A low-moment supercritical natural laminar flow airfoil for mid-span of high-speed coaxial twin-rotor helicopter blades

Abstract

The application provides a low-moment supercritical natural laminar flow airfoil for a middle part of a high-speed coaxial dual-rotor helicopter blade, wherein a wing profile leading edge radius is 0.78%, a maximum thickness of the wing profile is 9.0% and is located at 40.6% of a chord length of the wing profile, a maximum camber is 0.69% and is located at 14.0% of the chord length of the wing profile, and a trailing edge angle is 1.46 degrees.The application can obviously reduce a zero-lift moment ratio of the OA309 wing profile in a transonic state under the condition of ensuring that a low-speed aerodynamic performance loss is not large, so that a usable Mach number range controlled by the moment of the wing profile is obviously increased; the drag divergence characteristics of the wing profile are obviously superior to those of the OA309 wing profile, and the Mach number reaches 0.844, which is 0.013 higher than that of the OA309 wing profile; the wing profile can weaken an upper surface shock wave of the wing profile in a transonic state, and meets the robustness requirement of a resistance coefficient of a middle and outer section wing profile of the coaxial dual-rotor helicopter rotor, and lays a foundation for the wing profile design of the middle part of the coaxial dual-rotor helicopter blade.

Description

TECHNICAL FIELD

[0001] The present application belongs to the field of rotor airfoil design, and particularly relates to a low-moment supercritical natural laminar flow airfoil for the middle part of a high-speed coaxial dual-rotor helicopter rotor blade. BACKGROUND

[0002] The rotor is the core component of helicopter hovering and forward flight. Since the rotor is in two different flow fields of forward flight and self-rotation at the same time, it will face two flow field characteristics of forward blade shock wave and rear blade stall, which also leads to the difficulty of greatly improving the forward flight speed of the helicopter. In order to break through this bottleneck, the coaxial dual-rotor helicopter concept emerges as the times require. The coaxial dual-rotor of this helicopter serves as the lift system when hovering vertically, and uses a thrust propeller as the propulsion system when forward flying, fully exerting the respective advantages and maintaining the high hovering efficiency and autorotation gliding ability of the coaxial dual-rotor and the high efficient propulsion ability of the propeller.

[0003] Due to the improvement of the forward flight speed of the coaxial helicopter, the actual Mach number of the rotor is higher. In the airfoil design, in addition to the constraint of moment control, the drag divergence characteristic becomes particularly important. In order to reduce the drag, the natural laminar flow design is the key to the airfoil design. The traditional natural laminar flow airfoil achieves sufficient laminar flow area by delaying pressure recovery, but at the same time, it also leads to a small pressure recovery range. At high speed, it is easy to form a shock wave, which increases the shock wave drag. This is the contradiction between the natural laminar flow design and the supercritical design. Moreover, for the coaxial helicopter, the shock wave intensity and position of the forward blade middle and outer segments are very sensitive to the Mach number, and the aerodynamic force coefficient changes dramatically. Therefore, the airfoil at this position also needs to maintain the robustness of the drag coefficient within a certain range. SUMMARY

[0004] In order to meet the requirements of the rotor airfoil drag divergence characteristic for the improvement of the forward flight speed of the coaxial dual-rotor helicopter and overcome the contradiction between the supercritical airfoil and the natural laminar flow airfoil design, and meet the robustness requirements of the rotor middle and outer segment airfoil drag coefficient of the coaxial dual-rotor helicopter, the present application proposes a low-moment supercritical natural laminar flow airfoil for the middle part of a high-speed coaxial dual-rotor helicopter rotor blade. Compared with the classic forward blade OA309 with a thickness of 9%, the airfoil has a larger leading edge radius, a smaller maximum camber, and the maximum thickness position is moved from 33% chord length to 40.6% chord length. In the transonic state, it can effectively weaken the shock wave intensity and improve the aerodynamic characteristics of the airfoil. At the same time, the airfoil trailing edge protrusion can further balance the moment, so that it has better moment characteristics than OA309 in the transonic range.

