1.1 HISTORY OF FLIGHT Use your browsers 'back' function to return to synopsisReturn to Query Page
On July 17, 2004, at an undetermined time after 0830 Pacific daylight time, a Beech 58, N807Q, impacted a glacier at an approximate elevation of 12,270 feet mean sea level (msl) about 17 nautical miles southwest of Bishop, California. The pilot was operating the airplane, registered to a private company, under the provisions of 14 CFR Part 91. The commercial pilot and two passengers sustained fatal injuries; the airplane was destroyed. Visual meteorological conditions prevailed along the general route of flight; however, low level clouds were reported in the mountains. The planned route of flight was from Eastern Sierra Regional Airport, Bishop, to Paso Robles Municipal Airport, Paso Robles, California; the flight departed Bishop at 0815. The flight was scheduled to terminate at Sacramento, California.
The airplane became a subject of an Alert Notice (ALNOT) when the pilot's family notified authorities that the flight had not reached Paso Robles the evening of July 17. At 2235, local authorities initiated a search of the airports that the pilot intended to land. The Air Force Rescue Coordination Center (AFRCC) received the first emergency locator transmitter (ELT) beacon signal by satellite at 1956. The accident site was located on July 18.
A radar plot of the accident airplane's presumed flight track to Bishop, provided by to the Civil Air Patrol (CAP), indicated that the radar track proceeded in a direct line to Bishop and maintained a continuous climb until the last radar return about 11,500 feet msl at 0711, in surrounding terrain that rose to about 12,500 feet msl. At that point, radar contact with the airplane was lost.
According to the co-owner of the airplane, the pilot departed Sequoia Field, Visalia, California, about 0630 on July 17, with a passenger. They flew to Bishop where they picked up an additional passenger. The pilot intended to fly to Paso Robles where he would drop off the Bishop passenger, and then continue to Sacramento.
A friend of the Bishop passenger reported that they met the airplane at the airport at 0800 on July 17. The airplane departed to the south, approximately 0815. After a couple of minutes, the airplane turned westward toward the Sierra Nevada mountains and the accident site. The airplane's arrival had initially been planned for 1200 on July 17; however, on July 16, the pilot notified the passenger that the arrival time had changed to 0800.
The National Transportation Safety Board investigator-in-charge (IIC) plotted a direct route flight from Bishop to Paso Robles using a topographical mapping program. The location of the wreckage site was approximately 1.3 nautical miles southeast of the plotted point-to-point course.
1.2 PERSONNEL INFORMATION
A review of Federal Aviation Administration (FAA) airman records revealed that the pilot held a commercial certificate in airplanes, with ratings for single and multiengine land and instrument. The pilot had private pilot privileges for airplane single engine sea. The pilot held a second-class medical certificate issued on April 27, 2004, with no limitations or waivers.
No personal flight logs were located for the pilot. The pilot reported the following total pilot time hours on his previous FAA medical applications: April 26, 2004 - 6,500 hours; May 6, 2003 - 5,000 hours; March 27, 2002 - 2,500 hours; March 26, 1999 - 21 hours. According to a co-owner of the airplane, the pilot owned the accident airplane about 3 years and acquired 30 to 40 hours of flight time in the airplane during that period.
The pilot attended SimCom, a flight training facility in Scottsdale, Arizona, in June 2003. On the pilot's application dated June 6, 2003, he reported the following flight times: total flight time, 5,280 hours; total multiengine, 3,100 hours; total instrument 315 hours; total hours flown in the past 12 months, 1,474 hours.
1.3 AIRPLANE INFORMATION
The twin-engine airplane was a Beech 58 (Baron), serial number TH-823. A review of maintenance records indicated that the airplane and engine had an annual inspection on May 12, 2004. At that time, the airplane had a total airframe time of 4,093.7 hours.
The right engine (SN 244295-R) had accumulated 499.6 hours since major overhaul. The engine underwent a factory rebuild on February 10, 1988. The left engine (SN 827270-R) was factory rebuilt on August 3, 2002, and was removed and reinstalled prior to the annual inspection for a sudden stop inspection. The left engine had accumulated 77.2 hours since major overhaul.
