阅读说明：本技术 一种适用于宽速域的跨音速导弹弹翼 (Transonic missile wing suitable for wide speed range ) 是由 胡雪垚 屈可朋 韩璐 肖玮 唐娇姣 吴翰林 于 2021-08-27 设计创作，主要内容包括：本发明提供了一种适用于宽速域的跨音速导弹弹翼,包括上翼面、下翼面、翼尖面、翼根面、前缘面、后缘面和导弹翼面结构空间,导弹翼面结构空间内设置有导弹翼面结构,该导弹翼面结构包括前翼梁、后翼梁和翼肋,其中翼肋为水滴状结构,翼肋的最大厚度d为0.09C,翼肋最大厚度d所在的位置为0.4C,翼肋的最大弯度ω为0.031C。本发明的适用于宽速域的跨音速导弹弹翼,在宽速域(马赫数0.3～1.0的速域)流动情况下,减阻效果显著,升力特性明显提高,失速特性缓和,且具有较高的升阻比,即具有优良的气动特性；其导弹翼面结构中的翼肋相较于常规翼面结构中的翼肋,特点是前端钝圆,上表面平坦,下表面接近前后缘处有反凹,该结构能够有效提高导弹弹翼的升力特性。(The invention provides a transonic missile wing suitable for a wide speed range, which comprises an upper wing surface, a lower wing surface, a wing tip surface, a wing root surface, a front edge surface, a rear edge surface and a missile wing surface structural space, wherein a missile wing surface structure is arranged in the missile wing surface structural space and comprises a front wing beam, a rear wing beam and a wing rib, the wing rib is of a water-drop-shaped structure, the maximum thickness d of the wing rib is 0.09C, the position of the maximum thickness d of the wing rib is 0.4C, and the maximum camber omega of the wing rib is 0.031C. The transonic missile wing suitable for the wide-speed range has the advantages that under the condition of flowing in the wide-speed range (the speed range with the Mach number of 0.3-1.0), the drag reduction effect is obvious, the lift characteristic is obviously improved, the stall characteristic is mild, and the transonic missile wing has a high lift-drag ratio, namely has excellent aerodynamic characteristics; compared with the wing rib in the conventional wing surface structure, the wing rib in the missile wing surface structure has the characteristics that the front end is blunt and round, the upper surface is flat, the lower surface is provided with the reverse concave part close to the front edge and the rear edge, and the structure can effectively improve the lift characteristic of the missile wing.)

1. A transonic missile wing suitable for a wide speed range comprises a left missile wing (1) and a right missile wing (2) which are oppositely arranged, wherein the left missile wing (1) and the right missile wing (2) have the same structure;

the left missile wing (1) is of an irregular hexahedral structure, the vertical top surface of the left missile wing (1) is an upper wing surface (101), and the vertical bottom surface of the left missile wing (1) is a lower wing surface (102); the horizontal outer surface of the left missile wing (1) is provided with a wing tip surface (103), the wing tip surface (103) is of a cambered surface structure, the maximum curvature position of the wing tip surface (103) is a wing tip (3), and the horizontal inner surface of the left missile wing (1) is a wing root surface (104); the longitudinal front surface of the left missile wing (1) is a front edge surface (105), and the longitudinal rear surface of the left missile wing (1) is a rear edge surface (106); spaces in the upper wing surface (101), the lower wing surface (102), the wing tip surface (103), the wing root surface (104), the leading edge surface (105) and the trailing edge surface (106) are missile wing surface structural spaces (107), and missile wing surface structures (4) are arranged in the missile wing surface structural spaces (107);

the intersection of the front edge surface (105) of the left missile wing (1) and the front edge surface (105) of the right missile wing (2) is a front edge (5), and the intersection of the rear edge surface (106) of the left missile wing (1) and the rear edge surface (106) of the right missile wing (2) is a rear edge (6);

the method is characterized in that:

the rear edge surface (106) is provided with a front break point (10601), and the front break point (10601) is positioned at 60% of the length of the left missile wing (1) by taking the position of the wing tip surface (103) as a starting point; a trailing edge surface (106) between the wing tip surface (103) and the front break point (10601) is an outer trailing edge surface (10602), and a trailing edge surface (106) between the front break point (10601) and the wing root surface (104) is an inner trailing edge surface (10603);

