N9575T – Piper Aircraft Corp PA-38-112 – 1999-09-17 NTSB Accident Report

HISTORY OF FLIGHT

On September 17, 1999, at 1248 central daylight time, a Piper PA-38-112, N9575T, operated by Decatur Aviation Inc., Decatur, Illinois, was destroyed during impact with the terrain following an in-flight loss of control, near Warrensburg, Illinois. Visual meteorological conditions prevailed at the time of the accident. The Federal Aviation Administration (FAA) private pilot checkride was operating under the provisions of 14 CFR Part 91 and was not on a flight plan. The private pilot candidate and FAA Designated Examiner were fatally injured. The local flight originated from the Decatur Airport, Decatur, Illinois, at 1217.

A witness to the accident was working in an agricultural field approximately 1/4 statute mile from the accident site. The witness reported that, "I saw the last 200 foot of the plane crash. I think [the] plane was spinning counter clockwise dropping like a rock from the sky. It crashed nose first in the field just east of where I was sitting."

The witness reported that approximately two minutes prior to the accident he, "Saw large whirl wind in S.E. [southeast] corner of our corn field. Approx[imate] size 100' tall by 40-50 ft wide with thousands of corn stalks leaves swirling in it." The witness reported that the whirlwind's size was described by one of his fellow workers, located approximately 3-statute miles from the accident site, as a "small tornado".

The witness' written statement is appended to this factual report.

Aircraft radar track data for the period before and after the reported accident time was obtained from FAA Air Traffic Control. The obtained data indicated there was a single aircraft, which was transmitting a visual flight rules (VFR) 1200 transponder beacon code, maneuvering near the accident location around the time of the accident. According to the data, the aircraft's Mode-C transponder altitude encoding function was transmitting erroneous information, and accurate altitude data was not obtained.

A plot of the aircraft radar track data was generated, and is appended to this factual report along with a copy of the source data.

PERSONNEL INFORMATION

According to FAA records, the private pilot candidate was the holder of a private pilot certificate with a glider rating, which was issued on September 13, 1997. FAA records show the pilot's last medical examination date was March 3, 1999, and the pilot was issued a third-class medical certificate with the limitation, "MUST WEAR CORRECTIVE LENSES."

The private pilot candidate reported on FAA Form 8710-1 "Airman Certificate and/or Rating Application", dated September 16, 1999, that his total flight time was 132 hours. The private pilot candidate listed his total time in the accident make and model as 71 hours and his total time in single engine land airplane as 71 hours. The private pilot candidate stated that 37 of the 71 hours were as pilot-in-command (PIC). The private pilot candidate reported that he had 61 hours of flight time in a glider, of which 39 hours were as PIC.

According to the private pilot candidate's flight logbook, he had flown 45.7 hours the last 90 days and 12.1 hours in the last 30 days, all of which were in the accident make and model of airplane. The pilot had flown 2.1 hours during the previous 24 hours and all 2.1 hours were flown in the accident make and model of airplane.

According to the flight logbook, the private pilot candidate had an endorsement for the private pilot practical exam, dated September 17, 1999. According to the logbook, spin entry, spins, and spin recovery techniques were not demonstrated to, or executed by, the private pilot candidate in an airplane. According to Federal Aviation Regulations (FAR) 61.105 Aeronautical Knowledge, private pilot candidates are required to have satisfactory aeronautical knowledge of, "stall awareness, spin entry, spins, and spin recovery techniques for airplane and glider category ratings." The private pilot candidate took the required aeronautical knowledge test on August 3, 1999, and received a passing score. According to FAR 61.107 Flight Proficiency, a private pilot candidate does not need to have logged any flight training in spin entry, spins, and spin recovery techniques in order to be eligible for a private pilot certificate.

