As you may know, I have been designing a flying wing. In the absence of any name I like, I will refer to it as the Chevron. At this point, the initial design of the aerodynamic surfaces is wrapping up, and I am comfortable with the wing that has been chosen. The basic layout of the airplane is a swept, low wing, low-taper-ratio flying wing with inboard flaps and outboard elevons, and rudder/winglets. The airplane is to be a single seater, target weight 900 - 1045 lbs, 80 - 95 hp, with stall ~60 kts and cruise ~160 kts IAS. High speed performance is obtained by use of a relatively high (for a flying wing) wing loading.
The wing span is 25 ft, with a wing trapezoidal area of 60 ft^2, giving an aspect ratio just over 10.4. Taper ratio is 0.85 (higher than initially expected), giving a centerline chord of 31.16 inches, and a tip chord of 26.49 inches. Reynolds number is thus in the 1.4E6 to 4.5E6 range. Leading edge sweep is 23.50°, with quarter-chord sweep therefore 23.12°. Geometrical anhedral (negative dihedral) is just over 1°; the line formed by the trailing edges of the airfoils will provide 1° of anhedral, and the spanwise washout of the airfoils provides a small amount more. The fuselage is just under 22" wide (projected, not finalized) at the wing/fuselage junction, and is faired to the wing until the 12" spanwise station.
The airfoils selected are both traditional turbulent-flow low-moment airfoils that achieve high CLmax and low Cm by having forward camber. The airfoils are defined at the 8% (inboard) and 97% (outboard) spanwise positions, and flat-wrap interpolated/extrapolated to other positions. The inboard airfoil uses a NACA 5-digit mean line with maximum camber adjusted to 2.4% (23000 camber line scaled by 129%). The outboard airfoil uses a NACA 4-digit mean line with 1.1% maximum camber at 20% of chord. Both airfoils use the NACA 4-digit thickness distributions; the inboard airfoil is 18% thick, and the outboard airfoil is 12% thick. Mean lines and thickness distributions are combined using the Riblett method. Relative to cruise attitude, the inboard airfoil is rotated nose-up by 2.28°, and the outboard airfoil is rotated nose down by -0.39°. Total geometric twist of the extrapolated wing (0% to 100% spanwise) is 3°.
The wing has two control surfaces, both of 30% chord. Inboard, a flap extends from 8% (12") spanwise position, right at the fuselage, to 53% (79.5") spanwise position. 46.4% of the wing area is thus flapped; this rather low number is chosen to allow effective elevons without excessive sweep, and to minimize the pitching moment associated with flap deflection. Outboard of the flaps, elevons extend from the 53% spanwise position to the 97% (145.5") spanwise position. The flaps will be electrically driven with support for at least 0° (cruise), 20° (take-off), and 40° (landing) positions; the elevons will be mechanically actuated.
Under static conditions of interest, maximum aileron deflection needed reaches 9° antisymmetric; the control system will be built to allow 13° maximum deflection. Necessary elevator deflection ranges from 2° downward deflection to 12° upward deflection; deflections from 6° downward to 16° upward will be possible. The mechanical elevon mixer linearly combines commanded aileron and elevator deflection, giving the elevons a movement range from 15° downward to 29° upward.
Tip winglets act both as Whitcomb winglets for drag reduction, and as all-flying dual rudders. Winglet geometry is similar but not identical to that used for the upper winglet in the Whitcomb design; no surfaces project below the wing. The leading edge of the base of the winglet is aligned with the 40% chord position of the wing tip, but the base of the winglet extends behind the wing tip by an additional 5.13", giving a base chord of 21.02". A winglet taper ratio of 0.6 gives the winglet tip chord of 12.61", with a winglet height of 30.00". Winglet aspect ratio is just under 1.8. To match the wing visually, the winglet leading edge is swept by 23.50°, with an aerodynamic sweep of 20.03°. Winglet area is 3.5 ft^2 each, or 7.0 ft^2 total.
To ease construction, the winglet will use a linear twist distribution and a constant airfoil section. The selected airfoil is the PSU-90-125, at 12.5% thickness. The winglet is twisted outward 1.60° from base to tip. The toe-out of the base of the winglet is nominally 2.08°; however, the all-flying nature of the winglets makes this somewhat moot. To balance aesthetics and adverse yaw concerns, the winglet cant angle is set to 3.00°*outward.
The outer portion of the wing (4.5" outboard of the aileron) is faired/blended into the bottom 3" of the winglet. The portion of the winglet above this fairing is an all-flying control surface. As in other designs with tip rudders, outward-only deflection of a single rudder is used to induce yaw, and symmetric outward-only deflection is used to increase drag for steeper approaches and speed control. Relative to the nominal toe-out angle, the highest statically needed rudder deflection is 24° deflection. A further 18° of outward deflection capability is provided for symmetric application, giving a maximum outward deflection of 42°. In addition, up to 3° of deflection inward beyond the nominal toe-out is provided to allow adjusting winglets to a wider range of cruise conditions.
The winglets will be electrically driven. Control inputs to the winglets will be via a fly-by-wire system, implementing at a minimum a yaw damper and rudder deflection limits as a function of speed to control loads. Flight without a yaw damper is predicted to be possible but unpleasant. The exact scope of the fly-by-wire system has not yet been determined, but it will definitely be restricted to only the yaw axis for this first-generation craft.
The exact CG range and static margin will be pending final analysis, but design assumed a 12% nominal static margin and a total CG range of 3" (7% to 17% static margin range).
While I hope the wing described will need no boundary layer control devices, due to the critical nature of separation on a tailless wing two will be considered based on further analysis (full CFD). The stall is predicted to occur (flaps up or flaps down) at roughly the 30% spanwise position; sprad of the stall much beyond the 50% spanwise position increases the risk of pitch-up at stall. The simplest fix for this problem, should it evidence itself, is a wing fence placed at the 53% spanwise position. Alternately, a full-span array of vortex generators can be considered.
