On November 4, 2006, about 1139 Pacific standard time, an Aircraft Manufacturing & Development Co. (AMD), CH601XL SLSA, N158MD, experienced an in-flight breakup while cruising approximately 11 miles south-southeast of Yuba City, California. The special light sport airplane (S-LSA), referred to as a "Zodiac," was destroyed. The private pilot and passenger were killed. The pilot operated the airplane using the Federal Aviation Administration (FAA) registered name "Zodiac LSA, Inc." Visual meteorological conditions prevailed, and no flight plan was filed. The personal flight was performed under the provisions of 14 Code of Federal Regulations Part 91 and originated from Lincoln, California, about 1129.

The pilot's son reported to the National Transportation Safety Board investigator that his father and he jointly operated the airplane. The son was listed as the president of Zodiac LSA, Inc., which was the name that the pilot and his son gave to their company.

The family purchased the newly manufactured airplane directly from AMD in July 2006. Following their purchase and the son's introductory flight provided by AMD personnel, the son flew the airplane from AMD's factory in Georgia to his California home. The pilot's son also stated to the Safety Board investigator that in California he and his father flew the airplane together on several occasions. They became familiar with the airplane's operating characteristics, which according to the son were somewhat different than those of the other airplanes their family had owned or operated.

The accident flight commenced after the pilot refueled the airplane at the Lincoln Regional Airport. Upon departure, the pilot headed in a westerly direction toward his intended destination, the Willows-Glenn County Airport, Willows, California. Willows is located about 63 miles west-northwest of Lincoln. The pilot's son reported that his parents planned to fly to Willows for lunch.

While cruising toward Willows, about 10 minutes after takeoff and 16 miles west of Lincoln, five ground-based witnesses heard the airplane. Two of the witnesses also observed the airplane prior to the breakup.

In summary, three of the auditory witnesses reported hearing a "bang" sound. One of these witnesses reported that for about 10 seconds prior to hearing the "bang," the engine was misfiring or sputtering. The sound increased in loudness and ended with a loud "bang."

The fourth witness reported to a Sutter County Sheriff's deputy that he had been working in his field when he heard an airplane overhead. The witness opined that the airplane's engine was missing really badly. A few seconds later, as he was looking at it flying an estimated 800 to 1,200 feet above ground level, the airplane "blew up." The wings flew off, parts went everywhere, and the cockpit turned in circles as it descended.

The fifth witness reported to the Safety Board investigator that he was a retired United States Air Force mechanic and was familiar with light airplanes. The witness was standing outside his residence, about 0.5-mile southeast of the accident site. In summary, the witness reported that he heard the sound of the airplane's engine, and it sounded fine. It was operating smoothly, its rpm sounded steady in that it was producing a constant tone, and it was not backfiring or sputtering. Then, the witness looked upward in the direction of the engine sound and immediately observed the airplane. The airplane was northwest of his location, and it was cruising in a northwesterly direction. Its wings were level. The airplane was not turning, climbing, or descending. The witnesses further stated that he could clearly see the airplane and saw no evidence of fire or smoke trailing from it.

The fifth witness additionally reported that, after a few seconds, he stopped looking at the airplane and started talking on his cell phone. The witness estimated that he looked away from the airplane for about 5 seconds. Then, he heard the sound of the airplane's engine rapidly change rpm. Within about 0.5 seconds, it decreased and then increased, as if the pilot had retarded the throttle and then suddenly changed his mind and restored the power. When the rpm came back up, it did not sound like the engine had over revved. The tone sounded the same as before the power had decreased. The witness reported that immediately after the engine power came back up, he heard the sound of an explosion, which was followed by the sound of metal scraping.

Thereafter, he saw what he believed were three distinctive large components falling. The components were the wings and the fuselage. The witnesses stated that it took perhaps 6 to 8 seconds for the airplane to fall, and it fell straight down. As the components descended, the fuselage spiraled around. There was no fire or smoke.

The witness opined that, originally, he thought there had been a mid-air collision. However, there was no other aircraft in the area.



The pilot, age 79, held a private pilot certificate with an airplane single-engine land rating. The pilot's last aviation medical certificate was issued in the third class on February 16, 2005. The certificate bore the following limitation: "Must wear glasses for near and distant vision." Also, the certificate stated "Not valid for any class after February 28, 2006." (The FAA does not require that pilots operating a S-LSA under the accident flight conditions hold any class of medical certificate.)

