Military

F118 Engine

A derivative of the F110, the F118, powered the B-2 bomber. The Air Force Integrated High Performance Turbine Engine Technology (IHPTET) program provided the key fan technology for the F118 engine that powers the B-2. The F101 engine was originally developed for the Advanced Manned Strategic Aircraft program, which became the B-1 bomber. Utilizing the same core design as the F101, the F110 and F118 derivative engines were created by developng new low pressure systems to tailor engine performance to the desired aircraft application. Because essentially the same engineering and parts are also used in the F101 and F118 engines, the knowledge they have gained in the development of the service life extension program for the F110 engine will be very transferable to other areas and engines.

In 1988, the U.S. Air Force unveiled the Northrop B-2 Stealth Bomber, powered by four non-augmented F118 engines. The Stealth aircraft was very similar to an experimental aircraft that was developed in the 1940s, the Northrop YB-49A, which had been powered by eight GE J35 engines. Flight testing of the F118 began in mid-1989 and in 1991, the B-2 was awarded the Collier Trophy for the greatest achievement in aeronautics and astronautics.

By 1998 the F118-GE-101 engine was mounted in all U-2 aircraft. It proved a reliable replacement to the J-75 with only one uncommanded in-flight shutdown from its addition to the fleet to the end of 1998. In 1994, Lockheed Martin delivered to the Air Force the first U-2S high-altitude, long-endurance reconnaissance and surveillance system with state-of-the-art sensors able to collect intelligence in all weather and light conditions. The U-2R was in the process of being converted to the U-2S, and the majority of that process was the addition of a new, high-efficiency F118 engine.

The U-2 Directorate managed the U-2 Production Engine Improvement Program (PEIP) for U-2 conversion and deployment. As of 31 December 1996, 20 USAF U-2s and one National Aeronautics and Space Administration (NASA) ER-2 aircraft underwent conversion to the F-118-GE-101 powered "S" model configuration. By the end of FY97, six additional U-2R aircraft were in assorted stages of the process. Nine more U-2Rs and one ER-2 waited to undergo the conversion. In August 1996, Air Force officials sent 3 U-2S aircraft to Taif AB, Saudi Arabia. Other locations for U-2 deployment had been Osan AB, South Korea, in October 1995, and RAF Akrotiri, Cyprus, in January 1996. In February 1997, three U-2S aircraft deployed from Beale AFB, California, to Istres, France. This action completed the deployment of the U-2S to overseas operating locations.

During FY97, the PEIP re-engined two NASA owned/operated ER-2 aircraft (Tail Numbers 1063 and 1097). Tail Number 1063 was completed on 30 October 1996, with T/N 1097 finished on 13 March 1997. Under the terms of a Memorandum of Agreement with NASA AMES, Moffett Field, California, WR-ALC would support the ER-2/F118 as had been done with the ER-2/J75. Plans called for the PEIP to continue through December 1998. As 1997 ended, 26 of 37 U-2 aircraft had been re-engined in the U-2S/F118 configuration and delivered to Air Combat Command, 9 Reconnaissance Wing (RW), Beale AFB, California. Five aircraft were in PDM/PEIP conversion while 4 U-2Rs remained in service at Beale AFB. The last U-2R would enter PDM/PEIP conversion on 30 March 1998, with delivery scheduled for 22 December 1998.

AEDC's unique capabilities and teamwork are validating the performance of critical control systems in the USAF B-2 Spirit bomber. The General Electric F118-GE-100, which serves as the B-2 power plant, completed Digital Electronic Control (DEC) testing in AEDC's Propulsion Development Test Cell J-2 on 28 June 2002. While at AEDC, the 19,000-pound-thrust engine underwent 35 hours of simulated flight testing to ensure the new DEC is functionally interchangeable with the existing engine fan temperature control and engine monitoring system processor. GE officials credited the program's success to the teamwork atmosphere at AEDC. The engine was first tested in the production configuration (with the engine fan temperature control and engine monitoring system processor installed) to gather baseline engine operating data. The baseline testing was then repeated with the new DEC installed and the data was reviewed to verify proper functionality.

A basic gas turbine engine for an aircraft includes a compressor that compresses air entering the engine, a combustor section where the compressed air is mixed with fuel and combusted as a hot gas, a turbine where energy is extracted from the hot gaseous stream to turn the engine shaft on which the compressor is mounted and an exhaust where the remaining hot gaseous stream is used to propel the aircraft. Turbine engines used to propel large aircraft such as passenger aircraft and transport aircraft may include a fan mounted on the shaft in front of the compressor, which may direct some of the air around the compressor. Turbine engines used to propel military aircraft may include augmentors or afterburners in the exhaust to inject and burn additional fuel into the exhaust gas stream for additional thrust. Typically, advanced turbine engines used in high performance military aircraft include variable inlet guide vanes, although there is nothing to preclude their use on commercial aircraft or low performance military aircraft.

