N321AD

On September 7, 2023, at 1625 central daylight time, an Alisport SRL Silent 2 Electro glider, N321AD, sustained substantial structural damage while in flight near Giddings, Texas. The pilot was uninjured. The glider was operated as a Title 14 Code of Federal Regulations Part 91 personal flight.

The pilot stated that the glider was in normal straight-ahead cruise flight at approximately 81 kts and 4,800 ft when the cockpit began to shake violently and noisily side-to-side. He said it felt as if the empennage was rocking/twisting from side-to-side. He said there were no unusual conditions and no significant control inputs prior to the shaking, which lasted about 10-15 seconds. The maximum speed of the glider during the occurrence was 90 kts and slowed to about 60 kts toward the end of the shaking. There was an associated altitude loss of about 300 ft. Once the shaking ceased, the glider felt controllable, and the pilot returned to the departure airport without further incident.

The glider sustained substantial damage due to cracks in the fuselage about 30 inches forward of the vertical stabilizer. The cracks were oriented approximately 45° relative to the horizontal plane on both the left and right sides of the fuselage, in addition to a longitudinal crack on the upper fuselage surface.

The glider manufacturer was no longer in business and information received by the investigative staff was limited. The glider had unbalanced horizonal and vertical control surfaces. The ailerons were balanced. A verification of proper flight control cable rigging, elevator trim tab free play, and a static balance of the control surfaces was not performed due to the lack of available instructions for continued airworthiness.

An examination of the aft fuselage section forward of empennage was conducted by the NTSB materials laboratory. The examination revealed no preexisting damage or weakness that could have led to reduced stiffness of the fuselage structure. The

submitted piece was 33 1/2-inch long from the aft end of the fuselage forward of the empennage. Cracks on the upper side of the fuselage were located up to approximately 30 inches forward of the vertical stabilizer.

Skin cracks were observed on the sectioned piece. The skin cracks were oriented approximately 45° relative to the transverse plane on both the left and right side. Additionally, longitudinal cracks were observed on the upper and lower sides of the fuselage. The 45° cracks intersected the longitudinal crack at the lower side of the skin as indicated by the continuous boundary around the cracked areas. On the upper side of the fuselage, the cracks oriented at 45° were separate from the longitudinal crack. The longitudinal crack on the upper side of the fuselage was located between 7 and 11 inches away from the cut end and was difficult to see visually. To aid in detection and photography, a fine black powder was rubbed over the surface to help reveal the cracks in the white paint.

The fuselage skin was made from fiber reinforced polymer composite materials with both glass fiber and carbon fiber fabric reinforcement. The fuselage skin was constructed of two halves (left half and right half) that were spliced together at the upper and lower sides. At each splice location, the skin layers butted together, and strips of glass-fiber reinforced fabric were bonded to the interior surfaces to form single strap butt joints.

A portion of the upper skin containing the longitudinal crack and a section of the upper splice was cut from the remainder of the fuselage skin using a hand-held cutting tool with an oscillating diamond-coated blade. On the interior side, a portion of the strap had a lighter appearance within the region. The change in color was consistent with a disbond between the strap and the skin.

The aft end of the piece shown was sectioned approximately ¾ inch away from the end using a water-cooled abrasive saw. The aft face of the sectioned piece was then rough polished by hand using silicon-carbide paper at successively finer grit size up to 1200 grit to prepare the surface for materialographic examination. Images of the prepared cross-section as viewed using optical microscopy, and images using a scanning electron microscope (SEM).

The middle of the splice between the left and right skin halves. Resin filled the gap between the butt ends of the left and right skin layers and the area between the strap and the skins. The resin layer was approximately 0.10 inch thick on the right side of the splice and became thinner toward the left side. The skin and strap layers were closely bonded to each other at the left side of the splice. A vertical crack was observed at the butt joint as indicated with an unlabeled bracket. The upper end of the vertical crack intersected a smaller transverse crack at the lower side of the strap.

The SEM images of the cross-section were obtained using a backscatter detector. The displayed images are collections of individual SEM images that were stitched together to show overall views of the layup in the single strap splice to the left and right of the butt joint. Glass fibers appear lighter gray relative to the carbon fibers and resin, and voids appear relatively dark. Fiber orientations were determined by relative elongation of fibers ends where they intersected the transverse plane with little or no elongation to 0-degree fibers, intermediate elongation to ±45-degree fibers, and relatively high elongation associated with 90-degree fibers. All fibers were observed in bundles consistent with fabric construction. No evidence of glass fiber roving was observed, including in the resin pocket between the strap and the skin.

Information about the construction of the fuselage for the accident aircraft was limited. Alisport engineering drawing T-SE-X37017, revision 4, dated August 6, 2004, titled Silent Ellisse (Silent 2) fuselage lamination scheme was obtained and reviewed. The drawing indicated each half of the fuselage skin at the damaged location forward of the empennage would consist of 1 layer ±45° glass fiber, 1 layer 0/90° glass fabric, 1 layer ±45° carbon fiber, and 1 layer 0/90° glass fiber starting from the outside surface. The longitudinal joint between the halves was shown as an overlap joint with 6 threads of glass roving in the dished portion of the joint. Details of how the fuselage halves were to be bonded together were not included in the available information.

The accident aircraft was serial number 2077 manufactured in 2016. The as-built construction at the location of the materialographic section as illustrated in the lower image in figure 9 showed skins with one additional ±45° glass fabric layer on the exterior, one additional ±45° carbon fabric layer in the middle, and one additional 0°/90° layer on the inner surface of the skin on the right side only, relative to the reference build drawing. Additionally, the joint between the two halves was a single strap butt joint instead of the overlap splice shown in the reference drawing. The strap consisted of four layers at the centerline and to the right of the butt joint. To the left of the joint, the layers of the strap tapered to one layer ending approximately 4 inches from the butt splice. No evidence of 0° glass roving was observed in the as-built splice. Although the skin layers had a mix of fabric orientations with some layers in the 0/90° orientation and others in the ±45° orientation, all four layers of the strap were oriented in the 0/90° orientation.

The interior surface of the fuselage near the forward cut end, where two areas of the upper splice area are shown circled. These areas had a distinctly lighter appearance relative to the surrounding areas, consistent with areas of voids or disbonds. The longitudinal crack on the upper side of the fuselage was located in this area of the fuselage at a location.