On the February 2005 medical certificate application form the pilot reported that his total pilot time was 1,184 hours. In previous years, when the pilot was issued aviation medical certificates, he reported the following total pilot times: 1,025 hours in February 2003, and 783 hours in February 2001.

The pilot's personal flight record logbook indicates that he commenced flying the accident airplane on August 18, 2006. His logbook was endorsed by a certified flight instructor as having satisfactorily completed a flight review and "Zodiac checkout" in the accident airplane on October 28, 2006. According to the logbook, by the accident date the pilot's total logged flight time in all airplanes was about 1,281 hours.

The Safety Board investigator's review of the airplane's utilization flight logbook found in the wreckage revealed that between August 18, 2006, and November 4, 2006 (accident date), the pilot was listed as having flown the airplane on nine dates for a total of 20.2 hours. More recently, between October 4, 2006, and November 4, 2006, the logbook indicated that the pilot flew the airplane on six dates for a total of 15.3 hours.


Family members reported that the accident occurred during the wife's first flight in the airplane. The wife was not a pilot and had not received flight instruction. She had previously flown in other family-owned airplanes. The wife was reportedly in good physical condition, swam frequently, and was not handicapped.



The accident airplane was purchased as a factory-assembled S-LSA from AMD. The manufacture of the accident airplane began in Ontario, Canada. The accident airplane's components were then transported to Eastman, Georgia, where the airframe was assembled. The assembly included attaching the wings, stabilizers, flight controls, avionics, and engine.


AMD reported that the airplane, serial number 601-016S, was designed and manufactured in accordance with the American Society for Testing and Materials (ASTM) industry consensus standard for light sport aircraft. AMD's Director of Quality Assurance issued the following consensus standard certification statement to the FAA:

"I hereby certify that aircraft serial number 601-016S complies with the Consensus Standard(s) identified on this statement of compliance and that the Manufacturer's Continued Airworthiness System will be adhered to support the aircraft throughout its life. This aircraft (1) was manufactured following the consensus standard(s) procedures and Manufacturer's Quality Assurance System identified on this statement, (2) conforms to the manufacturer's design data, (3) was ground and flight tested successfully, and (4) is in condition for safe operation."

AMD contracted with a Designated Airworthiness representative (DAR) to issue the airplane its airworthiness certificate, and the certificate was issued on July 10, 2006. The DAR indicated to the Safety Board investigator that he followed industry procedures that included, in part, looking at the airplane's exterior and cockpit, and examining applicable paperwork presented by AMD. Finding nothing outstanding with the airplane and no safety issues, the DAR accepted AMD's statement of compliance with the ASTM standards. He issued the airplane a special airworthiness certificate on behalf of the FAA. The DAR was neither required to nor did he fly the airplane during the certification process.

A few days later, on July 18, 2006, the consensus standard certification statement was issued. According to the airplane log, the airplane was flown by a test pilot on July 19, 2006, and he certified "that its performance flying qualities, performance of controls, powerplant and landing gear, etc., were equivalent to the standard of the type."

Documents found in the wreckage indicate an FAA special airworthiness certificate was issued in the Light Sport Category on July 20, 2006.

Registration and Usage

On July 20, 2006, an airplane registration certificate was issued indicating the owner applicant was Zodiac LSA, Inc.

The pilot's son reported that in Eastman, Georgia, he flew the airplane with an AMD employee. Anomalies were noted with the airplane and, according to AMD, they were corrected prior to the pilot's departure from its facilities.

The pilot's son reported that the airplane was operated by his family for its personal use. The airplane was not rented to other pilots. By the time of the accident, the airplane's total flight time was about 98 hours.

Design Strength

In Section 3 of the "Pilot Operating Handbook" for the accident model of Zodiac airplane, AMD reports that the airplane's load factor (limit) is +4 positive G, and -2 negative G. A note states that the ultimate load factor is 1.5 times the limit.

In AMD's internet advertising, the company provides specifications for the Zodiac airplane. In part, AMD states the aforementioned load factors by indicating that at a gross weight of 1,320 pounds, the design load (ultimate) is +6 G and -3 G.

