Pacoima, CA, USA
N160LA
Sikorsky S-70A
The twin-engine helicopter's number 2 engine experienced an uncontained engine failure and liberated debris punctured the helicopter's firewall while in cruise flight. The pilot made an uneventful landing at a nearby airport. Examination of the engine revealed segments of the stage 2 aft cooling plate (S2ACP) exited the power turbine case. Further investigation determined that the gas generator rotor's forward curvic seal was not installed during the operator's replacement of the gas generator turbine (GGT) module 92 hours prior to the event. Manufacturer's analysis of the missing curvic seal revealed it would have caused the reversal of the secondary airflow in the B-sump, which eventually led to auto ignition of the B-sump oil and exposure of the B-sump hardware to elevated temperatures. The elevated temperatures resulted in the separation of several B-sump components, including the separation of the compressor shaft. The compressor shaft separation resulted in the GGT rotor to rub into the stator and extensive secondary damage, including the separation of the S2ACP. The GGT module was replaced 209 hours prior to the event. According to the manufacturer, the forward curvic seal is part of the engine assembly rather than part of either the GGT rotor assembly or the high pressure compressor rotor assembly. According to General Electric (GE), if the curvic seal was removed from the engine during removal of the previous GGT rotor-stator assembly, and either discarded, or returned to GE's service facility with the rotor, a new seal would not have been included with the new GGT rotor-stator assembly, and the GGT could have been reinstalled without it. The manufacturer also reported that the maintenance manuals did not call for the verification of curvic seal installation during GGT replacement. This military engine was not a Federal Aviation Administration (FAA) certificated engine, but did share similar design features of other FAA certificated engines. The manufacturer reported that the civilian certificated engine maintenance manuals do call for the verification of curvic seal installation following a GGT replacement. The engine was removed from the event helicopter 92 hours prior to the event and was shipped to an engine service center for an internal leak at the number 3 labyrinth seal (located in the front frame of the engine). The service center repaired the inlet guide vanes, which, according to the manufacturer, was a typical repair when oil is observed leaking in the flow path in front of the compressor. However, their analysis of the missing curvic seal indicates that it too would have resulted in an oil leak in the front frame.
HISTORY OF FLIGHT On August 30, 2004, approximately 1240 Pacific daylight time, a Sikorsky S-70A twin-engine helicopter, N160LA, experienced an uncontained number 2 engine failure upon reaching cruise flight. The helicopter landed uneventfully with minor damage at Whiteman Airport, Pacoima, California. The airline transport pilot and two paramedics were not injured. The helicopter was registered to, and operated by, the County of Los Angeles Fire Department (LACoFD) as a public-use flight under the provisions of 14 CFR Part 91. The helicopter departed an unimproved landing area approximately 5 minutes prior to the event, and was destined for Palmdale, California. Day visual meteorological conditions prevailed, and a company visual flight rules flight plan had been filed for the flight. According to the LACoFD pilot, the helicopter departed on its second flight of the day at an operating weight of 16,900 pounds. It departed the landing area, which was located on top of a mountain at the 3,700-foot level. The observed temperature was 90 degrees Fahrenheit. Upon reaching 100 knots with 80 percent torque, the pilot heard a loud bang followed by a slight yaw variation. The pilot observed that the number 2 engine parameters were indicating a power loss. The pilot performed the emergency procedures, and diverted to the Whiteman Airport where he performed a run-on landing. The helicopter's number 2 engine (serial number 763397) sustained damage from uncontained engine debris at its 8 o'clock position, and the helicopter's titanium firewall was punctured in two locations. One piece of uncontained debris punctured the firewall within 1.5 inches from a fuel line. AIRCRAFT INFORMATION Engine Information The incident helicopter was a public-use version of Sikorsky's Black Hawk helicopter. It utilized two General Electric (GE) T700-701C turboshaft engines, each rated at 1,890 shaft horsepower. The T700-701C engine is not a Federal Aviation Administration (FAA) certificated engine, but shares numerous similarities with the CT-7 turboshaft/turboprop engines, such as the ones utilized on the SAAB 