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CN-122009469-A - High pneumatic efficiency fuses formula wingtip winglet

CN122009469ACN 122009469 ACN122009469 ACN 122009469ACN-122009469-A

Abstract

The invention discloses a high pneumatic efficiency fusion type wingtip winglet which comprises a fusion wingroot section, a torsion main section and a wingtip rectifying section which are sequentially connected, wherein the wingtip winglet adopts a spanwise variable thickness modified NACA laminar wing profile, the relative thickness of the wingtip is 12% -15%, the relative thickness of the wingtip is 6% -8%, the wingtip is linearly graded along the spanwise direction, the radius of the front edge is 0.8% -1.2% of the chord length, the trailing edge is in forward pressure gradient transition, the maximum thickness is positioned at the position of 35% -40% of the chord length, the wingtip winglet body is of a multi-wall multi-closed-chamber integrated bearing structure and comprises a front beam, a rear beam, an upper skin and a lower skin, at least three main closed chambers are formed, and the high pneumatic efficiency fusion type wingtip winglet has the characteristics of excellent pneumatic performance, light weight, quick assembly and high universality and is suitable for civil aircraft, transportation aircraft and high-end fixed-wing unmanned aerial vehicles.

Inventors

  • DOU XIAOBO
  • YUAN BO

Assignees

  • 北京栖云通航科技有限公司

Dates

Publication Date
20260512
Application Date
20260327

Claims (10)

  1. 1. The high pneumatic efficiency fusion type wingtip winglet is characterized by comprising a fusion wingroot section, a torsion main section and a wingtip rectifying section which are connected in sequence; The wingtip winglet adopts a spanwise variable thickness modified NACA laminar flow wing profile, the relative thickness of a wingroot (2) is 12% -15%, the relative thickness of a wingtip (1) is 6% -8%, the wingtip winglet is linearly graded along the spanwise direction, the radius of a front edge (3) is 0.8% -1.2% of the chord length, a rear edge (4) is in forward pressure gradient transition, and the maximum thickness is positioned at a position of 35% -40% of the chord length; The wingtip winglet main body is of a multi-wall multi-closed chamber integrated bearing structure and comprises a front beam (5), a rear beam (11), an upper skin (8) and a lower skin (9) to form at least three main closed chambers (13); The fusion wing root section is provided with a fast assembly interface consisting of an upper butt joint rib (6), a lower butt joint rib (7), a positioning pin (10) and a locking bolt (12) and is used for being matched with a positioning hole and a bearing joint of the wing box section to realize fast assembly and disassembly; The whole wingtip winglet is a compound curved surface with a dihedral angle of 15-25 degrees plus an external skimming angle of 8-15 degrees plus a spanwise aerodynamic torsion of 3-6 degrees, the forward-edge sweepback angle is gradually changed from 35 degrees of a wingroot (2) to 25 degrees of a wingtip (1), and the rear edges (4) are smoothly fused to form an asymmetric torsion rectification form.
  2. 2. The high aerodynamic efficiency fusion winglet of claim 1 wherein the spanwise thickening satisfies a linear taper rate constraint, the maximum thickness t (y) of the airfoil at any spanwise location y is determined by the formula: t(y)=t root (1-η·(1-τ)) wherein eta=y/s is the dimensionless relative spanwise coordinate, eta is more than or equal to 0 and less than or equal to 1 (wing root eta=0, wing tip eta=1, S is the small span length, t root is the maximum thickness of the section of the wing root, t tip is the maximum thickness of the section of the wing tip, and τ=t tip / t root is the thickness root-tip ratio, and τ is more than or equal to 0.2 and less than or equal to 0.6; And the relative thickness t (y)/c (y) synchronously linearly changes gradually, so that the following conditions are satisfied: t (y)/c (y)=( t root / c root )・(1-η(1-τ・λ)) Where c (y) is the chord length at the spanwise location y, c root is the root chord length, c tip is the tip chord length, λ=c tip / c root , and is the chord length to root ratio.
  3. 3. The high aerodynamic efficiency fusion wingtip winglet of claim 1 wherein the dihedral angle Γ, the skim angle Λw, and the spanwise aerodynamic twist angle θt (η) satisfy a vortex suppression synergistic relationship, and the synergistic constraint formula under linear twist is: θt (η)=-k・(Г+Λw・cosΛ 1/4 )・(1-η) The boundary conditions are the root (eta=0) twist angle θt (0) =0, the tip (eta=1) twist angle θt (1) = -k (f+Λw seed cos Λ 1/4 ); Where k is the vortex suppression coefficient and Λ 1/4 is the winglet 1/4 chord line sweep.
  4. 4. The high aerodynamic efficiency fusion winglet of claim 3, wherein the vortex suppression coefficient k is: when the winglet aspect ratio ar=8, k=0.812; When AR >8, k=1.2 to 1.5.
  5. 5. A high aerodynamic efficiency fusion winglet according to claim 1, characterized in that the number n of primary closure chambers (13) and the maximum chord width C max satisfy the following constraint: n≤(C max /2·(bw+ts))· Wherein bw is the spar spacing, bw is approximately 0.8 seed troot; ts is the skin thickness; tw is the spar wall thickness; g is the shear modulus of the carbon fiber composite material, and the shear modulus is 25-30 GPa; E is the elastic modulus of the carbon fiber composite material, and the elastic modulus is 140-160 GPa; Delta is the allowable torsion angle; And the number of main closing chambers (13) is selected so as to satisfy the following conditions: when C max <0.3m, n=1; when 0.3m is equal to or less than C max is equal to or less than 0.6m, n=2; When C max >0.6m, n=3, and the rib spacing is 0.2-0.3 m.
  6. 6. The high aerodynamic efficiency fusion winglet of claim 5, wherein the skin thickness ts=0.5 to 1.5mm, and the spar wall thickness tw is equivalent to the skin thickness ts.
  7. 7. A high aerodynamic efficiency fusion winglet according to claim 5, characterized in that the allowable twist angle delta is less than or equal to 0.01rad (i.e. 0.57 °).
  8. 8. The high aerodynamic efficiency fusion wing tip winglet of claim 2, wherein the chord length tip root ratio lambda is in the range of 0.3-0.7.
  9. 9. The high aerodynamic efficiency fusion wingtip winglet according to claim 1, wherein the upper butt joint rib (6) and the lower butt joint rib (7) are in surface contact fit with the bearing joint of the wingbox section, and the fit clearance between the positioning pin (10) and the positioning hole is less than or equal to 0.02mm.
  10. 10. A high aerodynamic efficiency fusion wing tip winglet according to claim 3, characterized in that the 1/4 chord line sweep angle Λ 1/4 = 39 ° -50.

