0007] Embodiments of the present invention can be implemented in the following ways:

[0008] An inlet guide vane for an aero-engine duct, the inlet guide vane comprising a blade body, the blade body having a root section and a tip section at both ends along the blade height direction;

[0009] The blade body also has an intermediate section located between the blade root section and the blade tip section along the blade height direction. The portion of the blade body located between the blade tip section and the intermediate section forms a thickened portion, and the ratio of the maximum thickness at the blade tip section to the maximum thickness at the intermediate section is 2 to 3.

[0010] Optionally, the relative blade height at which the intermediate section is located is 50% to 80%.

[0011] Optionally, the intermediate section is located at 70% of the relative blade height.

[0012] Optionally, the blade body has a cross-sectional airfoil with a leading edge, and the thickness distribution from the leading edge to the position of maximum thickness satisfies the following formula:

[0013]

[0014] TR L =T L / T MAX

[0015] TR X =T X / T MAX

[0016] Where X is the relative position from the leading edge to the maximum thickness position, and its value ranges from 0 to 1. The closer X is to the maximum thickness position, the closer it is to 1, and the closer it is to the leading edge, the closer it is to 0; T X The thickness at position X; T MAX T is the maximum thickness of the airfoil section; L K1 is the thickness of the leading edge; K1 is an adjustment parameter.

[0017] Optionally, along the blade height direction of the blade body, the adjustment parameter K1 of the cross section at different blade heights in the thickened portion varies linearly.

[0018] Optionally, the adjustment parameter K1 can range from 1 to 18.

[0019] Optionally, the blade body has a trailing edge in its cross-sectional shape, and the thickness distribution between the trailing edge and the position of maximum thickness satisfies the following formula:

[0020]

[0021] TR T =T T / T MAX

[0022] TR Y =T Y / T MAX

[0023] Wherein, Y is the relative position from the leading edge to the maximum thickness position, and its value ranges from 0 to 1. The closer Y is to the maximum thickness position, the closer it is to 1, and the closer it is to the trailing edge, the closer it is to 0; T Y The thickness at position Y; T MAX T is the maximum thickness of the airfoil section; T K1 represents the thickness of the trailing edge; K2 is an adjustment parameter.

[0024] Optionally, along the blade height direction, the adjustment parameter K2 of the cross section at different blade heights in the thickened portion varies linearly.

[0025] Optionally, the value range of the adjustment parameter K2 is 1 to 18.

[0026] An aero-engine, the aero-engine comprising a plurality of inlet guide vanes as described in claims 1 to 9.

[0027] The beneficial effects of the inlet guide vanes of the aero-engine duct and the aero-engine provided by the embodiments of the present invention include:

[0028] An embodiment of the present invention provides an inlet guide vane for an aero-engine duct, comprising a blade body. The blade body has a root section and a tip section at both ends along the blade height direction. Simultaneously, the blade body also has an intermediate section located between the root section and the tip section along the blade height direction. The portion of the blade body between the tip section and the intermediate section forms a thickened portion, and the ratio of the maximum thickness at the tip section to the maximum thickness at the intermediate section is 2 to 3. By increasing the maximum thickness from the intermediate section to the tip section, the thickness is increased over a wide range of blade height conditions, thereby improving the adaptability of the inlet guide vane to incoming flow conditions, and thus contributing to improved aerodynamic performance and stability.

[0029] Embodiments of the present invention also provide an aero-engine comprising a plurality of aero-engine duct inlet guide vanes. Since the aero-engine includes the aforementioned aero-engine duct inlet guide vanes, it also has the beneficial effect of improving the adaptability of the duct inlet guide vanes to incoming flow conditions, thereby contributing to improved aerodynamic performance and stability. Attached Figure Description

[0030] The above-described features and advantages of the present invention will be better understood after reading the following detailed description of embodiments of the present disclosure in conjunction with the accompanying drawings. In the drawings, components are not necessarily drawn to scale, and components having similar related characteristics or features may have the same or similar reference numerals.

[0031] Figure 1A partial structural schematic diagram of an aircraft engine provided according to one aspect of the present invention is shown;

[0032] Figure 2 An overall diagram of the flow field near the split ring according to one aspect of the present invention is shown;

[0033] Figure 3 A partial diagram of the flow field near the splitter ring provided according to one aspect of the invention is shown;

[0034] Figure 4 A structural diagram of the blade body in the inlet guide vane of an aero-engine according to one aspect of the present invention is shown;

[0035] Figure 5 A comparison diagram is shown between the maximum thickness variation law of the inlet guide vane of an aero-engine according to one aspect of the present invention and that of a conventional blade;

[0036] Figure 6 A comparison diagram of the cross-sectional airfoil of the inlet guide vane according to one aspect of the present invention and a conventional airfoil is shown;

[0037] Figure 7 A comparison diagram is shown between the front section blade profile thickness distribution curve of the inlet guide vane provided according to one aspect of the present invention and the front section blade profile thickness distribution curve of a conventional blade.