the missile airfoil structure (4) comprises a front wing beam (401), and the front wing beam (401) and a front edge surface (105) are arranged in parallel; the wing frame also comprises a rear wing beam (402), wherein an inflection point (40201) is arranged on the rear wing beam (402), the transverse outer end of the rear wing beam (402) is taken as a starting point, and the inflection point (40201) is positioned at 60% of the rear wing beam (402); the rear spar (402) between the transverse outer end of the rear spar (402) and the inflection point (40201) is an outer rear spar (40202), and the outer rear spar (40202) and the outer rear edge surface (10602) are arranged in parallel; the rear wing beam (402) between the inflection point (40201) and the transverse inner end of the rear wing beam (402) is an inner rear wing beam (40203), and the inner rear wing beam (40203) and the inner rear edge surface (10603) are arranged in parallel;

a plurality of front rib mounting holes (403) are formed in the front wing beam (401), and a plurality of rear rib mounting holes (404) are formed in the rear wing beam (402); a plurality of ribs (405) are transversely distributed between the front wing beam (401) and the rear wing beam (402), the longitudinal front parts of the ribs (405) extend out of the front rib (405) mounting holes (504) and are mounted in the front rib mounting holes (403), and the longitudinal rear parts of the ribs (405) extend out of the rear rib mounting holes (404) and are mounted in the rear rib mounting holes (404);

the bottom surface (40501) of the wing rib (405) is arranged on the inner wall of the lower wing surface (102), and the top surface (40502) of the wing rib (405) is closely adjacent to the inner wall of the upper wing surface (101);

the wing rib (405) is of a water-drop structure, the longitudinal front end of the wing rib (405) is of an arc-surface structure, the maximum curvature position of the longitudinal front end of the wing rib (405) is a wing rib front edge (40503), and the end point of the longitudinal rear end of the wing rib (405) is a wing rib rear edge (40504);

the maximum distance between the bottom surface (40501) and the top surface (40502) is the maximum thickness d, and the distance between the front edge (40503) and the rear edge (40504) is the chord length C; the maximum thickness d of the rib (405) is 0.09C; taking the position of the wing rib front edge (40503) as a starting point, and taking the position of the maximum thickness d of the wing rib (405) as 0.4C; the maximum camber omega of the wing rib (405) is 0.031C.

2. The transonic missile wing of claim 1 adapted for use in a wide range of speeds, wherein the leading edge face (105) is angled in the transverse direction at a leading edge sweep angle X0, and the leading edge sweep angle X0 is 30 °; the included angle between the outer rear edge surface (10602) and the transverse direction is a first rear edge sweepback angle X1, and the first rear edge sweepback angle X1 is 20 degrees; the connecting line between the end point of the transverse outer end of the outer trailing edge surface (10602) and the trailing edge (6) is a trailing edge line (7), the included angle between the trailing edge line (7) and the transverse direction is a second trailing edge sweepback angle X2, and the second trailing edge sweepback angle X2 is 13 degrees.

3. The transonic missile wing suitable for use in the wide velocity range of claim 1 wherein the distance between the wing tip (3) of the left missile wing (1) and the wing tip (3) of the right missile wing (2) is the span L1; the longitudinal length of the wing root surface (104) is wing root chord length B0; the longitudinal length of the wing tip surface (103) is the chord length B1 of the wing tip (3); and taking the position of the front edge (5) as a starting point, and taking one fourth of the chord length B0 of the wing root as an aerodynamic center (8).

4. The transonic missile wing suitable for use in the wide velocity range of claim 3, having an aspect ratio A of 4 and a tip to root ratio λ of 0.48.

