N58RV

HISTORY OF FLIGHT

On May 24, 1998, at 0630 hours Pacific daylight time, a factory-built kit, experimental airplane, Van's RV-8, N58RV, experienced an in-flight structural separation and crashed 1 mile south of Ripley, California. The aircraft was destroyed and the pilot and pilot-rated passenger sustained fatal injuries. The aircraft was being operated as a business flight by Van's Aircraft, Inc., under the provisions of 14 CFR Part 91 when the accident occurred. The flight originated from a private agricultural strip in Blythe, California, at 0618. Visual meteorological conditions prevailed at the time and no flight plan had been filed.

A relative of the passenger reported that, on departure, the pilot was seated in the front seat, while the passenger was in back.

The purpose of the flight was for the pilot to demonstrate the aircraft's flight characteristics to the passenger, who was believed by the kit manufacturer to be a potential kit buyer.

An agricultural pilot who was inbound to the strip in the opposite direction passed the southbound aircraft. He described the aircraft as being in straight and level flight at an estimated altitude of 500 feet agl. He said he did not notice anything unusual at the time.

An eyewitness, standing about 1.5 miles northwest of the crash site, reported that he heard the sound of an engine surging and looked in the direction of the sound. He saw a yellow aircraft flying straight and level, about 1,000 feet agl. The aircraft was on a southerly heading, about 1 mile east of his location. He estimated that he watched the flight for about a minute when he saw something fall from the aircraft. This was followed almost immediately by a loud boom that he described as sounding like a "shotgun." The aircraft's nose suddenly pitched up about 45 degrees then abruptly nosed over as it began a left roll. The aircraft entered a nose-down spin to the left, continuing in a vertical descent until impact.

A farmer, flying over the accident site a few days after the accident, reported that he observed crop discoloration in an arc-shaped pattern between the location of the left wing and the main wreckage.

PERSONNEL INFORMATION

The pilot had been employed as a factory demonstration pilot for the manufacturer about 2.5 years. According to the kit manufacturer, a factory demonstration flight consists of turns, climbs, descents, slow flight, stalls and lazy 8's. There are no parachutes onboard and no aerobatic maneuvers are planned.

The pilot had built and flown his own RV-4 and had flown aerobatic maneuvers in that aircraft.

In addition, he was an experienced agricultural pilot and a commercially rated rotorcraft pilot, having served two tours in Vietnam as an Army helicopter pilot.

AIRCRAFT INFORMATION

The aircraft was the second factory-built kit aircraft of this design, having been completed on March 28, 1997, and received its airworthiness certificate on April 2, 1997. It had flown over 400 hours since that time. As part of its flight test program, the aircraft had performed loops, rolls, immelmann turns, horizontal eight's, spins, and other similar positive g aerobatic maneuvers without incident. The kit was designed to meet the design standards of 14 CFR Part 23.

The aircraft was designed for flight loads that were within +6 g's and -3 g's. The wing design has been static tested to +9 g's. Since the wing spar is nearly symmetrical in cross section, there were no negative g static tests or flight test maneuvers performed. The maximum allowable gross weight is 1,800 pounds; however, the maximum allowable gross weight for aerobatic performance is 1,550 pounds. A g-meter was not installed at the time of the accident nor were there parachutes on board.

The wing spar specifications are for 2024T3 sheet (web), 2024T351 (caps) and 2024T3 sheet (stiffeners).

The aircraft was equipped with dual controls; however, the brakes can only be operated from the front seat. The elevator trim is actuated by an electric servo, which the pilot controls with a 4-way switch that was mounted on top of the front seat control stick.

The aircraft was equipped with an electronic engine management system that has a nonvolatile memory.

A newly designed right-wing aileron was installed on May 8, 1998. The design change was intended to decrease burbling that had been experienced with full deflection control movements. A flight test that was completed on May 15, 1998 included full control deflection up to the maneuvering speed of 142 mph. No adverse effects were noted during the test. The aircraft had flown about 8 to 9 hours since the installation.

The aircraft is equipped with two main fuel tanks, each with a capacity of 21 gallons. There is no flow through between the tanks. The fuel selector has three positions: left, right, and off. There was one inverted pickup tube in the left tank. The tanks are vented separately.

