On April 7, 2008 at 1701 eastern daylight time, a Czech Aircraft Works SPOL SRO CH 601 XL RTF, N357DT, was substantially damaged when it impacted trees and terrain following an uncontrolled descent near Polk City, Florida. The certificated private pilot/owner was fatally injured. Visual meteorological conditions prevailed, and no flight plan was filed for the flight which departed Williston Municipal Airport (X60), Williston, Florida and was destined for Lakeland Linder Regional Airport (LAL), Lakeland, Florida. The personal flight was conducted under 14 Code of Federal Regulations Part 91.

According to a friend of the accident pilot, they both shared a hangar at Lumpkin County Airport (9A0) in Dahlonega, Georgia, and they planned to fly their respective airplanes to an air show in Lakeland, Florida. The accident pilot departed 9A0 about 1240, and the friend departed about 1300. During the flight the friend passed the accident pilot, landed at X60 for a planned fuel stop, and subsequently departed. He last spoke to the accident pilot about 1500, as the accident pilot approached X60. The accident pilot planned to land about 1530, and stated that his airplane was "running fine."

Review of fuel receipts revealed that the pilot purchased fuel in X60 at 1607.

According to Federal Aviation Administration (FAA) radar data, a radar target correlated to be the accident airplane departed Williston at 1614, and proceeded southeast. The radar target transmitted a transponder code of 1200, but did not transmit any altitude data. About 1620, transmissions from the target transponder ceased, and only primary radar returns were observed. The radar target continued southeast before turning south, and the final radar target was observed at 1701 in the vicinity of where the wreckage came to rest.

About 1700, a witness, located about 1/2-mile southeast of the accident site, observed the accident airplane as it approached him from the north. He noted that the airplane was noticeably lower than other airplanes that generally flew in the area, and that it was in a 3- to 5-degree left bank. The airplane then banked right, about the same magnitude, before it returned level for a moment. About 1 to 3 seconds later, the airplane banked to the left and to the right, significantly steeper than during the previous banks.

After returning to level flight, the airplane yawed to the right, and the right wing "folded up." The airplane banked to the right and the left wing "went up" just before the airplane entered a nose dive. The witness described that the wings of the airplane looked perpendicular to their normal position, and that the engine sound changed, as the airplane descended downward. The witness heard the sound of the engine pitch change to the point where it sounded like the engine "over sped" as the airplane descended. He lost sight of the airplane shortly thereafter as it descended behind a building.

Another witness, located about 1/3-mile west of the accident site, stated that he heard a "pop" and that when he looked up he saw the accident airplane. He described that the right wing was folded "back and to the side," and that the accident airplane was spinning as it descended. About 8 to 10 seconds later he heard the sounds of an impact.


The weather conditions reported at LAL, located about 17 miles southwest of the accident site, at 1658, included winds from 030 degrees at 4 knots, 15 statute miles visibility, scattered clouds at 3,000 feet, broken clouds at 25,000 feet, a temperature of 27 degrees Celsius (C), a dewpoint of 16 degrees C, and an altimeter setting of 29.90 inches of mercury.


The pilot held a private pilot certificate with a rating for airplane single engine land. His most recent FAA second class medical certificate was issued on February 29, 1980. Examination of the pilot's logbook revealed that he had accumulated 514 total hours of flight experience, 71 hours of which were in the accident airplane make and model. According to the pilot's friend, the pilot originally obtained a pilot certificate in the 1970's, and had recently began flying again after purchasing the accident airplane in October 2007.


The accident airplane was manufactured in 2005 and was classified by the FAA as a Special Light Sport Aircraft. According to maintenance records, the airplane was factory test flown on June 16, 2005. The airplane's most recent annual conditional inspection was completed on September 1, 2007. At the time of the inspection, the airplane had accumulated 142 total hours of operation.


The airplane came to rest in a heavily wooded area and exhibited various degrees of damage throughout. The trees in and around the accident site had been removed to facilitate access to the wreckage. There was no evidence of either pre or post-impact fire. No evidence or any patterns like those typically associated with a moving in-flight fire were identified. No soot patterns were identified. No melted or splattered aluminum was observed on any of the structure.

The entire fuselage, less the canopy, was located at the main wreckage site. The forward fuselage was buried in a 6-foot deep crater, and the remainder of the fuselage aft of the firewall was protruding from the ground at about a 45 degree angle. The fuselage forward of the firewall was completely destroyed by impact forces. The fuselage aft of the firewall displayed evidence of impact damage and skin wrinkling. The upper left hand side of the fuselage in the area just below the canopy exhibited black transfer marks. The marks were in a 5 inch by 5 inch area. The area extended from about 5 inches below the canopy to about 5 inches forward of a fuselage rivet line beginning at the flap actuation mechanism. The canopy was not recovered, and only several small fragments of the canopy were located at the impact site.