The right propeller (Hub SN 972299) and the left propeller (Hub SN 972302) were overhauled. Review of the propeller logbooks indicated that the logbooks were labeled opposite of the propeller serial number indications; however, the contents of the logbooks coincided with the propeller installation location on the airplane.
The airplane was last fueled at the Visalia airport on July 16, 2004, with the addition of 104 gallons of fuel. According to the airport manager, the fuel tanks were topped-off.
1.4 METEOROLOGICAL INFORMATION
Using geostationary satellite (GOES-10) infrared and visible satellite imagery, a Safety Board meteorologist observed a layer of stratiform clouds covering the mountains west of Bishop during the time period following the airplane's departure from Bishop. The imagery indicated that the accident site was under the northern edge of the low cloud layer. Based on cloud temperature, the cloud tops rose to 16,000 feet msl. Based on pilot reports and upper air soundings, the cloud bases were between 10,000 and 14,000 feet msl. There was no evidence of turbulence or mountain waves, nor did the specialist observe lenticular clouds.
1.5 WRECKAGE AND IMPACT INFORMATION
The airplane wreckage was located on the southern side of a glacier-fed mountain lake at an approximate elevation of 12,270 feet msl. The approximate global positioning satellite coordinates were 37 degrees 08 minutes 18 seconds north latitude by 118 degrees 38 minutes 34 seconds west longitude. The entire airplane structure was at the accident site.
The glacier was positioned just above a glacier-fed mountain lake. The mountain lake was in a bowl-shaped valley surrounded by higher mountain peaks and ridges. An opening toward downsloping terrain and the greater Owens Valley (location of Bishop) was at the northeastern end of this valley. The valley started about 8 miles west-southwest of Bishop. The glacier sloped approximately 40 degrees downward toward the lake. Except for an area immediately around the wreckage, and just upslope of the airplane, the glacier surface had a hardened, crusty, undisturbed, and uniform appearance. The area of the glacier immediately around and just behind the airplane was noticeably disrupted with a narrow rectangular shaped impression in the ice dimensionally similar to the wing span of the airplane that was roughly parallel to the lateral axis of the wings. In the center of the rectangular shape was an oval impression; the main fuselage mass was just downslope of the oval impression with the empennage still over the impression. There was no evidence of pre- or post-impact fire.
The airplane came to rest upright with the nose pointed down slope in a north-northeasterly direction. The entire forward fuselage structure of the airplane from the nose back past the cabin rear bulkhead was crushed rearward and upward, and had collapsed downward. The right wing tip and green navigational light were displaced from the wing and located nearby. The 3-foot outboard portion of the leading edge of the right wing was crushed aft. The leading edge of the left wing was not crushed. The structure of the fuselage was compromised midsection, from the underside, and upward to the location of the passenger window.
Due to the compromise of the fuselage midsection, the empennage rested on the right elevator. The elevator and rudder control surfaces appeared undamaged and remained secured to their respective attachment points.
The right propeller separated from the engine crankshaft flange, and was just forward of the center point of the wing. The left engine and propeller were inverted under the left wing.
1.6 MEDICAL AND PATHOLOGICAL INFORMATION
The Inyo County Coroner completed an autopsy on the pilot. The FAA Bioaeronautical Sciences Research Laboratory, Okalahoma City, Oklahoma, completed toxicological testing on specimens of the pilot. The results were negative for carbon monoxide, cyanide, volatiles, and tested drugs.
1.7 TESTS AND RESEARCH
The IIC, the FAA inspector, and a representative from Raytheon, a party to the investigation, examined the wreckage on July 28, 2004, after its recovery from the glacier.
