Subsonic high-lift low-drag laminar flow airfoil for mid-span of helicopter blades

Abstract

The application provides a subsonic high-lift low-drag laminar flow airfoil for the middle part of a helicopter blade, wherein the airfoil has a front edge radius of 0.00588, a maximum thickness of 13.10% at 30.5% of the chord length, a maximum camber of 0.0126 at 27.7% of the chord length, and a trailing edge included angle of 6.78°; the front edge radius, the maximum thickness and the maximum camber are described by using dimensionless quantities and taking the chord length c of the airfoil as the reference. Compared with the classical NACA 62(3)-213 laminar flow airfoil, the airfoil can push back the transition position, maintain a larger laminar flow area and has better lift-drag characteristics. Meanwhile, the increase of the camber of the airfoil can balance the nose-up moment, so that the airfoil has a lower pitch moment peak value than the NACA 62(3)-213 airfoil.

Description

TECHNICAL FIELD

[0001] The present application relates to the field of helicopter blade airfoil design, in particular to a subsonic high-lift low-drag laminar airfoil for the middle part of a helicopter blade. BACKGROUND

[0002] The helicopter rotor is in a variable aerodynamic environment, the rotor blade rotates around the rotation axis while also having a forward speed with the forward flight of the helicopter, the incoming flow velocity and the angle of attack of the profile airfoil change dramatically during rotation. With the increase of the flight speed of the helicopter, the aerodynamic environment of the blade becomes more and more complex, various aerodynamic phenomena such as forward compressibility, rear dynamic stall, reverse flow and the like exist in rotation, which limits the further improvement of the flight speed of the helicopter, making it difficult to exceed 300km / h.

[0003] The aerodynamic performance of the airfoil has a decisive effect on the pros and cons of the rotor performance, and significantly affects the forward flight speed, rapid maneuvering performance, take-off and landing performance, handling quality and flight efficiency of the helicopter. At the same time, the flow characteristics and working state of different positions in the span direction of the rotor also put forward different requirements for the airfoil design. For example, the strong reverse flow area within 35% of the span direction needs to be arranged with blunt trailing edge airfoils to reduce separation resistance, the transition area from blunt trailing edge airfoil to sharp trailing edge airfoil at the position of 35%-40% of the span direction needs to be arranged with normal sharp trailing edge airfoil, the laminar flow area at the position of 55% of the span direction mainly needs high lift-drag ratio laminar airfoil and generally does not need to consider the moment divergence characteristics, the area from 60% to 85% of the span direction needs to consider high lift-drag ratio, low moment and the like, and the tip part needs to consider the divergence Mach number due to the proximity to the supersonic speed. Therefore, the design of the rotor airfoil is a comprehensive optimization design problem with multiple design points, multiple design indexes and complex constraint conditions. SUMMARY

[0004] With the increasing demand for the performance of the helicopter, the aerodynamic performance requirements for the rotor airfoil are also getting higher and higher, and the traditional airfoil has been unable to meet the new aerodynamic performance requirements. For example, according to the performance indexes of a new type of helicopter, a new airfoil is required in the middle section of the blade, on the basis of effectively balancing the lifting moment of the airfoil, the laminar flow area is expanded as much as possible, and the lift-drag ratio is improved as much as possible at an attack angle of not more than 10°.

[0005] To this end, the present application proposes a subsonic low-resistance laminar airfoil for the mid-section of the helicopter blade at about 55% of the spanwise position, which takes the classic NACA 62(3)-213 laminar airfoil with a thickness of 13% as a comparative airfoil, has a relatively small leading edge radius, can push back the transition position and expand the laminar flow area, has a larger lift-drag ratio under the same incoming flow at an attack angle of not more than 10°, and has increased camber, with the maximum thickness position moved from 35% of the chord line to 30.5% of the chord line and the maximum camber position moved from 51.9% of the chord line to 27.7% of the chord line, so as to balance the lifting moment and improve the trim characteristics of the rotor, in addition, the slope of the upper surface of the airfoil is more gentle, which can maintain a larger laminar flow range, thereby facilitating the reduction of friction resistance.