5. The transonic missile wing suitable for use in the wide speed range of claim 1 wherein the XOY coordinate system is established with the leading rib edge (40503) at the location as the origin of coordinates O, wherein the (1, 0.0027) point is the trailing rib edge (40504) at the location;

the data point coordinates of the bottom surface (40501) of the wing rib are as follows:

the data point coordinates of the rib top surface (40502) are as follows:

6. the transonic missile launcher adapted for use in the wide speed range of claim 1, wherein the distance between rib leading edges (40503) of adjacent ribs (405) is a leading edge spacing L2, and the distance between rib trailing edges (40504) of adjacent ribs (405) is a trailing edge spacing L3;

the number of the ribs (405) is 28, wherein the rib (405) nearest to the tip surface (103) is the 1 st rib, and the rib (405) nearest to the root surface (104) is the 28 th rib; at an angle of attack α of 0 °, the airfoil standing geometry parameters of the 28 ribs (405) are as follows:

7. the transonic missile wing suitable for use in the wide speed range of claim 1, wherein the distance between the leading rib edge (40503) and the front spar (401) is the front spar rib pitch L4, and the distance between the trailing rib edge (40504) and the rear spar (402) is the rear spar rib pitch L5; the rib distance L4 of the front beam is 20% of the chord length C of the rib; the rib distance L5 of the back beam is 20% of the chord length C of the rib.

Technical Field

The invention belongs to the technical field of aerodynamics, relates to missile wing section design, and particularly relates to a transonic missile wing suitable for a wide speed range.

Background

Along with the development of military science and technology, the mission requirements of the missiles are increasingly complicated and diversified, so that the missiles are required to have good performance in a larger speed range, and therefore different combat missions are met, the wing ribs (also called wing profiles in the aspect of aerodynamics, and referred to as wing profiles in the following pneumatic analysis) have increasingly increased functions, the transonic wing profiles of the conventional missiles have better performance only under the transonic working condition, and the performance is seriously reduced under other working conditions, so that the military mission requirements in future wars are difficult to meet. Under the background, the wing section design of the wide-speed-range missile is urgently needed to be introduced, so that the combat adaptability of the tactical missile is improved, and the task requirement of future wars is met. The wide-speed-range missile wing profile not only improves transonic aerodynamic performance compared with the conventional wing profile, but also has good aerodynamic characteristics under subsonic working conditions.

At present, the missile wing profiles specially aiming at wide-speed-range missiles are rarely researched, and in the prior art, the realized classical scheme is the SC-XXX supercritical series wing profiles proposed by the Lanli research center of the national space agency. When approaching the speed of sound, the SC-XXX wing profile can delay the occurrence of the phenomenon of resistance sharp increase, so that the missile has better transonic flight performance. Although the SC-XXX wing profile can delay the occurrence of the resistance surge phenomenon, the wing surface structure of the SC-XXX wing profile is not specially designed for the aerodynamic characteristics of the missile wing transonic wing profile of the wide-speed-range missile, and the aerodynamic performance of the SC-XXX wing profile still has a space for further improvement.

Disclosure of Invention

Aiming at the defects in the prior art, the invention aims to provide a transonic missile wing suitable for a wide-speed range, and solve the technical problem that the aerodynamic characteristics of the missile wing are insufficient under the flow of the wide-speed range in the prior art.

In order to solve the technical problems, the invention adopts the following technical scheme:

a transonic missile wing suitable for a wide speed range comprises a left missile wing and a right missile wing which are oppositely arranged, and the left missile wing and the right missile wing have the same structure;

the left missile wing is of an irregular hexahedral structure, the vertical top surface of the left missile wing is an upper wing surface, and the vertical bottom surface of the left missile wing is a lower wing surface; the transverse outer surface of the missile wing of the left missile is a wing tip surface, the wing tip surface is of a cambered surface structure, the maximum curvature position of the wing tip surface is a wing tip, and the transverse inner surface of the missile wing of the left missile is a wing root surface; the longitudinal front surface of the missile wing is a front edge surface, and the longitudinal rear surface of the missile wing is a rear edge surface; the space in the upper wing surface, the lower wing surface, the wing tip surface, the wing root surface, the front edge surface and the rear edge surface is a missile wing surface structure space, and a missile wing surface structure is arranged in the missile wing surface structure space;

the intersection of the front edge surface of the left missile wing and the front edge surface of the right missile wing is a front edge, and the intersection of the rear edge surface of the left missile wing and the rear edge surface of the right missile wing is a rear edge;

