N59JKCessna 188B · 1973

On September 4, 2013, about 1835 eastern daylight time, a Cessna 188B, N59JK, operated by Hopkins Flying Service was substantially damaged during landing roll, after a flight control malfunction, at a private airport in Woodrow, South Carolina. The commercial pilot received minor injuries. Visual meteorological conditions prevailed for the aerial application flight conducted under the provisions of Title 14 Code of Federal Regulations Part 137.

According to the pilot, he was spraying a cotton field near Lydia, South Carolina when he completed a pass and at the end of the field, pulled up to turn to the right. During the climb he applied right rudder as usual but this time, the right rudder pedal assembly broke off, and went up against the back of the hopper tank.

The pilot then decided to return to a private airport where the airplane was based in Woodrow, South Carolina as it was larger than the other spray strips in the area. Upon arrival at the airport, he flew over the top and used his cell phone to call and let personnel at the field know he would be landing without rudder control.

The landing was uneventful, but during the landing roll, the airplane began to turn left. The pilot however had no way to correct for the left turn. The airplane then exited the left side of the runway and entered a soybean field. The right main landing gear then collapsed, and the the right wing impacted the ground.

According to Federal Aviation Administration (FAA) and pilot records, the pilot held a commercial pilot certificate with a rating for airplane single-engine land. His most recent FAA second-class medical certificate was issued on January 30, 2013. He reported that he had accrued 2,249 total hours of flight experience, of which, 1,392 hours were in the accident airplane make and model.

The accident aircraft was a single seat, strut braced, low wing monoplane specifically designed by the manufacturer for aerial application use. The tail cone and empennage assemblies were of semi-monocoque construction. From the tailcone forward, a welded, tubular-steel structure was incorporated, which was covered with aluminum skin panels. It was equipped with a liquid dispersal system, which included a fiber glass reinforced plastic hopper which was located immediately forward of the cockpit, and a conventionally configured heavy duty landing gear system, which consisted of chrome-vanadium steel main landing gear springs, and a spring-steel tubular tail wheel spring, with a steerable tailwheel. It was powered by a normal aspirated, 230 horsepower, six-cylinder, horizontally opposed, air cooled, engine.

The interior of the airplane was internally corrosion-proofed, and the external finish was acid resistant.

According to FAA and maintenance records the airplane was manufactured in 1973.

On April 27, 1973, the airplane was sold by the Cessna Aircraft Company to an operator in Montana and was registered as N21796.

On July 24 1973, the airplane was sold to an operator in Canada, and was removed from the United States Civil Aircraft Registry and registered in Canada as C-FFZI.

On July 12, 1982 the airplane was involved in an accident during landing which damaged the landing gear, one of the wings, and the vertical stabilizer.

On March 24, 1994 it was sold to an operator in Colorado, and reinstated on the United States Civil Registry as N59JK.

On July 16, 1997, the airplane was modified by the installation of Aeronautical Testing Service, Inc. vortex generators on the wings and "Split Flaps" on the lower trailing edge of the inboard stub wing.

On April 2, 2008, it was sold by the operator in Colorado and over the next year was operated by a succession of operators.

On September 14, 2009, it was purchased by Hopkins Flying Service.

On April 7, 2013 a Knisley Welding aftermarket modified exhaust system was installed.

On April 7, 2013, the airplane received its most recent annual inspection.

At the time of the accident, the airplane had accrued 5,393.2 total hours of operation.

The recorded weather at Shaw Air Force Base (SSC), Sumter, South Carolina, located 8 nautical miles southwest of the accident site, at 1858, included: winds 330 at 4 knots, 10 miles visibility, few clouds at 5,500 feet, broken clouds at 10,000 feet, temperature 31 degrees C, dew point 19 degrees C, and an altimeter setting of 29.99 inches of mercury.

Examination of the wreckage revealed that the right wing and landing gear structure was substantially damaged.

Examination of the right rudder pedal arm and right rudder bar were constructed of steel tubing. Further examination revealed the presence of a plug in the top of the tubing which made up the right rudder pedal arm. This plug appeared to be made of a flexible material similar to silicone sealant. The left rudder pedal arm also contained a plug of this same silicone like material. Both the right rudder pedal arm, and the right rudder bar exhibited significant amounts of corrosion, and it was discovered that, the right rudder pedal arm had fractured just above its mounting location on the right rudder bar.

The pilot who was not wearing a helmet received minor injuries during the accident when his shoulder came into contact with the right side of the cockpit.