Weight and Balance

The pilot's family reported that the probable weight of the pilot and passenger was 375 pounds. The Safety Board investigator estimated 7 pounds of baggage was on board. AMD estimated that, based upon 10 pounds of fuel burning off since takeoff, the fuel weight was 170 pounds. The airplane's empty weight was 836 pounds. In total, the airplane weighed about 1,388 pounds at the time of the accident.

This is about 68 pounds over the maximum authorized gross weight of 1,320 pounds. The airplane's calculated center of gravity was 15.264 inches aft of datum, near the center of the balance envelope.

Maintenance and Operation

AMD provided the airplane's owner with documents specifying the operation, maintenance, and inspection procedures that were to be followed.

The pilot's son holds a private pilot certificate with an airplane single-engine land rating. The son enrolled in a training program to obtain FAA certification as a repairman, light sport airplane. He received certification following the accident.

The pilot's son reported to the Safety Board investigator that he was aware some of the maintenance he performed on the airplane as a private pilot-owner (prior to his being repairman-certified), was permissible under the Federal Aviation Regulations. However, he was also aware that some of the maintenance he had performed was not permissible, and it had not been recorded in the airplane's maintenance records.

According to the airplane's FAA "Operating Limitations: Light Sport Aircraft" form issued July 20, 2006, non compliance with the limitations "...will render the airworthiness certificate invalid..." and "[a]ny change, alteration, or repair not in accordance with the manufacturer's instructions and approval will render the airworthiness certificate invalid...." Also, any maintenance must be recorded in the aircraft's maintenance records.


At 1139, an aviation routine weather report (METAR) for the Yuba County Airport (about 9 miles north of the accident site) indicated the wind was from 340 degrees at 7 knots; the visibility was 7 miles; and there was a ceiling 1,300 feet above ground level (agl). The temperature and dew point were 18 and 13 degrees Celsius, and the altimeter was 30.21 inches of Mercury. About 14 minutes later, at 1153, the cloud condition had changed, and the sky condition was reported as few clouds at 1,300 feet agl. The Yuba County Airport's elevation is 62 feet mean sea level (msl).

At 1153, a METAR for the Sacramento International Airport (about 19 miles south of the accident site) indicated the wind was from 360 degrees at 6 knots; the visibility was 8 miles; and there were few clouds at 1,500 feet agl. The temperature and dew point were 19 and 14 degrees Celsius, and the altimeter was 30.20 inches of Mercury. The preceding hour, at 1053, there were scattered clouds at 1,300 feet agl. The Sacramento International Airport's elevation is 27 feet msl.

Three witnesses, who were located within 1.5 miles of the accident site, reported that at the time of the accident they noted the weather condition. One of the witnesses, who was located southeast of the accident site, reported that he observed the airplane cruising (its altitude was about 2,800 feet agl). The witness did not report experiencing any difficulty observing the airplane. He stated that the visibility was good, and the sky was broken and in some places it was overcast. The sun was peaking through the clouds.

The second witness, who was about 1.5 miles west of the accident site, reported that the sky was "very, very, clear," and there was a light wind. The third witness reported that it was "clear and sunny," and there was little wind.


The FAA reported that a search of facilities near the accident site did not reveal evidence that any air to ground communications or services had been provided to the pilot or the accident airplane.


The airplane was equipped with a Dynon electronic flight display. The display was sent to the Safety Board's Research and Engineering Vehicle Recorder Division laboratory, Washington, D.C., for examination. The laboratory reported that, by design, the display did not record data or have memory.


The Safety Board investigator's examination of the accident site and airplane wreckage, in conjunction with witness statements, revealed that the airplane experienced an in-flight breakup. All of the airplane's structural components were located and found in sparsely vegetated level fields, having an estimated elevation of 30 feet msl.

The wreckage was scattered in an oval shaped area about 1,900 feet long by 500 feet wide. The heaviest components were found at the northern end of the wreckage path. These components (referred to as the main wreckage) consisted of the cockpit, engine, and the main landing gear. The two occupants were found within 125 feet of the main wreckage and about 100 feet from each other. The lightest components were found in the middle and southern end of the wreckage path. These components consisted of Plexiglass fragments, a carpet fragment, one seat cushion, and wheel fairings.