the rear edge surface is provided with a front break point which is positioned at 60 percent of the length of the missile wing by taking the position of the wing tip surface as a starting point; the rear edge surface between the wing tip surface and the front break point is an outer rear edge surface, and the rear edge surface between the front break point and the wing root surface is an inner rear edge surface;

the missile airfoil structure comprises a front wing beam, and the front wing beam is arranged in parallel with a front edge surface; the rear wing beam is provided with an inflection point, the transverse outer end of the rear wing beam is taken as a starting point, and the inflection point is positioned at 60% of the rear wing beam; the rear wing beam between the transverse outer end and the inflection point of the rear wing beam is an outer rear wing beam, and the outer rear wing beam is arranged in parallel with the outer rear edge surface; the rear wing beam between the inflection point and the transverse inner end of the rear wing beam is an inner rear wing beam, and the inner rear wing beam is arranged in parallel with the inner rear edge surface;

the front wing beam is provided with a plurality of front wing rib mounting holes, and the rear wing beam is provided with a plurality of rear wing rib mounting holes; a plurality of wing ribs are transversely arranged between the front wing beam and the rear wing beam, the longitudinal front parts of the wing ribs extend out of the front wing rib mounting holes and are mounted in the front wing rib mounting holes, and the longitudinal rear parts of the wing ribs extend out of the rear wing rib mounting holes and are mounted in the rear wing rib mounting holes;

the bottom surface of the wing rib is arranged on the inner wall of the lower airfoil surface, and the top surface of the wing rib is closely adjacent to the inner wall of the upper airfoil surface;

the wing rib is of a water-drop structure, the longitudinal front end of the wing rib is of an arc surface structure, the maximum curvature position of the longitudinal front end of the wing rib is a wing rib front edge, and the end point of the longitudinal rear end of the wing rib is a wing rib rear edge;

the maximum distance between the bottom surface of the rib and the top surface of the rib is the maximum thickness d, and the distance between the front edge of the rib and the rear edge of the rib is the chord length C of the rib; the maximum thickness d of the rib is 0.09C; taking the position of the leading edge of the wing rib as a starting point, and taking the position of the maximum thickness d of the wing rib as 0.4C; the maximum camber omega of the wing rib is 0.031C.

The invention also has the following technical characteristics:

the included angle between the front edge surface and the transverse direction is a front edge sweepback angle X0, and the front edge sweepback angle X0 is 30 degrees; the included angle between the outer rear edge surface and the transverse direction is a first rear edge sweepback angle X1, and the first rear edge sweepback angle X1 is 20 degrees; the connecting line between the end point of the outer trailing edge surface transverse outer end and the trailing edge is a trailing edge line, the included angle between the trailing edge line and the transverse direction is a second trailing edge sweepback angle X2, and the second trailing edge sweepback angle X2 is 13 degrees.

The distance between the wing tip of the left missile wing and the wing tip of the right missile wing is L1; the longitudinal length of the wing root surface is wing root chord length B0; the longitudinal length of the wing tip surface is wing tip chord length B1; the position of the front edge is taken as a starting point, and one quarter of the chord length B0 of the wing root is taken as an aerodynamic center.

The missile wing has an aspect ratio A of 4 and a tip-root ratio lambda of 0.48.

Establishing an XOY coordinate system by taking the position of the leading edge of the wing rib as a coordinate origin O, wherein the point (1, 0.0027) is the position of the trailing edge of the wing rib;

the data point coordinates of the bottom surface of the rib are as follows:

the data point coordinates of the rib top surface are as follows:

the distance between the rib leading edges of the adjacent ribs is the leading edge spacing L2, and the distance between the rib trailing edges of the adjacent ribs is the trailing edge spacing L3;

the number of the wing ribs is 28, wherein the wing rib closest to the wing tip surface is the 1 st wing rib, and the wing rib closest to the wing root surface is the 28 th wing rib; at an angle of attack α of 0 °, the airfoil standing geometry parameters of the 28 ribs are as follows:

the distance between the rib front edge and the front wing beam is L4, and the distance between the rib rear edge and the rear wing beam is L5; the rib distance L4 of the front beam is 20% of the chord length C of the rib; the rib distance L5 of the back beam is 20% of the chord length C of the rib.