1.7.1 Airframe and Flight Controls
The fuselage structure collapsed downward and was vertically compressed. The nose section was crushed aft at a 50-degree angle relative to the longitudinal axis of the airplane. Looking forward from the aft of the airplane, the structure was buckled and distorted to the right. The right wing leading edge was crushed upward from the wing tip, inboard to the engine, and the wing tip separated from the wing. The left wing was intact and displayed vertical crushing on its entire surface. At station 151.00, the fuselage structure was circumferentially buckled and the structure opened. The empennage section remained intact. The 4-inch outboard leading edge section of the right elevator was crushed upward.
The flight control system of the Beech Baron is through combinations of tubes, cables, pushrods, pulleys, and bell cranks. The aileron cables traveled from the control column assembly, forward of the instrument panel, and aft to the forward passenger seating area. These cables then traveled outboard to the wings where they attached to bell crank assemblies. An aileron pushrod extended from the bell cranks to the ailerons. A balance cable ran through the system to complete the loop. The left pushrod extending from the bell crank was slightly bent, but moved when investigators pulled the aileron cable. No control anomalies were noted with the aileron control system.
The elevator cables extended from the control column aft through a series of pulleys and fair leads. The cables ran to the empennage area where they connected with the aft elevator bell crank. Pushrods attach to the bell crank and travel aft, where they connect to the elevators. The rudder cables ran from the left rudder pedals, aft though a series of pulleys and fair leads, and ended with a bell crank in the empennage section. All of the cables remained intact, and investigators noted no evidence of rubbing or chafing on any of the cables, pulleys, or fair leads. The rudder and elevator control surfaces moved when investigators pulled their respective cables.
Investigators measured the trim actuators for the control surfaces. The aileron trim actuator measured 1.25 inches, which corresponded to 4.75 degrees up tab. The elevator trim actuator measured 1.65 inches, which corresponded to a full down position, indicating nose up trim. Although the on-scene measurements of the actuators were not obtained, a review of photographs taken on-scene indicated that the rudder trim tab was deflected slightly to the left. The rudder actuator measurement following the recovery of the airplane coincided with a full right deflection.
The flap actuator measured 2.1 inches, which corresponded to a flap up position.
The airplane has two elevator down tension springs in the empennage, which are installed for certification requirements and for assistance in flight control feel. The right spring was secured; the left spring's aft hook was detached. The general conditions of both springs were the same, with dirt along the lower hook of the springs, and an oily film both on the outer surface of the springs and on the interior surface between the coils. According to the airplane manufacturer, the spring adds stability to the airplane and is primarily only needed in a power-on climb, at low speed, at the aft center of gravity (CG) limit. The loss of the spring would result in a nose-up pitch change. According to the manufacturer, this would not result in a controllability issue, but it may place the airplane out of its regulatory acceptable limits.
1.7.2 Fuel Selectors
The fuel selector valves were examined in the wings to determine which fuel tank was selected. The fuel selector valves are located in the inboard wing area and are controlled through cable linkage to the cockpit fuel selector. The cockpit fuel selectors allow the pilot to position the selectors in the ON, OFF, or CROSSFEED position for each tank. The fuel selector levers in the cockpit indicated that both fuel selectors were in the ON position. The left tank fuel selector valve was to the OFF position at the fuel selector valve. The right tank valve was to the ON position at the fuel selector valve. Investigators noted that the wings had shifted and deformed during the accident sequence, which placed the cable linkage between the cockpit selector lever and the fuel valve under tension.
Investigators tested the electric fuel pumps using a 12-volt charger. Both motors operated when supplied with electricity. The fuel strainers were removed and examined. The left fuel strainer had about a 1/2 teaspoon of sand on the outside of the screen. The right fuel strainer did not contain debris.
Following the post accident examination of the engines, they were sent to Teledyne Continental Motors (TCM), Mobile, Alabama, for further examination and testing. The examinations commenced in the presence of the IIC, the Raytheon representative, and a Teledyne Continental engine representative, also a party to the investigation.
18.104.22.168 Right Engine
The right engine remained attached to the wing; however, the propeller separated at the engine crankshaft flange. The number 1 blade bent 90 degrees aft, 5 inches from the blade root. The number 2 blade had scratches on its cambered side, and the blade twisted slightly toward the tip. The number 3 blade bent aft 5 inches from the root. The propeller spinner was bent and crushed around the hub.
