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On April 19, 1994, at 1515 central daylight time, an Experimental VK-30, N108P, impacted the terrain during a visual approach to Lake in the Hills Airport, Lake in the Hills, Illinois. The airplane was substantially damaged. The pilot and one passenger were fatally injured. Visual meteorological conditions prevailed. An IFR flight plan had been filed. The airplane was owned and operated by Remote Sensing Services, LTD., McHenry, Illinois. The flight originated in Lake in the Hills, Illinois, at approximately 1430.
Several witnesses to the accident reported they heard no engine noise. One witness reported the engine lost power during the visual approach.
The pilot, age 59, held Airline Transport Pilot certificate No. 1335633 issued on August 13, 1970, with airplane ratings for multiengine land aircraft and commercial pilot privileges for single engine land aircraft. The second class medical certificate was issued to the pilot on September 9, 1993, and contained the limitation that the pilot must wear corrective lenses while exercising the privileges of his airman's certificate. According to the records on file in the Aeromedical Certification Division at the Mike Monroney Aeronautical Center in Oklahoma City, Oklahoma, the airman had accrued 21,000 flight hours. 234 hours were in the Cirrus VK-30.
The Federal Aviation Administration (FAA) aircraft registration certificate for N108P was issued to Remote Sensing Services, Inc., on September 29, 1992. The airplane builder listed on the Special Airworthiness Certificate was Steven C. Baldwin, who was also the passenger in the airplane at the time of the accident.
An hour meter was recovered from the right wheel well area. It indicated 234.2 hours. Representatives of the registered owner, Remote Sensing Services, and Cirrus Corporation indicated this was the actual total flight time of the airplane.
WRECKAGE AND IMPACT INFORMATION
During approach to landing, the airplane impacted the west rim of an active gravel pit, coming to rest on a bearing approximately 100 degrees and 3,600 feet from the approach end of runway 27. Based on ground scars, the wreckage stopped within 50 feet of initial impact on a approximate magnetic heading of 300 degrees. The left wing leading edge was split open outward from the fuselage to the fuel filler opening at the wing tip. The right wing leading edge was split open outward from the fuselage to approximately one half its length. At this point, the wing was fractured chordwise. The remaining outboard section of wing was laying downward against the ground contour of the gravel pit. The nose of the airplane was fractured. The windshield and airplane entrance door separated from the airframe upon impact and came to rest adjacent to the right front side of the wreckage. The empennage was buckled on the right side, just aft of the engine compartment access door.
There was no evidence of any preimpact structural failure.
The instrument panel was bent forward along the lower row of instruments and had sheared the structure which fastened it to the control console. The throttle control cable was broken away from its mount. The throttle, propeller, and mixture controls were pushed to the full forward position.
No apparent damage could be seen to the rudder, elevators, or their respective stabilizers.
The three wood/composite propeller blades were broken at the hub. The wood splinters on the blades and hub were pointing in the opposite direction of the propeller rotation. The propeller was in the low pitch, high RPM position. The blades were recovered in an area approximately 30 feet aft of the airplane's tail and in line with the path of flight. The distance between the first and second ground slash scars was approximately 13 inches, between the second and third scars was approximately 31 inches.
The engine mount was bent downward from impact, with the right upper mount tube suffering a total fracture.
The landing gear position selector switch was in the down position and the landing gear was extended. The main landing gear were bent back. The nose landing gear was bent forward. The wing flap position selector and wing flaps were both in the 15 degree down position.
Control system continuity was established for all flight and engine controls.
Airplane engine and airframe logbooks contained identification numbers for the engine and airframe. Additionally, the airframe record contained an entry dated October 20, 1993, indicating the static, altimeter, and transponder tests were completed in accordance with the applicable regulations. This entry was endorsed by an appropriately rated certified repair station. An unsigned entry on the same date indicated an inspection of the engine fuel pump drive was performed in accordance with Continental Service Bulletin M93-9. Airworthiness Directive (AD) 93-16-15 was not applicable to the fuel pump, by serial number.
According to the date of issue of the Special Airworthiness Certificate on September 29, 1992, the next airplane conformity inspection required in accordance with the Operating Limitations issued on that date would be September 30, 1993. Without any additional sets of maintenance records, it cannot be established if the airplane met the requirements of the certificate issued by completion of the annual conformity inspection.
The engine is a six cylinder, opposed, Teledyne Continental, model TSIO-550-A, serial number 801005. All equipment and accessories installed on the engine were in accordance with the manufacturers recommendations.
The engine mount legs were fractured. Some of the mounting studs were bent or pulled from their threaded area; however, no fractures were visible on the crankcase. The oil sump was dented inward at the rear of the engine.
The engine fuel pump, part number 649368-17, serial number H299002B, was removed from the engine. The drive coupling, part number 631263, was found intact. By part number, AD 93-16-15, and Teledyne Continental Mandatory Service Bulletin M93-9 was not applicable.
A compression test of each cylinder was performed, with the following results:
#1-62/80, #2-68/80, #3-62/80, #4-52/80, #5-60/80, #6-60/80
After the engine test run and the engine had returned to ambient air temperature, cylinders #4 and 6 were retested with the following results:
The spark plugs were tested. All plugs fired evenly at the same test range.
The oil filter had been changed on the same day as this flight. The oil filter container was cut open. No metal particles or unacceptable levels of contamination were evident.