Compared with the prior art, the invention has the following technical effects:

compared with the wing rib in the conventional wing surface structure, the wing rib in the wing surface structure of the transonic missile suitable for the wide speed range has the characteristics that the front end is blunt and round, the upper surface is flat, and the lower surface is provided with the reverse concave part close to the front edge and the rear edge.

The transonic missile wing suitable for the wide-speed range has the advantages that the drag reduction effect is obvious, the lift characteristic is obviously improved, the stall characteristic is mild, the lift-drag ratio is high, and the transonic missile wing has excellent aerodynamic characteristics under the condition of flowing in the wide-speed range (the Mach number is 0.3-1.0).

Drawings

Fig. 1 is a schematic overall structure diagram of a transonic missile wing suitable for a wide speed range.

FIG. 2 is a schematic view of the overall structure of a missile airfoil structure.

FIG. 3 is a schematic view of the geometric parameters of a transonic missile wing suitable for a wide speed range.

FIG. 4 is a schematic representation of the geometric parameters of a missile airfoil configuration.

FIG. 5 is a graph of the geometric profile of the missile airfoil configuration of example 1 and comparative example 1, wherein the solid line represents the HU-SC90 airfoil profile of example 1 and the dashed line represents the SC1095 airfoil profile of comparative example 1.

FIG. 6 is a graph of the lift characteristics of the missile airfoil configurations of example 1 and comparative example 1 in a first state of calculation, wherein the solid line represents the HU-SC90 airfoil profile of example 1 and the dashed line represents the SC1095 airfoil profile of comparative example 1.

FIG. 7 is a graph of the lift-drag characteristics of the missile airfoil configurations of example 1 and comparative example 1 in a first state of calculation, wherein the solid line represents the HU-SC90 airfoil profile of example 1 and the dashed line represents the SC1095 airfoil profile of comparative example 1.

FIG. 8 is a plot of lift-to-drag ratios for the missile airfoil configurations of example 1 and comparative example 1 at a first calculated condition, wherein the solid line represents the HU-SC90 airfoil profile of example 1 and the dashed line represents the SC1095 airfoil profile of comparative example 1.

FIG. 9 is a graph of the moment characteristics of the missile airfoil configurations of example 1 and comparative example 1 in a first state of calculation, wherein the solid line represents the HU-SC90 airfoil profile of example 1 and the dashed line represents the SC1095 airfoil profile of comparative example 1.

FIG. 10 is a graph of the lift characteristics of the missile airfoil configurations of example 1 and comparative example 1 in a second calculated condition, wherein the solid line represents the HU-SC90 airfoil profile of example 1 and the dashed line represents the SC1095 airfoil profile of comparative example 1.

FIG. 11 is a graph of the lift-drag characteristics of the missile airfoil configurations of example 1 and comparative example 1 at a second calculated condition, wherein the solid line represents the HU-SC90 airfoil profile of example 1 and the dashed line represents the SC1095 airfoil profile of comparative example 1.

FIG. 12 is a plot of lift-to-drag ratios for the missile airfoil configurations of example 1 and comparative example 1 at a second calculated condition, wherein the solid line represents the HU-SC90 airfoil profile of example 1 and the dashed line represents the SC1095 airfoil profile of comparative example 1.

FIG. 13 is a graph of the moment characteristics of the missile airfoil configurations of example 1 and comparative example 1 in a second state of calculation, wherein the solid line represents the HU-SC90 airfoil profile of example 1 and the dashed line represents the SC1095 airfoil profile of comparative example 1.

The meaning of the individual reference symbols in the figures is: 1-left missile wing, 2-right missile wing, 3-wingtip, 4-missile airfoil surface structure, 5-leading edge, 6-trailing edge, 7-trailing edge line and 8-aerodynamic center;

101-upper wing surface, 102-lower wing surface, 103-wing tip surface, 104-wing root surface, 105-leading edge surface, 106-trailing edge surface, and 107-missile wing surface structure space;

401-front spar, 402-rear spar, 403-front rib mounting hole, 404-rear rib mounting hole, 405-rib;

10601-front break point, 10602-outer trailing edge surface, 10603-inner trailing edge surface;

40201-inflection point, 40202-outer rear wing spar, 40203-inner rear wing spar;

40501-rib bottom, 40502-rib top, 40503-rib leading, 40504-rib trailing.