The cockpit was surrounded by a welded tubular steel structure and featured an emergency door release system to aid in egress, and a fire extinguisher. Cockpit seals and two cockpit pressurization scoops helped to lessen the possibility of contamination of the pilot by chemicals. The instrument panel crash pad was covered in ¼ inch black Ensolite, and urethane bumper pads were bonded to the tubular steel structure above the cockpit door, down the corner doorposts, across the tubular structure behind the pilots head, and across the lower door seal.

Wire cutter blades were installed on the landing gear struts and in front of the cockpit windshield. A steel deflector cable was installed between the top of the cockpit canopy and the top of the vertical tail.

The pilot restraint system consisted of seat belts and a double-strap shoulder harness. The lower ends of the harness were permanently attached to the seatbelt.

According to Cessna Single Engine Service Bulletin SEB96-9, Service experience indicated that the seat belt and shoulder harness assembly could become worn, frayed and/or deteriorated. To assist in maintaining optimum seat belt and shoulder harness performance and to minimize the potential for failure of the pilot restraint system, the service bulletin required that the seat belt and shoulder harness assembly should be inspected and replaced at specific intervals.

This service bulletin required that some seat belt and shoulder harnesses be replaced before flight and some within 100 hours or six months, however, ultimately all the existing seat belt and shoulder harness restraint systems should have been replaced within one year.

For airplanes equipped with a four point seat belt and shoulder harness restraint system such as the accident airplane, an initial inspection was required and then an operational inspection was to be accomplished within the next 50 hours of operation or 3 months whichever occurred first.

Repetitive inspections were then required which included an operational inspection 50 hours of operation after the initial inspection or a seat belt and shoulder harness replacement.

The replacement of the seat belt and shoulder harness with a five point restraint system was then required before the next flight, for harnesses that failed an operational inspection, or within the next 12 months, for seat belt and shoulder harnesses that passed the initial operational inspection.

Examination of the seat belts and double strap shoulder harness in the airplane revealed however, that they were manufactured in 1972. They displayed areas of broken stitching, wear, fading, and furthermore, graying, which indicated that they were also deteriorating due to exposure to sunlight and ultra violet rays.

Rudder, Brakes, and Tail Wheel

The primary flight control surfaces (ailerons, elevator, and rudder) were controlled by a conventional control stick and rudder pedal arrangement.

The rudder system consisted of individual rudder pedal assemblies, with return springs, rudder, rudder bellcrank, cables, and pulleys.

The hydraulic brakes on the main wheels were conventionally operated by applying toe pressure to the top of the rudder pedals, with the rotation of the pedals actuating the brake master cylinders, which would result in braking action on the main wheels.

The tailwheel steering was controlled through the tailwheel steering arms by cables and bellcranks attached to the rudder control cables. Tailwheel steering of 24 degrees left and right was available, and for tighter turns, application of toe pressure on either rudder pedal would cause the tailwheel to free swivel, enabling the airplane to be pivoted around the wheel being braked. The tailwheel was also equipped with an anti-swivel locking system which could be engaged by the pilot to limit steering to 2.5 degrees left and right.

Examination of the rudder system, brake system, and tailwheel steering systems revealed that with the right rudder pedal arm separated from the right rudder pedal bar, that right rudder, right brake, and right tailwheel steering, would not have been available.

Interior Care

According to the Cessna 188 Owner's Manual, care of the interior of the airplane was as important as the care of the exterior. The primary factors to be considered being cleanliness of the cockpit area and freedom from dirt and corrosion throughout the entire airframe, advising that "Some dirt and toxic chemicals will find its way into the fuselage through long periods of use; these hazards must be minimized if the pilot is to operate the airplane with safety and if the airplane is to give the long service it was designed to give. It was also good practice, before cleaning, to check the interior for signs of leaking fittings and corrosion and to note any areas where further investigation is needed; however, to not make any repairs until the airplane was thoroughly cleaned to prevent contamination from toxic chemicals.

To facilitate cleaning and inspection of the interior, the fuselage was equipped with removal panels. Two large panels on each side of the fuselage were completely removable for access to the interior structure, hopper, and cockpit area. A large door which was hinged at the top, was also located just aft of the firewall on each side of the airplane for access to the forward fuselage components. Smaller removal panels on the sides of the fuselage tailcone provided access to the control system cables and fuselage structure. The engine cowling was also completely removable for access to the engine. In general, the entire fuselage structure could be exposed for cleaning and inspection.