The engine was run in a test cell for an estimated total time of 25 minutes. The engine developed 2,800 RPM and 36 inches of manifold pressure. The fuel pressure was 16.5 PSI. At 2,000 RPM, the magneto drop was 20 RPM on each magneto, and the drop was smooth. The oil pressure was 45 PSI.
During the running of the engine, the airplane electric fuel pump, Dukes, part number EXPERIMENTAL, serial number -none-, was turned on, both in the high and low position, at various power settings. The fuel pressure did not raise in excess of .5 PSI, and engine performance did not change.
All monitored systems on the engine performed satisfactory during the test running.
The propeller clutch is a dry clutch, filled with stainless steel shot, which "locks" the drive ring to the hub. The propeller drive shaft is inserted into this clutch assemble. Disassembly of the clutch was performed. No slippage was evident.
The propeller drive shaft was intact from the clutch assembly mounted to the engine crankshaft flange, to the propeller hub. The rear drive shaft bearing and support showed no visible signs of previous damage or impending failure.
The fuel system was modified from the manufacturers specifications. The manufacturers configuration consists of a wet wing fuel system with a filler point at each wing tip. The wing is divided into two separate fuel tanks by a solid wing rib at the center of the wing structure. Each tank had a capacity of 53 gallons, 52.5 gallons usable. A fuel sump and fuel pick-up point is added by the builder to the lower side of the wing, on either side of this rib. The fuel flows to a three position selector valve (LEFT-RIGHT-OFF), through an airframe auxiliary fuel pump, and to the engine fuel system.
The accident airplane fuel system was modified in the following manner. A fuel tank of approximately 35 gallons was bonded to the top of the wing in the fuselage cabin areas. Two holes, approximately 28 square inches each, were cut in the bottom of this tank and top of the wing tank, over respective locations which would have allowed fuel to drain into each wing tank. The center dividing tank wing rib was partially removed, in effect, making a single 140 gallon fuel tank, allowing uninterrupted fuel flow between the tanks. A submersible fuel pump was placed in the right wing sump area, and was plumbed to a single pick-up point in the left sump. From this point, the fuel went through a two position (ON-OFF) fuel selector to the auxiliary fuel pump and to the engine fuel system.
The modified fuel system could be filled from either wing tip, and fuel would seek its own level in the modified tank area. A pencil change in the Pilots Operating Handbook, Section 3- Systems, indicated this change to the fuel system.
No devices, such as hinged flapper valves, were present in the fuel tanks, which would allow fuel to enter the sump area, but prevent rapid exiting of the fuel from the sump area.
MEDICAL AND PATHOLOGICAL INFORMATION
Autopsy of the pilot was conducted by the Cook County Coroner's Office, Cook County, Illinois. Toxicological tests were negative for all tests conducted.
The interior of the airplane sustained substantial damage. The carry through wing assembly leading edge was split open across the entire cabin width. A large aerial camera and mount, which was installed immediately aft of the left front pilot seat, was ripped from the airplane floor. The camera viewfinder had broken loose from its mount and was resting on the pilot's right leg. The pilot's shoulder harness and mount were torn from the cabin ceiling. The pilot and rear passenger seats were collapsed to the cabin floor. The passenger seat, located in the right rear portion of the cabin, was torn from its structure. The pilot and passenger seat belts were fastened to their respective locations on the floor. Both occupants used their seat belts.
The pilot used the only installed shoulder harness. The cabin floor was broken loose, however, the composite fibers kept it in place.
TESTS AND RESEARCH
Approximately one month before the accident, the president of Cirrus Design, the kit manufacturer, had verbally requested the owners of Cirrus airplanes voluntarily cease flying their airplanes. That request stemmed from questions about the strength of materials used in the wing kits.
Available fueling records indicate the airplane was last refueled at Benton Harbor, Michigan, on April 17, 1994, at 1541 local time with the addition of 40 gallons of aviation fuel. The straight line distance from Benton Harbor, across Lake Michigan to Lake in the Hills, is 85 nautical miles. The distance from Benton Harbor, around the southern end of Lake Michigan, to Lake in the Hills is approximately 135 miles.
The total elapsed flight time from the last fueling until the accident could not be determined from existing records.
The airplane weight and balance records were reviewed, however the actual weight of the airplane at the time of the accident could not be established. On September 23, 1992, six days before certification, a weight and balance form indicated a total empty weight of 2,729 pounds and the center of gravity to be 151.28 inches. Comments on this form indicate no fuel, no oxygen, only required flight instruments, no speed brakes, no seat belts. At the bottom of that page was a notation of weights with 02 (oxygen) and Inst (instruments), and a weight increase of 76 pounds, resulting in the airplane empty weight of 2,805 pounds. The airplane's four oxygen bottles were weighed, and totaled 44 pounds.
The camera and related equipment were removed and weighed. Following are the weight and balance calculations based on the airplane stations listed in the pilots operating handbook and the weight and balance report:
ITEM WEIGHT ARM
Aircraft empty weight 2,805
Camera 258 124.0
Viewfinder 65 89.5
Pilot 170 89.5
Passenger 170 124.0
Total weight 3,468
Based on the above calculations, 132 pounds of fuel, or twenty two gallons, would bring the airplane to its maximum allowable takeoff weight of 3,600 pounds.
Using the above loading calculations, it was determined the airplane was most likely operating within its center of gravity (CG) range when the accident occurred.
The wreckage was released to the representatives of the owner on May 20, 1994, via NTSB Form 8020.15.