The present invention will be explained in further detail with reference to examples.

Detailed Description

The conventional supercritical missile wing surface is applicable to a narrow speed range, and has good aerodynamic performance mostly in the Mach number range of 0.7-1.0, however, with the continuous improvement of military requirements, a tactical missile develops towards the direction of a wide speed range, and the tactical missile inevitably experiences multiple flight stages such as subsonic speed, transonic speed and the like in actual flight. The invention designs transonic missile wings suitable for wide speed range, which are mainly designed according to the following technical indexes: first, the drag coefficient is as small as possible; second, the maximum thickness is 40% of the chord length; third, the lift-to-drag ratio is as large as possible; fourth, the resistance dispersion characteristics are good.

According to the technical index requirements, the invention designs the transonic missile wing suitable for the wide speed range, the missile wing surface structure is named as an HU-SC90 wing type, and the HU-SC90 wing type has the characteristics of low resistance, high lift-drag ratio and mild stall characteristic. Compared with the wing rib in the conventional wing surface structure, the wing rib in the missile wing surface structure has the characteristics that the front edge is blunt and round, the upper surface is flat, and the lower surface is provided with the reverse concave part close to the front edge and the rear edge, so that the wing profile has the characteristics of front loading and rear loading, and the lift characteristic of the wing profile can be effectively improved. The HU-SC90 wing profile has higher aerodynamic efficiency and better resistance divergence characteristic in a wide speed range (the speed range with the Mach number of 0.3-1.0), and the missile wing surface structure can be suitable for a subsonic state, improves the performance of a transonic missile, and can meet the requirements of a new-generation tactical missile.

In the invention:

the wide speed range refers to a speed range with Mach number of 0.3-1.0.

The attack angle alpha is an included angle between the advancing direction of missile wings and the chord length direction C of wing ribs.

The SC1095 airfoil is of a missile airfoil construction known in the art for comparison with the HU-SC90 airfoil of the present invention.

The following embodiments of the present invention are provided, and it should be noted that the present invention is not limited to the following embodiments, and all equivalent changes based on the technical solutions of the present invention are within the protection scope of the present invention.

Example 1:

the embodiment provides a transonic missile wing suitable for a wide speed range, the transonic missile wing is of a structure named as HU-SC90 airfoil profile, as shown in FIGS. 1 to 5, and comprises a front wing beam 401, and the front wing beam 401 is arranged in parallel with a front edge surface 105; the rear wing beam is characterized by further comprising a rear wing beam 402, wherein an inflection point 40201 is arranged on the rear wing beam 402, the transverse outer end of the rear wing beam 402 serves as a starting point, and the inflection point 40201 is located at 60% of the rear wing beam 402; the rear spar 402 between the transverse outer end of the rear spar 402 and the inflection point 40201 is an outer rear spar 40202, and the outer rear spar 40202 is arranged in parallel with the outer rear edge surface 10602; the rear spar 402 between the inflection point 40201 and the transverse inner end of the rear spar 402 is an inner rear spar 40203, and the inner rear spar 40203 is arranged in parallel with an inner rear edge surface 10603;

a plurality of front rib mounting holes 403 are formed in the front wing beam 401, and a plurality of rear rib mounting holes 404 are formed in the rear wing beam 402; a plurality of ribs 405 are transversely distributed between the front wing beam 401 and the rear wing beam 402, the longitudinal front parts of the ribs 405 extend out of the front rib 405 mounting holes 504 and are mounted in the front rib mounting holes 403, and the longitudinal rear parts of the ribs 405 extend out of the rear rib mounting holes 404 and are mounted in the rear rib mounting holes 404; the bottom surface 40501 of rib 405 is mounted on the inner wall of the lower airfoil surface 102 and the top surface 40502 of rib 405 is in close proximity to the inner wall of the upper airfoil surface 101.