According to FAA records, the FAA Designated Examiner was the holder of a commercial pilot certificate with ratings for airplane single-engine land, airplane multi-engine land, and instrument airplane operations. The examiner also held a certified flight instructor certificate with ratings for single-engine land, multi-engine land, and instrument airplane operations. According to FAA records, designated examiner authority was granted to the examiner on March 16, 1999, by the Springfield Flight Standards District Office, Springfield, Illinois. FAA records show the pilot's last medical examination date was October 12, 1998, and the pilot was issued a second-class medical certificate with the limitation, "MUST WEAR CORRECTIVE LENSES FOR NEAR AND DISTANT VISION".

The FAA Designated Examiner reported on FAA Form 8710-1 "Airman Certificate and/or Rating Application", dated March 16, 1999, that his total flight time was in excess of 29,250 hours. According to NTSB Form 6120.1/2, completed by a representative of Decatur Aviation Inc., the examiner had accumulated more than 2,000 hours in the accident make and model of airplane as pilot-in-command. The examiner was reported to have flown more than 18,000 hours in single-engine land airplanes and 14,000 hours in multi-engine airplanes. The examiner had flown 31.5 hours in the last 90 days, of which 6.3 were in the accident make and model of airplane. The examiner had flown 17.2 hours in the last 30 days, of which 3.4 were in the accident make and model of airplane. According to company flight logs and NTSB Form 6120.1/2, the examiner did not fly within 24 hours of the accident.

AIRCRAFT INFORMATION

The aircraft was a Piper PA-38-112, Tomahawk, serial number 38-78A0135. The Tomahawk is a single-engine, low wing monoplane of all metal construction equipped with a fixed landing gear, and can accommodate a pilot and a passenger in a side-by-side seating arrangement. The Tomahawk has a T-tail empennage configuration with a fixed horizontal stabilizer mounted on top of the vertical fin.

The FAA issued the airplane a Standard Airworthiness Certificate on April 25, 1978, and the airplane was certificated for both normal and utility categories. The airplane had accumulated a total-time of 6019.25 hours at the time of the accident. The last annual inspection was completed on July 16, 1999, and had accumulated 13.35 hours since the inspection. According to the aircraft maintenance logbooks, all applicable FAA Airworthiness Directives were complied with at the completion of the last annual inspection.

The FAA issued Airworthiness Directive (AD) 78-22-01 "Rudder or Elevator Binding" for the Piper PA-38-112 Tomahawk aircraft on October 30, 1978. The purpose of AD 78-22-01 was, "To prevent possible binding of the rudder or elevator in flight... ." A one-time compliance with the AD could be accomplished by completing all portions of Piper Service Bulletin 607A. According to the maintenance logbook, AD 78-22-01 was fully complied with on December 22, 1978. Copies of AD 78-22-01, Piper Service Bulletin 607A, and the maintenance logbook entry are attached to this factual report.

Piper Service Bulletin 876, entitled "Additional Flow Strip Installation", was issued on April 12, 1979. The purpose for issuance of Piper Service Bulletin 876 was reported as, "The installation of two (2) additional flow strips to the leading edges of the wings has been found to improve operational characteristics during stall maneuvers. This Service Release announces the availability of a Flow Strip Installation Kit." A copy of Piper Service Bulletin 876 is attached to this factual report.

The FAA issued Airworthiness Directive (AD) 83-14-08 "Stall Characteristics" for the Piper PA-38-112 Tomahawk aircraft on September 27, 1983. The purpose of AD 83-14-08 was to mandate that all areas addressed in Piper Service Bulletin 876 to be accomplished within 100 hours time-in-service of the AD effective date. According to the airframe maintenance logbook, AD 83-14-08 was complied with on September 1, 1983. The logbook entry states, "... C/W [complied with] AD 83-14-08 installed kit 763-930 and had airspeed [indicator] remarked see maintenance release." Copies of AD 83-14-08, the maintenance logbook entry, and airspeed indicator maintenance release tag are attached to this factual report.

