A low-moment, low-subsonic, high-lift-to-drag ratio 18% thickness laminar airfoil based on a forward blade scheme

Abstract

The present invention proposes a low-moment, low-subsonic, high lift-to-drag ratio 18% thickness laminar airfoil based on a forward blade solution, the leading edge radius of the airfoil is 0.01682, the maximum thickness of the airfoil is 0.184728, located at 33.9% of the chord length of the airfoil, the maximum camber is 0.020937, located at 16.6% of the chord length of the airfoil, and the trailing edge angle is 7.64 degrees. Compared with the NACA 63(3)‑218 airfoil design, the leading edge radius of the airfoil is increased, the maximum thickness position is moved backward, the maximum camber is increased, its position on the airfoil is also moved forward, and the trailing edge angle is reduced. Within the design requirement range of 0.3‑0.5Ma, the lift-to-drag ratio of the designed airfoil is improved compared with the basic airfoil, the maximum lift-to-drag ratio is improved, the moment characteristics are improved, and the designed airfoil achieves the design purpose.

Description

Technical Field

[0001] The invention relates to the technical field of coaxial twin-rotor helicopter blade airfoil design, in particular to a low-moment, low-subsonic, high-lift-to-drag ratio, 18%-thickness laminar airfoil based on a forward blade solution. Background Art

[0002] In order to surpass the speed limit of helicopters, countries around the world are exploring new helicopter concepts and configurations. Existing high-speed helicopter configurations include compound, tilt-rotor and stall-rotor types, among which compound types include coaxial twin-rotor helicopters and helicopters with conventional rotors / forward blades equipped with auxiliary propulsion or lift devices. Coaxial rigid rotors can make full use of the high dynamic pressure of the forward blades to provide sufficient pulling force to maintain the helicopter's hovering, maneuvering and forward flight operations. Compared with traditional single rotors, the requirements for the aerodynamic performance of the airfoil of coaxial rigid rotors are also quite different. During high-speed forward flight, the rearward blades can be unloaded, thereby avoiding the resistance and noise surge caused by the reversal zone, which is conducive to increasing the forward flight speed. In 1964, the ABC rotor system proposed by Sikorsky Corporation of the United States, in the form of a coaxial rigid hingeless rotor, provides the main lift when flying forward at high speed by the forward blades of the upper and lower rotors, slowing down or eliminating the airflow separation in the reversal zone of the rearward blades, thus ensuring the ability to fly forward at high speed. Over the past 50 years, Sikorsky has been continuously developing this configuration. In 1973, the XH-59A achieved a maximum level flight speed of 238kt (about 441km / h) in its test flight. In 2008, the X2 achieved a maximum level flight speed of 250kt (about 460km / h). In 2015, the S-97 compound high-speed helicopter made its first flight, with a design speed of 260kt (480km / h). A large number of tests and verifications have shown that the ABC rotor high-speed helicopter configuration takes into account both hovering and high-speed flight capabilities, has a compact structure, and has good aerodynamic performance, maneuverability and controllability. It is the current trend in the development of high-speed helicopters.

[0003] In the process of helicopter rotor design, airfoil selection is the most critical step. It determines the aerodynamic performance of the rotor and the overall performance of the helicopter, and is the basis for determining the flight performance indicators of the helicopter, such as flight speed, range, noise, vibration, and commercial load. Therefore, airfoil analysis is a very important step in the rotor design process, and factors such as aircraft type, aerodynamic performance, structural dynamics performance, weight, cost, and anti-interference need to be considered to determine the best airfoil. When the helicopter is flying forward, the rotor blades are in an unsteady and asymmetric aerodynamic environment, and the flow characteristics near them are very different from those of fixed wings. Due to the continuous improvement of rotor performance requirements, it has become increasingly difficult to design advanced rotor airfoils. The development of new high-speed helicopters has also brought higher requirements for airfoil design. Therefore, airfoil design is a key part of the helicopter technology field and the basis for helicopter development. Taking the flight state as an example, when the rotor blades are in a hovering state, the airfoil is required to have a higher lift coefficient and lift-to-drag ratio to reduce power consumption. During high-speed flight and overload maneuvering flight, the trailing side blades need to have a higher maximum lift coefficient to delay the occurrence of stall. When flying forward at high speed, the leading blade needs to have a higher drag divergence Mach number to reduce the tip compressibility effect. In addition, in order to reduce the control load and blade torsional deformation, the rotor airfoil should also have a lower pitch moment coefficient. Summary of the invention

[0004] A certain type of coaxial twin-rotor helicopter requires the rotor to have a high maximum lift coefficient in the low Mach number to mid-subsonic state, a small zero-lift drag coefficient and high drag divergence characteristics in the transonic state, and needs to have a small pitch moment and low torque and control load. In addition, in order to ensure the hovering characteristics of the helicopter, it should also have a high lift-to-drag ratio in the hovering state.

