TEST PANEL DESCRIPTIONS

The two panels were removed from the crown of the airplane as shown in Figure 3.1. Each panel is 125″ in the longitudinal direction, 73″ in the circumferential direction, with a radius of 74″. The panels are structurally similar, except for the location of the circumferential butt joint. The inner and outer surfaces of Panel No. 1 are shown in Figure 3.2. The panels are structurally similar, except for the location of the circumferential butt joint.

Panel No. 2 Panel No. 1

Figure 3.1. Location of panels on fuselage crown

Inner Surface Outer Surface

Frame

Stringer

Lap Joint

Butt Joint

Figure 3.2. Photographs of the inner and outer surfaces of Panel No. 1

The structural details and dimensions of the two panels are shown in Figure 3.3.

Each panel consists of six frames, six stringers (S-2L through S-7L), a longitudinal lap joint along S-4L, and a circumferential butt joint. The location of each frame on the actual fuselage is identified by a fuselage station (FS) number, which indicates the distance in inches from a reference point located near the forward end of the fuselage. A higher fuselage station number indicates a position further away from the reference point, meaning that higher station numbers are further aft. Panel No. 1 spans from FS 620 through FS 720, and has a circumferential butt joint along FS 680. Panel No. 2, is forward of Panel No. 1, and starts from FS 480 through FS 580, and has a butt splice at FS 480.

BS 620 BS 640 BS 660 BS 680 BS 700 BS 720

Figure 3.3. Panel No. 1 drawing, showing substructure

The skin of the panels is made of 2024-T3 aluminum alloy 0.04 inches thick. A longitudinal lap joint is located slightly off the centerline along stringer S-4L. Each panel has six frames and six stringers as shown in Figure 3.3. The spacing between each two adjacent frames is 20″ while the stringers are spaced 9.5″ apart. The stringers have a

hat-shaped cross section and are riveted to the skin with one row of rivets (middle rivet row of the lap joint).

S-2L

S-7L S-4L S-3L

S-5L S-6L

FS 480 500 520 540 560 FS 580

73"

Test Section

Row C Row B Row A

125"

up

aft

Figure 3.4. Panel No. 2 drawing, showing substructure

The frames are not attached directly to the skin, but rather, riveted to the stinger clips, that are in turn, riveted to the skin. This arrangement is known as a “floating frame” configuration. The Z-shaped frames, stringer clips, and hat-shaped stringers are made of 7075-T6 aluminum alloy. The thickness of the frames is 0.050″ while the stringers and stringer clips are 0.032″ thick.

There is a tear strap on the inside of the fuselage skin beneath every frame. The tear straps are simply aluminum 2024-T3 strips attached circumferentially to the skin of the fuselage. The purpose of a tear strap is to arrest longitudinal cracks and to contain the damage between the two tear straps. These tear straps are fabricated from 0.040″ thick 2024-T3 aluminum alloy and are bonded to the fuselage skin. A solid model of the panel, showing the tear straps, which are 2.4″ wide, is shown in Figure 3.5. The figure also shows the other inner structural details of the panel. The lower skin and upper skin are so named as discussed in the next paragraph.

Frame

Stringer Clip

Stringer

Tear Strap Lower Skin

Upper Skin

Figure 3.5. Solid model showing inner surface of panel

The longitudinal lap joint consists of the two 0.040 thick 2024-T3 aluminum skin layers joined by a 0.020 inch thick 2024-T3 aluminum doubler layer sandwiched between the two skin layers, which are connected together by three rows of rivets labeled A, B, and C, Figure 3.6. The three rows of 5/32″ diameter Al 2017-T4 rivets are spaced about 0.90 inches apart. The doubler is cold bonded to the outer skin while there is an elastomeric sealant between the doubler and the inner skin of the lap joint. The middle rivet row (row B) is common to both the joint and the underlying hat-shaped type stringer, while rivet rows A and C join the two skins to the double as shown in Figure 3.6.

The lap joint consists of fifteen equally spaced rivets between neighboring frames in each row for a total of forty five rivets in the test section. The nominal spacing between the rivets, in each rivet row is 1.06″. From the results of the teardown inspection performed on the lap joint on the right-hand side of the same airplane at S-4R [3.1], it was anticipated in these tests that fatigue cracks would develop in the inner skin of the lap joint at the lower rivet row A. A detailed discussion of this particular issue is given in Chapter 2, section 2.2.5. The skin layer that continues upwards from the lap joint is known as the outer skin, while the skin layer that continues downwards from the lap joint is called the lower skin, because of their locations in the airplane relative to the lap joint, Figure 3.6. At the lap joint, the lower skin is called the “inner” skin since it is the inside layer of the joint stack-up. Similarly, the upper skin is called the “outer” skin, at the lap joint, as shown in Figure 3.6.

Lower Skin Inner Skin

Row C Upper Skin

Row B

Row A Outer Skin

Stringer Doubler

Figure 3.6. Schematic of the longitudinal lap joint

The panels were modified and prepared for testing in the Full-Scale Aircraft Structural Test Evaluation and Research loading fixture (see next section). The four edges of the curved panels, along which the hoop and longitudinal loads are applied, were reinforced by bonding 0.050″ thick aluminum alloy doublers to the skin. This was done in order to ensure a uniformly distributed load transfer. The reinforcing doublers along the longitudinal sides of the panel were 112″ long while those along the hoop sides were 56″ long. Holes, 0.5″ in diameter were spaced approximately 4″ apart along the longitudinal doublers and 3.5″ apart along the hoop doublers to attach the whiffle tree assemblies, which apply the load. Altogether, there were 28 load application points along each side of the panel and 16 load application points along each end. In order to reinforce the load application points at the frames, doublers were added to the frame

ends, where they attach to the frame loaders. The test section of the panels, the area in which the effects of the fixture attachment points could be neglected, is approximately one frame bay and one stringer bay from the reinforcement area, Figure 3.3 and Figure 3.4.

Each panel was instrumented with a total of over 100 uniaxial, biaxial, and rosette strain gages to monitor and record strain distribution during the test. The Strain gages were installed on the skin, frames, and stringers. Several back-to-back strain gage sets were installed at various locations on the skin to measure skin bending. The locations of the strain gages are shown in Figure 3.7 and Figure 3.8 respectively for Panel No. 1 and Panel No. 2. The procedures for strain gage installation, the types of strain gages used, and the exact locations of the strain gages in the two panels are given in Appendix A and reference [3.3]. The strain gages were connected to strain gages bridges, and shunt calibrated before the tests. The calibration of the gages was performed at once before each test, and remained stable throughout the test.