The magnetic bearing between the heaviest and lightest components was about 170 degrees. The wreckage consisted of the following components that were found separated from each other:

1. Left wing with attached flap (but without the aileron). Note: A rubber-like smudge, consistent with a main landing gear tire transfer mark, was found on the bottom of the left wing, near the aft spar attachment support bracket.
2. Aileron (left wing);
3. Right wing with attached flap and aileron;
4. Main landing gear assembly;
5. Cockpit (forward of occupant restraint system anchor points) with attached instrument panel, firewall, engine, and propeller;
6. Aft cockpit wall with attached fuselage. Note: The left side seatbelt floor attachment extrusion (fitting) in the cockpit was deformed (twisted) in an upward and forward direction. The right side seatbelt fitting was broken in an upward and forward direction.
7. Empennage with attached rudder;
8. Horizontal stabilizer with attached elevator and trim tab (assembly remained connected to empennage only with control cables); and
9. Wheel fairings, a seat cushion, and Plexiglass fragments (found in several locations). Note: The canopy was shattered. The canopy's latch mechanism, on the side of the fuselage, was found bent in a position consistent with the latch having been closed.)

No evidence of an oil spray residue was noted on any of the components. There was no evidence of pre- or post impact fire.

No visible evidence of corrosion was noted on the fracture surfaces of any structural components that separated. (See the wreckage distribution diagram for additional details.)


The Sutter County, California, coroner performed an autopsy on the pilot. The cause of death was listed as "extreme blunt force trauma, total body." The autopsy findings noted there was evidence of "moderate to severe coronary atherosclerosis."

The FAA’s Civil Aerospace Medical Institute (CAMI), Oklahoma City, Oklahoma, performed forensic toxicology on specimens from the pilot. No ethanol or drugs of abuse were detected.

A review of the pilot's aviation medical records revealed that on February 8, 2005, CAMI's Manager, Aerospace Medical Certification Division, issued the pilot a letter entitled "Authorization for Special Issuance of a Medical Certificate." In part, the letter stated that the FAA had reviewed the information submitted by the pilot in support of his request for a medical certificate. The evidence revealed a history of "angina pectoris, myocardial infarction and coronary artery disease requiring percutaneous transluminal coronary angioplasty with stenting and hypertension with the use of medication." The FAA manager stated that although the pilot was "ineligible for third-class medical certification," based on the complete review of the available medical evidence, the FAA would grant authorization for special issuance of the requested third-class airman medical certificate.


Radar Data and Flight Path

Recorded radar data was obtained for the accident airplane's anticipated route of flight, and for the time of the flight. An examination of the data revealed that only one target exhibited a flight profile matching the accident airplane's anticipated flight path and it disappeared over the accident site.

The Safety Board's Office of Research and Engineering, Vehicle Performance Division, reviewed the radar data. The Office provided a report describing the target airplane's flight path and certain performance characteristics including the airplane's speed, altitude, and rates of climb and descent.

In summary, the radar data indicated that at 1129 the airplane was climbing through about 200 feet msl (as indicated by the airplane's Mode C altitude encoding transponder). The airplane initially climbed in a northerly direction from the 127-foot msl Lincoln Regional Airport. Thereafter, the airplane turned and continued climbing while proceeding in a westerly direction (toward the destination airport).

By 1137:30, the airplane was in steady, level flight, at 2,600 feet and cruising about 106 knots (ground speed), on a heading of about 295 degrees (true). About 1137:50, the airplane entered a climb about 700 feet per minute to 2,800 feet, and subsequently entered a rapid descent (estimated at 2,000 feet per minute). The last transponder (secondary) radar return from the airplane was recorded at 1138:27, at an altitude of 2,600 feet. A primary radar return about 360 feet to the south of the airplane’s track appeared at 1138:18, and several other primaries to the south of the track appeared between 1138:37 and 1138:50. As shown in the Safety Board's radar plots, the wreckage distribution lies along a north-south line about 0.1 to 0.4 nautical mile (600-2,400 feet) southeast of the last secondary return. (See the Radar Data memo for additional details.)

Engine History, Design and Modification

The accident airplane was equipped with a Teledyne Continental Motors (TCM) engine, model O-200-A (82), bearing serial number 256194. TCM's examination of the accident engine revealed it was equipped with automotive-type spark plugs and a compatible magneto wiring harness. TCM reported to the Safety Board investigator that its company has not tested the engine with AMD's modifications. The modifications were contrary to TCM's tested design.

