Airfoil for lifting element of aircraft

Abstract

The invention relates to an airfoil for a lifting element of an aircraft. The airfoil includes upper and lower contours formed by convex curves and intersection points of the convex curves, the upper and lower contours having given coordinates relative to a chord, the airfoil having a rounded leading edge. For the given chord length and maximum relative thickness of the airfoil, the relative coordinates # imgabs1 # # imgabs2 # of the convex curve of the airfoil are calculated from the midline of the airfoil at a position up to a relative distance # imgabs0 # from the leading edge. The maximum relative thickness of the airfoil is 10% to 25%. The airfoil of a lifting element of an aircraft allows a tail tab to be mounted behind the trailing edge of the airfoil, the tail tab having a rectangular or trapezoidal cross-sectional shape, including a curvilinear shape. The minimum resistance of the airfoil is reduced, the value of the maximum aerodynamic mass is increased, and a smaller value of the pitch moment at zero lift is provided.

Description

Technical Field

[0001] The invention relates to the field of aviation engineering, and in particular to an airfoil used for the middle and root parts of a main rotor blade and a tail rotor of an aircraft.

[0002] The airfoil of the NCV-2 lifting element of an aircraft, like other known airfoils having a similar purpose, consists of the profiles of an upper surface and a lower surface described by a set of geometrical coordinates of the upper surface and the lower surface.

[0003] The geometry of the upper and lower surfaces of the airfoil (i.e. the shape of the profile) determines the flow properties of the airfoil (i.e. the distribution of air pressure and velocity over these surfaces), and therefore the distribution of aerodynamic forces and moments that occur on the airfoil as air flows around it. Therefore, the shape of the airfoil profile determines the aerodynamic properties of the airfoil as air flows around it.

[0004] The main aerodynamic characteristics of the airfoil (i.e., maximum lift coefficient, minimum drag coefficient, maximum aerodynamic mass value, pitching moment coefficient at zero lift, position of the aerodynamic focus) have a significant influence on the maximum lifting capacity of the rotor, the power consumed by the rotor in various flight modes, the load level in the control system and the stable movement of the blades during rotor operation. Background Art

[0005] Known rotorcraft blade airfoils (CN106314791A, B64C27 / 467, published on January 11, 2017) have a maximum relative thickness of 8%, 9%, 12%. These airfoils have a modified centerline, the concavity of which increases the laminar part of the flow around the airfoil, which makes it possible to reduce aerodynamic drag.

[0006] A wing design is known (CN103693187A, B64C3 / 14, published on April 2, 2013), which consists of a set of airfoils, and the ratio of the wingspan to the chord length is 5.50-7.24; the leading edge is a parabolic curve; the maximum relative curvature of each wing airfoil is 7.5% and is located at the level of 17%-33% of the chord length; the maximum relative thickness of each wing airfoil is 13.1% and is located at the level of 11%-24% of the chord length. The design of the wing provides a high lift coefficient value, thereby delaying the stall of the airflow at high angles of attack and reducing noise in flight.

[0007] A helicopter blade airfoil is known (US6164918A, B64C27 / 46; B64C27 / 467, published on December 26, 2000), the geometry of which is determined by the radii of the upper and lower surfaces and the leading edge, which provide a large critical Mach number and lift coefficient and a reduced noise level.

[0008] The mentioned airfoils do not offer an optimal combination of aerodynamic properties and have specific application within the model range of aircraft for solving a limited range of tasks.

[0009] Unbalanced airfoils of the NACA 230XX series with five-digit designations are known, where the last two digits XX are the maximum relative thickness of the airfoil as a percentage of its chord (see, for example, NACA Report No. 586, 1937, p. 236; Kravets AS, Characteristics of aircraft airfoils, M.-L.: Oborongiz, 1939, pp. 206-217; Ushakov BA et al., Atlas of aerodynamic characteristics of wing airfoils, Moscow: BNT NKAP at TsAGI, 1940, pp. 165-196; NACA Report No. 824, 1945, pp. 101, 146-150). These airfoils are widely used in the design of helicopter blades.

[0010] The NACA230XX series of airfoils have an elongated teardrop shape with a rounded leading edge and a pointed or blunt trailing edge, the leading and trailing edges being connected by contoured portions of the upper and lower airfoil surfaces. The profiles of these airfoils are formed by a continuous smooth line with a smoothly varying curvature, by superimposing the profiles of symmetrical airfoils (with corresponding thicknesses) on a median line (perpendicular to the median line), which median line does not change for the entire series of airfoils. The maximum relative thickness of the airfoil series is located at 30% of the airfoil chord length, and the median line has a maximum relative concavity of 2% (along the axial arc) and is located at 15% of the airfoil chord length. The profile shape of the NACA23015 prototype airfoil obtained in this way determines the aerodynamic characteristics of the airfoil when the airflow flows around the airfoil.