[0005] In order to achieve the above design requirements, the rotor blade root needs to have anti-flow separation and low drag characteristics, so it is necessary to design a suitable low torque with anti-flow separation and low drag characteristics. By analyzing the velocity distribution of different sections during the forward flight of the coaxial twin rotor, it is determined that the maximum incoming flow velocity at the blade root airfoil section is 0.62Ma, the incoming flow velocity during hovering is 0.3Ma, and the inner wing section where the rotor blade root is located has a large range of laminar flow areas. Therefore, the final design requires the airfoil to have a high lift-to-drag ratio and low torque characteristics in the range of 0.3-0.6Ma, especially a high lift-to-drag ratio in the small angle of attack state of 0.5-0.6 Mach.

[0006] The initial airfoil of the design adopts the NACA 63(3)-218 airfoil. On the basis of the initial airfoil, the Mach number Ma obeys the N(0.4,0.052) distribution as the design state, the minimum drag coefficient is the optimization target, and the lift coefficient CL=0.7, the relative thickness of the airfoil does not decrease, and the pitching moment coefficient is greater than the initial airfoil moment coefficient are used as constraints for robust optimization. Finally, a low-moment, low-subsonic, high lift-to-drag ratio 18% thickness laminar airfoil based on the advancing blade scheme is obtained.

[0007] The leading edge radius of the airfoil is 0.01682, the maximum thickness of the airfoil is 0.184728, located at 33.9% of the chord length of the airfoil, the maximum camber is 0.020937, located at 16.6% of the chord length of the airfoil, and the trailing edge angle is 7.64 degrees. It should be noted that in the field of airfoil design, parameter descriptions are all described using dimensionless quantities, so the leading edge radius, maximum thickness, maximum camber, and subsequent airfoil coordinate descriptions are all described using dimensionless quantities, and the dimensionless process is based on the airfoil chord length c.

[0008] Furthermore, the geometric coordinate expressions of the upper and lower surfaces of the airfoil are

[0009]

[0010]

[0011] where z u (x) and z l (x) are the ordinate positions of the upper and lower surfaces of the unit airfoil, x is the abscissa position of the unit airfoil profile point, A u,i and A l,i (i=0,1,2…,7) are the fitting coefficients of the upper and lower surfaces of the airfoil in the above airfoil expression, respectively. te =0.00100; the fitting coefficients of the upper and lower surfaces of the airfoil are

[0012]

[0013] Furthermore, the fitting coefficients of the upper and lower surfaces of the airfoil are optimized as follows:

[0014]

[0015] Further, the upper and lower surface data of the coaxial twin-rotor helicopter airfoil with low moment, low subsonic speed and high lift-to-drag ratio are given in the following table:

[0016] Airfoil upper surface data

[0017]

[0018]

[0019] Airfoil lower surface data

[0020]

[0021] Beneficial Effects

[0022] The present invention provides a low-moment, low-subsonic, high lift-to-drag ratio 18% thickness rotor airfoil based on a forward blade solution based on the NACA 63(3)-218 airfoil. The design results show that compared with the basic airfoil, the leading edge radius of the designed airfoil increases, the maximum thickness position moves backward, the maximum curvature increases, its position on the airfoil also moves forward, and the trailing edge angle decreases. Within the design requirement range of 0.3-0.5Ma, the lift-to-drag ratio of the designed airfoil is improved compared with the basic airfoil, the maximum lift-to-drag ratio is improved, the moment characteristics are improved, and the designed airfoil achieves the design purpose.

[0023] Additional aspects and advantages of the present invention will be given in part in the following description and in part will be obvious from the following description, or will be learned through practice of the present invention. BRIEF DESCRIPTION OF THE DRAWINGS

[0024] The above and / or additional aspects and advantages of the present invention will become apparent and easily understood from the description of the embodiments in conjunction with the following drawings, in which:

[0025] Figure 1 This is a comparison diagram of the airfoil design and the airfoil shape.