AMD reported to the Safety Board investigator that it modified and approved TCM's engine assembly and installation of automotive-type spark plugs with a compatible magneto wiring harness.

Initial Engine Examination

The impact-damaged accident engine was initially examined by the TCM participant and the Safety Board investigator immediately following its recovery. In pertinent part, the following was noted:

One of the four engine mounts was broken. The carburetor was also broken from the engine. The oil sump and the oil filter adapter were both damaged. The exhaust and induction system exhibited impact damage.

The crankshaft propeller flange was bent aft. The crankshaft was rotated by hand at the propeller flange. Rotational continuity was established throughout the valve train and to the rear of the engine. Thumb compression was noted on all four cylinders. Examination of the crankshaft gear cluster revealed that at least one gear tooth was broken.

The left magneto was not the magneto installed on the engine by TCM. According to AMD, in response to communications from the pilot's son that the left magneto had a problem, AMD sent him a replacement magneto.

The starter was observed to be an aftermarket accessory. It remained attached to the engine. The drive coupling Bendix was worn and was loose inside the starter. It fell out when the starter was examined. The seal/gasket around the starter mount to the starter adapter appeared to be displaced. The alternator remained attached and was impact damaged.

No entries were found in the engine logbook pertaining to the replacement of the left magneto or starter.

Magneto timing was verified at 28 degrees before top dead center (BTDC) for the right magneto. Timing for the left magneto was found to be beyond 32 degrees BTDC. Both the right and left magneto impulse couplings snapped when the crankshaft was rotated. Neither magneto was removed nor examined further. The ignition harness was impact damaged.

The top spark plugs were removed and examined. When compared to the Champion Aviation Check-a-Plug chart, the top spark plugs exhibited signatures that were consistent with normal wear.

All cylinders remained attached to the engine and visually appeared to be undamaged except for the number four cylinder, which had fin damage. The cylinders were examined using a lighted borescope. A normal amount of combustion deposits was observed in the combustion chamber on all cylinders, according to the TCM participant.

The TCM participant additionally reported that the intake and exhaust rocker arms were oil coated and moved when the crankshaft was rotated. The intake and exhaust valve faces exhibited normal signatures. The piston heads were examined using a lighted borescope. The piston head exhibited a normal amount of combustion deposits.

TCM Engine Teardown Results and Assembly Error

After the engine's initial examination, the engine was transported to TCM's manufacturing facilities in Mobile, Alabama, where a complete teardown examination was performed, also under the Safety Board's observation. In pertinent part, the following anomalies were observed during this examination:

1. The gasket for the starter's mount was found displaced. It was not in complete alignment with the starter's base;
2. The left magneto to engine timing was 35 degrees BTDC; and the right magneto's timing was 28 degrees BTDC. (The TCM specification is for 28 degrees, BTDC.) Scrape marks on the magneto-to-engine mounting surfaces were consistent with magneto rotation in both clockwise and counterclockwise directions;
3. The magnetos' ignition harness caps exhibited red silicone sealant protruding at each lead;
4. The crankshaft cluster gear (starter drive interface) had a fractured gear tooth. The gear tooth was found in the engine's oil sump. The gear's fracture surface and material properties were examined. TCM's laboratory reported that the gear fractured in overload, and its material properties met specifications; and
5. During manufacture, TCM's factory personnel had assembled the engine's crankcase without installing a starter adapter plug assembly (TCM part number 654502). This plug assembly blocks the flow of oil within a portion of the engine. However, TCM's engine build book record indicated that this part had been installed during the engine build process. TCM investigated the engine's history and reported to the Safety Board investigator that the required plug assembly had never been installed in the engine despite the build book record. TCM acknowledged that assembling the engine without the required plug assembly was contrary to the engine's prescribed type design.

Subsequently, TCM reported to the Safety Board investigator that this plug assembly blocks the pressurized flow of oil to an area of the engine where its presence is undesirable. TCM opined that its assembly error neither compromised the functionality of the engine nor jeopardized safety. TCM acknowledged that absent the required part, the engine was prone to leak oil from the vicinity of the starter.