As a specific scheme of this embodiment, the rib 405 is a drop-shaped structure, the longitudinal front end of the rib 405 is an arc-shaped structure, the maximum curvature of the longitudinal front end of the rib 405 is a rib front edge 40503, and the end point of the longitudinal rear end of the rib 405 is a rib rear edge 40504;

the maximum distance between the rib bottom surface 40501 and the rib top surface 40502 is the maximum thickness d, and the distance between the rib leading edge 40503 and the rib trailing edge 40504 is the rib chord length C; the maximum thickness d of rib 405 is 0.09C; taking the position of the rib leading edge 40503 as a starting point, and the position of the maximum thickness d of the rib 405 is 0.4C; the maximum camber ω of the rib 405 is 0.031C.

In this embodiment, the camber ω is calculated according to the following formula: ω | y1| - | y2| - | S1|, where y1 represents the ordinate of the rib top surface 40502; y2 represents the ordinate of rib bottom surface 40501 on the same abscissa; s1 represents the ordinate of the geometric chord length on the same abscissa. The maximum camber ω is the maximum of all camber ω values.

In the embodiment, the position of the rib leading edge 40503 is taken as a starting point, the part of the rib 405 between 0 and 0.1C is an obtuse circular structure, the rib top surface 40502 is a flat curve at 0.3C to 0.8C, and the rib bottom surface 40501 is thinner than the conventional rib at 0.95C to 1.0C; the rib 405 with the structure has the characteristics of front loading and rear loading, and can effectively improve the lift characteristic of missile wings.

As a specific solution of this embodiment, the included angle between the leading edge surface 105 and the transverse direction is a leading edge sweep angle X0, and the leading edge sweep angle X0 is 30 °; the outer trailing edge face 10602 makes an angle with the transverse direction of a first trailing edge sweep angle X1, the first trailing edge sweep angle X1 being 20 °; a connecting line between the end point of the transversely outer end of the outer trailing edge surface 10602 and the trailing edge 6 is a trailing edge line 7, the included angle between the trailing edge line 7 and the transverse direction is a second trailing edge sweep angle X2, and the second trailing edge sweep angle X2 is 13 °.

As a specific scheme of the embodiment, the distance between the wing tip 3 of the left missile wing 1 and the wing tip 3 of the right missile wing 2 is the wingspan L1; the longitudinal length of the root surface 104 is a root chord length B0; the longitudinal length of the tip surface 103 is the chord length B1 of the wing tip 3; starting from the position of the leading edge 5, one quarter of the chord length B0 of the wing root is the aerodynamic center 8.

As a specific scheme of this embodiment, the aspect ratio a of the missile wing is 4, and the aspect ratio a is calculated according to the following formula: a ═ L1/[0.5(B0+ B1) ]; the missile wing has a tip-root ratio lambda of 0.48, and the tip-root ratio lambda is calculated according to the following formula: λ is B0/B1.

As a specific scheme of this embodiment, an XOY coordinate system is established with the position of the leading edge 40503 of the rib as the origin O, where the point (1, 0.0027) is the position of the trailing edge 40504 of the rib;

the data point coordinates for rib floor 40501 are:

the data point coordinates for rib top surface 40502 are:

as a specific solution of this embodiment, the distance between the rib leading edges 40503 of the adjacent ribs 405 is a leading edge spacing L2, and the distance between the rib trailing edges 40504 of the adjacent ribs 405 is a trailing edge spacing L3;

the number of ribs 405 is 28, wherein the rib 405 closest to the tip face 103 is the 1 st rib, and the rib 405 closest to the root face 104 is the 28 th rib; at an angle of attack α of 0 °, the airfoil standing geometry parameters of the 28 ribs 405 are as follows:

as a specific solution of this embodiment, the distance between the leading edge 40503 of the rib and the front spar 401 is a front spar rib distance L4, and the distance between the trailing edge 40504 of the rib and the rear spar 402 is a rear spar rib distance L5; the rib distance L4 of the front beam is 20% of the chord length C of the rib; the trailing spar rib pitch L5 is 20% of the rib chord length C.