[0038] Figure 8 A comparison diagram is shown of the thickness distribution curve of the rear section of the guide vane of the inner channel provided according to one aspect of the present invention and the thickness distribution curve of the rear section of a conventional blade.

[0039] Figure 9 A comparison diagram of the low-loss angle of attack range of the inlet guide vane provided according to one aspect of the present invention and conventional blades is shown.

[0040] Figure label:

[0041] 10-Aircraft engine; 11-Fan casing; 12-Fan blade; 13-Split ring; 14-Outer bypass exhaust guide vane; 15-Inlet cone; 16-Inner duct inlet guide vane; 17-Axis shaft; 18-Inner duct; 19-Outer bypass duct;

[0042] 110-Blade body; 111-Blade root section; 112-Intermediate section; 113-Blade tip section; 114-Thickened part; 115-Leading edge; 116-Maximum thickness position; 117-Leading section blade shape; 118-Tail edge; 119-Tail section blade shape. Detailed Implementation

[0043] The present invention will now be described in detail with reference to the accompanying drawings and specific embodiments. It should be noted that the aspects described below with reference to the accompanying drawings and specific embodiments are merely exemplary and should not be construed as limiting the scope of protection of the present invention in any way.

[0044] In the description of this invention, it should be noted that if terms such as "upper," "lower," "inner," "outer," or "vertical" appear, the orientation or positional relationship indicated is based on the orientation or positional relationship shown in the accompanying drawings, or the orientation or positional relationship in which the product of this invention is usually placed when in use, and does not indicate or imply that the device or element referred to must have a specific orientation or be constructed and operated in a specific orientation, and therefore should not be construed as a limitation of this invention.

[0045] At the same time, it should be noted that the terms "first" and "second" are used only for distinguishing descriptions and should not be interpreted as indicating or implying relative importance.

[0046] In the description of this invention, it should also be noted that, unless otherwise explicitly specified or limited, the terms "installation," "connection," and "linking" should be interpreted broadly. For example, they can refer to a fixed connection, an integral connection, or a detachable connection; they can refer to a mechanical connection or an electrical connection; they can refer to a direct connection or an indirect connection through an intermediate medium, or a connection within two components, etc. Those skilled in the art can understand the specific meaning of the above terms in this invention according to the specific circumstances.

[0047] Figure 1 This is a partial structural schematic diagram of the aircraft engine 10 provided in this embodiment. Figure 2 This is an overall diagram of the flow field near split ring 13. Figure 3 This is a partial diagram of the flow field near split ring 13. Please refer to the reference diagram. Figures 1-3The aero-engine 10 includes a fan casing 11, fan blades 12, a flow divider ring 13, an outer bypass outlet guide vane 14, an inlet cone 15, and an inner bypass inlet guide vane. The inlet cone 15 is located at the center of the fan casing 11, thus forming an intake channel between the inlet cone 15 and the fan casing 11. The fan blades 12 are arranged around the axis 17 of the inlet cone 15 (i.e., the axis 17 of the aero-engine 10) in the intake channel, thereby guiding outside air into the intake channel. The flow divider ring 13 is located behind the fan blades 12, and an inner bypass 18 is formed on the radially inner side of the flow divider ring 13, while an outer bypass 19 is formed on the radially outer side of the flow divider ring 13. The airflow entering the aero-engine 10 through the fan blades 12 is divided into two paths by the flow divider ring 13, one path flowing along the inner bypass 18 and the other path flowing along the outer bypass 19. The outer bypass outlet guide vane 14 is disposed in the outer bypass duct 19 to guide the airflow entering the outer bypass duct 19; the inner bypass inlet guide vane 16 is disposed in the inner bypass duct to guide the airflow entering the inner bypass duct 18, and multiple inner bypass inlet guide vanes 16 are distributed around the axis 17 of the aero-engine 10.

