Wing construction

main wing assembly jig. This used
every clamp I own

The wing construction started in April of 2009. The spar mold was built up using sections of MDF board. The first project was to lay up a test wing spar panel to verify the design concept. The spar itself was a "C" section shear web with .125" carbon pultrusion "Graphlite" rods laid up as a secondary bond in the corners of the "C". The rods themselves were completely encased or "bedded" with a mixture of cotton flox and epoxy mixed to a paste consistency. This is an area that is a source of additional unneeded weight in the wing design. All future spars will eliminate the use of the round rods. The voids between the rods add considerable weight when multiplied by the 44 feet of total spar length (there is a 2 foot overlap in the root). The first test panel was also designed to static design criteria. Later it was discovered that the test wing exhibited excessive deflection during test and prototype wing was revised to limit deflection by the addition of more Graphite pultrusions.

One of the design problems associated with the use of carbon pultrusions is load fitting termination. I debated a number of design concepts for the wing root attachments, but finally decided on overlapping the root spars. Because of this, the Robin Prototype wing is actually asymmetrical in configuration. The left wing sits 1.5 inches ahead of the right wing. The two spars overlap and sit between a box frame section in the fuselage. Two large shear pins pass through both walls of the box frame and both spars. The advantage here is that the wing bending moment is internally reacted by the overlapping spars and not in excessive structure built into the carry through fuselage box. The disadvantage that I later discovered during construction is having to drill the two holes!! This problem is so great that the overlapping of the wing spars has been eliminated on the production design and an alternative spar attach scheme is being used. The other great disadvantage in the original design is ease of assemble and disassembly in the field. The final design is far easier to field assemble.

aft ribs being assembled

 In the years I have been designing ultralight airplanes I am amazed at how many designers are actually ignorant of the design load conditions of the wings. There is a belief by many amateur designers that the worst case design condition regarding "Drag" is a rear acting load that is trying to tear the wings off the plane in a terminal dive!! Consequently these designs have all featured some kind of "drag" wire that attaches from the fwd fuselage to the wing spars. In reality the aft acting shear load or drag on an aircraft is quite low. In terms of actual numbers consider the condition of equilibrium of an ultralight pointing straight down in the zero lift condition. The total drag acting on the aircraft is exactly equal to the gross weight. In the case of the Robin this would be 550 lbs.

 Now divide that number between all of the components and you will find that the actual load on the wings themselves is in the area of 80 lbs per panel. The actual design criteria that determines the horizontal shear load is the abrupt pitch up maneuver at maximum maneuvering speed. When the aircraft is flying straight and level at the max maneuvering speed, the forces of lift and weight are balanced, as are the thrust and drag. But when the nose attitude is abruptly increased the lift vector shifts forward and increases the load acting in the fwd direction on the wing. 

Leading edge skin being trimmed



This induced vector in the case of the Robin is around 375 lbs acting mid span in the fwd direction. So you can see that the actual horizontal loading is around 8 times greater in the fwd direction as it is in the aft direction. This is why the early aircraft used to fail by the wings folding fwd. To react this load in a weight efficient manner, I took advantage of the deep leading edge "D" cell box. There is a leading edge spruce spar and leading edge fitting designed to react this load. The load is reacted by the fuselage in a couple between the nose spar and the main wing attach pins. Horizontal shear is reacted by bearing the wings spars into the main attach box.


Both Wings D cells laid out


I chose the the configuration of the wing to try and achive laminar flow over the first 33% of the wing. The airfoil is a Wortmann FX 170, low speed, high lift laminar airfoil. This airfoil is designed to be laminar over the first 32% of the upper surface, because of this I placed the wing spars at the 33% chord point. The nose "D" cell also acts as a closed torsion member to react wing twist due to aileron deflection and airfoil pitching moments. The aft section of the wing is covered in light weight Dacron fabric. I intend to bond the fabric with a heavy "pinked" edge right at the 32% chord point. This is designed to emulate sailplane turbulator tape that is usually placed at this point on modern sailplanes. The idea here is to force rapid transition of the laminar flow into turbulent flow that will be contained in a small but attached boundary layer. This is one of the riskiest concepts in the prototypes design. I am banking on a benign stall characteristic because of the low wing loading. Its not called "Experimental Aviation" for nothing!! As I said in my opening mission statement, I want to push the design beyond the low drag of my old Wren.

main aileron bay rib. All bonds between

 foam and wood are with 5 minute epoxy

The remainder of the wing is constructed very similar to my old Wren and that of The Sky Pup, I am using Styrofoam core for the rib shear panels and spruce wood for the rib caps. The big advantage here, besides ease of construction is a wide rib cap. This greatly increases the fabric bond area and eliminates the need for rib stitching. If I were to redesign the Robin into a LSA or a  higher speed homebuilt, these ribs would become wooden trusses and require rib stitching. You can get away with this in an ultralight when your maximum level speed is only 63 mph.

Lower Aileron Cove after forming with ammonia



The current prototype wings weigh 48 lbs per panel; This is higher than my original estimate. Because of this, the wing spar is being redesigned to eliminate the round carbon rods and the fiberglass shear web. I have identified a 9 lb per panel weigh savings in this area alone.




Root of wing showing pre cured fiberglass stiffeners





In the next post I will show pictures of the wing test and the test configuration