340 transport category turboprop airplanes. The turboshaft engine consists of four modules; the cold section, hot section, power turbine section, and accessory section. The cold section module includes an inlet particle separator and variable inlet guide vanes (IGV) and variable stage 1 and stage 2 vanes. The cold section also includes the compressor section, which utilizes a five-stage axial low-pressure compressor (LPC) and one centrifugal stage high-pressure compressor (HPC). The hot section module consists of three subassemblies; the gas generator turbine (GGT), stage 1 nozzle assembly, and combustion liner. The GGT consists of a gas generator stator assembly and a two-stage, air-cooled turbine rotor assembly, which drives the compressor and the accessory gearbox. The nozzle assemblies direct gas flow to the gas generator turbine. The power turbine (PT) section module includes a two-stage power turbine, an exhaust frame, and the shaft and C-sump assembly. The PT rotors drive the input to the main rotor transmission. Engine History Engine serial number (SN) 763397 was shipped from GE's manufacturing facility located in Lynn, Massachusetts, in July 1998, and entered operation with the LACoFD in 2000, when it was installed in the number 2 engine position of a sister ship with zero hours since new. On July 20, 2003, at an engine total time of 1,077 hours, LACoFD maintenance personnel removed and replaced the GGT after a temperature exceedance during an overspeed check. The replacement GGT was new. An endorsement in the engine maintenance logbook dated March 1, 2004, (engine total time of 1,284 hours) indicated that the GGT was removed from SN 763397 after 207 hours of operation due to low performance and returned to GE's engine service center. On March 19, 2004, (at an engine total time of 1,284 hours) LACoFD personnel installed another GGT, which had been received from GE in November 2003. This rotor had 850 hours on the GGT disks and stator assemblies, 1,077 hours on the stage 1 aft cooling plate (S1ACP) and stage 2 forward cooling plate (S2FCP), and a new stage 1 forward cooling plate (S1FCP) and stage 2 aft cooling plate (S2ACP). On June 8, 2004, the number 2 engine was removed from the sister ship at an engine total time of 1,401 hours, and shipped to GE's engine service center due to a front frame oil leak (number 3 labyrinth seal). The service center removed the front frame and repaired the IGV, which according to GE is a typical repair when oil is observed leaking in the flow path in front of the compressor. The service center did not expose or remove any of the hot section components. The engine was returned to the LACoFD and installed in the number 2 position of N160LA (the event helicopter). Maintenance personnel conducted an engine performance calibration. The performance check showed a low, but acceptable performance margin. LACoFD personnel began keeping an oil consumption log on August 19, 2004, after noting an increase in oil consumption. The following is a listing of that oil consumption log: - 08/19/2004; Aircraft total time 1,048 hrs; 38 oz added - 08/24/2004; Aircraft total time 1,062 hrs; 22 oz added; (estimated oil consumption per engine hour 36 cubic centimeters - cc) - 08/28/2004; Aircraft total time 1,066 hrs; 22 oz added; (estimated oil consumption per engine hour 124 cc) - 08/29/2004; Aircraft total time 1,069 hrs; 24 oz added; (estimated oil consumption per engine hour 180 cc) - 08/30/2004; Event occurred at an Aircraft total time of 1,072 hrs According to GE, the maximum limit for oil consumption is 133 cc/hr. At the time of the event, the number 2 engine accumulated a total of 1,493 hours, and had not yet undergone an overhaul. WRECKAGE AND IMPACT INFORMATION Engine Examination The engine, along with the debris liberated from the engine case, was shipped to GE Transportation's Development Assembly Building 42, Lynn, Massachusetts, where it underwent a teardown examination under the supervision of FAA Engine Certification Office personnel. The following information was gleaned from the GE and FAA teardown examination reports. Damage on the engine is referenced via a clock face when viewed from the front of the engine looking aft. The PT case was punctured about the 9:00 o'clock position just forward of the T4.5 harness. About the 3:30 position, the case outside diameter was also bulged outward, but not punctured. Extensive damage was observed on all turbine blades, vanes, T4.5 probes, shrouds, honeycomb seals and seal teeth. The stage 3 disk forward face and PT shaft bolts and nuts displayed heavy rubs and wear damage. Large amounts of debris were recovered from the cavity between the GGT and PT rotors, as well as in the transition duct. The PT shaft was cut in two after it became trapped in the engine's B-sump/mid-frame