CN114194373A

CN114194373A - Aircraft wing and aircraft

Info

Publication number: CN114194373A
Application number: CN202111465506.6A
Authority: CN
Inventors: 曾锐; 刘毅; 李洪淼; 周义; 赵新新; 林琳; 冷崇富; 杨蕊姣
Original assignee: Cetc Wuhu Diamond Aircraft Manufacture Co ltd; Cetc Wuhu General Aviation Industry Technology Research Institute Co ltd
Current assignee: Cetc Wuhu Diamond Aircraft Manufacture Co ltd; Cetc Wuhu General Aviation Industry Technology Research Institute Co ltd
Priority date: 2021-12-03
Filing date: 2021-12-03
Publication date: 2022-03-18
Anticipated expiration: 2041-12-03
Also published as: CN114194373B

Abstract

The application discloses aircraft wing, its section structure includes wing section leading edge, wing section middle part upper segment, wing section middle part hypomere and wing section trailing edge to the extreme point that the camber line is close to wing section leading edge in the wing section is as the origin of coordinate system, and the line between the camber line both ends point is as the X axle of coordinate system in the wing section, then the shape and the size of wing section satisfy following relation: f/C is 1.875% -5.625%; Xf/C is 36.71%; t/C is 6.925% -20.775%; Xt/C is 32.31%; wherein f is the maximum value of the longitudinal coordinate of the camber line of the airfoil, Xf is the abscissa numerical value corresponding to the f point, t is the maximum thickness value of the airfoil, Xt is the abscissa value corresponding to the maximum thickness position of the airfoil, and C is the chord length of the airfoil. The specific shape of the airfoil is defined by a fairing line passing through a given coordinate point. The wing disclosed in the application has advantages such as the lift coefficient is big, drag coefficient is low when big lift coefficient in the transition reynolds number scope, the pitching moment of wing is lower, and inner space is big, and the improvement of airfoil trailing edge appearance has also reduced the degree of difficulty of manufacturing simultaneously.

Description

Aircraft wing and aircraft

Technical Field

The application relates to the technical field of wing airfoils, in particular to an aircraft wing and an aircraft.

Background

For an aircraft, the range of flight Reynolds numbers from 5 to 10 ten thousand is commonly referred to as a low Reynolds number, more than 600 ten thousand is a high Reynolds number, and the range of 100 to 300 ten thousand is between the high and low Reynolds numbers and may be referred to as a transition Reynolds number. In the range of the transitional Reynolds number, the flow characteristic of the airflow is closer to the state of high Reynolds number, but special flow states such as large laminar boundary layer area, local separation of bubbles in the laminar flow area and the like exist, and the key aerodynamic characteristic data of the airfoil, such as the maximum lift coefficient, the drag coefficient and the like, are often greatly changed in the area.

At present, a large number of general airplanes or unmanned planes and the like usually enter a transitional Reynolds number range in small-speed flight states such as take-off, landing and the like. However, common general-purpose aircraft mostly adopt NACA four-digit or five-digit airfoil profiles or NACA6 family airfoil profiles, and when the airfoil profiles operate in a transition reynolds number range, large performance loss can be generated, while low-reynolds-number airfoil profiles often have the problems of overlarge camber, small relative thickness, large low-head pitching moment and the like, so that the common general-purpose aircraft is not suitable for the transition reynolds number range.

Disclosure of Invention

Based on the problem of the above-mentioned background art, the application provides an aircraft wing and an aircraft, solves the problem that the low head pitching moment is big, the performance loss is serious when traditional wing airfoil shape moves in low reynolds number scope.