Piper Service Bulletin 661, entitled "Control Column Installation Modification", was issued on August 22, 1979. The purpose for issuance of Piper Service Bulletin 661 was reported as, "It has been reported that roughness or binding of the control shaft in the instrument panel cutout hole can occur when the control wheel is full aft and direct forward or upward-and-forward pressure, rather than linear pressure along the control shaft, is applied. This Service Release contains instructions for the modification of the control wheel shaft bushing assembly to provide a smoother forward movement of the control shaft from a fully aft and up position." A copy of Piper Service Bulletin 661 is attached to this factual report.

Piper Service Bulletin 800, entitled "Elevator Control Travel Modification", was issued on November 19, 1984. The purpose for issuance of Piper Service Bulletin 800 was reported as, "It has been reported that when either control wheel is full aft and an upward-and-forward pressure is applied, the potential exists that the pilot may be unable to move the control yoke out of the full aft position. This Service Bulletin provides instructions to modify the elevator up travel to alleviate the above described problem." A copy of Piper Service Bulletin 800 is attached to this factual report.

The FAA issued Airworthiness Directive (AD) 85-11-06 "Control Loss" for the Piper PA-38-112 Tomahawk aircraft on May 20, 1985. The purpose of AD 85-11-06 was to mandate that Piper Service Bulletins 661 and 800 to be accomplished within 50 hours time-in-service of the AD effective date. According to the airframe maintenance logbook, AD 85-11-06 was complied with on July 25, 1985. The logbook entry states, "Complied with AD 85-11-06 Ele. [Elevator] Control, dated 5-30-85 by compling [complying] with Piper S/B's [Service Bulletins] 661 & 800 per AD." Copies of AD 85-11-06 and the maintenance logbook entry are attached to this factual report.

The aircraft Pilot Operating Handbook (POH) included a "Weight and Balance Supplement" sheet, dated February 24, 1989, which listed the following information:

New Empty Weight 1,166.5 lbs New Empty Weight CG 73.51 inches New Useful Load 503.5 lbs New Empty Weight Moment 8,5753.12 in/lbs

The POH listed the PA-38-112 certificated maximum gross weight as 1,670-lbs.

FAA records show the private pilot candidate's weight, as of the last medical examination dated March 3, 1999, was 214 lbs.

FAA records show the FAA Designated Examiner's weight, as of his last medical examination dated October 12, 1998, was 206 lbs.

According to fueling records and written statements provided by employees of Decatur Aviation Inc. the accident airplane was fully fueled prior to the accident flight. The Piper PA-38-112 POH lists the maximum usable fuel capacity as 30 gallons or 180 lbs.

Based on the above information, the aircraft's approximate weight was determined to be 1,766.5 lbs at the beginning of the flight. The center-of-gravity location was determined to be 76.789-inches aft of the datum point.

A time log was located in the accident aircraft and indicated that the accident flight began at 2,726.4 hours, and the hour-meter indicated 2,727.2 hours at the accident site. The duration of the accident flight, from engine start-up to the time of the accident, was calculated to be 0.8 hours.

The weight and balance supplement sheet, the aircraft time log, and the weight and balance section from the Piper PA-38-112 POH are appended to this factual report.

The engine was a 112-horsepower Avco-Lycoming, O-235-L2C, serial number L-15518-15, and at the time of the accident had accumulated 6,019.25 total hours since new. The engine had accumulated 697.55 since the last major overhaul, which was completed on June 1, 1995.

The propeller was a Sensenich, 72CK-0-56, serial number K6382, and at the time of the accident had accumulated 71.5 hours since overhaul. The last propeller overhaul was completed on November 30, 1998. The propeller was inspected on July 16, 1999, and had accumulated 13.35 hours since inspection.

