N12508

HISTORY OF FLIGHT

On March 18, 1997, approximately 1332 central standard time, a Douglas DC-9-32, N12508, operated by Continental Airlines (COA), under Title 14 CFR Part 121, as scheduled domestic passenger flight 1682, from Houston Intercontinental Airport (IAH), Houston, Texas, to Indianapolis, Indiana, experienced a total loss of power in the No. 2 (right) engine during initial takeoff climb. The flight returned to IAH and landed without further incident at 1355. Emergency crews extinguished a fire in the No. 2 engine with foam after the airplane stopped on the runway. There was no emergency evacuation, and there were no injuries to the 96 passengers or the 5 crewmembers. The airplane sustained minor damage. Instrument meteorological conditions prevailed, and an IFR flight plan was filed.

In written statements provided to the NTSB investigator-in-charge (IIC), the captain and the first officer of flight 1682 reported that no discrepancies were noted during completion of the pre-flight checklists. The captain stated that power indications for the takeoff on runway 14L were "normal" until "just after liftoff" when he heard "a sound similar to a compressor stall," the airplane yawed, and the No. 2 engine readings "dropped." The first officer stated that "shortly after rotation, [he] heard a loud pop with subsequent power loss in the right (#2) engine." Both crewmembers reported that the first officer, who was the pilot flying, "flew the engine out profile." According to the captain, the airplane "had a problem climbing," which he believed was "due to the drag of a large piece of cowling hanging from the engine." The first officer stated that the airplane was "very close" to maximum gross takeoff weight and "very sluggish to climb."

According to FAA personnel at the Houston Intercontinental Air Traffic Control Tower, flight 1682 was cleared for takeoff at 1330 and, at 1332, declared an emergency and requested a return to IAH for landing. At 1337, the captain advised the controller that the No. 2 engine had failed, and at 1340, he further advised that the right hydraulic system had also failed. After the flight was vectored around thunderstorms in the area, the first officer executed a single-engine approach to runway 27, landing at 1355. According to the captain, "at no time" during the flight was there a cockpit fire warning, and the crew was not aware of the No. 2 engine fire until the airplane was stopped on the runway and contact was made "with the fire captain by radio who advised the fire was out and not to evacuate."

FLIGHT RECORDERS

The cockpit voice recorder (CVR) was not readout. At the request of the NTSB IIC, the Digital Flight Data Recorder (DFDR) was readout by COA. The tabular data indicated that the takeoff roll began at FDR count 17913. During the next 45 seconds (17914 to 17959), the engine pressure ratio (EPR) of both engines increased to and stabilized at 2.10, and the indicated airspeed increased to 151 knots. At FDR count 17959, the pitch began to increase. Between FDR count 17959 and 17965 (6 seconds), the pitch increased from 2 to 10 degrees nose up, and the indicated airspeed reached 165 knots. At FDR count 17966, the No. 2 engine EPR dropped to 1.63, and the pitch reached 11 degrees nose up. At FDR count 17967, the No. 2 engine EPR further dropped to 1.11, the pitch reached 12 degrees nose up, and the altitude began to increase. During the next 60 seconds (17968 to 18028), the airplane climbed to 636 feet MSL. Using the IAH field elevation of 98 feet MSL, an initial takeoff rate of climb of 538 feet per minute was calculated for the airplane.

TESTS AND RESEARCH

Following the incident, the No. 2 engine, a Pratt & Whitney (P&W) JT8D-15, serial number 701016, was removed from the airplane and shipped to Greenwich Air Services, Dallas, Texas. On April 3, 1997, it was examined at Greenwich by representatives from the NTSB, FAA, P&W, and COA. The engine was not disassembled. For a detailed description of the damage to the engine refer to the Powerplants Group Chairman's Factual Report.

The exterior of the thrust reverser was sooted with most of the soot concentrated on the upper thrust reverser bucket. The exterior of the fan exhaust outer duct was discolored blue, and there was approximately half a shovel full of loose, solidified molten metal and debris at the bottom of the duct. A mound of solidified molten metal and debris (approximately 3 to 4 shovels full) lined the lower portion of the turbine nozzle rear case with airfoil segments and fir tree roots visible in the debris. All of the 2nd, 3rd, and 4th-stage low pressure turbine blades and vanes, as well as almost all of the low pressure turbine disk rims, were burned away. All of the 1st-stage high pressure turbine blades and almost all of the 1st-stage high pressure turbine disk rim were burned away. All of the 1st-stage vane trailing edges were burned with all the remaining vane material blackened and granular.