[0048] The inventors discovered that the airflow state near the splitter ring 13 changes depending on the operating state of the aero-engine 10, such as... Figure 2 and Figure 3 The airflow state near the split ring 13 is shown, where, Figure 2 and Figure 3 Point A represents the incoming flow field at the % blade height of the inlet guide vane 16 of the inner duct. This incoming flow condition affects the aerodynamic performance and stability of the booster stage, and the flow field pattern changes with the operating state of the aero-engine 10. Therefore, in order to improve the inner duct inlet guide vane's ability to cope with changes in incoming flow conditions caused by changes in the operating state of the aero-engine 10 and to improve its aerodynamic performance and stability, an embodiment of the present invention provides an aero-engine inner duct inlet guide vane, hereinafter referred to as the inner duct inlet guide vane.

[0049] Figure 4 This is a structural diagram of the blade body 110 in the inlet guide vane of the aero-engine provided in this embodiment. Figure 5 This diagram compares the maximum thickness variation of the inlet guide vane of the aero-engine's internal duct provided in this embodiment with that of a conventional blade. The solid line represents the maximum thickness variation curve of the inlet guide vane provided in this embodiment, while the dashed line represents the maximum thickness variation curve of a conventional blade. The horizontal axis represents a gradual increase in maximum thickness to the right, and the vertical axis represents a gradual increase in blade height. Please refer to the diagram for further information. Figures 1-5 This embodiment provides an aero-engine 10 and its inner duct inlet guide vanes. The aero-engine 10 provided in this embodiment has the following structure: Figure 1 The structure shown is adopted and the following is used: Figure 4The diagram shows an inlet guide vane with an inner duct. The inlet guide vane has a blade body 110. The blade body 110 has a root section 111 and a tip section 113 at both ends along the blade height direction. Simultaneously, the blade body 110 also has an intermediate section 112 located along the blade height direction between the root section 111 and the tip section 113. The portion of the blade body 110 between the tip section 113 and the intermediate section 112 forms a thickened portion 114, meaning the thickened portion 114 is thicker than a conventional blade. Furthermore, the ratio of the maximum thickness at the tip section 113 to the maximum thickness at the intermediate section 112 is 2 to 3. By increasing the maximum thickness from the intermediate section 112 to the tip section 113, the thickness is increased over a wide range of blade height conditions, thereby improving the adaptability of the inlet guide vane to incoming flow conditions and thus contributing to improved aerodynamic performance and stability.

[0050] It should be noted that, in the description of this embodiment, the cross section of the blade body 110 at a certain blade height refers to the surface obtained by the plane perpendicular to the blade height direction B at that blade height position. Specifically, the blade root cross section 111 is the cross section of the blade body 110 at 0% of the blade height, and the blade tip cross section 113 is the cross section of the blade body 110 at % of the blade height.

[0051] like Figure 5 As shown, in this embodiment, the relative blade height of the intermediate section 112 is Y%, and the maximum thickness of the inlet guide vane of the inner duct provided in this embodiment at Y% is X. 12 The maximum thickness at the blade tip section 113 (i.e., at a relative blade height of %) is X2. Thus, the thickness of the inner channel inlet guide vane provided in this embodiment is X2 / X. 12 The value is 2 to 3. In this embodiment, X2 / X 12 =2.4. The maximum thickness of a conventional blade at Y% is X. 12 The maximum thickness at the leaf tip section 113 (i.e., at a relative leaf height of %) is X1. Generally, in conventional blades, X1 / X 12 The value range is 1 to 1.5. Figure 5 The curve shown shows X1 / X 12 =1.1, thus passing Figure 5 It is evident that the maximum thickness of the inlet guide vane provided in this application is significantly increased in the thickened portion 114 compared to traditional blades.

[0052] In this embodiment, the relative blade height of the intermediate section 112 is 50% to 80%, that is, the value of Y is in the range of 50 to 80. Optionally, in this embodiment, the relative blade height of the intermediate section 112 is 70%, that is, Y = 70. It is understood that in some other embodiments, Y% can also be set to 50% or 80%, or it can be set to a value between 50% and 80% as needed.

[0053] Currently, the thickness distribution pattern of traditional blades comes from traditional blade types, such as the NACA series and C4 series. The inner channel inlet guide vane provided in the embodiment of the present invention increases the maximum thickness of the thickened part 114 to improve the adaptability of the inner channel inlet guide vane. For the non-thickened part 114 of the blade body 110 (i.e., the part from the blade root section 111 to the middle section 112), the thickness distribution of traditional blades can be adopted, and there is no restriction on its specific thickness distribution pattern here.

