NTSB Factual Report

History of Flight

On April 10, 2015, about 1447 Pacific daylight time, a Boeing 707-338C, N624RH, operated by Omega Aerial Refueling Services, Inc., experienced a nacelle uncontained release of turbine material from the No. 1 engine, a Pratt & Whitney (P&W) JT3D-3B turbofan engine, during the climb after departing from the Pt. Mugu Naval Air Station (NTD), Port Hueneme, California. During the takeoff roll after the airplane had accelerated past V1, the No. 1 engine's exhaust gas temperature (EGT) started to increase towards the red line limit. The flight engineer had to retard the No. 1 engine's throttle several times to keep the EGT below the red line limit during the takeoff roll and during the climb. As the airplane was climbing through 17,000 feet, the No. 1 engine's low oil pressure light illuminated. The crew stopped the climb at flight level 210 to reduce the power of the No. 1 engine. In accordance with the checklist procedures, the Captain retarded the No. 1 engine's throttle to idle. While the throttle was being retarded, a vibration began that became severe within a few seconds. The pilots shutdown the No. 1 engine, declared an emergency, and requested to divert to the Victorville Airport (VCV), Victorville, California. The mechanics who were on board the airplane could see the No. 1 engine vibrating on the pylon as well as smoke and debris trailing out of the engine's exhaust. The airplane landed at VCV without further incident. The post-landing examination of the No. 1 engine revealed a hole in the bottom of the nacelle and several holes in the turbine exhaust case. The airplane was operating on an instrument flight rules flight plane under the provisions of 14 Code of Federal Regulations (CFR) Part 91 from NTD to Brunswick Golden Isles Airport (BQK), Brunswick, Georgia. The airplane was being ferried from NTD to BQK to replace the No. 1 engine because of excessive oil consumption.

Injuries to Persons

There were no reported injuries to the two pilots, the flight engineer, and the two mechanics on board the airplane.

Damage to Airplane

There was a hole in the bottom of the No. 1 engine's nacelle and there was a small hole in the underside of the left outboard aileron trim tab. The remainder of the airplane did not have any damage.

Other Damage

No other damage was reported.

Fire

There was no fire damage.

Airplane information

The airplane was a Boeing 707-338C, serial number (SN) 19624, registered as N624RH, owned by Omni Air International, San Antonio, Texas and operated by Omega Aerial Refueling Services, Inc, Alexandria, Virginia. The Boeing 707 is a four-engine transport category airplane. The incident airplane was operating on an experimental airworthiness certificate because it had been modified for air-to-air in-flight refueling of military airplanes. The airplane has a maximum takeoff gross weight of 333,600 pounds and the airplane's takeoff gross weight for the incident flight was 286,900 pounds. According to Omega's records, the airplane had been manufactured in 1968 and at the time of the incident, had accumulated 52,852 hours since new.

The No. 1 engine was a JT3D-3B, SN 667883. The JT3D-3B is a dual-spool, axial flow, low-bypass turbofan that features a two-stage fan, a six-stage low-pressure compressor (LPC), a seven-stage high-pressure compressor (HPC), a can-annular combustor with eight fuel nozzle assemblies and combustion chambers, a single-stage high-pressure turbine that drives the HPC, and a three-stage low-pressure turbine (LPT) that drives the fan and LPC. The JT3D-3B engine has a takeoff thrust rating of 18,000 pounds that is flat-rated to 84 degrees F. According to Omega's records, at the time of the incident, the engine had accumulated 49,599 hours and 15,340 cycles since new. The records show that the engine was last repaired by Aero Engines Ireland, Ltd. (AEI), Dublin, Ireland in March 2011, when it underwent a hot section inspection (HSI). The engine had been removed from service in June 2010 for compressor stalls and sent to AEI for repair. According to the maintenance records, the 10th and 13th stage compressor disks as well as the 1st stage turbine blades were replaced during the HSI. The engine was installed on the incident airplane in December 2013, and had accumulated 985 hours and 284 cycles since it was installed.

The review of the airplane's maintenance records for the 30 days prior to the incident that involved 23 flights and about 60 hours of flying showed that the No. 1 engine had to serviced with 49 quarts of oil. In comparison, combined total of oil added to the other three engines was 19 quarts.

Prior to the incident flight, the airplane had been conducting in-flight refueling operations with military airplanes making it a public-use aircraft. Prior to the incident flight, the airplane's logbook was annotated to show that the airplane was removed from the public-use category to permit the airplane to be operated under 14 CFR Part 91 for the repositioning flight from NTD to BQK.