area during PT module removal. The fragments that exited the engine case were similar in geometry to sections of the S2ACP (normally located in the GGT section). The S2ACP was found with a 230-degree section of the web completely separated above its rabbet. The remaining circumference of the cooling plate was separated just below the seal wire-retaining groove. Heavy rub damage was noted on the S2ACP seal teeth and aft side of the remaining web material. The tiebolts, which provide axial retention of the stage 1 and 2 assemblies, were rubbed and worn. The S2FCP seal teeth were heavily rubbed and broken. The stage 2 turbine disk displayed significant rubs on the aft face of the bore, and deposited metal and light rubs on the aft face (adjacent to the area where the cooling plate rim was missing). The stage 2 nozzle assembly had heavy rubs on the forward face, and the heat shield was completely destroyed, with only debris remaining. The S1ACP was heavily rubbed on the aft side, and the rim section above the hammerhead was missing 360 degrees. When the GGT rotor was removed from the compressor rotor, the forward curvic seal was not found attached to either the stage 1 disk, or the GGT shaft. A search throughout the remaining engine was unable to locate either the missing seal, or debris that might have been the remnants of a broken seal. Two identical curvic seals are installed under the stage 1 disk-to-GGT shaft curvic and below the stage 2 disk-to-stage 1 disk curvic. Both seals prevent secondary T3 cooling air from leaking through the spaces between the curvic teeth. The diameter below the stage 1 disk forward curvic teeth and GGT shaft curvic teeth displayed the same uniform dirty and discolored appearance as the remainder of the part. For comparison, the diameter below the stage 2 disk curvics (where the curvic seal was found properly installed) had a sharp scalloped pattern of discoloration that matched the curvic seal's scalloped pattern. Coking was observed in the B-sump forward buffer cavity, the B-sump drain line, the line running through the B-sump, and the external tube and fitting. Pressure test of the B-sump drain line revealed that it was mostly clogged. The HPC shaft was severed between the forward compressor discharge pressure (CDP) seal pilot and the forward air holes. The CDP static and rotating seals were also severed with only the aft most tooth remaining (the middle rub land on the static seal was completely destroyed and the middle set of the rotating seal teeth were mostly destroyed). In addition, the inside diameter of the forward section of the CDP seal displayed heavy rub damage and smeared metal. The location of the CDP damage is directly above the location where the HPC shaft was severed. The HPC exhibited heavy rubs on the blisk airfoil tips, vanes and impeller. The bores of the stage 1-4 blisks and the vortex spoiler all had coking and oil varnish (none of these surfaces are exposed to oil during normal operation). The number 3 labyrinth seal was heavily coked and the IGVs and scroll case were coated with oil sludge and coked oil. During teardown, pooled oil was observed in the front frame. All of the oil sump screens were clean except for the B-sump screen, which contained a piece of non-metallic debris. All of the oil quantity levels from each sump were typical of normal operational levels. Metallurgical examinations of the following components were conducted at GE's facility with National Transporation Safety Board review and approval: HPC shaft, rotating CDP seal, static CDP seals, B-sump hardware, and GGT cooling plates. In addition, oil samples from the various engine sumps were analyzed. The following is a summary of those examinations: Scanning electron microscope (SEM) examination of the HPC shaft's aft fracture surface displayed an intergranular surface. Metallography of shaft sections revealed significant grain growth and hardness measurements, which showed a significant reduction in hardness, indicating the CDP seal operated above 1,800 degrees Fahrenheit (F). The areas of reduced hardness on the HPC shaft were localized to within 0.1 inch of either side of the fracture surface. The fracture surface on the forward section of the rotating CDP seal displayed multiple cracks around the circumference, and the entire aft end of the forward section was deflected outward (normally this section of the part would taper inward). The aft section of the CDP seal was locked together with the number 4 static oil seal and could not be disassembled due to axial interference. The number 4 oil seal was cut in two locations and removed. The fracture surface on the aft section of the CDP seal appeared similar to the forward section, with multiple radial cracks around the circumference. Metallography from both the forward and aft