[0054] After determining the maximum thickness distribution pattern of the blade body 110 along the blade height direction, for the blade profile of the blade body 110, it is also necessary to set the thickness distribution pattern of each cross-section. Typical thickness distribution curves include circular arc curves, cubic curves, and quartic curve distributions. However, these distribution patterns require adjustment of many parameters. If the given variables are not appropriate, problems such as uneven thickness distribution will occur, leading to increased blade profile loss and complex blade profile design. To improve this problem, this embodiment determines the thickness distribution pattern of the blade profile at the cross-section in the following way:

[0055] Figure 6 This is a comparison diagram of the cross-sectional airfoil of the inlet guide vane provided in this embodiment and a conventional airfoil. Figure 6 The solid line represents the cross-sectional airfoil of the inlet guide vane provided in this embodiment, while the dashed line represents the traditional airfoil. Figure 7 This diagram shows a comparison between the thickness distribution curve of the leading section 117 of the guide vane at the inlet of the internal duct provided in this embodiment and the thickness distribution curve of the leading section 117 of a conventional blade. The horizontal axis represents relative position, with values ​​closer to the maximum thickness (approximately 1) and closer to the leading edge (approximately 115) (approximately 0). The solid line represents the thickness distribution curve of the leading section 117 of the guide vane at the inlet of the internal duct provided in this embodiment, and the dashed line represents the thickness distribution curve of the leading section 117 of a conventional blade. Please refer to... Figure 6 and Figure 7 In this embodiment, the blade body 110 has a cross-sectional airfoil with a leading edge 115 and a trailing edge 118. The airfoil portion between the leading edge 115 and the maximum thickness position 116 is called the leading section airfoil 117, and the airfoil portion between the trailing edge 118 and the maximum thickness position 116 is called the trailing section airfoil 119. The thickness distribution of the leading section airfoil 117 satisfies the following formula:

[0056]

[0057] TR L =T L / T MAX

[0058] TR X =T X / T MAX

[0059] Where X is the relative position from the leading edge 115 to the maximum thickness position 116, and its value ranges from 0 to 1. That is, X can be regarded as Figure 7 The x-coordinate in T; X The thickness at position X; T MAX T is the maximum thickness of the airfoil section; L K1 is the thickness of the leading edge 115; K1 is an adjustment parameter.

[0060] The value of parameter K1 can be preset according to requirements. Generally, the value range of parameter K1 is 1 to 18. Optionally, the value of parameter K1 can be set to 1, 6, 9, 12 or 18, etc., according to requirements.

[0061] Since the thickness distribution of the front section airfoil 117 of the blade body 110 satisfies the above formula, when designing the blade cross section airfoil, the thickness distribution of the front section airfoil 117 can also be designed by selecting an appropriate adjustment parameter K1. The design process is simple and the calculation is convenient.

[0062] Furthermore, along the blade height direction of the blade body 110, the adjustment parameter K1 of the cross-section at different blade heights in the thickened portion 114 changes linearly. Specifically, the linear change of the adjustment parameter K1 at different blade heights means that the coordinate relationship formed by the two variables, adjustment parameter K1 and blade height, is linear. In other words, the linear change of the adjustment parameter K1 at different blade heights means that the adjustment parameter K1 at different blade heights is the same, increases, or decreases.

[0063] Figure 8 A comparison diagram is shown between the thickness distribution curve of the rear section blade profile 119 of the inlet guide vane provided in this embodiment and the thickness distribution curve of the rear section blade profile 119 of a conventional blade. The horizontal axis represents relative position, with values ​​closer to the maximum thickness being closer to 1, closer to the leading edge 115 being closer to 0. The solid line represents the thickness distribution curve of the rear section blade profile 119 of the inlet guide vane provided in this embodiment, and the dashed line represents the thickness distribution curve of the rear section blade profile 119 of a conventional blade. The thickness distribution of the rear section blade profile 119 satisfies the following formula:

[0064]

[0065] TR T =T T / T MAX

[0066] TR Y =T Y / T MAX

[0067] Where Y is the relative position from the leading edge 115 to the maximum thickness position 116, and its value ranges from 0 to 1. That is, Y can be regarded as Figure 8 The x-coordinate in T; Y The thickness at position Y; T MAX T is the maximum thickness of the airfoil section; T K1 represents the thickness of the leading edge 115; K2 is an adjustment parameter.

[0068] The value of parameter K2 can be preset according to requirements. Generally, the value range of parameter K2 is 1 to 18. Optionally, the value of parameter K2 can be set to 1, 6, 9, 12 or 18, etc., according to requirements.

[0069] Since the thickness distribution of the front section airfoil 117 of the blade body 110 satisfies the above formula, when designing the blade cross section airfoil, the thickness distribution of the rear section airfoil 119 can also be designed by selecting an appropriate adjustment parameter K2. The design process is simple and the calculation is convenient.