[0026] Figure 2 The lift-to-drag ratio curve of the designed airfoil and the comparative airfoil varies with the lift coefficient (Ma=0.3).

[0027] Figure 3 The following is a graph showing the pitch moment coefficient versus lift coefficient for the designed airfoil and the comparative airfoil (Ma=0.3).

[0028] Figure 4 This is the lift coefficient ratio curve of the designed airfoil and the comparative airfoil as a function of the angle of attack (Ma=0.3).

[0029] Figure 5 This is the curve of the pitch moment coefficient ratio of the designed airfoil and the comparison airfoil changing with the angle of attack (Ma=0.3).

[0030] Figure 6 The lift-to-drag ratio curve of the designed airfoil and the comparative airfoil varies with the lift coefficient (Ma=0.4).

[0031] Figure 7 This is the curve diagram of the pitch moment coefficient of the designed airfoil and the comparison airfoil versus the lift coefficient (Ma=0.4).

[0032] Figure 8This is the lift coefficient ratio curve of the designed airfoil and the comparative airfoil as a function of the angle of attack (Ma=0.4).

[0033] Fig. 9 This is the curve of the pitch moment coefficient ratio of the designed airfoil and the comparison airfoil as a function of the angle of attack (Ma=0.4).

[0034] Fig.10 The lift-to-drag ratio curve of the designed airfoil and the comparative airfoil varies with the lift coefficient (Ma=0.5).

[0035] Fig.11 This is the curve diagram of the pitch moment coefficient of the designed airfoil and the comparison airfoil versus the lift coefficient (Ma=0.5).

[0036] Fig.12 This is the lift coefficient ratio curve of the designed airfoil and the comparative airfoil as a function of the angle of attack (Ma=0.5).

[0037] Fig.13 This is the curve of the pitch moment coefficient ratio of the designed airfoil and the comparison airfoil changing with the angle of attack (Ma=0.5). DETAILED DESCRIPTION

[0038] The present invention is aimed at a certain type of coaxial twin-rotor helicopter, which requires the rotor to have a higher maximum lift coefficient in a low Mach number to mid-subsonic state, a small zero-lift drag coefficient and a high drag divergence characteristic in a transonic state, and needs to have a smaller pitch moment and a low torque and control load. In addition, in order to ensure the hovering characteristics of the helicopter, the invention also has a design requirement of a high lift-to-drag ratio in the hovering state, and designs the airfoil at the root of the rotor.

[0039] Firstly, the velocity distribution of different sections of the coaxial twin rotor during forward flight was analyzed, and it was determined that the maximum incoming flow velocity at the airfoil section at the blade root was 0.62 Ma, the incoming flow velocity during hovering was 0.3 Ma, and the inner wing section where the rotor blade root was located had a large laminar area. Therefore, the final design required the airfoil to have a high lift-to-drag ratio and low torque characteristics in the range of 0.3-0.6 Ma, especially a higher lift-to-drag ratio at small angles of attack of 0.5 to 0.6 Mach.

[0040] Secondly, the NACA 63(3)-218 airfoil is used as the basic airfoil. Robust optimization is performed starting from the basic airfoil to ensure robustness within the design range. The design state, design objectives and constraints are shown in the following table:

[0041]

[0042] Finally, the low-moment, low-subsonic, high lift-to-drag ratio, 18% thickness laminar airfoil based on the advancing blade scheme was obtained and recorded as Opt-1.

[0043] The leading edge radius of the airfoil is 0.01682, the maximum thickness of the airfoil is 0.184728, located at 33.9% of the chord length of the airfoil, the maximum camber is 0.020937, located at 16.6% of the chord length of the airfoil, and the trailing edge angle is 7.64 degrees. It should be noted that in the field of airfoil design, parameter descriptions are all described using dimensionless quantities, so the leading edge radius, maximum thickness, maximum camber, and subsequent airfoil coordinate descriptions are all described using dimensionless quantities, and the dimensionless process is based on the airfoil chord length c.