sections of the CDP seal showed significant grain growth and reduction in hardness values, indicating that the CDP seal was exposed to operational temperatures in excess of 1,800 degrees F. The static CDP seal was sectioned at two locations and examined via SEM. Both sections displayed significant softening and grain boundary separation due to over-temperature exposure. The S1ACP was fractured at the rim with multiple radial cracks around the circumference. Heavy rubs were observed on the remainder of the rim adjacent to the fracture surface. SEM examination of the fracture surface found non-oxidized intergranular features indicating stress rupture. Metallographic examination of a radial section revealed an area of coarse grains along the cooling plate rim, just below the rub. The S2ACP fractures (one which was flat and extended over 230 degrees of the circumference, and one which was at a 45-degree angle to the web, just below the seal wire retaining groove) were examined under SEM. Heavy rubs were observed near each fracture area. The flat fracture surface displayed non-oxidized intergranular features indicative of a stress rupture. The 45-degree fracture surface exhibited fractographic features, indicative of shear tensile overload. Microstructure and hardness values away from the rim and fracture surfaces for both plates were consistent with the manufacturer's specifications. Oil samples were determined to be Royco 500 synthetic aviation oil, with no trace of oil from any other manufacturers or dissolved contaminants. Lower boiling components that are present in new oil were missing from all samples, indicating exposure to elevated temperatures above typical engine operation. All samples contained high concentrations of particulate matter (most common being copper and zinc). The external B-sump drain tube was sectioned. The heaviest blockage was found on the forward (B-sump) end of the tube, but the tube was not completely blocked. Most of the material in the tube was carbon residue consistent with coked oil. TESTS AND RESEARCH According to GE, with the forward curvic seal installed, P25 air from the vortex spoiler flows aft, passing through the holes in the impeller, HPC shaft and CDP rotating seal. This air pressurizes the B-sump forward buffer cavity, and flows further aft over the bearing cage, surrounding the B-sump and pressurizing the #4 bearing with relatively cool T25 air. Oil from the B-sump is removed through the B-sump oil scavenge line, while the air vents inwards from the B-sump through the holes in the GGT shaft and moves both forward and aft outside of the PT shaft. T3 air from the accelerator flows either into the GGT, or across the IBP seal and through the aft holes in the GGT shaft, joining the T25 air underneath the GGT. When an engine operates without the forward curvic seal, T3 air normally flows into the GGT passes through the forward curvic, and into the cavity under the GGT, increasing the pressure. The higher pressure air overpowers the T25 air, and goes through both sets of holes in the GGT shaft. This air enters the B-sump and flows forward into the B-sump forward buffer cavity. This allows oil to leave the sump and enter the B-sump forward buffer cavity. Oil and T3 air continue forward through the CDP seal, HPC shaft, and flow between the HPC tie rod and blisk bores, into the front frame and into the flow path. It should be noted that the LACoFD report of a front frame oil leak is consistent with the secondary flow analysis of engine operation without a curvic seal. Without the forward curvic seal installed, the air in the B-sump forward buffer cavity is not hot enough to ignite oil spray. Oil is normally prevented from pooling in the B-sump forward buffer cavity by the 6:00 o'clock B-sump drain line, which drains to the engine exhaust frame. If that line should coke shut, or be prevented from draining, the B-sump wall metal temperature is sufficient to cause auto ignition of pooled oil in the B-sump forward buffer cavity. Because of the missing curvic seal, T3 air, which normally flows into the GGT, instead flows into the B-sump. This change causes a reduction in stage 1 blade cooling flow, resulting in an increase in blade metal temperature. This effect would cause increased oxidation of the stage 1 blades, eventually resulting in
the failure of the operator's mechanics to install the turboshaft engine's forward curvic seal during a gas generator turbine (GGT) replacement. Contributing factors to the incident were the failure of the manufacturer to provide a forward curvic seal with the replacement GGT, and the lack of a curvic seal installation verification inspection in the maintenance manuals following the GGT installation.
Source: NTSB Aviation Accident Database
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