At the conclusion of the TCM teardown examination, TCM reported the following: the exhaust system was clear; spark was observed upon rotation of the magnetos; the oil pump drive gear was intact; the oil filter and screen were not obstructed; the combustion chambers and valves had normal amount of deposits; and no evidence of any internal component failure, malfunction, or abnormal wear signature was observed. TCM reported no evidence of abnormalities that would have prevented normal operation and production of rated horsepower.

Engine Test Run Data and Corrective Action During Manufacture

TCM had tested the improperly assembled engine at the completion of its manufacture in May 2006. The engine had successfully operated, but during the test run TCM personnel had observed a "nonconformity." The engine leaked oil.

TCM personnel indicated that oil leaked from the vicinity of the starter assembly oil seal and from the fuel pump cover plate area. The indicated corrective action taken by TCM during manufacture was to replace the starter assembly and the cover and gasket. The fact that the engine had been improperly assembled was not recognized, and the engine was delivered to AMD.

Comparison Between Structure and Design Specifications

The thickness of a few structural components was measured to determine conformity with construction drawing specifications. The following components were measured:

1. Right and left wing rear spar attachment plates;
2. Rear spar root doublers;
3. Wing rear channel (spar);
4. Center spar web, front and rear;
5. Center spar cap (main wing spar), center section; and
6. Rear spar attachment for stabilizer.

The thickness measurements of all examined components were found consistent with the design specifications. No anomalies were noted by the Safety Board investigator or the AMD participant.

Airplane Manufacturing Anomalies, Reported by Co-Owner

The pilot's son reported to the Safety Board investigator that he observed the following anomalies during his operation of the newly purchased Zodiac:

1. The pitot-static system did not function, and the airspeed indicator did not work;
2. The autopilot did not function; its case was broken at the pitot tube connection;
3. Upon application of brakes, the airplane pulled right, and the brakes felt "spongy." Eventually, the right brake failed;
4. The oil pressure sending unit was cross threaded onto the oil bracket and was leaking oil;
5. The right fuel gauge did not function;
6. A fitting associated with a fuel transducer was attached in such manner as to vibrate;
7. The right magneto failed;
8. On the bottom of the left wing, adjacent to and trailing from rivets located near the main landing gear, evidence of dark colored debris was present. The presence of the debris next to the rivets was termed "smoking rivets;"
9. The engine assembly leaked oil, and the leak appeared to emanate from the starter;
10. An engine exhaust valve was burnt in one cylinder;
11. The magneto timing was improperly set;
12. A rudder cable rubbed against an aluminum fairlead on the side of the fuselage and appeared improperly rigged;
13. Occasionally, a vibration became evident in the floor area of the cockpit; and
14. The hub of the propeller was found cracked in September 2006, at an engine tachometer total time of 37 hours.

Regarding the cracked propeller, during the manufacturer of the airplane, AMD used propeller attachment bolts to secure the wood propeller to the engine crankshaft's mounting flange that were the incorrect length. The bolts did not conform to the design specification set by AMD. After the crack was brought to the attention of AMD, AMD provided the owner with a replacement propeller using the correct length attachment bolts.

To remedy the other anomalies, the pilot's son communicated with and received guidance from personnel at AMD, TCM, and independent mechanics. The above-listed anomalies were addressed, and flight operation of the airplane continued.

Airframe Manufacturing Anomalies, Observed by the Safety Board

1. Fuel tank sender unit cover plate installation

The AMD drawing (6-N-11) shows the specification for attaching the wing fuel tank sender unit cover plate to the wing's upper surface. An examination of the right wing revealed that the rivet pattern for the plate's attachment to the wing was not consistent with the drawing. Also, the cut out hole in the wing skin panel was offset from the rivets, and four of the eight pop rivets (shank portions) that were installed into the cover plate were not secured into the wing's skin panel.

The examination revealed that the heads of eight pop rivets were imbedded in the surface of the cover plate, and it appeared that the rivets were securing the plate to the wing. The rivet pattern consisted of one rivet at each of the four corners of the square plate, and one rivet midway between each of the corners. To remove the plate from the wing, the four rivets at the four corners of the plate were cut off. When this action was completed, the cover plated separated from the wing. The other four rivets, that were observed attached to the cover plate, were located such that they missed the wing skin panel and went into the wing tank sender unit cut out hole (see photographs).