Investigators removed the top spark plugs; they were dark gray colored, which was consistent with a slightly rich operation when compared to a Champion Check-A-Plug chart. All of the electrodes were elliptical in shape and had similar gapping. Thumb compression was obtained on all six cylinders and the valves produced the same amount of lift, in firing order. The ignition leads produced sparks with crankshaft rotation.
The right engine examination occurred at TCM on October 25, 2004. The magneto to engine timing was 23 degrees on the right and 22 degrees on the left. According to TCM, the magneto to engine timing should be between 21-23 degrees. The engine driven fuel pump and coupling were intact; however, the line had ruptured.
TCM removed the fuel screen from the throttle control quadrant. Prior to its removal, they noted that the bolt was safety wired. The screen was clean.
TCM personnel removed the following items from the engine in preparation for an engine run: cooling baffles; crankcase breather tube and associated hoses and clamps; crankcase heater and associated wiring harness; propeller deice brush block and portion of contact ring; propeller governor; tachometer generator; vacuum pump and associated filter; airframe mount; left and right exhaust riser/collector; oil temperature thermocouple; and the number 2 cylinder head temperature thermocouple.
TCM personnel installed the following substitute or repaired parts in preparation for an engine run due to damage sustained during the accident sequence and during the airplane's recovery: Fuel system: fuel pump inlet fitting; hose, fuel pump to fuel control; hose, fuel control to fuel manifold valve; ignition system harness; induction system balance tube; induction system risers (all cylinders); oil sump; oil pickup tube; oil filter cap; induction system (left and right elbows); and all engine mounts.
22.214.171.124 Left Engine
The left engine and propeller remained attached to the wing. The number 1 and number 2 blades bent aft 90 degrees, 6 inches from the blade root. The number 1 blade had scratches along its leading edge. The number 3 blade bent slightly forward. The spinner was intact.
Investigators removed the top spark plugs. Spark plugs from cylinders one, three, and five were light gray in color, which corresponded to a normal operation when compared to a Champion Check-A-Plug chart. Spark plugs from cylinders two, four, and six were rust-colored. Spark plugs electrodes from cylinders one, three, five, and six were circular in shape; spark plug electrodes from cylinders two and four were elliptical in shape. All of the spark plug electrode gapping was similar.
Investigators removed the vacuum pump to assist with the manual rotation of the engine. The lower, right-hand portion of the gasket was not compressed and the nuts were finger-tight on both of the lower bolts. There was no oil seepage in the area, and vacuum pump was otherwise unremarkable. Investigators established mechanical continuity throughout the engine. Manual rotation of the engine through the accessory section produced thumb compression in each cylinder in proper firing order, with accessory gear and valve train continuity established. The ignition leads produced sparks upon crankshaft rotation.
The left engine examination occurred on October 26, 2004, at TCM. The magneto to engine timing was 23 degrees on the right and 22 degrees on the left. The engine driven fuel pump and coupling were intact and the pump manually rotated.
TCM personnel removed the fuel screen from the throttle control quadrant. TCM personnel noted that the bolt had been safety wired prior to its removal and the screen was clean.
TCM personnel removed the following items from the engine in preparation for an engine run: cooling baffles; crankcase breather tube and associated hoses and clamps; crankcase heater and associated wiring harness; propeller de-ice brush block and portion of contact ring; propeller governor; tachometer generator; vacuum pump and associated filter; airframe mount; left and right exhaust riser/collector; oil temperature thermocouple; and the number 2 cylinder head temperature thermocouple.
TCM personnel installed the following substitute or repaired parts by in preparation for an engine run due to damage sustained during the accident sequence: exhaust stacks, all cylinders; fuel system: fuel pump inlet fitting; cover, fuel pump vapor ejector due to a fractured fitting; hose, fuel pump to fuel control; hose, fuel control to fuel manifold valve; induction system: throttle body; balance tube; risers, all cylinders; left and right elbows; all engine mounts; oil sump; oil pick-up tube; and fitting, oil pressure reference.
