CN-122009470-A - Pneumatic appearance structure of supercritical wing

CN122009470ACN 122009470 ACN122009470 ACN 122009470ACN-122009470-A

Abstract

The application belongs to the technical field of aviation aircraft design, and particularly relates to a supercritical wing aerodynamic profile structure, wherein the wing is generated by controlling the supercritical wing profile of five specific stations along the span direction, wherein 45% of the stations are reference wing profiles. The front edge radius, thickness and installation angle of the inner side (0% and 20% of the station positions) wing profile are sequentially larger than those of the reference wing profile, the front edge radius of the outer side (70% and 100% of the station positions) wing profile is obviously larger than that of the reference wing profile, and the installation angle is negative and the absolute value is sequentially increased. The front edge and the rear edge of each section are respectively connected by secondary spline lines to form the front edge and the rear edge of the wing, so that a three-dimensional molded surface is generated. According to the application, through the differential design of the spanwise airfoil parameters, the large-area separation area under a large attack angle can be actively controlled to appear in the middle of the airfoil, so that the manipulation failure caused by stall of the outer airfoil and the lift force dip caused by stall of the inner airfoil are effectively avoided, and the stall safety of the upper single-wing layout conveyor is remarkably improved while the transonic cruising efficiency is maintained.

Inventors

LIAO ZHENRONG
WEI JIANLONG
GUO ZHAODIAN
HOU YINZHU
FENG HAIYONG

Assignees

中国航空工业集团公司西安飞机设计研究所

Dates

Publication Date: 20260512
Application Date: 20260331

Claims (12)

1. The aerodynamic profile structure of the supercritical wing is characterized by comprising an integral wing which is controlled and generated by supercritical wing profile sections at five specific spanwise stations; Wherein the five airfoil profile sections have the station positions of 0%, 20%, 45%, 70% and 100% half-span length along the machine span direction respectively, and a third airfoil profile section (3) positioned at the 45% half-span length is taken as a reference airfoil profile; The first airfoil section (1) and the second airfoil section (2) at 0% half-extension and 20% half-extension are all incremental in terms of leading edge radius, relative thickness and local mounting angle with the airfoil leading edge point as a base point relative to the reference airfoil (3); the front edge radius of the fourth airfoil profile (4) and the fifth airfoil profile (5) which are positioned at 70% half-extension and 100% half-extension is larger than that of the first airfoil profile (1) and the second airfoil profile (2) relative to the reference airfoil profile (3), the relative thickness is smaller than or equal to that of the reference airfoil profile (3), the local installation angle taking the airfoil front edge point as a base point is a negative value and the absolute value of the local installation angle is increased; The leading edge points of the five airfoil sections are connected through one secondary spline line to form an airfoil leading edge line, and the trailing edge points of the five airfoil sections are connected through the other secondary spline line to form an airfoil trailing edge line; the three-dimensional profile of the airfoil is generated from the five airfoil sections, the leading edge line and the trailing edge line as control lines.
2. The aerodynamic profile structure of a supercritical wing according to claim 1, characterized in that the airfoil parameters of the reference airfoil (3) are a leading edge radius of 0.0122, a relative thickness of 11.5%, a local mounting angle of 0.1 °, a maximum thickness position at 40.4% chord, a relative camber of 1.45% and a maximum camber position at 79.4% chord.
3. The aerodynamic profile of a supercritical wing according to claim 2, characterized in that the airfoil parameters of the first airfoil section (1) are a leading edge radius of 0.0132, a relative thickness of 13%, a local mounting angle of 3 °, a maximum thickness position at 35.2% chord length, a relative camber of 1.24%, and a maximum camber position at 77.6% chord length.
4. The aerodynamic profile structure of a supercritical wing according to claim 2, characterized in that the second airfoil section (2) at 20% half-span has airfoil parameters of leading edge radius 0.0136, relative thickness 12%, local mounting angle 1.3 °, maximum thickness position at 40.3% chord, relative camber 1.44%, maximum camber position at 80.3% chord.
5. The aerodynamic profile structure of a supercritical wing according to claim 2, characterized in that the fourth airfoil section (4) at 70% half-span has airfoil parameters of leading edge radius 0.0280, relative thickness 11.5%, local mounting angle-2.2 °, maximum thickness position at 36.5% chord length, relative camber 1.84%, maximum camber position at 76.5% chord length.
6. A supercritical wing aerodynamic profile structure according to claim 2, characterized by a fifth airfoil section (5) at 100% half-span as a wing tip airfoil with airfoil parameters of leading edge radius 0.0280, relative thickness 11.5%, local mounting angle-3.8 °, maximum thickness position at 36.5% chord, relative camber 1.84% and maximum camber position at 76.5% chord.
7. The supercritical wing aerodynamic profile structure according to claim 1, wherein the leading edge point spatial coordinates of the five airfoil sections are determined by the wing overall parameters such that the resulting wing has a sweep angle of 25 °, an aspect ratio of 8 and a root ratio of 4.
8. The supercritical wing aerodynamic profile structure according to claim 7, wherein the five airfoil sections have leading edge point coordinates ：（0, 0, 0）,（0.1083R, 0.2R, -0.0105R）,（0.2436R, 0.45R, -0.0236R）,（0.3789R, 0.7R, -0.0367R）,（0.5413R, 1.0R, -0.0524R）,, respectively, where R is the wing half-span.
9. The supercritical wing aerodynamic profile structure according to claim 7 or 8, wherein the overall torsion angle of the wing is 6.8 ° and no dihedral.
10. The aerodynamic profile structure of a supercritical wing according to claim 1, characterized in that the spanwise standing of the wing root airfoil (1) is flush with the plane of symmetry of the aircraft, the upper airfoil thereof protrudes from the fuselage, and the lower airfoil is in fusion connection with the fuselage.
11. The supercritical wing aerodynamic profile structure according to claim 1, characterized in that by the wing profile parameter matching according to any of claims 1 to 6 and the geometrical layout according to any of claims 7 to 9, the wing is in a large angle of attack state with its large area airflow separation area controlled to first appear in the midsection of the spanwise direction.
12. The supercritical wing aerodynamic profile structure according to claim 1, wherein the wing is applied to a transport vehicle of an upper mono-wing configuration.