METEOROLOGICAL INFORMATION

A weather observation station, located at the Decatur Airport (DEC), 14.2 nautical miles (nm) southeast of the accident site, reported the weather two minutes after the accident as:

Observation Time: 1250 cdt Wind: Variable Direction at 5 knots Visibility: 10 statute miles Sky Condition: 4,900 Scattered Temperature: 20-degrees centigrade Dew Point Temperature: 08-degrees centigrade Pressure: 30.26-inches of mercury

Wind data was obtained from a weather station located approximately 19 nautical miles northwest of the accident site. The upper atmospheric winds were reported to be:

Reporting Time: 0700 cdt Feet mean sea level (msl) Wind Direction/Velocity (degrees true/knots) 584 350/05 669 360/05 997 030/06 1,998 065/07 2,818 050/06 2,999 045/06 3,999 025/08 5,112 015/09

Reporting Time: 1900 cdt Feet mean sea level (msl) Wind Direction/Velocity (degrees true/knots) 584 040/03 669 050/02 997 055/01 1,998 040/03 2,818 045/03 2,999 045/05 3,999 040/05 5,112 055/03

WRECKAGE AND IMPACT INFORMATION

The National Transportation Safety Board's (NTSB) on-scene investigation began on September 17, 1999.

A global positioning system (GPS) receiver reported the accident site position as 39-degrees 59.332-minutes north latitude, 89-degrees 06.037-minutes west longitude. The accident site was approximately 0.14 nm south of Hampshire Road and 0.2 nm east of Lincoln Memorial Parkway near the town of Warrensburg, Illinois.

The aircraft impacted in an open, level, recently harvested cornfield. There were no trees or other ground obstructions in the vicinity of the accident site. The angle between the terrain and the longitudinal centerline of the engine was measured at 35-degrees. No wreckage propagation was noted from the initial impact location. The heading of the aircraft wreckage was measured with a compass along the aircraft's longitudinal axis and was 110-degrees magnetic. All components of the aircraft were located at the accident site and all flight control surfaces remained attached at their respective airframe positions.

The left wing remained attached to the fuselage and exhibited full span aft directional bending. Leading edge compression damage was noted along the entire wingspan. The leading edge skin was crushed aft to the forward wing spar and the bottom cap of the forward spar was deformed in the aft direction. The aft spar was found fractured approximately 1-3/4 feet outboard of the wing root and the flap actuation torque tube was bent in the aft direction. The left wing flap remained attached to the flap actuation torque tube, and the flap was in a fully retracted position. The left wing flap was buckled along its entire span. The left aileron remained attached to the wing and the left aileron counterbalance weight was found with the aileron. The left aileron was buckled along its entire span. Both aileron cables were found attached to the left aileron bellcrank and to the aileron torque tube located in the fuselage. The aileron actuation push/pull rod was found attached to the aileron bellcrank and the flight control surface. The left wing fuel tank was ruptured and no residual fuel was noted. The left main gear was found attached to the wing structure and was bent in an aft and outboard direction. Both inboard and outboard stall strips were installed on the left wing.

The right wing remained attached to the fuselage and exhibited full span forward directional bending. Leading edge compression damage was noted along the entire wingspan. The leading edge skin was crushed aft to the forward wing spar. The aft spar was found fractured approximately 1-3/4 feet outboard of the wing root and was bent forward. The right wing flap remained attached to the flap actuation torque tube, and the flap was in a fully retracted position. The right wing flap was buckled along its entire span. The right aileron remained attached to the wing. The right aileron counterbalance weight was torn from its mounting and was located at the wreckage site. The right aileron was buckled along its entire span. Both aileron cables were found attached to the right aileron bellcrank and to the aileron torque tube located in the fuselage. The aileron actuation push/pull rod was found attached to the aileron bellcrank and the flight control surface. The right wing fuel tank was ruptured and no residual fuel was noted. The right main gear was found attached to the wing structure and was bent in an aft and outboard direction. Both inboard and outboard stall strips were installed on the right wing.

