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CN-122009471-A - Pneumatic design method suitable for controlling stall of central part of supercritical machine in wingspan direction

CN122009471ACN 122009471 ACN122009471 ACN 122009471ACN-122009471-A

Abstract

The application belongs to the technical field of aviation aircraft design, and discloses a pneumatic design method suitable for controlling stall in the middle of a supercritical aircraft span. The method comprises the steps of configuring five specific supercritical airfoil profiles along the wing spanwise direction, wherein the spanwise stations are respectively 0%, 20%, 45%, 70% and 100% half-spanwise lengths. Wherein, 45% half-span length is the reference airfoil. The control of the wing airflow separation position under a large attack angle is realized by enabling the inner (0%, 20% standing position) wing profile to increase the front edge radius, thickness and installation angle relative to the reference wing profile and enabling the outer (70%, 100% standing position) wing profile to greatly increase the front edge radius and reduce the installation angle relative to the reference wing profile. The method can control the large-area separation area to occur in the middle of the wing in the expanding direction at a large attack angle, so that the risk of airplane roll and manipulation failure caused by stall of the outer wing is avoided, and meanwhile, the sudden drop of lift caused by stall of the inner wing is also prevented. The application can complete the design by only using five airfoil profiles, and obviously improves the stall safety of the upper single-wing layout conveyor while guaranteeing the transonic cruising efficiency.

Inventors

  • LIAO ZHENRONG
  • WEI JIANLONG
  • FENG HAIYONG
  • HOU YINZHU
  • LI MIN

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. 1. A pneumatic design method suitable for supercritical machine span to middle stall control, comprising the following steps: Step one, selecting five airfoil profile sections along the wing span direction, wherein the span-direction station positions of the five airfoil profile sections are respectively 0%, 20%, 45%, 70% and 100% half-span positions, the design profile sections positioned at the 45% half-span positions are reference airfoil sections, and the other four design profile sections are design airfoil sections; step two, carrying out differential configuration on airfoil geometric parameters of the five airfoil sections: the two inner side design airfoils at the half-spread points of 20% and 0% are larger than the corresponding parameters of the reference airfoil, and the radius of the leading edge, the relative thickness and the local installation angle taking the leading edge point of the airfoil as the base point; the two outer design airfoils are positioned at 70% and 100% half-extension points, the front edge radius of the two outer design airfoils is larger than that of the reference airfoil, the relative thickness of the two outer design airfoils is smaller than or equal to that of the reference airfoil, and the local installation angle taking the front edge point of the airfoil as a base point is smaller than that of the reference airfoil; determining the chord length and the space coordinates of the front edge points of each profile airfoil according to the overall design parameters of the airfoil, and enabling the sweepback angle of the front edge of the airfoil determined by the five front edge points to be between 20 and 30 degrees, and enabling the ratio of the root tip of the airfoil to be between 3 and 4; fourthly, connecting the front edge points of the five airfoil sections through spline curves to form airfoil front edge lines, and connecting the rear edge points of the airfoil sections through spline curves to form airfoil rear edge lines; And fifthly, taking the five airfoil profile sections and the leading edge line and the trailing edge line as geometric inputs to generate a continuous wing three-dimensional curved surface.
  2. 2. A aerodynamic design method for spanwise mid-stall control of a supercritical machine according to claim 1, wherein in step two the leading edge radius of the design airfoil at 20% half-span is increased by 8% to 15% over said reference airfoil.
  3. 3. A aerodynamic design method for spanwise mid-stall control of a supercritical machine according to claim 1, wherein in step two the leading edge radius of the design airfoil at 0% half-span is increased by 5% to 10% over said reference airfoil.
  4. 4. A aerodynamic design method for a supercritical spanwise mid-stall control according to claim 1, wherein in step two the leading edge radius of the design airfoil at 70% and 100% half-span is increased by 100% to 150% over the reference airfoil.
  5. 5. The aerodynamic design method for a spanwise mid-stall control of a supercritical machine according to claim 1, wherein in step two, the airfoil is designed at 20% half-span and rotated forward by 0.5 ° to 2 ° with the airfoil leading edge point as a base point to increase its local mounting angle.
  6. 6. The aerodynamic design method for a supercritical machine spanwise mid-stall control according to claim 1 or 5, wherein in step two, the airfoil is designed at 0% half-span and rotated forward by 2 ° to 4 ° with the airfoil leading edge point as a base point to increase its local mounting angle.
  7. 7. A aerodynamic design method for spanwise mid-stall control of a supercritical machine according to claim 1, wherein in step two, the airfoil is designed at 70% half-span and rotated negative-1 ° to-3 ° about the airfoil leading edge point as a base point to reduce its local mounting angle.
  8. 8. The aerodynamic design method for a spanwise mid-stall control of a supercritical machine according to claim 1 or 7, wherein in step two, the airfoil is designed at 100% half-span and rotated negative-3 ° to-5 ° about the airfoil leading edge point as a base point to reduce its local mounting angle.
  9. 9. A aerodynamic design method for spanwise mid-stall control of a supercritical machine according to claim 1, wherein in step two the relative thickness of the design airfoil at 20% half-span is increased by 0.3% to 0.8% over said reference airfoil.
  10. 10. The aerodynamic design method for a spanwise mid-stall control of a supercritical machine according to claim 1, wherein in step two, the relative thickness of the design airfoil at 0% half-span is increased by 1% to 2% over the reference airfoil and its maximum thickness position is advanced by 3% to 7% over the reference airfoil.
  11. 11. A aerodynamic design method for spanwise mid-stall control of a supercritical machine according to claim 1, wherein in step two the relative camber of the design airfoil at 70% and 100% half-span is increased by 0.2% to 0.6% over the reference airfoil and both its maximum thickness position and maximum camber position are advanced by 2% to 5% over the reference airfoil.
  12. 12. A aerodynamic design method for a supercritical spanwise mid-stall control according to claim 1, wherein in step four, the spline is a quadratic spline.