Description

Pneumatic appearance structure of supercritical wing Technical Field The application belongs to the technical field of aviation aircraft design, and particularly relates to a supercritical wing aerodynamic profile structure. Background In order to solve the problem, the traditional method is to control the large-area separation to appear in the inner wing, but for the upper single-wing layout aircraft, when the large-area separation appears in the inner wing, the aircraft lift force is greatly lost, the aircraft rapidly falls to the height, and the danger is easy to generate. Disclosure of Invention In order to solve the problems, the application provides a supercritical wing aerodynamic profile structure, which comprises an integral wing, a plurality of wing units and a plurality of wing units, wherein the integral wing is generated by controlling the supercritical wing profile at five specific spanwise stations; Wherein the five airfoil profile sections have the station positions of 0%, 20%, 45%, 70% and 100% half-span length along the machine span direction respectively, and a third airfoil profile section (3) positioned at the 45% half-span length is taken as a reference airfoil profile; The first airfoil section (1) and the second airfoil section (2) at 0% half-extension and 20% half-extension are all incremental in terms of leading edge radius, relative thickness and local mounting angle with the airfoil leading edge point as a base point relative to the reference airfoil (3); the front edge radius of the fourth airfoil profile (4) and the fifth airfoil profile (5) which are positioned at 70% half-extension and 100% half-extension is larger than that of the first airfoil profile (1) and the second airfoil profile (2) relative to the reference airfoil profile (3), the relative thickness is smaller than or equal to that of the reference airfoil profile (3), the local installation angle taking the airfoil front edge point as a base point is a negative value and the absolute value of the local installation angle is increased; The leading edge points of the five airfoil sections are connected through one secondary spline line to form an airfoil leading edge line, and the trailing edge points of the five airfoil sections are connected through the other secondary spline line to form an airfoil trailing edge line; the three-dimensional profile of the airfoil is generated from the five airfoil sections, the leading edge line and the trailing edge line as control lines. Preferably, the airfoil parameters of the reference airfoil (3) are that the front edge radius is 0.0122, the relative thickness is 11.5%, the local installation angle is 0.1 degrees, the maximum thickness position is at 40.4% chord length, the relative bending degree is 1.45%, and the maximum bending degree position is at 79.4% chord length. Preferably, the airfoil parameters of the first airfoil section (1) are a leading edge radius of 0.0132, a relative thickness of 13%, a local mounting angle of 3 °, a maximum thickness position at 35.2% chord, a relative camber of 1.24%, and a maximum camber position at 77.6% chord. Preferably, the second airfoil section (2) at 20% half-span has airfoil parameters of a leading edge radius of 0.0136, a relative thickness of 12%, a local mounting angle of 1.3 °, a maximum thickness position at 40.3% chord length, a relative camber of 1.44%, and a maximum camber position at 80.3% chord length. Preferably, the fourth airfoil section (4) at 70% half-span has airfoil parameters of leading edge radius 0.0280, relative thickness 11.5%, local mounting angle-2.2 °, maximum thickness position at 36.5% chord, relative camber 1.84%, maximum camber position at 76.5% chord. Preferably, the fifth airfoil section (5) at 100% half-span is used as a tip airfoil with airfoil parameters of a leading edge radius of 0.0280, a relative thickness of 11.5%, a local mounting angle of-3.8 °, a maximum thickness position at 36.5% chord, a relative camber of 1.84%, and a maximum camber position at 76.5% chord. Preferably, the leading edge point spatial coordinates of the five airfoil sections are determined by the overall parameters of the wing such that the resulting wing has a sweep angle of 25 °, an aspect ratio of 8, and a root tip ratio of 4. Preferably, the coordinates of the leading edge points of the five airfoil sections are ：（0, 0, 0）,（0.1083R, 0.2R, -0.0105R）,（0.2436R, 0.45R, -0.0236R）,（0.3789R, 0.7R, -0.0367R）,（0.5413R, 1.0R, -0.0524R）, respectively, wherein R is the half-span length of the airfoil. Preferably, the overall torsion angle of the wing is 6.8 ° and no dihedral. Preferably, the span-wise standing position of the wing root airfoil (1) is flush with the plane of symmetry of the aircraft, the upper airfoil surface of the wing root airfoil protrudes out of the fuselage, and the lower airfoil surface of the wing root airfoil is in fusion connection with the fuselage. Preferably, the wing is controlled to appear first