The aft fuselage, including the empennage, was buckled approximately nine feet aft of the rear edge of the main cabin door. The aft fuselage was deformed in the forward direction approximately 120 degrees and to the left side approximately 40 degrees. Compression buckling was noted on the upper and left side surfaces where the aft fuselage was buckled forward. The horizontal stabilizer, elevator, vertical stabilizer, and rudder were found inverted above and left of the main cabin. The leading edge of the vertical stabilizer was found adjacent to and in contact with the left side of the aft fuselage. The horizontal stabilizer exhibited buckling along its entire span. The elevator remained attached to the horizontal stabilizer and exhibited buckling along its entire span. Elevator flight control continuity was established from the flight control to the control column tee-bar. The elevator control stops were inspected and no deformation and/or repetitive striking was noted. The rudder remained attached to the vertical stabilizer and had minor impact damage. Rudder flight control continuity was established from the flight control surface to the rudder pedals. The rudder control stops were inspected and no deformation and/or repetitive striking were noted.

The main cabin was destroyed. Both cabin entry doors were found separated from the cabin and were located forward of the main wreckage. The top of the main cabin, including the upper main cabin structural frame and doorframes, was found detached and forward of the main wreckage. The left and right seat frames were bent to the left and forward. The left and right seatbacks were bent to the left and forward. The left seat was positioned and pinned in the third hole from the aft most position on the seat rails. The right seat position was found between the second and third hole from the aft most position on the seat rails. All lap-belt ends remained connected to the airframe. According to initial responders to the accident, the private pilot candidate was seated in the left seat and FAA Designated Examiner was seated in the right seat.

Flight instrument readings:

Airspeed Indicator - 151 knots Directional Gyro - 125 degrees Altimeter - 1,860 feet Altimeter Setting - 30.35 inches-of-mercury Vertical Speed Indicator - +50 feet/minute Engine Tachometer - 2,000 rotations/minute

Additional cockpit documentation can be found in Supplement B, which is attached to this factual report.

The right control column was found in an aft position and was bent upward approximately 22 degrees. The bend began approximately 6.25 inches from the forward end of the control column. The right yoke was bent down and forward. Both yoke handles were bent inward approximately 35 degrees from their normal vertical position. The chain sprocket was found still attached to the forward end of the control column. The nylon guide bushing, part number 68273-02, and O-ring, part number 484 792, were found fragmented on the main cabin floor. A visual inspection of the right control column assembly revealed that Piper Service Bulletin 661 had been complied with.

The left control column was found in an intermediate position and no bending of the control column was noted. The left yoke was bent down and forward. The left yoke handle was bent inward approximately 16 degrees from the normal vertical position. The right yoke handle was bent inward approximately 20 degrees from the normal vertical position. The chain sprocket was found on the main cabin floor, detached from the forward end of the control column. The nylon guide bushing, part number 68273-02, and O-ring, part number 484 792, were found on the control column. Semicircular impressions were noted on the nylon guide bushing. A spacer, part number 67193-02V, was found installed on the left control column. A visual inspection of the left control column assembly revealed that Piper Service Bulletins 661 and 800 had been complied with.

There were multiple indentations and deformations noted on the lower surface of the directional gyro case. These indentations and deformations were consistent with damage caused by the chain sprocket impacting the lower portion of the directional gyro case.

The engine was partially buried in the ground at a measured angle of 35 degrees to the surrounding terrain. The engine remained attached to the engine mounts and to the firewall. The engine was displaced downward and to the right with the firewall impacted up against the engine accessory section. Engine continuity was established by rotating the crankshaft at the propeller flange. Valve train continuity was established and there was compression on all cylinders. Both the left and right magnetos produced spark on all leads when rotated by hand. The engine driven fuel pump was removed from the engine and the fuel pump was found to function when operated by hand. The upper spark plugs were removed and their electrodes were light gray in color.

The propeller remained attached to the engine propeller flange with one blade completely buried in the terrain and the other blade partially buried. The propeller blades had chordwise scratching, burnishing of the blade face and back, and leading edge impact damage. One blade was bent aft at mid-span approximately 90 degrees. The other blade exhibited minor aft bending at the blade tip.