Description

Pneumatic design method suitable for controlling stall of central part of supercritical machine in wingspan direction Technical Field The application belongs to the technical field of aviation aircraft design, and particularly relates to a pneumatic design method suitable for controlling stall in the middle of a supercritical aircraft span. Background The modern transport plane generally adopts a supercritical wing with a sweepback angle to adapt to transonic flight, so that shock wave resistance is reduced conveniently, however, when the attack angle is larger, large-area separation occurs on the wing, the outer wing boundary layer is piled up, the large-area separation occurs at first, the outer wing stalls, the airplane is enabled to have a roll phenomenon, the aileron of the wing is invalid, the airplane loses safety, in order to solve the problem, the traditional method is to control the large-area separation to occur on the inner wing, but for the upper single-wing layout airplane, when the large-area separation occurs on the inner wing of the wing, great loss occurs on the lift force of the airplane, the airplane drops sharply, and dangers are easy to occur. Disclosure of Invention In order to solve the problems, the application provides an aerodynamic design method suitable for controlling the stall of a central part of a supercritical machine span, which comprises the following steps: Step one, selecting five airfoil profile sections along the wing span direction, wherein the span-direction station positions of the five airfoil profile sections are respectively 0%, 20%, 45%, 70% and 100% half-span positions, the design profile sections positioned at the 45% half-span positions are reference airfoil sections, and the other four design profile sections are design airfoil sections; step two, carrying out differential configuration on airfoil geometric parameters of the five airfoil sections: the two inner side design airfoils at the half-spread points of 20% and 0% are larger than the corresponding parameters of the reference airfoil, and the radius of the leading edge, the relative thickness and the local installation angle taking the leading edge point of the airfoil as the base point; the two outer design airfoils are positioned at 70% and 100% half-extension points, the front edge radius of the two outer design airfoils is larger than that of the reference airfoil, the relative thickness of the two outer design airfoils is smaller than or equal to that of the reference airfoil, and the local installation angle taking the front edge point of the airfoil as a base point is smaller than that of the reference airfoil; determining the chord length and the space coordinates of the front edge points of each profile airfoil according to the overall design parameters of the airfoil, and enabling the sweepback angle of the front edge of the airfoil determined by the five front edge points to be between 20 and 30 degrees, and enabling the ratio of the root tip of the airfoil to be between 3 and 4; fourthly, connecting the front edge points of the five airfoil sections through spline curves to form airfoil front edge lines, and connecting the rear edge points of the airfoil sections through spline curves to form airfoil rear edge lines; And fifthly, taking the five airfoil profile sections and the leading edge line and the trailing edge line as geometric inputs to generate a continuous wing three-dimensional curved surface. Preferably, in step two, the leading edge radius of the design airfoil at 20% half-span is increased by 8% to 15% over the reference airfoil. Preferably, in step two, the leading edge radius of the design airfoil at 0% half-span is increased by 5% to 10% over the reference airfoil. Preferably, in step two, the leading edge radius of the design airfoil at 70% and 100% half-extension is increased by 100% to 150% over the reference airfoil. Preferably, in the second step, the airfoil is designed at a half-spread position of 20%, and the airfoil is rotated forward by 0.5 to 2 ° with the leading edge point of the airfoil as a base point, so as to increase the local installation angle thereof. Preferably, in the second step, the airfoil is designed at a half-extension position of 0%, and the airfoil is rotated forward by 2 ° to 4 ° with the airfoil leading edge point as a base point, so as to increase the local installation angle thereof. Preferably, in the second step, the airfoil is designed at the half-extension position of 70%, and the airfoil is rotated by-1 DEG to-3 DEG in the negative direction by taking the airfoil front edge point as a base point so as to reduce the local installation angle. Preferably, in the second step, the airfoil is designed at a half-extension of 100%, and the airfoil is rotated in the negative direction by-3 DEG to-5 DEG with the leading edge point of the airfoil as the base point, so as to reduce the local installation angle thereof. Preferably, in step two, the