Forward of the turbine section, there was no visible heat damage or soot on the exterior of the fan ducts, wire bundles, or engine components. The terminal ends of one fire detector loop were separated on each side of the J-flange support bracket at the 5:30 clock position. (All references to the clock are as viewed from aft looking forward.) The terminal ends of the second fire detector loop, which shared the same bracket, were intact. The hydraulic quick-connect junction-box from the 3 o'clock position was hanging underneath the engine. The flexible hydraulic lines from the hydraulic pump to the junction-box were intact. One of the rigid hydraulic lines to the junction-box at the 3 o'clock position and two of the rigid hydraulic lines to the junction-box at the 9 o'clock position were severed at the B-nuts.

Between H- and J-flanges, the diffuser outer duct was fractured from the 2 to 8:30 clock locations. The crack extended circumferentially around the duct, a maximum of 5 inches forward of the J-flange. Between J- and K-flanges, the combustion chamber outer upper fan duct was circumferentially torn from the 2 to 3 o'clock location, approximately 11 inches aft of the J-flange. The combustion chamber outer lower fan duct was fractured from 6 to 9 o'clock. The fracture extended circumferentially around the duct then traversed forward, across J-flange, into the diffuser outer fan duct at approximately the 8:30 clock location. Additionally, the lower fan duct was missing the forward-most 26 inches (approximately 40 percent of the duct's length) from the 3 to 6 o'clock location.

The combustion chamber outer case (CCOC) was axially cracked along the 6 o'clock location from the rear flange to the weld seam located 1 1/2 inches aft of the forward flange. The axial crack extended around the right side of both the front and rear drain bosses in the vicinity of their respective weld seams. At the weld seam aft of the forward flange, the crack split in two directions, passing circumferentially around the CCOC adjacent to the forward weld seam. The circumferential cracks on the right and left sides continued to approximately the 2 o'clock and 8 o'clock locations respectively.

Visual examination of the CCOC crack in the vicinity of the forward drain boss revealed approximately 3.5 inches of reddish (rust-colored) discoloration along the fracture surface. The discoloration extended partially around the circumference of the boss-to-case weld seam and axially aft approximately 1.25 inches. The two adjacent fracture surfaces in the vicinity of the forward drain boss and a 10 inch segment of the CCOC forward flange, which included the split-point at the 6 o'clock location, were machine-cut from the CCOC. The three segments were hand-carried to the NTSB Materials Laboratory, Washington, DC, for metallurgical examination. For a detailed description of the metallurgical findings refer to the Metallurgist's Factual Report.

According to the NTSB metallurgist, scanning electron microscope examination of the fractures adjacent to the forward drain boss revealed "features characteristic of a fatigue crack that emanated from multiple origins on the inside diameter surface" in the area where the reddish discoloration (identified as an oxide layer) was noted. Several origins were located at corrosion pits on the inside diameter surface. The metallurgist reported that fatigue propagation was from the inside to the outside diameter surface and was through the wall thickness for a distance of approximately 0.5 inch. He further reported that the depth of the fatigue crack decreased with increasing distance extending away from the through-the-wall portion of the crack, and that the areas of the fracture outside the fatigue region "contained features typical of overstress separation."

The metallurgist performed a striation count at one end of the through-the-wall portion of the fatigue crack. The accumulated striations from the inside diameter (fatigue origin) to a point .004 inches from the outside diameter (total distance .069 inches) totaled 6,299. The accumulated striations from the inside diameter to the outside diameter (total distance .073 inches) could not be determined due to mechanical damage to the outside diameter edge.

ADDITIONAL INFORMATION

Review of maintenance records by representatives from the NTSB, FAA, COA, P&W, and the Independent Association of Continental Pilots (IACP) revealed that the last overhaul of engine serial number 701016 was performed by COA at their engine buildup facility in Los Angeles, California, (LAX EBU) on August 31, 1994. At the time of the incident, the engine had accumulated 4,665 cycles and 5,462 hours since this overhaul.

The maintenance records indicated that the combustion chamber outer case (CCOC) installed in the engine was a part number 767197 bearing serial number KU-4547. Visual examination of the CCOC revealed that the serial nu...