MEDICAL AND PATHOLOGICAL INFORMATION

An autopsy was performed on the private pilot candidate at the Memorial Medical Center, Springfield, Illinois, on September 18, 1999.

A Forensic Toxicology Fatal Accident Report was prepared by the FAA Civil Aeromedical Institute, Oklahoma City, Oklahoma, for the private pilot candidate.

The toxicology results for the private pilot candidate were:

* No Carbon Monoxide detected in Blood * No Cyanide detected in Blood * No Ethanol detected in Urine * No Drugs detected in Liver

An autopsy was performed on the FAA Designated Examiner at the Memorial Medical Center, Springfield, Illinois, on September 18, 1999.

A Forensic Toxicology Fatal Accident Report was prepared by the FAA Civil Aeromedical Institute, Oklahoma City, Oklahoma, for the FAA Designated Examiner.

The toxicology results for the FAA Designated Examiner were:

* No Carbon Monoxide detected in Blood * No Cyanide detected in Blood * No Ethanol detected in Kidney * No Ethanol detected in Muscle * No Drugs detected in Kidney

TESTS AND RESEARCH

Written statements were collected from a sample of individuals who had completed a FAA checkride with the FAA Designated Examiner and none of the prior checkrides included any spin maneuvers.

According to the PA-38-112 maintenance manual, the description and operation of the flight control system was reported as, "The aircraft is controlled in flight by the use of the three standard primary control surfaces, consisting of the ailerons, elevator, and rudder. Operation of these controls is through the movement of the dual control wheels through the control column and dual rudder pedals. The individual surfaces are connected to their control components through the use of cables and push-pull or torque tubes." The maintenance manual describes the elevator control system as, "The elevator controls are also actuated by the movement of the control column. Cables are connected to the lower end of the tee bar and routed around a series of pulleys through the center of the fuselage to the aft end then up through the leading edge of the fin, and terminate at the elevator bellcrank located within the elevator mounted atop the fin. An elevator trim cable is also connected to the bellcrank and routed along with the standard elevator cables to a trim control wheel located between the seats." Copies of the PA-38-112 maintenance manual's description of the control systems and supporting diagrams are attached to this factual report.

According to the PA-38-112 Pilot Operating Handbook (POH), the stalling speed at 1,670 pounds gross weight with power off, with both outboard and inboard flow strips installed, and full flaps is 49 knots indicated airspeed (KIAS). The stall speed is increased by 3 knots when the flaps are in a fully retracted position. The PA-38-112 states, "Loss of altitude during stalls can be as great as 320 feet, depending on configuration and power."

According to the PA-38-112 POH, the airplane is designed for intentional spinning when the flaps are fully retracted and the weight and balance of the airplane is within the utility category limitations. The PA-38-112 POH states in part, "...Spins should only be started at altitudes high enough to recover fully by 4,000 feet AGL [above ground level], so as to provide an adequate margin of safety. A one-turn spin, properly executed, will require 1,000 to 1,500 feet to complete and a six-turn spin will require 2,500 to 3,000 feet to complete... ."

The PA-38-112 POH describes the spin entry, spin recovery, and dive out techniques as follows:

"SPIN ENTRY

Application of full aft control wheel and full rudder before the airplane stalls is not recommended as it results in large changes in pitch attitude during entry and the first turn of the spin. Consequently the initial 2-3 turns of the spin can be more oscillatory than when the spin is entered at the stall.

SPIN RECOVERY

The recommended procedure has been designed to minimize turns and height loss during recovery. If a modified recovery is employed (during which a pause of about 1 second - equivalent to about one half turn of the spin - is introduced between the rudder reaching the stop and moving the control column forward) spin recovery will be achieved with equal certainty. However the time taken for recovery will be delayed by the length of the pause, with corresponding increase in the height lost.

In all spin recoveries the control column should be moved forward briskly, continuing to the forward stop if necessary. This is vitally important because the steep spin attitude may inhibit pilots from moving the control column forward positively.