Tests and Research

The engine was removed from the airplane and shipped to AEI for disassembly and examination. The disassembly of the engine in the presence of the investigation team revealed one 1st stage turbine blade part number 819501 was fractured transversely across the airfoil about an inch above the blade root platform. The blade's fracture surface from the leading edge back about 0.48-inches was smooth and planar and the remainder of the fracture surface was coarse and grainy. (Photo No. 1)

The fractured blade was removed from the engine for a metallurgical examination that was conducted at the NTSB's and P&W's Materials Laboratories. (The NTSB and P&W laboratory reports are in the NTSB's public docket for this incident.) The metallurgical examinations revealed the blade had fractured from a high cycle fatigue (HCF) fracture that had initiated from multiple origins along the leading edge at the interface between the coating and the blade's base metal. The metallurgical examinations showed that the crack had progressed rearward in HCF for about 0.22-inches and then a mixed mode of HCF and overload for about 0.26-inches with the remainder of the fracture being in overload. The metallurgical examination also revealed that the blade's material and coating conformed to the requirements of the engineering drawing and the engine manual, respectively. The metallurgical examination also showed that the blade's leading edge radius at the location of the fracture conformed to the requirements of the engineering drawing. In addition, the metallurgical examination of the blade revealed that it had not been exposed to an over-temperature condition. The metallurgical examination revealed a secondary craft that was directly below the primary fracture. The cross-section of the secondary crack, which traversed the coating into the base metal of the blade had a similar morphology to that of the primary fracture. In addition, the energy dispersive spectroscopy of the fatigue portion of the primary fracture and secondary crack produced a spectra that was rich in phosphorous. However, the overload portion of the fracture surface did not have any phosphorous. Phosphorous is not a constituent of either the blade's alloy or the coating, but is associated with turbine engine oil.

The examination of the engine revealed extensive damage throughout the LPT. All of the 2nd, 3rd, and 4th stage turbine blades were in their respective disks; but all of the blades had some degree of damage, mostly nicks and dents on the airfoils. Of the 114 2nd stage turbine blades, there was 1 blade that was missing the outer end of the blade. Of the 108 3rd stage turbine blades, all were broken about 6-inches from the blade platform and were missing the tip shrouds except for 5 blades that were broken between 1.75- and 5-inches above the blade platform. Of the 80 4th stage turbine blades, 20 were full length with the tip shroud still in place. The remaining 4th stage turbine blades were broken about 7.5-inches from the blade platform and were missing the tip shroud and many of the broken ends were bent opposite the direction of rotation. The LPT case did not have any holes in it. But the turbine exhaust case had four splits and three holes around its circumference that were all about 3.5- to 4-inches aft of the front flange and the plane of rotation of the 4th stage turbine blades. The holes in the turbine exhaust case were all adjacent to and on the counter-clockwise side of the Pt7, or exhaust total pressure, probes. (Photo No. 2) The largest hole that was at the bottom of the turbine exhaust case and was about 8-inches long. (Photo No. 3) The hole at the bottom of the turbine exhaust case was adjacent to a through hole on the side and bottom of the right hand thrust reverser actuator. The tab of metal that was bent into the hole in the side of the thrust reverser actuator had the imprint of a turbine blade tip shroud on it. (Photo No. 4) The examination of the remainder of the engine did not reveal any other damage.

Five randomly located 1st stage turbine blades that were removed from the 1st stage turbine disk were submitted for a dimensional inspection of the blades' tip shroud twist angle and cross notch and cross shroud dimensions. All of the blade had untwisted from the overhaul manual's limits and were 1°22', 0°26', 0°12', 0°20', and 0°26' below the minimum. The blades' tip sh roud cross shroud dimensions vaview above and below the overhaul manual limits and were 0.0025- and 0.0015-inches above the maximum limit and were 0.0015-, 0.0005, and 0.0025-inches below the minimum limit. The blades' tip shroud cross notch dimensions were mostly below the minimum limit with one blade just within the lower limit by 0.0007-inch and of the remaining four blades, one was 0.0063- and three were 0.0023-inch below the minimum limit.

During the HSI, AEI replaced all of the 1st stage turbine blades. The records from the HSI shat the blades that had been installed in the engine were beyond economic repair. AEI purchased 130 1st stage turbine blades for installation in the incident engine from Western Aero Services, a Denver, Colorado FAA-approved Part 145 repair station that repairs airplane parts that also acts as a broker for other parts such as out-of-production engine parts. Western had purchased the parts from the Royal Australian Air Force and sent the blades to EC Technologies, San, Antonio, Texas for inspection and repair in January 2000. EC Technologies is an FAA-approved Part 145 repair station that repairs turbine blades and vanes. The records show that the blades underwent a preliminary inspection, coating removal, fluorescent penetrant inspection, shroud rotation [twist] restoration, cross shroud and cross notch surface weld repair, coating application and shot peen the blade root fir tree in accordance with the JT3D overhaul manual. EC Technologies currently operates as MT Texas and continues to repair turbine blades and vanes, although it no longer works on JT3D turbine blade and vanes because of a lack of demand.

NTSB Probable Cause

the fracture of one 1st stage turbine blade from a high cycle fatigue crack that originated from a break in the coating on the leading edge of the blade. The cause for the break in the coating could not be determined. Contributing to the uncontained release of turbine material was the yielding and rupture of the turbine exhaust case wall after turbine debris collected at probes in the case.