[0011] The disadvantages of NACA23015 airfoil are:

[0012] - The value of the minimum drag coefficient of the airfoil is relatively high;

[0013] - low aerodynamic mass values ​​in the Mach number range of 0.15 to 0.25;

[0014] - In the Mach number range of 0.15 to 0.25, the proposed airfoil has lower values ​​of the pitching moment coefficient at zero lift than the NACA23012 airfoil. Summary of the invention

[0015] The technical problem to be solved by the present invention is to improve the basic aerodynamic characteristics of the airfoil of the root and middle part of the rotor blade, which airfoil operates at low Mach numbers of 0.15 to 0.45 and has a counterflow from the tail of the airfoil when the airfoil is used in the root of the blade; in the development of the geometry of the upper and lower surfaces of the profile of the airfoil NCV-2 (hereinafter referred to as the wing) of the lifting element of the aircraft, while:

[0016] - Compared to the prototype, it has a lower minimum drag coefficient C in the Mach number range from 0.15 to 0.45 xa min value;

[0017] - Compared to the prototype, it has the maximum lift coefficient K in the range of Mach numbers from 0.15 to 0.25 max The larger value of

[0018] - Compared to the prototype, it has a zero lift pitch moment coefficient C in the Mach number range from 0.15 to 0.25 mo The modulus value is smaller.

[0019] The specific aerodynamic characteristics of the airfoil (when installed in the root of the blade) have a significant influence on the power consumed by the main and tail rotors, the load levels on structural elements in the thruster control components, and the efficiency of the main rotor, tail rotor and other lifting surfaces of the aircraft.

[0020] The technical result of the application of the invention is the improvement of the basic aerodynamic characteristics of the airfoil for the root and middle portion of the rotor blade by using a developed airfoil profile which ensures the highest possible quality of the blade and rotor as a whole in hover mode and forward flight and an acceptable low level of variable and constant loads in the lifting system.

[0021] In order to achieve the technical effect, an airfoil of a lifting element of an aircraft is proposed, the airfoil comprising an upper contour and a lower contour formed by a convex curve and an intersection of a convex curve, the upper contour and the lower contour having given coordinates relative to the chord, the airfoil having a circular leading edge, at a relative distance from the leading edge At the position of For a given chord length and maximum relative thickness of an airfoil, measured from the airfoil centerline, Determined by equations (1) and (2):

[0022]

[0023]

[0024] in:

[0025] -Relative abscissa of the airfoil contour point, %;

[0026] -Relative ordinate of the point on the upper contour line of the airfoil, %;

[0027] -Relative ordinate of the point on the lower contour line of the airfoil, %;

[0028] - Relative ordinate of the point on the midline of the airfoil, %;

[0029] - relative ordinate of the points in the symmetric components of the upper and lower profiles of the airfoil, %;

[0030] - Maximum relative thickness of airfoil, %;

[0031] c MAX - Maximum thickness of airfoil, m;

[0032] X - the horizontal coordinate of the contour point of the airfoil, m;

[0033] Y B - the ordinate of the point on the upper contour of the airfoil, m;

[0034] Y Η - the ordinate of the point on the lower contour of the airfoil, m;

[0035] Β - chord length of airfoil, m,

[0036] At the same time, the following lists the distance from the leading edge The ordinate of the airfoil at the position The value associated with the chord length B, the distance For the maximum relative thickness Defines the vertical axis Range:

[0037]

[0038] Furthermore, the maximum relative thickness of the airfoil is in the range of 10% to 25%.

[0039] In this case, the airfoil of the lifting element of the aircraft may have a pointed, blunt or rounded trailing edge.

[0040] Furthermore, the airfoil of the lifting element of the aircraft allows the installation of a tail tab behind the trailing edge of the airfoil, said tail tab having a rectangular or trapezoidal cross-sectional shape, including a curved shape, a length not exceeding 15% of the chord without the tab, a thickness not exceeding 1.5% of the chord, the deflection angle of the tail tab relative to the chord being in the range of -5° to +10°.