The left and right wings were independent of each other, and joined to a spar carry through structure in the center fuselage, just forward of the main landing gear. Both the left and right wings remained attached to the fuselage. The left wing was bent 90 degrees up at the side of body and was resting on the fuselage. The left lower spar chord, where the left wing attaches to the spar carry through, was fractured and bent upward about 20 inches outboard of butt line zero outboard of the three AN5-15A airframe bolts.

All of the bolts remained attached and were recovered, but the bolt hole at the location of the fracture was deformed out of round. The chord pieces were both fractured vertically through the third hole, outboard of the manufactured edge. The fractures exhibited significant elongation and through the thickness yielding consistent with overstress separations. Deformation at the fractures was indicative of upward bending of the left wing tip. The rivets connecting the two pieces of the left wing lower spar chord were sheared. The left hand upper spar chords were bent upwards at about a 90 degree angle about 7 inches from the inboard end of the outboard spar chord at the 6th fastener hole. The upper spar chord was also bent aft at the 13th fastener hole from the inboard end of the outboard spar chord. The left hand rear spar attachment was fractured and bent aftward and the attachment bolt remained attached to the wing structure.

The right wing was intact, and all of the spar carry through structure remained attached (3 upper and 3 lower bolts) along with the rear spar to fuselage attachment bolt. The right hand side of the wing displayed impact damage and was bent forward, about 60 inches outboard of the fuselage at the flap to aileron intersection. All of the fracture surfaces examined exhibited features consistent with static overload with no evidence of metal fatigue.

The horizontal stabilizer, elevator and rudder remained attached to the fuselage by the elevator and rudder control cables, and were located at the main wreckage site. The aft 20 inches of the empennage (aft fuselage) were bent to the right when looking forward. The left hand side of the horizontal stabilizer exhibited evidence of a tree strike. The horizontal stabilizer to aft fuselage attachments were all fractured. The horizontal stabilizer right forward attachment angle remained attached to the aft fuselage and was bent inboard (to the left). The left forward attachment had separated from the fuselage and remained attached to the horizontal stabilizer attachment angle. The aft horizontal stabilizer attachment was fractured at the fuselage interface and bent aft. A small portion of the attachment flanges remained on the fuselage side with the attachment bolts. The rear attachments on the fuselage side were both bent to the right, when looking forward. Both the upper and lower surfaces exhibited evidence of impact damage. All of the fracture surfaces examined exhibited features consistent with static overload with no evidence of metal fatigue.

All of the movable control surfaces were located and identified at the accident site. The elevator had fractured about 3.5 inches to the left of butt line zero into two sections. The section left of the fracture separated from the horizontal stabilizer at the hinge line, and the other remained attached to the horizontal stabilizer. The elevator trim tab remained attached to the left hand portion of the elevator. Both elevators had impact marks on both the upper and lower surfaces. The rudder was separated from the empennage at both the upper and lower attach points at the fastener locations. Both the right and left sides of the rudder exhibited evidence of impact damage and the leading edge exhibited evidence of a tree strike about 36 inches above the lower attach point.

Control cable continuity was confirmed from the rudder pedals to the rudder horn attachment points. Elevator control continuity was confirmed from the cockpit to the elevator control horn attachment point, which had been disconnected by first responders. The electrically actuated elevator trim was in a neutral position. Continuity of the aileron control cables was confirmed from the cockpit to both ailerons.

The electrically actuated aileron trim was positioned at the tab trailing edge down limit. Disassembly of the aileron trim actuator revealed that the mechanism was positioned aft, in contact with a forward-facing micro-switch. The aileron trim servo, switch, and indicator were retained for testing. Functional testing of all three components revealed that they operated within the manufacturer's specifications.

The wing flaps were in the up position with evidence of impact damage. The flap linear actuator was in the retracted position, with about a half inch of the pushrod extending from the actuator housing. The left wing flap exhibited evidence of impact damage and a tree strike. It remained attached by the outboard 10 inches of the hinge. The right wing flap remained attached to the wing with only the outboard 12 inches of the hinge separated from the wing. Both ailerons remained attached to the wing, and exhibited evidence of impact damage. The left aileron pushrod access hole on the wing was fractured on the upper side, and the right pushrod aileron access hole had a corresponding tear on the lower side. The aileron trim tab on the right aileron was deflected to the full trailing edge down position, which was confirmed through examination of the trim tab linear actuator. All of the fracture surfaces examined exhibited features consistent with static overload with no evidence of metal fatigue.

The nose landing gear had separated from the fuselage structure and was located in the crater. It remained attached to the firewall structure with the power plant and propeller .The main landing gear remained attached to the airplane.