[0006] The technical scheme of the present application is:

[0007] A subsonic high-lift low-resistance laminar airfoil for the mid-section of a helicopter blade has the following airfoil characteristics:

[0008] The airfoil has a leading edge radius of 0.00588, a maximum thickness of 13.10% located at 30.5% of the chord length, a maximum camber of 0.0126 located at 27.7% of the chord length, and a trailing edge angle of 6.78°. It should be noted that in the field of airfoil design, all parameters are described using dimensionless quantities, so the above-mentioned leading edge radius, maximum thickness, maximum camber, and subsequent airfoil coordinate descriptions are all described using dimensionless quantities, and the dimensionless process takes the chord length c as the reference.

[0009] Further, the geometric coordinate expressions of the upper surface and the lower surface of the airfoil are:

[0010]

[0011] Where x represents the transverse coordinate of the upper surface or the lower surface of the airfoil, y represents the corresponding longitudinal coordinate of the upper surface or the lower surface of the airfoil, n represents the order of the CST parameterization method, y tail represents the y coordinate of the root step of the airfoil;

[0012] The fitting coefficients of the upper surface of the airfoil are:

[0013]

[0014] The fitting coefficients of the lower surface of the airfoil are:

[0015]

[0016]

[0017] Further, the fitting coefficients of the upper and lower surfaces of the airfoil are preferably:

[0018] The fitting coefficient for the upper surface of the airfoil is:

[0019]

[0020] The fitting coefficient for the lower surface of the airfoil is:

[0021]

[0022] Furthermore, the coordinates of the upper and lower surfaces of the subsonic, high-lift, low-drag laminar airfoil used in the middle of the helicopter rotor blade are as follows:

[0023] Top surface coordinates:

[0024]

[0025]

[0026] The coordinates of the lower surface are as follows:

[0027]

[0028]

[0029] Beneficial effects

[0030] Compared with the prior art, the present invention has the following technical effects:

[0031] This invention provides a subsonic, high-lift, low-drag laminar airfoil for the mid-section of helicopter rotor blades. For helicopter rotors, the mid-section airfoil operates within a relatively large laminar flow region. This invention improves drag characteristics at angles of attack below 10° while maintaining minimal change in the pitching moment coefficient, and simultaneously reduces the peak pitching moment coefficient. Compared to the classic laminar airfoil NACA 62(3)-213, this invention can push back the transition position, maintain a larger laminar flow region, and exhibit better lift-drag characteristics. Simultaneously, the increased airfoil camber balances the pitching moment, resulting in a lower peak pitching moment than NACA 62(3)-213. This invention lays the foundation for the design of mid-section airfoils for helicopter rotor blades.

[0032] Additional aspects and advantages of the invention will be set forth in part in the description which follows, and in part will be obvious from the description, or may be learned by practice of the invention. Attached Figure Description

[0033] The above and / or additional aspects and advantages of the present invention will become apparent and readily understood from the description of the embodiments taken in conjunction with the following drawings, in which:

[0034] Figure 1 This is a geometric shape diagram of the airfoil of the present invention.

[0035] Figure 2 Comparison of airfoil geometry of the present invention with NACA 62(3)-213 airfoil geometry

[0036] Figure 3 Camber line of the present invention airfoil and NACA 62(3)-213 camber line

[0037] Figure 4 Pressure distribution pattern of the present invention airfoil and NACA 62(3)-213 airfoil at design point (Ma=0.5)

[0038] Figure 5 Comparison of lift-drag ratio characteristic curve of the present invention airfoil and NACA 62(3)-213 airfoil at low subsonic state (Ma=0.4)

[0039] Figure 6 Comparison of lift-drag ratio characteristic curve of the present invention airfoil and NACA 62(3)-213 airfoil at low subsonic state (Ma=0.5)

[0040] Figure 7 Pressure distribution of the present invention airfoil at design point

[0041] Figure 8 Pressure distribution of NACA 62(3)-213 airfoil at design point

[0042] Figure 9 Comparison of low speed moment characteristic curve of the present invention airfoil and NACA 62(3)-213 airfoil (Ma=0.4)

[0043] Figure 10 Comparison of low speed moment characteristic curve of the present invention airfoil and NACA 62(3)-213 airfoil (Ma=0.5) DETAILED DESCRIPTION

[0044] The embodiments of the present invention are described in detail below, which are exemplary and intended to explain the present invention, and cannot be understood as a limitation of the present invention.

