means the components may bend, and take a permanent set at 6g's but not rupture at less than 9g's. The outer panel fitting will be made of a single 2024T351 aluminum plate which will be attached to the front flat surface of the main spar with bolts loaded in single shear, that is, they join two parts at a single interface surface. We now have all the data necessary to design the fitting.

Designing and Building For Safety by

Lu Sunderland (EAA 5477, Designee No. 60) 5 Griffin Dr.

Apalachin, N.Y. 13732

J- HE OCCURRENCE OF two serious accidents at Oshkosh in 1974 has made it evident that builders and designers of custom built aircraft need to be made aware of certain important design and construction principles. One case involved the substitution of materials for wing fitting attachment and the other an apparent failure to insert a structural pin in a roadable aircraft wing. Perhaps one of the biggest problems encountered by the sellers of aircraft construction plans is the inevitable situation where a builder makes deviations from the specified plans, either in the basic structural design or in the substitution of materials. Most designers are forced, by practical necessity, to establish a policy of approving only the design specified on the plans. To do otherwise would require the expenditure of a considerable amount of engineering effort and the broadening of his liability exposure without financial compensation. Builders must, therefore, accept full responsibility for the design of the entire aircraft when they deviate from the plans in any way because a change in one part can affect the function of another part or the flight characteristics of the aircraft in a way that is not apparent to the layman. For instance, increasing the skin thickness of a control surface requires an attendant increase in the weight of counterbalance weights to maintain the same flutter stability margin. One important structural component which should not be modified without a thorough background in mechanical design is the wing main spar fitting. To show why this part is designed with certain features, let us, as an example, go through some of the steps necessary to design a main spar fitting for the T-18 outer wing panel. FUSELAGE C/L

50"

FIG. 2 LOADING DIAGRAM

The first step in sizing the fitting and its attachment hardware is to determine the forces acting on the fitting. It is necessary to know the wing lift distribution for a rectangular wing in order to make this calculation. NACA Technical Report 572 describes a means to precisely determine the lift distribution. The loading curve resembles an 80% second degree curve. This is too complicated for our simple example, so we shall use a rule-of-thumb used for years by the FAA for a rectangular wing. It is a bit on the conservative side. It says the lift is constant out to 1 chord length from the tip, then it tapers in a straight line to one half this value at the tip. See Figure 2. To calculate wing loading y, we must use some simple algebra breaking the loading diagram into 3 parts as in Figure 2. Lift = loading times span. Total lift required, assuming no lift on the tail is 1250 lb x 9 = 11,250 lbs. Set this equal to the sum of the three areas under the loading diagram.

y

WING TP

lb inch

p- 1 CHORD*

y = 50 T

t

t I

I

y

11,250 = (75.6y + 50 — + 50 — x 1/2) x 2 wings 2 2

^ i ! I

}~

To determine the tension force in the bottom fitting pin, sum the moments about the top pin, again using simple algrebra. A moment is just a force times a distance. The force on a section of the wing acts through the centroid of the loading diagram for that section.

•5y

_JL

FIG. 1 SIMPLIFIED LIFT DISTRIBUTION FOR A RECTANGULAR WING, y = LIFT PER UNIT OF SPAN.

Example Wing Fitting Design A standard T-18 wing is constructed of four 4-foot long panels plus two 10.8" wing tips. The 8° dihedral begins at buttock line (BD 65.625 or 65.625 inches from the fuselage center line. The wing has a constant 50 inch chord with the tip at BL 125.6. A fitting will be designed to join the main spar with two pins at BL 65.625. Sufficient depth is available at the main spar to permit the pins to be located 4.559" apart vertically. First choice of pin vertical spacing is somewhat arbitrary and may need to be changed after the fitting is sized. Assume a gross weight of 1250 pounds and an ultimate load factor of 9g's. To meet FAA aerobatic requirements, an aircraft must withstand a 9g ultimate and a 6g yield load factor. This

F2 = 50 x 10 = 500 lbs. F3

50

2

x 25 = 625 lbs.

F< = 25 x 50 = 1250 lbs 50#/in

— 35"——i -26.6"