The aircraft's fuel system is equipped with a fuel boost pump that is intended to remain on during takeoff and landing.

The aircraft was last refueled on May 23, 1998, at Payson Aviation in Payson, Arizona. At that time, it was serviced with 31.4 gallons of 100-octane low lead aviation fuel.

METEOROLOGICAL INFORMATION

The localized weather conditions at the time of the accident were uniformly described by witnesses near the crash site as being calm and clear.

WRECKAGE AND IMPACT INFORMATION

A postaccident inspection of the aircraft by the Safety Board found an outboard section of the left wing about 0.17 miles northeast of the main wreckage on a bearing of 240 degrees to the main wreckage site. The coordinates were 33 degrees 30.018 minutes north longitude and 114 degrees 38.734 minutes west latitude.

The main spar of the left wing was fractured at a point inboard of the aileron and outboard of the flap. The main spar of the right wing was also fractured about the same location, but remained attached by the wing's outer skin.

Fragments of the canopy were found between the separated wing and the main wreckage. Each fragment was examined for evidence of foreign material transfers; however, none were found.

The main wreckage was located at 33 degrees 29.967 minutes north longitude and 114 degrees 38.934 minutes west latitude. The elevation of the accident site was estimated at 380 feet msl.

The aircraft was found in an agricultural field, crushed, and buried in the ground to a depth of 5 to 6 feet so that only the empennage and the remaining wing structure were visible. The leading edges of both horizontal stabilizers were crushed. The vertical stabilizer had "scorpioned" over, with crushing visible on its leading edge. The remaining left wing structure and the right wing showed leading edge damage along with aft bending of the spars.

The elevator trim tab control arm was separated at the clevis from the control surface. The trim actuator control rod was extended 1.5625 inches (full extension) tab up. The arm has a clevis on both ends. The clevis is made of DuPont nylon, a composite material, comprised mainly of nylon and fiberglass. The rod is threaded with a jam nut fixing the position of the clevis. The control arm itself is made from stainless steel. The control arm had separated at the aft clevis. Further examination revealed a notch had been cut, making the clevis deeper. After a discussion with the kit manufacturer, it was established that the extended notch was not a design feature. However, without the extended notch, it was evident that the clevis would have contacted the control horn on the elevator trim tab during it range of movement.

The aircraft was removed from the impact point with a backhoe, in order to recover the occupants. This prevented Safety Board investigators from examining the wreckage while it was still in an undisturbed condition. Photographic evidence provided by the Riverside County Sheriff's Office was the basis for the wreckage condition and impact description.

The wreckage was recovered to an inspection site in Phoenix, Arizona, and laid out in a two-dimensional manner on May 27, 1998. All control surfaces were identified. Due to the degree of impact damage, it was not possible to establish control continuity.

The forward control stick was located; however, the electric elevator trim switch was destroyed. Engine control continuity was established from the front seat. Rear seat controls, though destroyed, were identified as being installed.

The fitting connecting the right tank fuel line to the fuel selector was less than finger tight and was separated from the selector with less than three full turns. The fuel selector valve exhibited multiple internal fractures. The selector handle was displaced but remained closer to the left position than it was to the off or right position. No fuel selector position is specified in the aircraft POH for takeoff or landing but the kit manufacturer suggested that it is common practice to select the fuller of the two tanks during those phases of flight.

An engine teardown was also conducted on this date. When the No. 1 and 2 cylinders were removed from the case, it was noted that the front of the case was loaded with compacted dirt. The cooling fins to both cylinders exhibited aft crushing. Crankshaft rotation was limited to about 20 to 30 degrees with binding. Disassembly continued with the remaining cylinders being removed. When the crankshaft was removed, it was visually noted that there was an apparent bend at the No. 1 rod journal. The oil sump was fractured into several pieces. The ring gear and support housing had fractured into several pieces. The rear crankshaft gear was in place with the safety intact. The top of the engine case was cracked but there was no evidence of an oil leak. The engine data plate was not recovered.

The top spark plugs, rocker box covers, and accessory case were removed. The rocker box cov...