The Owner's Manual also advised that in order to thoroughly clean the fuselage or hopper interior to first hose it down with water, and then wash with warm soapy water. A hose rinse should then follow to flush away the soapy water. It further went on to say that, when hosing down the interior it was best to remove the pilot's seat and precautions should be made to keep water away from the instrument panel, radio, heater outlets, and map compartment, and that a protective waterproof covering for these items was recommended.

Service Manual and Illustrated Parts Catalog

Review of the Cessna 188 and 188T Service Manual, and Cessna Model 188 Series Illustrated Parts Catalog (IPC), revealed that the manufacturer had published inspection criteria for the airplane based on operating usage and operating environment, that provided mandatory time and inspection time intervals for components and airplane structures, as well as information on disassembly, overhaul, and parts breakdowns.

The manufacturer had also enacted a Corrosion Prevention and Control Program (CPCP) to help prevent or control corrosion in the airplane's primary structure so that it did not cause a risk to continued airworthiness as the airplane aged.

Further review of the Service Manual also revealed that the manufacturer had included expanded maintenance inspection items that were to be examined after the first 100 hours of operation. The inspection was then to be repeated every 600 hours of operation, or 12 months, whichever occurred first, after the initial inspection had been accomplished, and then every 600 hours of operation.

A Supplemental Inspection Document (SID) was also included in the Service Manual that listed items that were to be examined after 12,000 hours or 20 years, whichever occurred first, after the initial inspection had been accomplished, for airplanes operating in a typical usage environment. Furthermore, the Service Manual included items for airplanes that were operated in a severe usage environment (including aerial application) that were to be examined after the first 6,000 hours of operation or 10 years, whichever occurred first, with the inspection to be repeated every 1,000 hours of operation or 5 years, whichever occurred first.

Examination of the Service Manual, IPC, CPCP, and SID also revealed that guidance regarding corrosion inspection and corrosion control of the rudder system was included which contained information regarding rudder attachment (hinge brackets, hinge bolts, and hinge bearings), the rudder structure (rudder skins, ribs, forward and aft spars, and torque tube), the "rudder pedal torque tube", and cable attachments, and the rudder cable system (control cables and pulleys).

Materials Laboratory Examination

Examination of the rudder pedal assemblies by the NTSB Materials Laboratory revealed that on the outboard side of the right rudder pedal, the right rudder pedal arm was affixed to the pedal arm cross shaft using a cotter pin, and modified fender washer and bolt/nut assembly which was not specified in either the Service Manual or IPC. When the fender washer was removed from the right rudder pedal, an elongated hole due to adhesive wear was discovered.

The inboard side of the right rudder pedal was also discovered to be affixed to the pedal arm cross shaft using a cotter pin and bronze bushing. Neither of which was specified in either the Service Manual or IPC. Further examination also revealed that the bushing was not part of the originally manufactured rudder pedal assembly, and it had been inserted through a hole which had been drilled into the inboard side of the right rudder pedal.

Examination of the silicone plugs which had been applied and cured as a thixotropic paste in the top of the left and right rudder pedal arms revealed that, remnants of corrosion product had adhered to the silicone plugs indicating, that corrosion product was present before the application of the silicone resin occurred.

Examination of the right rudder pedal arm for corrosion also revealed that the external surfaces of the right rudder pedal arm exhibited areas with disbonded topcoat paint, and exposed steel with minor surface corrosion. Internally however, the right rudder pedal arm exhibited heavy rust scaling over all surfaces and rust scale which had sloughed from the surfaces had collected at the base of the right rudder pedal arm where it attached to the right rudder bar. Also though Measurements of the tube wall thickness at the top of the rudder pedal arm appeared to be full thickness with measurements of 0.054 inch, 0.053 inch, 0.056 inch, and 0.055 inch, measured at 90 degree intervals around the circumference, measurements of the tube wall thickness at the point of the fracture were only 0.030 inch at the thickest point, and 0.013 at the thinnest point.

Maintenance Records

Review of the maintenance records revealed no record of when the right rudder pedal had been modified by addition of the fender washer and bolt/nut assembly on the outboard side or of the bronze bushing assembly on the inboard side. Nor was any record discovered as to when the rudder pedal arms had the silicon sealant plugs added to them. Furthermore, no record of Supplemental Inspection Number: 27-20-01 being accomplished was discovered.