[0070] Furthermore, along the blade height direction of the blade body 110, the adjustment parameter K2 at different blade height sections in the thickened portion 114 changes linearly. Specifically, the linear change of the adjustment parameter K2 at different blade height sections means that the coordinate relationship formed by the adjustment parameter K2 and the blade height is linear. In other words, the linear change of the adjustment parameter K2 at different blade height sections means that the adjustment parameter K2 at different blade height sections is the same, increases, or decreases.

[0071] Specifically, Figure 6 For the medium-leaf type, the adjustment parameter K1 is set to 2, and the adjustment parameter K2 is set to 6.

[0072] The embodiments of the present invention provide an aero-engine inner duct inlet guide vane and an aero-engine 10. By increasing the maximum thickness of the inner duct inlet guide vane at the thickened portion 114, and simultaneously achieving a thickening design at the maximum thickness position 116 of the traction ring finger, and a thickness reduction design from the maximum thickness position 116 to the trailing edge 118, it is possible to expand the blade's operating condition / loss range while maintaining low losses in the inner duct inlet guide vane, thereby improving its adaptability to changes in incoming flow conditions. Figure 9 The embodiment demonstrates that the low-loss angle of attack range of the inlet guide vane provided in this embodiment is B2-A2, which is 11.0° in this example; the low-loss angle of attack range of the conventional airfoil is B1-A1, which is 8.8° in this example. The increased low-loss angle of attack range enhances the adaptability of the inlet guide vane to changes in incoming flow conditions.

[0073] The above description is merely a specific embodiment of the present invention, but the scope of protection of the present invention is not limited thereto. Any variations or substitutions that can be easily conceived by those skilled in the art within the scope of the technology disclosed in the present invention should be included within the scope of protection of the present invention.

Claims

1. An inlet guide vane for an aero-engine duct, the inlet guide vane comprising a blade body, the blade body having a root section and a tip section at both ends along the blade height direction; characterized in that, The blade body also has an intermediate section located between the blade root section and the blade tip section along the blade height direction. The portion of the blade body located between the blade tip section and the intermediate section forms a thickened portion, and the ratio of the maximum thickness at the blade tip section to the maximum thickness at the intermediate section is 2 to 3.

2. The inlet guide vane of the aero-engine inner duct according to claim 1, characterized in that, The intermediate section is located at 50% to 80% of the relative blade height.

3. The inlet guide vane of the aero-engine inner duct according to claim 2, characterized in that, The intermediate section is located at 70% of the relative blade height.

4. The inlet guide vane of the aero-engine inner duct according to claim 1, characterized in that, The blade body has a cross-sectional airfoil with a leading edge, and the thickness distribution from the leading edge to the position of maximum thickness satisfies the following formula: TR L =T L / T MAX TR X =T X / T MAX Where X is the relative position from the leading edge to the maximum thickness position, and its value ranges from 0 to 1. The closer X is to the maximum thickness position, the closer it is to 1, and the closer it is to the leading edge, the closer it is to 0; T X The thickness at position X; T MAX T is the maximum thickness of the airfoil section; L K1 is the thickness of the leading edge; K1 is an adjustment parameter.

5. The inlet guide vane of the aero-engine inner duct according to claim 4, characterized in that, Along the blade height direction of the blade body, the adjustment parameter K1 of the cross section at different blade heights in the thickened part changes linearly.

6. The inlet guide vane of the aero-engine inner duct according to claim 4, characterized in that, The value range of the adjustment parameter K1 is 1 to 18.

7. The inlet guide vane of the aero-engine inner duct according to claim 1, characterized in that, The blade body has a trailing edge in its cross-sectional shape, and the thickness distribution between the trailing edge and the position of maximum thickness satisfies the following formula: TR T =T T / T MAX TR Y =T Y / T MAX Wherein, Y is the relative position from the leading edge to the maximum thickness position, and its value ranges from 0 to 1. The closer Y is to the maximum thickness position, the closer it is to 1, and the closer it is to the trailing edge, the closer it is to 0; T Y The thickness at position Y; T MAX T is the maximum thickness of the airfoil section; T K1 represents the thickness of the trailing edge; K2 is an adjustment parameter.

8. The inlet guide vane of the aero-engine inner duct according to claim 7, characterized in that, Along the blade height direction, the adjustment parameter K2 of the cross section at different blade heights in the thickened part changes linearly.

9. The inlet guide vane of the aero-engine inner duct according to claim 7, characterized in that, The value range of the adjustment parameter K2 is 1 to 18.

10. An aircraft engine, characterized in that, The aero-engine includes multiple inlet guide vanes as described in claims 1 to 9.

Images