[0005] The technical scheme of the present application is as follows:

[0006] A low-moment supercritical natural laminar flow airfoil for middle part of high-speed coaxial twin-rotor helicopter blade, the airfoil has a leading edge radius of 0.78%, a maximum thickness of 9.0% at 40.6% of the chord length, a maximum camber of 0.69% at 14.0% of the chord length, and a trailing edge angle of 1.46 degrees; the leading edge radius, the maximum thickness, and the maximum camber are described by dimensionless quantities with the chord length c of the airfoil as the reference.

[0007] Further, the geometric coordinate expressions of the upper surface and the lower surface of the airfoil are:

[0008]

[0009] 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;

[0010] The fitting coefficients of the upper surface of the airfoil are:

[0011]

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

[0013]

[0014] Further, the fitting coefficients of the upper and lower surfaces of the airfoil are preferably:

[0015] The fitting coefficients of the upper surface of the airfoil are:

[0016]

[0017] The fitting coefficients of the lower surface of the airfoil are:

[0018]

[0019] Further, the coordinate point positions of the upper and lower surfaces of the low-moment supercritical natural laminar flow airfoil for middle part of high-speed coaxial twin-rotor helicopter blade are:

[0020] Upper surface coordinates:

[0021]

[0022]

[0023]

[0024] Lower surface coordinates:

[0025]

[0026]

[0027] Advantages

[0028] The application provides a low-moment supercritical natural laminar flow airfoil for a middle part of a high-speed coaxial dual-rotor helicopter blade. The application has obvious decline in zero lift moment ratio of the OA309 airfoil in a transonic state, so that the available Mach number range of the airfoil controlled by the moment is obviously increased; the drag divergence characteristics of the airfoil are also obviously superior to those of the OA309 airfoil, and the Mach number reaches 0.844, which is 0.013 higher than that of the OA309 airfoil, and the specific performance is that the airfoil can weaken the shock wave on the upper surface of the airfoil in a transonic state; and the airfoil meets the robustness requirement of the resistance coefficient of the middle and outer section airfoils of the coaxial dual-rotor helicopter rotor, and lays a foundation for the design of the middle part airfoil of the coaxial dual-rotor helicopter blade.

[0029] Additional aspects and advantages of the application will be set forth in part in the description which follows, and in part will become apparent to those skilled in the art upon examination of the following and / or can be learned by practice of the application. BRIEF DESCRIPTION OF DRAWINGS

[0030] The above and / or additional aspects and advantages of the application will become apparent and be readily understood by considering the following detailed description, including the accompanying drawings, in which:

[0031] Figure 1 It is a geometric shape diagram of the airfoil of the application.

[0032] Figure 2 It is a comparison diagram of the geometric shape of the airfoil of the application and the geometric shape of the OA309 airfoil.

[0033] Figure 3 It is a comparison diagram of the pressure distribution of the airfoil of the application and the OA309 airfoil at the investigation point.

[0034] Figure 4 It is a comparison diagram of the drag divergence curves of the airfoil of the application and the OA309 airfoil.

[0035] Figure 5 It is a comparison diagram of the moment curves of the airfoil of the application and the OA309 airfoil.

[0036] Figure 6 It is a comparison diagram of the lift coefficient curves of the airfoil of the application and the OA309 airfoil at a low speed state Ma=0.3.

[0037] Figure 7 It is a comparison diagram of the lift-drag ratio curves of the airfoil of the application and the OA309 airfoil at a low speed state Ma=0.3.

[0038] Figure 8Figure 1 is a lift coefficient curve comparison chart of the airfoil of the present application and the OA309 airfoil at a low speed state Ma=0.4.

[0039] Figure 9 Figure 2 is a lift-drag ratio curve comparison chart of the airfoil of the present application and the OA309 airfoil at a low speed state Ma=0.4.

[0040] Figure 10 Figure 3 is a lift coefficient curve comparison chart of the airfoil of the present application and the OA309 airfoil at a low speed state Ma=0.5.