[0041] Thus, the technical result of the application of the invention is achieved - the main aerodynamic characteristics of the airfoil of the root and middle of the rotor blade are improved by using the developed airfoil profile, which ensures the highest possible quality of the blade and the rotor as a whole in hover mode and translational flight, promoting an acceptable low level of variable and constant loads in the system. The use of the proposed airfoil allows to improve flight performance and economic characteristics, extend range and increase the competitiveness of the aircraft. BRIEF DESCRIPTION OF THE DRAWINGS

[0042] The invention is explained by the following drawings:

[0043] Figure 1 - For maximum thickness Plot of the variation of the airfoil profile and the dependence of the geometrical characteristics of the airfoil on the percentage of the airfoil chord length.

[0044] Figure 2 - Minimum drag coefficient С of the compared airfoils xa min Plot of the dependence on Mach number.

[0045] Figure 3 - Maximum aerodynamic quality coefficient K of the compared airfoils max Plot of the dependence on Mach number.

[0046] Figure 4 - The pitching moment coefficient C determined for the compared airfoil relative to the leading edge of the airfoil at zero lift m0 Plot of the dependence on Mach number. DETAILED DESCRIPTION

[0047] The proposed airfoil has a rounded leading edge and a pointed, blunt or rounded trailing edge, the leading edge and the trailing edge being interconnected by smooth profiles of the upper and lower surfaces of the airfoil, the leading edge being made with a rounded radius which increases smoothly along the chord length B of the airfoil so that the distance Y from the chord perpendicular to the chord upward to the upper part of the profile is B The upper part of the profile changes smoothly from convex to concave, increasing smoothly from the leading edge of the airfoil and reaching its maximum value in the coordinate range X=0.35B to 0.45B along the chord. The leading edge of the lower surface of the airfoil is rounded with a radius of curvature along the lower part of the profile. The distance Y from the chord to its lower surface is H It gradually increases from the leading edge and reaches its maximum value at X=0.3B to 0.4B.

[0048] The proposed airfoil refers to the root and the middle of the blade and comprises an upper and a lower profile formed by convex curves and their intersections, with given coordinates relative to the chord. The coordinates of the convex curves are measured from the airfoil centerline: the upper ones are plus, the lower ones are minus.

[0049] At a relative distance from the leading edge The relative coordinates of the convex curve of the airfoil at the position is measured from the airfoil centerline. For a given chord length and maximum relative thickness of the airfoil, Determined by equations (1) and (2):

[0050]

[0051]

[0052] in:

[0053] -Relative abscissa of the airfoil contour point, %;

[0054] -Relative ordinate of the point on the upper contour line of the airfoil, %;

[0055] -Relative ordinate of the point on the lower contour line of the airfoil, %;

[0056] - Relative ordinate of the point on the midline of the airfoil, %;

[0057] - Relative ordinate of the points in the symmetrical components of the upper and lower profiles of the airfoil, %.

[0058] - Maximum relative thickness of airfoil, %;

[0059] c MAX - Maximum thickness of airfoil, m;

[0060] X - the horizontal coordinate of the contour point of the airfoil, m;

[0061] Y B - the ordinate of the point on the upper contour of the airfoil, m;

[0062] Y Η - the ordinate of the point on the lower contour of the airfoil, m;

[0063] Β - chord length of airfoil, m,

[0064] Table 1 (columns 1, 2, and 3) lists the values ​​of the following items: Relative abscissas of points of the airfoil profile The corresponding relative ordinate of the airfoil centerline point and the ordinates of the points of symmetry of the upper and lower contours of the airfoil The relative ordinate of the upper and lower points of the airfoil The calculation of the range is performed based on the relations (1) and (2) respectively. Table 1 gives the maximum relative thickness of the airfoil. Example (See columns 4 and 5).

[0065] Table 1 Geometric characteristics of airfoil:

[0066]

[0067] When designing an airfoil, the value c MAX , X, Y Β , Y Η , the size of Β can be selected not only in meters, but also in other units.

[0068] The coordinates of a series of airfoils of different thicknesses can be plotted. In this case, The values ​​of remain the same as in Table 1 and for the maximum relative thickness of the airfoil in the range of 10% to 25% Based on equations (1) and (2), we calculate The value of .

[0069] By adding the relative value Multiply by the set value of the airfoil chord length Β to obtain the dimension values ​​X, Y CP , Y CИΜ , Y Β , Y Η .

[0070] The geometry of the airfoil profile ( Figure 1 ) defines the aerodynamics of the airfoil when air flows around it. The aerodynamic characteristics of the proposed airfoil obtained by computational fluid dynamics are Figures 2 to 4 The proposed airfoil is shown in FIG. 1 and compared with the characteristics of the NACA23015 prototype airfoil. In the graph, the dotted line represents the characteristics of the NACA23015 prototype airfoil, and the solid line represents the characteristics of the proposed airfoil.