In a first aspect, the present application discloses an aircraft wing, the profile shape of which includes an airfoil leading edge, an airfoil middle upper section, an airfoil middle lower section and an airfoil trailing edge, and which is characterized in that the end point of an airfoil mean camber line close to the airfoil leading edge is taken as the origin of a coordinate system, and the connecting line between the two end points of the airfoil mean camber line is taken as the X-axis of the coordinate system, and then the shape and size of the airfoil satisfy the following relationships:

f/C＝1.875％-5.625％；

Xf/C＝36.71％；

t/C＝6.925％-20.775％；

Xt/C＝32.31％；

wherein f is the maximum value of the longitudinal coordinate of the camber line of the airfoil, Xf is the abscissa numerical value corresponding to the f point, t is the maximum thickness value of the airfoil, Xt is the abscissa value corresponding to the maximum thickness position of the airfoil, and C is the chord length of the airfoil.

Optionally, the shape of the aerofoil section is defined by a given coordinate point, and when the chord length C of the aerofoil is defined to be 1, the upper and lower surface coordinates of the aerofoil include:

x	on y is	y is below	x	On y is	y is below
						0	0	0	0.45	0.10144	-0.02873
0.005	0.01578	-0.01096	0.5	0.09614	-0.02686
						0.0125	0.02571	-0.01531	0.55	0.08890	-0.02459
0.025	0.03698	-0.01954	0.6	0.08011	-0.02149
						0.05	0.05250	-0.02427	0.65	0.06997	-0.01710
0.075	0.06397	-0.02708	0.7	0.05781	-0.01139
						0.1	0.07373	-0.02896	0.75	0.04369	-0.00509
0.15	0.08795	-0.03138	0.8	0.03133	-0.00166
						0.2	0.09708	-0.03250	0.85	0.02354	-0.00145
0.25	0.10289	-0.03271	0.9	0.01601	-0.00128
						0.3	0.10592	-0.03230	0.95	0.00849	-0.00111
0.35	0.10647	-0.03148	1	0.00097	-0.00094
						0.4	0.10488	-0.03028

When the x coordinate is unchanged, the coordinates above and below y can be synchronously scaled by 0.5-1.5 times.

Optionally, the upper and lower surfaces of the airfoil trailing edge are flat surfaces.

Optionally, the upper and lower surfaces of the airfoil trailing edge are in smooth transition with the upper section of the airfoil middle part and the lower section of the airfoil middle part respectively.

In a second aspect, the present application discloses an aircraft comprising the aircraft wing of any of the first aspects.

The beneficial effect of this application is as follows:

the maximum lift coefficient of the airfoil profile in the transitional Reynolds number range is improved, the maximum pitching moment is reduced, and the trimming moment required by the empennage is obviously reduced; the high lift coefficient has obvious resistance reduction effect, so that the wing profile resistance is reduced in the takeoff or crawling stage, and the flight state performance is improved; straightening the trailing edge part of the airfoil profile, so that the difficulty in production and manufacturing of the airfoil profile is reduced; as the airfoil profile is within the range of 20-80% of chord length, the thickness of the airfoil profile is increased, and the loading volume of fuel oil in the interior of the airfoil is increased.

Drawings

In order to more clearly illustrate the technical solutions of the present application, the drawings needed to be used in the description of the embodiments are briefly introduced below, it is obvious that the drawings in the following description are only some embodiments of the present application, and it is obvious for a person skilled in the art to obtain other drawings based on these drawings without exceeding the protection scope of the present application.

Fig. 1 is a schematic structural diagram of an aircraft wing according to an embodiment of the present application, where a parameter corresponding to the configuration is f/C — 3.75%; t/C is 13.85%.

FIG. 2 is a comparison of the profile of an embodiment of the present application with the profile of the original airfoil Wortmann FX 63-137.

FIG. 3 is a graph comparing the lift curves of an original airfoil and a classic NACA2412 airfoil according to an embodiment of the present application.

FIG. 4 is a graph comparing the drag curves of an original airfoil profile and an NACA2412 airfoil profile in accordance with an embodiment of the present application.

FIG. 5 is a comparison graph of pitch moment curves of an original airfoil and an NACA2412 airfoil according to an embodiment of the present application.

FIG. 6 is a graph comparing the pressure distribution of the original airfoil profile at a typical lift coefficient of 0.7 in the embodiment of the present application.