In this embodiment, the aerodynamic characteristics of the HU-SC90 airfoil are analyzed and calculated, and the specific design state parameters are as follows: the attack angle alpha (same attack angle alpha) is 4 degrees, the Mach number is 0.8, and the Reynolds number is 6.4 multiplied by 10⁶. The aerodynamic characteristics are shown in table 1 at this design parameter. As can be seen from table 1, the HU-SC90 airfoil of the present embodiment has excellent aerodynamic performance, significant drag reduction, and excellent torque characteristics in the above design state.

In this embodiment, a computational fluid dynamics CFD numerical simulation method is adopted, a k- ω SST model is adopted to perform turbulence simulation, aerodynamic performance evaluation is performed on the HU-SC90 airfoil profile, the aerodynamic performance evaluation adopts two calculation states, wherein parameters of the first calculation state are: mach number 0.5, Reynolds number 4.0X 10⁶The results are shown in fig. 6 to 9. The parameters of the second calculation state are: mach number 0.8, Reynolds number 6.4X 10⁶The results are shown in fig. 10 to 13.

Comparative example 1:

the present comparative example provides a missile wing, the missile wing surface structure of which is designated as SC1095 wing profile, as shown in fig. 5, and is known in the prior art.

In this comparative example, the aerodynamic characteristics of the SC1095 airfoil were analyzed and calculated, the specific design state parameters were the same as those of example 1, and the aerodynamic characteristics are shown in table 1. In this comparative example, the simulation method, model, and calculation transition parameter used in the aerodynamic performance evaluation were the same as those of example 1, and the results are shown in fig. 6 to 13.

TABLE 1 Main aerodynamic characteristics of the HU-SC90 airfoil of example 1 and the SC1095 airfoil of comparative example 1

The following conclusions can be drawn from example 1 and comparative example 1:

(A) as can be seen from Table 1, the HU-SC90 airfoil profile of example 1 meets design requirements of low resistance, high lift-drag ratio and stalling characteristic alleviation, and the HU-SC90 airfoil profile of example 1 is superior in aerodynamic performance, obvious in drag reduction and better in torque characteristic compared with the SC1095 airfoil profile of comparative example 1.

(B) As can be seen from fig. 6, the HU-SC90 airfoil of example 1 and the SC1095 airfoil of comparative example 1 both have lift coefficients that vary substantially linearly with increasing angle of attack. The lift line slope and the maximum lift coefficient of the HU-SC90 airfoil are both larger than those of the SC1095 airfoil, and the lift characteristic of the HU-SC90 airfoil is more advantageous.

(C) As can be seen from fig. 7, 8, 11 and 12, the maximum lift-drag ratio of the HU-SC90 airfoil of example 1 is higher than that of the SC1095 airfoil of comparative example 1, the usable lift coefficient range of the HU-SC90 airfoil is larger than that of the SC1095 airfoil, which indicates that the lift-drag characteristic of the HU-SC90 airfoil is better than that of the SC1095 airfoil, and the HU-SC90 airfoil has the optimal wide operating condition characteristic, has a higher lift-drag ratio in a larger lift coefficient range, and is comprehensively better than that of the SC1095 airfoil.

(D) As can be seen from fig. 9 and 13, the absolute value of the moment of the HU-SC90 airfoil of example 1 is smaller than that of the SC1095 airfoil of comparative example 1, indicating that the HU-SC90 airfoil has a more excellent moment characteristic.

(E) As can be seen from fig. 10, the stall angle of attack of the HU-SC90 airfoil of example 1 is greater than the stall angle of attack of the SC1095 airfoil of comparative example 1, indicating that the lift characteristics of the HU-SC90 airfoil are also better than the SC1095 airfoil at high angles of attack.

By combining the analysis, the transonic missile airfoil structure suitable for the wide-speed range has the advantages of obvious drag reduction effect, obviously improved lift force characteristic, mild stall characteristic and higher lift-drag ratio, namely excellent aerodynamic characteristic under the condition of wide-speed range flow, and can meet the requirements of a new-generation tactical missile.