The immediate effect of applying normal recovery controls may be an appreciable steepening of the nose down attitude and an increase in the rate of spin rotation. This characteristic indicates that the aircraft is recovering from the spin and it is essential to maintain full anti-spin rudder and to continue to move the control wheel forward and maintain it fully forward until the spin stops. The airplane will recover from any point in a spin in not more than one and one half additional turns after normal application of controls.

MISHANDLED RECOVERY

The airplane will recover from mishandled spin entries or recoveries provided the recommended spin recovery procedure is followed. Improper application of recovery controls can increase the number of turns to recover and the resulting altitude loss.

Delay of more than about 1-1/2 turns before moving the control wheel forward may result in the aircraft suddenly entering a very fast, steep spin mode which could disorient a pilot. Recovery will be achieved by briskly moving the control wheel fully forward and holding it there while maintaining full recovery rudder.

If such a spin mode is encountered, the increased rate of rotation may result in the recovery taking more turns than usual after the control column has been moved fully forward.

In certain cases the steep, fast spin mode can develop into a spiral dive in which the rapid rotation continues, but indicated airspeed increases slowly. It is important to recognize this condition. The aircraft is no longer auto-rotating in a spin and the pilot must be ready to centralize the rudder so as to ensure that airspeed does not exceed 103 kt [knots] (Va) [maximum maneuvering speed] with full rudder applied.

DIVE OUT

In most cases spin recovery will occur before the control wheel reaches the fully forward position. The aircraft pitches nose down quickly when the elevator takes effect and, depending on the control column position, it may be necessary to move the column partially back almost immediately to avoid an unnecessarily steep nose down attitude, possible negative "g" forces and excessive loss of altitude.

Because the aircraft recovers from a spin in a very steep nose-down attitude, speed builds up quickly in the dive out. The rudder should be centralized as soon as the spin stops. Delay in centralizing the rudder may result in yaw and 'fish-tailing'. If the rudder is not centralized it would be possible to exceed the maximum maneuver speed (Va) of 103 kt with the surface fully deflected."

The PA-38-112 POH states in section 6.1,

"In order to achieve the performance and flying characteristics which are designed into the airplane, it must be flown within the weight and center of gravity (C.G.) position within the approved operating range (envelope). Although the airplane offers flexibility of loading, it cannot be flown with the maximum number of adult passengers, full fuel tanks and maximum baggage. With the flexibility comes responsibility. The pilot must ensure that the airplane is loaded within the loading envelope before takeoff.

Misloading carries consequences for any aircraft. An overloaded airplane will not take off, climb or cruise as well as a properly loaded one. The heavier the airplane is loaded, the less climb performance it will have.

Center of gravity is a determining factor in flight characteristics. If the C.G. is too far forward in any airplane, it may be difficult to rotate for takeoff or landing. If the C.G. is too far aft, the airplane may rotate prematurely on takeoff or tend to pitch up during climb. Longitudinal stability will be reduced. This can lead to inadvertent stalls and even spins; and spin recovery becomes more difficult as the center of gravity moves aft of the approved limit."

The PA-38-112 POH states in section 6.7,

"IT IS THE RESPONSIBILITY OF THE PILOT AND AIRCRAFT OWNER TO INSURE THAT THE AIRPLANE IS LOADED PROPERLY."

ADDITIONAL INFORMATION

Written statements were collected from several employees of Decatur Aviation, Inc., who were responsible for the servicing of the fixed base operator's (FBO) fleet of aircraft. One employee reported, "Our normal fueling procedure with the company airplanes is to top off the main tanks, unless requested otherwise by the pilot." Another employee stated, "I have never heard of, or seen, or put less than fueling a Tomahawk [PA-38-112] full of fuel."

Parties to the investigation included the Federal Aviation Administration, The New Piper Aircraft Inc., and Textron-Lycoming.

The main wreckage was released to a representative of Decatur Aviation Inc. on September 28, 2000.