In the figure, 1, an airfoil leading edge; 2. an upper section of the middle of the airfoil; 3. the lower section of the middle part of the wing profile; 4. an airfoil trailing edge; 5. the airfoil mean camber line.

Detailed Description

The technical solutions of the present application will be described clearly and completely with reference to the accompanying drawings in the embodiments of the present application, and it is obvious that the described embodiments are some, not all, of the embodiments of the present application. All other embodiments, which can be derived by a person skilled in the art from the embodiments given herein without making any creative effort, shall fall within the protection scope of the present application.

The performance of the NACA four, five and NACA6 family airfoils is poor at lower Reynolds number ranges, and performance losses are also incurred when applied at transition Reynolds number ranges. Therefore, in the related art, one of the common schemes for selecting the airfoil profile in the range of the transitional reynolds number is a Wortmann FX63-137 airfoil profile (hereinafter referred to as an original airfoil profile), and wind tunnel tests show that the airfoil profile has a better lift-drag characteristic in the range of the transitional reynolds number, but the wind tunnel test also has the problems of large low head pitching moment, large trailing edge bending degree, difficulty in manufacturing, insufficient middle section thickness and the like, so that the wind tunnel test also has a great disadvantage when being directly applied to a general airplane or an unmanned aerial vehicle.

Referring to fig. 1, in the present application, an airfoil structure of an aircraft wing is represented by a cross-sectional view of the wing, and is divided into an airfoil leading edge 1, an airfoil middle upper section 2, an airfoil middle lower section 3, and an airfoil trailing edge 4 according to spatial positions.

In order to conveniently describe and present the shape and size of the airfoil, a two-dimensional coordinate system is established, the end point of the airfoil mean camber line 5 close to the airfoil leading edge 1 is taken as the origin of the coordinate system, the connecting line between the two end points of the airfoil mean camber line 5 is taken as the X axis of the coordinate system, the distance between the two end points of the airfoil mean camber line 5 is taken as the chord length C of the airfoil, the airfoil leading edge 1 is the front 20% part of the airfoil chord length C, the airfoil trailing edge 4 is the rear 20% part of the airfoil chord length C, and the part between the airfoil leading edge 1 and the airfoil trailing edge 4 is the middle part and is divided into an airfoil middle upper section 2 and an airfoil middle lower section 3.

After iterative optimization by the CFD method, the formed airfoil optimization scheme is as follows:

f/C is 1.875% -5.625%, wherein f is the maximum value of the mean camber line ordinate of the airfoil, and can also be called as the maximum camber of the airfoil; Xf/C is 36.71%, wherein Xf is an abscissa value corresponding to the f point in the coordinate system; t/C is 6.925% -20.775%, wherein t is the maximum thickness value of the airfoil; and Xt/C is 32.31 percent, wherein Xt is an abscissa value corresponding to the maximum thickness position of the airfoil.

As an alternative solution to the embodiment of the present application, the airfoil profile is determined by a fairing curve passing through given coordinate points, and the chord length C is defined as 1, and the reference parameter is f/C3.75%, and t/C13.85%. The coordinates of the upper and lower surfaces of the airfoil are then as shown in the following table:

It should be understood that the above shapes can be scaled in the y direction to form a series of airfoils, by scaling the y up and y down by 0.5-1.5 times simultaneously when the x coordinate is unchanged, and drawing a fairing line by the scaled coordinates.

For the airfoil profile, the typical transition Reynolds number is 150 ten thousand, the Reynolds mean NS equation is solved, the CFD calculation method of the gamma-theta transition model is combined, the aerodynamic characteristic difference of the airfoil profile, the original airfoil profile and the NACA2412 airfoil profile is compared and evaluated, and the obtained conclusion is as follows:

referring to fig. 3, the maximum lift coefficient of the present example is improved by 5.6% compared to the conventional NACA airfoil, and compared to the NACA2412 airfoil.