126.96.36.199 Results of TCM Examination and Test-Run
TCM personnel placed both engines in a test cell and test ran them successfully for approximately 20 minutes. The left engine produced 2,560 rpm with no anomalies noted.
During the initial test-run of the right engine the following discrepancies were noted:
1. A fuel leak at vapor separator cover. Removal of the fuel pump shroud and data plate revealed a crack from the inlet fitting to the front of the vapor separator body. The inlet fitting had been replaced prior to the run due to impact damage. There was no evidence of fuel staining in the area of the crack.
2. An oil leak at front of engine. TCM personnel tightened the propeller governor mount pad oil pressure reference plate.
3. An oil leak at rear of engine. Investigators discovered a crack in crankcase.
The engine was then test run to 2,700 rpm with no operational anomalies noted.
The IIC sent the propellers to the engineering lab of McCauley Propeller, Wichita, Kansas, for examination. The examination commenced on November 15, 2004, in the presence of the IIC. In addition, the Raytheon and McCauley Propellers representatives were present. The representative from McCauley Propellers was also a party to the investigation.
Both propellers were 3-bladed, constant speed models, with feathering capabilities. The manufacturer made the following findings during the propellers examinations:
1. Both propellers' damage resulted from impact. There were no indications of any propeller failures prior to impact.
2. Both propellers were rotating at impact. Neither propeller was at or near the feather position at impact.
3. Both propellers were being operated under low power conditions at impact.
4. While exact blade angles at impact were not determined, impact signature markings indicate both propellers were operating near low pitch at impact.
The Janitrol heater was uncovered in the crushed nose section. The heater can appeared to be intact, with no corrosion present on its outer surface. The fan and exhaust assemblies sustained impact damage. There were no cracks in the combustion chamber and no evidence of exhaust leakage from the system.
1.7.6 Performance Information
The airplane was modified from its type design 2-bladed propellers to 3-bladed propellers. Performance data was calculated using charts from the Beechcraft Baron 58 Pilot Operating Handbook for airplane serial numbers TH-1090 and after. Using the Bishop temperature of 26 degrees Celsius and a standard lapse rate of 2 degrees per 1,000 feet and an estimated weight of 4,500 pounds, performance charts indicated a climb rate of about 500 feet per minute at the accident site elevation.
From the glacier lake at an elevation of 11,600 feet, the accident site rose 620 feet in elevation over a distance of 1,760 feet. Based on the airplane's estimated rate of climb performance and best rate of climb airspeed, a distance of 1.73 miles (10,380 feet), would have been required to climb 500 feet.
The IIC calculated estimated turning performance using the measured widths of the bowl and Figure 2.29, General Turning Performance (Constant Altitude, Steady Turn) from Aerodynamics for Naval Aviators (NAVWEPS 00-80T-80). The approximate width of the entire bowl-shaped canyon was 2,450 feet when measured using a topological mapping program. The terrain immediately surrounding the wreckage measured an approximate width of 1,275 feet. With a turn radius of 1,225 feet and airspeed of 104 knots (best rate of climb), the airplane's bank angle would have been 40 degrees. At 156 knots (maneuvering airspeed), the airplane's bank angle would have been 65 degrees. Using a radius of 637.5 feet (the immediate surrounding terrain) and 104 knots, the bank angle would have been 60 degrees. At 156 knots, the airplane's bank angle would have been about 75 degrees to complete the turn.
According to stall speed versus angle of bank data supplied by Raytheon Aircraft, at the anticipated weight of the airplane for the ambient air conditions and pressure altitude of the accident site, the stall speed would be 85 knots at 40 degrees of bank varying to 105 knots indicated air speed at 60 degrees of bank. The chart does not go above 60 degrees.
1.8 ADDITIONAL INFORMATION
The IIC released the recovered airplane and its components to the owner's representative on November 18, 2004.