[0041] Figure 11 Figure 4 is a lift-drag ratio curve comparison chart of the airfoil of the present application and the OA309 airfoil at a low speed state Ma=0.5.

[0042] Figure 12 Figure 5 is a lift coefficient curve comparison chart of the airfoil of the present application and the OA309 airfoil at a low speed state Ma=0.6.

[0043] Figure 13 Figure 6 is a lift-drag ratio curve comparison chart of the airfoil of the present application and the OA309 airfoil at a low speed state Ma=0.6. DETAILED DESCRIPTION

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

[0045] In order to meet the requirement of the forward flight speed improvement of the coaxial dual-rotor helicopter on the rotor airfoil drag divergence characteristics and overcome the contradiction between the supercritical airfoil and the natural laminar airfoil design, the present application designs a low-moment supercritical natural laminar airfoil for the middle part of the blade of the high-speed coaxial dual-rotor helicopter, which is used for the middle section of the rotor and needs to maintain the robustness of the drag coefficient within a certain range.

[0046] The design idea of the airfoil is to constrain the drag of the rotor airfoil at a low speed state Ma=0.3, Ma=0.4, Ma=0.5 and Ma=0.6 based on the proxy optimization method, so as to ensure that the aerodynamic performance at low speed is not severely lost. On this basis, the transonic aerodynamic characteristics are further optimized to reduce the drag and moment of the airfoil and improve the drag divergence Mach number. The above is the optimization mode of the airfoil, which does not belong to the characteristic parameters of the airfoil, and therefore is not described here.

[0047] The low-moment supercritical natural laminar airfoil for the middle part of the blade of the high-speed coaxial dual-rotor helicopter mentioned in the present embodiment has a high-speed investigation Mach number of 0.78-0.87 Mach and an acoustic speed Reynolds number of 7.2e6. The airfoil is drawn according to the drawing paper with a horizontal and vertical coordinate ratio of 1 as shown in Figure 1 .

[0048] The airfoil leading edge radius is 0.78%, the airfoil maximum thickness is 9.0% at 40.6% chord length, the maximum camber is 0.69% at 14.0% chord length, and the trailing edge angle is 1.46 degrees. It should be noted that the parameters described in the airfoil design are all dimensionless, so the above geometric characteristics are all the results of the chord length C dimensionless.

[0049] The specific airfoil upper and lower surface coordinate expression formula is:

[0050]

[0051] Where x represents the horizontal coordinate of the airfoil upper surface or lower surface, y represents the corresponding longitudinal coordinate of the airfoil upper surface or lower surface, n represents the order of the CST parameterization method, for this embodiment, the 7th order CST parameterization method is used, so n is 7, y tail represents the y coordinate of the airfoil root step.

[0052] The specific fitting coefficients of the airfoil upper and lower surfaces are:

[0053] The upper surface related parameters of the airfoil are described as follows:

[0054]

[0055] The lower surface related parameters of the airfoil are described as follows:

[0056]

[0057] The upper and lower surface coordinates of the airfoil in this embodiment are:

[0058] Upper surface coordinates:

[0059]

[0060]

[0061] Lower surface coordinates:

[0062]

[0063]

[0064] This embodiment is compared with the classic 9% thickness OA309 airfoil:

[0065] Figure 2For the contrastive drawing of the geometry of the two airfoils, in order to better observe the difference, the coordinate axis aspect ratio is set to 0.25, and it can be seen that the airfoil of the embodiment (named as OPT in the drawing legend) has a larger leading edge radius relative to the reference airfoil (named as OA309 in the drawing legend); the embodiment has a smaller slope on the upper surface of the leading edge and a larger slope on the lower surface of the leading edge relative to the reference airfoil; the position of the maximum thickness of the embodiment is moved backward relative to the reference airfoil; the embodiment has a smaller maximum camber relative to the reference airfoil, and is more convex at the trailing edge of the airfoil.