[0044] The geometric coordinate expressions of the upper and lower surfaces of the airfoil are

[0045]

[0046]

[0047] where z u (x) and z l (x) are the ordinate positions of the upper and lower surfaces of the unit airfoil, x is the abscissa position of the unit airfoil profile point, A u,i and A l,i (i=0,1,2…,7) are the fitting coefficients of the upper and lower surfaces of the airfoil in the above airfoil expression, respectively. te =0.00100; the fitting coefficients of the upper and lower surfaces of the airfoil are

[0048]

[0049] The upper and lower surface data of the coaxial twin-rotor helicopter airfoil opt1 with low moment, low subsonic speed and high lift-to-drag ratio in this embodiment are given in the following table:

[0050] Airfoil upper surface data

[0051]

[0052] Airfoil lower surface data

[0053]

[0054]

[0055] The aerodynamic performance of the Opt-1 airfoil of this embodiment is compared with that of the NACA 63(3)-218 airfoil, and the lift-drag characteristics and moment characteristics of 0.3Ma to 0.5Ma are investigated under the conditions of Reynolds number 7.2e6, and the MSES program is used for evaluation, natural transition, and N = 6. As shown in the attached figure, in the range of 0.3-0.5Ma, the lift-drag ratio of the designed airfoil is improved compared with the basic airfoil, the maximum lift-drag ratio is improved, the moment characteristics are improved, and the designed airfoil achieves the design purpose.

[0056] Although the embodiments of the present invention have been shown and described above, it is to be understood that the above embodiments are exemplary and are not to be construed as limitations on the present invention. A person skilled in the art may change, modify, replace and modify the above embodiments within the scope of the present invention without departing from the principles and intent of the present invention.

Claims

1. A method for designing a low-moment, low-subsonic, high lift-to-drag ratio 18% thickness laminar airfoil based on a forward blade scheme, characterized in that: First, the design requirements of the rotor blade root are determined according to the airfoil application scenario; the airfoil application scenario is a coaxial twin-rotor helicopter, the rotor has a high maximum lift coefficient in the low Mach number to mid-subsonic state, a small zero-lift drag coefficient and a high drag divergence characteristic in the transonic state, and needs to have a small pitch moment and low torque and control load. In addition, in order to ensure the hovering characteristics of the helicopter, it should also have a high lift-to-drag ratio in the hovering state; the design requirements of the rotor blade root are that the rotor blade root has anti-flow separation and low drag characteristics; By analyzing the velocity distribution of different sections of the coaxial twin-rotor in forward flight, it is determined that the maximum incoming flow velocity at the airfoil section at the blade root is 0.62Ma, the incoming flow velocity during hovering is 0.3Ma, and the inner wing section where the rotor blade root is located has a large range of laminar flow area. Based on the NACA 63(3)-218 airfoil as the initial airfoil, the Mach number Ma obeys the N(0.4,0.052) distribution as the design state, the minimum drag coefficient is the optimization target, and the lift coefficient CL=0.7, the relative thickness of the airfoil does not decrease, and the pitch moment coefficient is greater than the initial airfoil moment coefficient are used as constraints for robust optimization. Finally, the low-moment, low-subsonic, high lift-to-drag ratio 18% thickness laminar airfoil based on the forward blade scheme is obtained as follows: The leading edge radius of the airfoil is 0.01682, the maximum thickness of the airfoil is 0.184728, located at 33.9% of the chord length of the airfoil, the maximum camber is 0.020937, located at 16.6% of the chord length of the airfoil, and the trailing edge angle is 7.64 degrees; the leading edge radius, maximum thickness, and maximum camber are described by dimensionless quantities, based on the chord length c of the airfoil; The airfoil is used for the rotor blade root of a coaxial twin-rotor helicopter; the maximum incoming flow velocity at the airfoil section at the blade root is 0.62 Ma, the incoming flow velocity during hovering is 0.3 Ma, and the inner wing section where the rotor blade root is located has a large laminar flow area; The geometric coordinate expressions of the upper and lower surfaces of the airfoil are where z u (x) and z l (x) are the ordinate positions of the upper and lower surfaces of the unit airfoil, x is the abscissa position of the unit airfoil profile point, A u,i and A l,i (i=0,1,2…,7) are the fitting coefficients of the upper and lower surfaces of the airfoil in the above airfoil expression, respectively. te =0.00100; the fitting coefficients of the upper and lower surfaces of the airfoil are 2. According to claim 1, a method for designing a low-moment, low-subsonic, high lift-to-drag ratio 18% thickness laminar airfoil based on a forward blade scheme, characterized in that: The fitting coefficients of the upper and lower surfaces of the airfoil are:

3. A method for designing a low-moment, low-subsonic, high lift-to-drag ratio 18% thickness laminar airfoil based on a forward blade scheme according to claim 1 or 2, characterized in that: The upper and lower surface data of the airfoil are given in the following table: Airfoil upper surface data Airfoil lower surface data

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