[0045] When the helicopter is in the forward flight state, the middle part of the blade requires a large laminar flow area to maintain the laminar flow in the middle part of the blade to improve the lift-drag characteristics of the rotor and improve the flight efficiency of the helicopter.

[0046] Therefore, according to the aerodynamic performance requirements of the helicopter blade, the present embodiment proposes a subsonic high-lift low-drag laminar airfoil for the middle part of the helicopter blade. The design idea of the airfoil is based on the robust optimization design method, and the aerodynamic performance of the rotor airfoil at low subsonic Ma=0.4 and Ma=0.5 is optimized to meet the design requirements of high lift-drag ratio and low moment characteristics.

[0047] The airfoil designed in this embodiment is shown in the drawing with the horizontal and vertical coordinates in proportion as 1. Figure 1 The leading edge radius of the airfoil is 0.00588, the airfoil area is 0.08075, the maximum thickness of the airfoil is 13.10% at 30.5% chord length, the maximum camber is 0.0126 at 27.7% chord length, and the trailing edge angle is 6.78°. The above-mentioned leading edge radius, maximum thickness, maximum camber and subsequent airfoil coordinate description are described by dimensionless quantities, and the dimensionless process is based on the chord length c of the airfoil.

[0048] The unified expressions of the upper surface and the lower surface of the airfoil fitted by using 7-order CST parameterization are as follows:

[0049]

[0050] Wherein x represents the horizontal coordinate of the upper surface or the lower surface of the airfoil, y represents the corresponding vertical coordinate of the upper surface or the lower surface of the airfoil, n represents the order of the CST parameterization method, y tail represents the y coordinate of the root step of the airfoil;

[0051] The fitting coefficients of the upper surface of the airfoil are as follows:

[0052]

[0053] The fitting coefficients of the lower surface of the airfoil are as follows:

[0054]

[0055] Further, the upper and lower surface coordinate positions of the subsonic high-lift low-drag laminar airfoil for the middle part of the helicopter blade are as follows:

[0056] The upper surface coordinates are as follows:

[0057]

[0058]

[0059]

[0060] The lower surface coordinates are as follows:

[0061]

[0062]

[0063] The NACA 62(3)-213 classic laminar airfoil is taken as a reference airfoil for comparison with this embodiment, Figure 2For the contrast of two airfoils, in order to better observe the difference, the coordinate axis aspect ratio is set to 0.15. From the geometric point of view: it can be observed that the present embodiment (the figure legend is named as OPT) has a relatively small leading edge radius compared with the reference airfoil (the legend is named as NACA 62(3)-213); the present embodiment has a smaller slope on the upper part of the trailing edge compared with the reference airfoil (A of the picture); the reference Figure 3 It can be observed that the present application has a larger maximum curvature compared with the reference airfoil, and the position of the maximum curvature is advanced, and the trailing end of the airfoil is more flat.