Referring to fig. 4, compared with the conventional NACA airfoil profile, the embodiment of the present invention has a significant drag reduction effect at a high lift coefficient, and the performance of the aircraft in the takeoff and climb flight states is improved. The drag is comparable to the NACA2412 airfoil at lift coefficients < 0.8, but the drag is reduced relative to the NACA2412 airfoil for climb when the lift coefficient > 0.8.

Referring to fig. 5, the absolute value of the pitch moment coefficient of the original wing profile of the embodiment of the application is reduced, the trimming moment required by the tail wing is obviously reduced, and the lift force and the drag force loss caused by trimming are improved. The absolute value of the pitching moment of the application example is equivalent to that of the NACA2412 wing profile, and the absolute value of the pitching moment of the original wing profile is reduced by 68 percent.

Referring to fig. 6, under the same general lift coefficient state, the pressure difference between the upper and lower airfoils of the embodiment of the present application is mainly concentrated on the front section, so as to improve the disadvantages of large load and large low head moment of the original airfoil rear section. In addition, as can be seen by combining the coordinates in the graph and fig. 2, in order to retain the advantage of the soft stall characteristic of the original airfoil in the range of the transition reynolds number, the profile and the coordinates of the leading edge and the middle upper section of the airfoil are kept unchanged.

According to the wing profile, the thickening area is designed at the lower section of the middle part, the 60% chord length is a control point of the maximum thickness increment, and the increase of the whole thickness of the wing profile within the range of 20% -80% chord length is driven, so that the loading volume of fuel oil in the wing is improved, and the moment arm of the wing for bearing bending moment is increased; the lifting amplitude is larger compared with the traditional NACA series wing profile loading volume.

Referring to fig. 1, as an alternative solution to the embodiment of the present application, the upper and lower surfaces of the trailing edge 4 of the airfoil in the present application are flat surfaces, and the trailing edge portion of the original airfoil is an arc portion that is curved downward and has a relatively narrow tip, so that compared with the large-curvature arc surface of the original airfoil, the straight section is changed from a sharp section, thereby reducing the difficulty of production and manufacturing.

Optionally, the upper and lower surfaces of the airfoil trailing edge 4 of the present application are in smooth transition with the airfoil middle upper section 2 and the airfoil middle lower section 3, respectively, so as to maintain the advantage of soft stall characteristics of the original airfoil. And the overall thickness of the airfoil trailing edge 4 is increased relative to the original airfoil.

The embodiment of the application also discloses an airplane, and the airplane wing in any one of the embodiments is adopted.

The embodiments of the present application are described in detail above. The principle and the implementation of the present application are explained herein by applying specific examples, and the above description of the embodiments is only used to help understand the technical solutions and the core ideas of the present application. Therefore, the person skilled in the art should, according to the idea of the present application, change or modify the embodiments and applications of the present application based on the scope of protection of the present application. In view of the above, the description should not be taken as limiting the application.

Claims

1. The cross section of the aircraft wing comprises a wing section leading edge, a wing section middle upper section, a wing section middle lower section and a wing section trailing edge, and is characterized in that the end point of a wing section mean camber line close to the wing section leading edge is used as the origin of a coordinate system, and the connecting line between the two end points of the wing section mean camber line is used as the X axis of the coordinate system, so that the shape and the size of the wing section meet the following relations:

f/C＝1.875％-5.625％；

Xf/C＝36.71％；

t/C＝6.925％-20.775％；

Xt/C＝32.31％；

2. An aircraft wing according to claim 1, characterized in that the shape of the aerofoil section is defined by a given coordinate point, and when defining the chord length C of the aerofoil to be 1, the upper and lower surface coordinates of the aerofoil comprise:

3. An aircraft wing according to claim 1, wherein the upper and lower surfaces of the trailing edge of the airfoil are plannar.

4. An aircraft wing according to claim 3, wherein the upper and lower surfaces of the trailing edge of the aerofoil are in smooth transition with the upper and lower mid-aerofoil sections respectively.

5. An aircraft comprising an aircraft wing as claimed in any one of claims 1 to 4.