I will attach some sketches/3D views of this wing as soon as I have some that are useful.
The wing span is 25 ft, with a wing trapezoidal area of 60 ft^2, giving an aspect ratio just over 10.4. Taper ratio is 0.85 (higher than initially expected), giving a centerline chord of 31.16 inches, and a tip chord of 26.49 inches. Reynolds number is thus in the 1.4E6 to 4.5E6 range. Leading edge sweep is 23.50°, with quarter-chord sweep therefore 23.12°. Geometrical anhedral (negative dihedral) is just over 1°; the line formed by the trailing edges of the airfoils will provide 1° of anhedral, and the spanwise washout of the airfoils provides a small amount more. The fuselage is just under 22" wide (projected, not finalized) at the wing/fuselage junction, and is faired to the wing until the 12" spanwise station.
The airfoils selected are both traditional turbulent-flow low-moment airfoils that achieve high CLmax and low Cm by having forward camber. The airfoils are defined at the 8% (inboard) and 97% (outboard) spanwise positions, and flat-wrap interpolated/extrapolated to other positions. The inboard airfoil uses a NACA 5-digit mean line with maximum camber adjusted to 2.4% (23000 camber line scaled by 129%). The outboard airfoil uses a NACA 4-digit mean line with 1.1% maximum camber at 20% of chord. Both airfoils use the NACA 4-digit thickness distributions; the inboard airfoil is 18% thick, and the outboard airfoil is 12% thick. Mean lines and thickness distributions are combined using the Riblett method. Relative to cruise attitude, the inboard airfoil is rotated nose-up by 2.28°, and the outboard airfoil is rotated nose down by -0.39°. Total geometric twist of the extrapolated wing (0% to 100% spanwise) is 3°.
The wing has two control surfaces, both of 30% chord. Inboard, a flap extends from 8% (12") spanwise position, right at the fuselage, to 53% (79.5") spanwise position. 46.4% of the wing area is thus flapped; this rather low number is chosen to allow effective elevons without excessive sweep, and to minimize the pitching moment associated with flap deflection. Outboard of the flaps, elevons extend from the 53% spanwise position to the 97% (145.5") spanwise position. The flaps will be electrically driven with support for at least 0° (cruise), 20° (take-off), and 40° (landing) positions; the elevons will be mechanically actuated.
Under static conditions of interest, maximum aileron deflection needed reaches 9° antisymmetric; the control system will be built to allow 13° maximum deflection. Necessary elevator deflection ranges from 2° downward deflection to 12° upward deflection; deflections from 6° downward to 16° upward will be possible. The mechanical elevon mixer linearly combines commanded aileron and elevator deflection, giving the elevons a movement range from 15° downward to 29° upward.
Tip winglets act both as Whitcomb winglets for drag reduction, and as all-flying dual rudders. Winglet geometry is similar but not identical to that used for the upper winglet in the Whitcomb design; no surfaces project below the wing. The leading edge of the base of the winglet is aligned with the 40% chord position of the wing tip, but the base of the winglet extends behind the wing tip by an additional 5.13", giving a base chord of 21.02". A winglet taper ratio of 0.6 gives the winglet tip chord of 12.61", with a winglet height of 30.00". Winglet aspect ratio is just under 1.8. To match the wing visually, the winglet leading edge is swept by 23.50°, with an aerodynamic sweep of 20.03°. Winglet area is 3.5 ft^2 each, or 7.0 ft^2 total.
To ease construction, the winglet will use a linear twist distribution and a constant airfoil section. The selected airfoil is the PSU-90-125, at 12.5% thickness. The winglet is twisted outward 1.60° from base to tip. The toe-out of the base of the winglet is nominally 2.08°; however, the all-flying nature of the winglets makes this somewhat moot. To balance aesthetics and adverse yaw concerns, the winglet cant angle is set to 3.00°*outward.
The outer portion of the wing (4.5" outboard of the aileron) is faired/blended into the bottom 3" of the winglet. The portion of the winglet above this fairing is an all-flying control surface. As in other designs with tip rudders, outward-only deflection of a single rudder is used to induce yaw, and symmetric outward-only deflection is used to increase drag for steeper approaches and speed control. Relative to the nominal toe-out angle, the highest statically needed rudder deflection is 24° deflection. A further 18° of outward deflection capability is provided for symmetric application, giving a maximum outward deflection of 42°. In addition, up to 3° of deflection inward beyond the nominal toe-out is provided to allow adjusting winglets to a wider range of cruise conditions.
The winglets will be electrically driven. Control inputs to the winglets will be via a fly-by-wire system, implementing at a minimum a yaw damper and rudder deflection limits as a function of speed to control loads. Flight without a yaw damper is predicted to be possible but unpleasant. The exact scope of the fly-by-wire system has not yet been determined, but it will definitely be restricted to only the yaw axis for this first-generation craft.
The exact CG range and static margin will be pending final analysis, but design assumed a 12% nominal static margin and a total CG range of 3" (7% to 17% static margin range).
While I hope the wing described will need no boundary layer control devices, due to the critical nature of separation on a tailless wing two will be considered based on further analysis (full CFD). The stall is predicted to occur (flaps up or flaps down) at roughly the 30% spanwise position; sprad of the stall much beyond the 50% spanwise position increases the risk of pitch-up at stall. The simplest fix for this problem, should it evidence itself, is a wing fence placed at the 53% spanwise position. Alternately, a full-span array of vortex generators can be considered.
I will attach some sketches/3D views of this wing as soon as I have some that are useful.