An examination of the diameter of the nearly circular cut out hole in the two wings revealed that their diameters were different. The diameter of the cut out in the right wing was measured between 4.17 and 4.19 inches. The diameter of the cut out in the left wing was between 3.81 and 3.50 inches. The square shaped cover plates were the same size.

2. Main Landing Gear

The lower saddle support bracket attachment fitting was separated from the main landing gear. No washer was found above the nut that was used to secure the aluminum fitting. The fitting was found deformed into a shape consistent in appearance with it having been pulled through the bracket.

The main landing gear assembly was found to have been secured to the fuselage without use of 1/8-inch rubber washers on its left and right side attachment bolts. This was contrary to the airplane's design drawing number 6-G-3.

3. Flight Control Connections

Three of the four flight control system connection rods in the cockpit, which are connected to the aileron bellcrank and elevator, had rod-end nuts that were loose. Two of the three nuts had backed off from their respective seated positions, and the third nut was found finger tight. None of the rod ends were found disconnected.

Metallurgical and Airframe Structural Examinations

The aft 4 feet of the airplane's rear empennage, horizontal stabilizer with attached elevator, vertical stabilizer, rudder, and the left aileron were examined by a metallurgist from the Safety Board's Materials Laboratory, Washington, D.C. The metallurgist did not report finding any evidence of preimpact cracking or damage in any of the examined parts. The metallurgist reported that the horizontal stabilizer was completely fractured from the empennage. The upper skin of the horizontal stabilizer was perforated and locally deformed downward consistent with contact by the up elevator control horn. Stabilizer spar damage in the form of small buckles and dents were consistent with a downward movement relative to the empennage.

Also, as indicated by deformation signatures in its skin (jagged holes), the stabilizer's front empennage attachments were ripped out of the stabilizer. Damage was also noted to skin on the stabilizer's lower surface. Skin around the forward attachment brackets indicated an upward movement of the stabilizer pulling the brackets through the skin.

The aft horizontal stabilizer attachments were fractured with features consistent with tearing overstress in clockwise direction, as if the right tip of the stabilizer moved rearward.

The elevator trim tab actuator rod was found bent. When it was straightened, the trim tab appeared to be in a neutral position.

The aileron control horn was separated at the riveted attachments with deformation around the rivet holes indicating that the control horn was pulled inboard through the aileron structure. Red and blue paint was transferred to the right leading edge of the horizontal stabilizer in an area that had been flattened. The paint pattern in the flattened area matched the paint pattern in a deformed area of the left aileron. (See the Materials Laboratory Factual Report for additional details.)

The airplane's entire structure was examined by an aerospace engineer from the Safety Board's Aviation Engineering Division, Washington, D.C. The engineer reported that all of the fracture surfaces examined exhibited features consistent with static overload.

The engineer noted that the left and right wings are independent of each other and join to a spar carry through structure in the center fuselage just forward of the main landing gear forward attachment. The engineer made the following observations regarding the wings' separation from the fuselage:

Both wings, including the entire spar carry through structure, separated entirely from the fuselage. The entire left wing lower spar chord remained attached to the right wing. The lower spar chord was twisted several times about itself in a clockwise direction when viewed looking outboard. Both the lower surface and the leading edge of the left wing had black and blue transfer marks. The left wing upper spar cap at the spar carry through connection fractured in a downward direction at the outboard-most of three AN-5 bolts. The left wing rear attachment on the wing side was fractured/sheared at the attachment bolt location, and the attachment bolt remained attached to the mating fuselage structure that was bent forward.

The right wing rear attachment on the wing side remained intact with the attachment bolt. The mating fuselage structure fractured/sheared and bent forward at the attachment bolt location. (See the Structures Group Chairman's Factual Report for additional details.)


Airplane Handling Quality, Pitch Sensitivity

An FAA certified flight instructor (CFI) reported to the Safety Board investigator that he currently provides instruction in the same model of Zodiac that crashed. The CFI stated that he has been an instructor for 35 years, and has flown a variety of light and heavy aircraft including helicopters, gyroplanes, gliders, weight-shift control trikes, and Boeing 727s, 757s and 767s. The CFI is a retired airline captain, and he currently is a sport pilot examiner in land and sea airplanes, gliders, and trikes.

Regarding the airplane's maneuverability, the CFI reported that none of the airplanes he has flown have been as responsive in pitch as the Zodiac 601. He stated that the Zodiac is "extremely pitch sensitive."



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