The engine was located at the forward portion of the wreckage, and was buried at the bottom of the impact crater. All three composite propeller blades were separated near the propeller hub. A portion of one blade was found outside and adjacent to the crater, while the remainder of the blades were found within the crater. Crankshaft, gearbox, and valvetrain continuity were confirmed from the propeller hub to the aft portion of the engine. The propeller hub could only be rotated through 180 degrees of motion, and valve motion was observed on all but the number 4 cylinder intake valve. All eight spark plug electrodes were intact and light gray in color. Examination of the oil filter revealed no evidence of any visible metal particles.

Both fuel tanks were ruptured and absent of fuel. There was a strong odor of fuel at the scene, and fuel was observed leaking from several fuel lines during the recovery process. The fuel selector knob was observed in the "left" position.


The FAA's Bioaeronautical Sciences Research Laboratory, Oklahoma City, Oklahoma, performed toxicological testing on the pilot. No traces of ethanol or drugs were detected.

An autopsy was performed on the pilot by the Office of the District Medical Examiner, 10th Judicial Circuit of Florida, Winter Haven, Florida. The autopsy report noted the cause of death as "blunt impact."


The inboard portion of the left wing spar was disassembled, cut from the wing, and shipped to the Safety Board Materials Laboratory for further examination.

According to the Materials Laboratory factual report, the spar contained three web layers, two upper chords, and two lower chords. The aft upper chord was installed between the aft and center web portions, and the forward upper chord was installed between the center and forward web portions. A bolt was inserted through one of the attachment holes for the lower chords and three web portions. This bolt kept the three layers of web and chord intact.

Upon removing the bolt, the three layers of web were separated by hand so that the internal surface of the web could be examined. Visual examination of the disassembled web pieces revealed the internal surface of the forward web portion was stencil marked in ink "6061-T6 Corus Alum", an indication that the sheet was made by Corus Aluminum and was distributed/sold as 6061 aluminum alloy in the T6 condition. Bench binocular microscope examination of the various spar pieces revealed the fracture faces contained fine texture features on a slant plane consistent with overstress separation with no evidence of fatigue cracking.

The web and chord portions reportedly were specified as 6061 aluminum alloy in the T6 condition. MIL-H-6088F, titled "Heat Treatment of Aluminum Alloys", provided typical electrical conductivity and hardness values for aluminum alloys. Small pieces of the web and spar portions were excised with a saw from the spar in order to determine the thickness, electrical conductivity and hardness. The measured electrical conductivity and hardness for the web and chord portions were consistent with 6061 aluminum alloy in the T6 condition.

Detailed examination of the fractured left wing lower spar chord pieces revealed that they consisted of two side-by-side 1 1/2 by 1/4-inch riveted aluminum bars. The pieces were marked with orientations and were reported to be manufactured from 60611-T6 aluminum alloy.

The chord pieces were both fractured vertically through the third hole outboard of the manufactured edge. The fractures exhibited significant elongation through the entire thickness, consistent with overstress separations. Deformation at the fractures was indicative of upward bending of the left wing tip. The rivets connecting the two pieces were sheared.

Hardness and conductivity measurements on both pieces were consistent with the specified material, 6061-T6 aluminum alloy.


Comments of Previous Owner

During a telephone interview, the previous owner of the airplane stated that when flying the airplane with a single occupant in the left seat, it was normal to trim the aileron fully in a right wing down direction in order to maintain lateral balance.

Airworthiness Information

This category of airplane was not certified to Federal Aviation Regulations (FAR) Part 23 Airworthiness Standards, but was instead designed and manufactured in accordance with an industry consensus standard. The manufacturer of an aircraft for airworthiness certification in the light-sport category must manufacture the aircraft to the design requirements and quality system of the applicable consensus standard that have been accepted by the FAA and published through a notice of availability in the Federal Register. To meet the intent of FAA 21.190, and to be eligible for a special airworthiness certificate for light sport aircraft (LSA) category, the applicant must present satisfactory evidence that the aircraft was manufactured and found acceptable to the provisions of the applicable consensus standard. Evidence of acceptability is provided by the light-sport aircraft manufacturer's statement of compliance, FAA Form 8130-15, attesting to compliance with the requirements of 21.190 of the FAR.

The Zenair CH 601 XL design and analysis documentation was assessed by a National Transportation Safety Board Structures Specialist. The design methods were based on a building-block approach for coupon, sub-component, and full-scale testing. The methods and assumptions noted in the documents were typical of industry accepted practices. The structural analysis reviewed for the wing, empennage and fuselage structure were consistent with industry accepted practices and American Society for Testing and Materials (ASTM) consensus standards. The documents and testing indicated that the airplane met the ASTM requirements for limit loads; however may not have fully met the 6g ultimate load requirements.

Use your browsers 'back' function to return to synopsis
Return to Query Page