[0066] The differences in geometry inevitably lead to changes in aerodynamics. First, the aerodynamic performance of the present application in the high-speed zero-lift design state (Ma=0.84, Re=7.2e6, CL=0.00) is analyzed. Figure 3 The pressure distribution pattern of the airfoil of the present application and the OA309 airfoil in the high-speed zero-lift state is given. The decrease in the slope of the upper surface of the leading edge and the increase in the slope of the lower surface of the leading edge reduce the suction peak on the lower surface of the leading edge of the airfoil of the present application, and at the same time, convert a strong shock wave on the upper surface of the airfoil into two weak shock waves, which makes the airfoil have excellent low-drag characteristics and drag divergence characteristics within a certain Mach number range. The space formed by the pressure distribution curves of the upper and lower surfaces of the rear part of the reference airfoil OA309 is large, and the large nose-down moment is not balanced. The convexity of the rear part of the present application provides a nose-up moment, thereby reducing the absolute value of the moment of the airfoil.

[0067] Figure 4 For the contrastive drawing of the drag divergence curves of the airfoil of the present application and the OA309 airfoil in the transonic state, it can be seen that the airfoil of the present application has a lower zero-lift drag than the OA309 airfoil between 0.78 Mach and 0.87 Mach, and the drag divergence Mach number is increased from 0.831 to 0.844, an increase of 0.013 Mach. Figure 5 For the contrastive drawing of the moment curves of the airfoil of the present application and the OA309 airfoil in the transonic state, it can be seen that the absolute value of the moment of the airfoil of the present application is greatly reduced relative to the OA309 airfoil, and has more excellent moment characteristics.

[0068] The lift coefficient curves and lift-drag ratio curves of the present application in the low-speed state Ma=0.3, Ma=0.4, Ma=0.5 and Ma=0.6 are respectively as follows Figure 6 Figure 7 Figure 8 Figure 9 Figure 10 Figure 11 Figure 12 and Figure 13 ​​​​​​As shown, it can be observed that the low speed lift characteristics of the present application are reduced relative to OA309, the low speed lift-drag ratio characteristics are comparable to OA309 at small angles of attack, and the lift-drag ratio characteristics are reduced at large angles of attack, but this is acceptable for a rotor; and the robustness of the drag coefficient over a range can be maintained, satisfying the requirements of the mid-span airfoil of the high speed coaxial twin-rotor helicopter blade.

[0069] Although the embodiments of the present application have been shown and described above, it should be understood by those having ordinary skill in the art that the above embodiments are exemplary and are not to be construed as limiting the present application, and that changes, modifications, substitutions and variations of the above embodiments can be made by those skilled in the art without departing from the principles and spirit of the present application.

Claims

1. A low-moment supercritical natural laminar flow airfoil for use in the mid-span of a high-speed coaxial twin-rotor helicopter blade, characterized in that: The airfoil leading edge radius is 0.78%, the airfoil maximum thickness is 9.0% and is located at 40.6% chord length, the maximum camber is 0.69% and is located at 14.0% chord length, and the trailing edge angle is 1.46 degrees; the leading edge radius, the maximum thickness and the maximum camber are described by dimensionless quantity and are based on the airfoil chord length c; The airfoil upper surface coordinate point position is: The upper surface coordinate is: The lower surface coordinate is: 。 2. A low moment supercritical natural laminar flow airfoil for use in the mid-span of a high speed coaxial twin rotor helicopter blade as claimed in claim 1, wherein: The airfoil upper surface and lower surface geometric coordinate expression is: where x represents the transverse coordinate of the upper or lower surface of the airfoil, y represents the corresponding longitudinal coordinate of the upper or lower surface of the airfoil, n represents the order of the CST parameterization method, y tail y-coordinate representing the airfoil root step; The airfoil upper surface fitting coefficient is: The airfoil lower surface fitting coefficient is: 。 3. The low-moment supercritical natural laminar flow airfoil for the middle part of the blade of a high-speed coaxial dual-rotor helicopter according to claim 1, characterized in that: The airfoil upper surface fitting coefficient is: The airfoil lower surface fitting coefficient is: 。

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