[0064] The change of the geometric shape will inevitably lead to the change of the aerodynamic performance. First, the aerodynamic performance of the present application under the condition of low subsonic flow (Ma=0.5, Re=2.16e6) is analyzed. Figure 4 The pressure distribution pattern diagram of the present application airfoil and NACA 62(3)-213 airfoil under the condition of zero-lift design near the design point is given. First, the optimization of the leading edge radius of the airfoil reduces the original airfoil, so that the present application airfoil has a late transition position and a larger laminar flow region, thereby having better resistance characteristics under a larger attack angle. Second, the slope change of the upper surface of the optimized airfoil is more gentle, which is beneficial to delay the flow transition and form a larger laminar flow region. In order to more intuitively reflect the pressure distribution characteristics of the airfoil, Figure 7 and Figure 8 The pressure distribution cloud diagram and streamline diagram of the present application airfoil and the reference NACA 62(3)-213 airfoil under the condition of Ma=0.5 and attack angle α=10° are given respectively, Figure 7 The range of low pressure area on the upper surface of the airfoil and the range of high pressure area on the lower surface are larger than those of the original airfoil, so the airfoil has a higher lift-drag ratio. At the same time, the transition position of the optimized airfoil is later, and the separation area is smaller, which helps to reduce the friction resistance and improve the resistance characteristics of the airfoil. Thirdly, from the aerodynamic point of view: Figure 4 From the comparison of the pressure distribution of the airfoils before and after optimization, it can be seen that the pressure distribution on the lower surface of the airfoil after optimization does not change obviously. By comparing the pressure distribution on the upper surface, the pressure in the middle section of the original airfoil NACA 62(3)-213 is higher, which causes the airfoil to have a larger pitching moment. After optimization, the pressure in the middle section of the airfoil is reduced, so that the pitching moment of the airfoil is balanced, thereby reducing the absolute value of the moment of the airfoil.

[0065] The aerodynamic force coefficients of the present application airfoil (OPT) and the reference airfoil (NACA 62(3)-213) near the design point are as follows:

[0066] Ma=0.4

[0067]

[0068] Ma=0.5

[0069]

[0070] The aerodynamic characteristics of the present application near the design state are analyzed, and the aerodynamic force coefficients are shown in the above table. The lift-drag ratio curves of the present application and the reference airfoil near the design point are plotted and compared, as shown in Figure 5 , Figure 6 . Near the design point, the present application has a higher lift-drag ratio than NACA 62(3)-213 at different angles of attack. The moment coefficient curve comparison chart of the present application is analyzed, as shown in Figure 9 , Figure 10 . It can be observed that the present application has a smaller peak value of the pitching moment coefficient compared with NACA 62(3)-213.

[0071] On the basis of effectively balancing the lifting moment of the airfoil, the present application expands the laminar flow region and has a higher lift-drag ratio in the design angle of attack range, meeting the performance requirements of the new generation of high-speed helicopters.

[0072] Although the embodiments of the present application have been shown and described above, it should be understood that the above embodiments are exemplary and should not be construed as limiting the present application, and those of ordinary skill in the art can make changes, modifications, replacements and variations to the above embodiments without departing from the principles and purposes of the present application within the scope of the present application.

Claims

1. A subsonic, high-lift, low-drag laminar airfoil for use in the middle section of a helicopter rotor blade, characterized in that: The airfoil is applied to the mid-span of the helicopter rotor blade and is designed to expand the laminar flow region and improve the lift-to-drag ratio under low subsonic inflow conditions. Its leading edge radius is 0.00588, the maximum thickness is 13.10% located at 30.5% chord length, the maximum camber is 0.0126 located at 27.7% chord length, and the trailing edge angle is 6.78°. The leading edge radius, maximum thickness, and maximum camber are described using dimensionless quantities, with the airfoil chord length c as the reference.

2. The subsonic high-lift, low-drag laminar airfoil for the midsection of a helicopter rotor blade according to claim 1, characterized in that: The geometric coordinate expressions for the upper and lower surfaces of the airfoil are as follows: Where x represents the x-coordinate of the upper or lower surface of the airfoil, y represents the y-coordinate of the corresponding upper or lower surface of the airfoil, and n represents the order of the CST parameterization method. tail The y-coordinate represents the step at the root of the airfoil; The fitting coefficient for the upper surface of the airfoil is: The fitting coefficient for the lower surface of the airfoil is:

3. The subsonic high-lift, low-drag laminar airfoil for the midsection of a helicopter rotor blade according to claim 2, characterized in that: The fitting coefficients for the upper and lower surfaces of the airfoil are: The fitting coefficient for the upper surface of the airfoil is: The fitting coefficient for the lower surface of the airfoil is:

4. The subsonic high-lift, low-drag laminar airfoil for the midsection of a helicopter rotor blade according to claim 1, characterized in that: The upper and lower surface data of the airfoil are given in the table below: Top surface coordinates: The coordinates of the lower surface are as follows:

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