Description

High pneumatic efficiency fuses formula wingtip winglet Technical Field The invention relates to the technical field of aerodynamic design and structural engineering of aircraft wings, in particular to a high aerodynamic efficiency fusion type wingtip winglet. Background The wingtip winglet is an auxiliary aerodynamic component arranged at the wingtip of an aircraft wing, has the core functions of weakening wingtip vortex, reducing induced resistance and improving the lift-drag ratio of the wing, and is a key structure for improving the aerodynamic performance and the fuel efficiency of the aircraft. The traditional wingtip winglet has the technical defects that (1) the wing profile design adopts a fixed thickness or a non-gradual change wing profile, has poor laminar flow characteristics, is easy to cause boundary layer separation and shock wave loss, and has limited induced resistance reduction effect, and the lift-drag ratio is not enough, (2) the structure is single-wall single-closed chamber or spliced design, has insufficient bearing rigidity, needs to increase the material consumption to meet the strength requirement, causes higher structural weight, and has single connection mode of a beam, a rib and a skin, and has low structural efficiency, (3) the mounting interface adopts gasket type bolt connection, has poor positioning precision and poor mounting matching performance, takes time and labor for mounting and is difficult to adapt to the requirements of quick overhaul and modularized replacement, and (4) the structure is single up-reverse or outer skimming structure, lacks asymmetric torsion and composite curved surface design, has single suppression mechanism on wingtip vortex, has poor vortex interference effect, and cannot realize the cooperative optimization of aerodynamic performance and structural efficiency. Disclosure of Invention The invention aims to overcome the existing defects, provides the high pneumatic efficiency fusion type wingtip winglet, solves the problems of high induced resistance, high structural weight, poor installation matching performance and insufficient vortex suppression of the traditional wingtip winglet, has the characteristics of excellent pneumatic performance, high structural efficiency, convenience in installation and adjustment and strong universality, and can effectively solve the problems in the background art. In order to achieve the aim, the invention provides the technical scheme that the high pneumatic efficiency fusion type wingtip winglet comprises a fusion wingroot section, a torsion main section and a wingtip rectifying section which are connected in sequence; The wingtip winglet adopts a spanwise variable thickness modified NACA laminar flow wing profile, the relative thickness of a wingroot is 12% -15%, the relative thickness of a wingtip is 6% -8%, the wingtip winglet is linearly graded along the spanwise direction, the radius of the front edge is 0.8% -1.2% of the chord length, the rear edge is in forward pressure gradient transition, and the maximum thickness is located at a position of 35% -40% of the chord length; the wingtip winglet main body is of a multi-wall multi-closed-chamber integrated bearing structure and comprises a front beam, a rear beam, an upper skin and a lower skin, so that at least three main closed chambers are formed; the fusion wing root section is provided with a fast assembly interface consisting of an upper butt joint rib, a lower butt joint rib, a positioning pin and a locking bolt, and is used for being matched with a positioning hole and a bearing joint of the wing box section to realize fast assembly and disassembly; The whole wingtip winglet is a composite curved surface with a dihedral angle of 15-25 degrees plus an external skimming angle of 8-15 degrees plus a spanwise aerodynamic torsion of 3-6 degrees, the forward-edge sweepback angle is gradually changed from a wingroot of 35 degrees to a wingtip of 25 degrees, and the rear edges are smoothly fused to form an asymmetric torsion rectification form. As a preferred technical solution of the present invention, the spanwise thickening ratio satisfies the linear gradual change rate constraint, and the maximum thickness t (y) of the airfoil at any spanwise position y is determined by the following formula: t(y)=troot(1-η·(1-τ)) wherein eta=y/s is the dimensionless relative spanwise coordinate, eta is more than or equal to 0 and less than or equal to 1 (wing root eta=0, wing tip eta=1, S is the small span length, t root is the maximum thickness of the section of the wing root, t tip is the maximum thickness of the section of the wing tip, and τ=t tip/ troot is the thickness root-tip ratio, and τ is more than or equal to 0.2 and less than or equal to 0.6; And the relative thickness t (y)/c (y) synchronously linearly changes gradually, so that the following conditions are satisfied: t (y)/c (y)=( troot / croot)・(1-η(1-τ・λ)) Where c (y) is the chord length at the spanwi