[0071] Figure 1 Shown for the maximum thickness Dependence of the airfoil profile variation and the airfoil geometrical characteristics as a percentage of the airfoil chord length, where:

[0072] Y CP - the ordinate of the midline point of the airfoil;

[0073] Y CИΜ- ordinates of the points of symmetry of the upper and lower profiles of the airfoil;

[0074] Y Β - the vertical coordinate of the point on the upper contour, m;

[0075] Y Η - ordinate of the point of the lower contour, m;

[0076] Β – chord length of the airfoil;

[0077] - Maximum relative thickness of airfoil, %;

[0078] -Relative abscissa of the airfoil contour point, %;

[0079] - Relative ordinate of the contour point of the airfoil, %;

[0080] -Relative ordinate of the midline point of the airfoil, %;

[0081] - Relative ordinates of the points of symmetry components of the upper and lower profiles of the airfoil, %.

[0082] Figure 2 is the minimum drag coefficient of the compared airfoil xa min Plot of the dependence on Mach number. The minimum drag values ​​of the proposed airfoil are on average 33% lower than the minimum drag of the NACA23015 airfoil in the range of Mach numbers from 0.15 to 0.45.

[0083] Figure 3 The maximum aerodynamic mass K of the compared airfoils max Plot of dependence on Mach number. The maximum aerodynamic value of the proposed airfoil is 7% higher than that of the NACA 23015 airfoil in the Mach number range of 0.15 to 0.25.

[0084] Figure 4 The pitching moment coefficient C of the compared airfoil determined relative to the leading edge of the airfoil at zero lift m0 Graph of the results of the dependence on Mach number. The pitching moment values ​​of the proposed airfoil at zero lift are lower than those of the NACA 23015 airfoil in the Mach number range of 0.15 to 0.25.

[0085] Therefore, the proposed airfoil NCV-2 of the lifting element of an aircraft designed according to the essence of the present invention has advantages over the prototype airfoil NACA23015 in certain aerodynamic characteristics within a certain range of Mach numbers, namely, the drag of the airfoil is reduced, the value of the maximum lift to drag ratio, the value of the pitching moment (whose absolute value at zero lift is smaller) is provided.

[0086] When manufacturing the blades using the developed family of aerodynamic airfoils, it is planned to use state-of-the-art materials and techniques.

Claims

1. An airfoil of a lifting element of an aircraft, the airfoil comprising an upper profile and a lower profile formed by a convex curve and the intersection of the convex curve, the upper profile and the lower profile having given coordinates relative to the chord, the airfoil having a rounded leading edge, characterized in that At a relative distance from the leading edge The relative coordinates of the convex curve of the airfoil at the position is measured from the airfoil centerline. For a given chord length and maximum relative thickness of the airfoil, the relative coordinate It is determined by the relations (1) and (2): in: -Relative abscissa of the airfoil contour point, %; -Relative ordinate of the point on the upper contour line of the airfoil, %; -Relative ordinate of the point on the lower contour line of the airfoil, %; - Relative ordinate of the point on the midline of the airfoil, %; - relative ordinate of the points in the symmetric components of the upper and lower profiles of the airfoil, %; - Maximum relative thickness of airfoil, %; c MAX - Maximum thickness of airfoil, m; X - the horizontal coordinate of the contour point of the airfoil, m; Y B - the ordinate of the point on the upper contour of the airfoil, m; Y Η - the ordinate of the point on the lower contour of the airfoil, m; Β - chord length of airfoil, m, At the same time, the following gives the distance from the leading edge The ordinate of the airfoil at the position The value associated with the chord length B, the distance The maximum relative thickness of the airfoil Defines the vertical axis Range:

2. The airfoil of a lifting element of an aircraft according to claim 1, characterized in that The maximum relative thickness of the airfoil is between 10% and 25%.

3. The airfoil of a lifting element of an aircraft according to any one of claims 1 and 2, characterized in that The airfoil has a pointed, blunt or rounded trailing edge.

4. The airfoil of a lifting element of an aircraft according to any one of claims 1 to 3, characterized in that The airfoil allows a tail tab to be installed behind the trailing edge of the airfoil, the tail tab having a rectangular or trapezoidal cross-sectional shape, including a curved shape, a length not exceeding 15% of the chord without the tail tab, a thickness not exceeding 1.5% of the chord, and a deflection angle of the tail tab relative to the chord in the range of -5° to +10°.

Images