The variable inlet guide vanes are assemblies that allow for realignment of vanes due to changing air angles that occur as the operating condition of the compressor or fan changes so that the air can be can be passed through the engine in the most efficient manner. The inlet guide vane assemblies are located radially in the engine and in the air flow path and can pivot about an axis substantially perpendicular to the flow of air through the engine by about 45.degree.. They are moved in response to power requirements so as to control capacity of the compressor and hence the power generated by the engine. These vanes also direct the flow of the air in the most efficient manner through the compressor. In addition, the movement of the inlet guide vanes can be used to avoid surge and stall that can occur in the engine. Because of the frequently changing power demands as determined by the pilot, the inlet guide vanes assemblies are constantly moving in response to the changing power demands.

The frequent movement of the inlet guide vane assemblies in response to pilot requirements for power and due to engine vibrations results in considerable wear to the inlet guide vane assemblies, which are designed to accommodate wear. The inlet guides are mounted in bearings that typically include bushings, which are designed to minimize wear between the vane and the bushing, and the bushing and the engine casing and shroud. The bushings also act to seal the leak path that otherwise exists between the case and the vane. The variable vane includes a vane stem that extends through the opening in the engine casing (hereinafter referred to as the "outer end") and through the bushing and a washer. The bushing and washer are referred to herein as a bearing assembly, the bearing assembly positioned radially outboard referred to as the first bearing assembly. The vane also includes a similar second bearing assembly at its inner radial end. The bearing assembly produces a low friction surface that prevents metal on metal contact. Typically, better wear performance is achieved by polymeric bushings that are made from thin material, thinner materials yielding longer life.

A lever arm is fixedly joined to the vane stem extending outwardly from the vane bushing or first bearing assembly. The distal end of the lever arm is operatively joined to an actuation ring that controls the angle of the vane. All of the vane lever arms in a single stage are joined to a common actuation ring for ensuring that all of the variable vanes are positioned at the same angular orientation relative to the airflow in the compressor stage.

Currently, bushings are made from NR150 material, which is a high temperature polymer. For the F118 engine, the bushings have a thickness of about 0.025 inches (25 mils). As is typical, the bushings have a limited wear life before requiring replacement, and it is always desirable to increase the wear life of the bushings to increase the mean time between replacement or repair. An improved version provides a bushing for use in a F118 engine which has improved wear characteristics. The bushing is manufactured from NR150 material but the bushing thickness has been increased from 0.025 inches (25 mils) to 0.055 inches (55 mils). The bushing thickness is generally in the range of about 0.045 to about 0.055 inches. An advantage is that the wear life of the bushing is improved by over 100%, which means that the mean time between replacement for the bushing has been doubled so that maintenance related to bushing wear in variable inlet guide vanes can be reduced, resulting in decreased maintenance costs.

During the B-2 development phase, the primary fuel for the air vehicle was changed from JP-4 to JP-8. This was desired due to the requirement to refuel the air vehicle within a hangar and due to limited air flow for ventilating air vehicle compartments adjacent to the fuel tanks. The low volatility of JP-8 greatly reduced the probability of a fuel explosion. JP-4 was re-designated as an emergency fuel. With the routine exposure to JP-4 removed, other fuel subsystem changes could be made. The fuel tank pressurization system, which was mandatory for the fuel subsystem for hot JP-4, was deleted providing a reduction in system complexity and air vehicle weight along with an improvement in system maintainability. Also, fuel tank lightning protection was reduced due to the reduction in risk of fuel tank explosion.

The B-2 program conducted skin panel lightning tests to evaluate the improvement in safety afforded by the conversion of the air vehicle to JP-8 as the primary fuel, with JP-4 designated as an emergency fuel only. A "worst case" batch of JP-8 with a flash point of approximately 100F was obtained for the test. A procedure was developed to provide an optimum mixture for the test chamber. The procedure was developed using a bomb sampler to measure the maximum pressure rise and the pressure rise time. A maximum pressure rise or a minimum rise time both indicate an near optimum mixture. This optimum mixture will ignite at a minimum energy level. It was determined that the near optimum mixture was obtained at a temperature of 150F for the test fuel. Higher flash point fuel will require a higher optimum mixture temperature. It is estimated that the optimum mixture temperature will occur approximately 50 degrees above the flash point. The prime contractor conducted additional tests to determine the ignition energy of JP-8 as a function of fuel temperature. As the temperature of the fuel is varied from the optimum temperature the energy required to ignite the fuel goes up.

The panel tests were all conducted at the optimum temperature; therefore, this is conservative with respect to the air vehicle fuel tank conditions. The panels were tested in the new condition. The effects of aging were not investigated. The air vehicle will only rarely be at the optimum temperature. Although there was light on several of the test strikes there was only one ignition at the 200 KA test level during the test. The test results were used to define a probability of ignition for use in a hazard analysis model.

Most new systems have a completely automatic control system which accomplishes engine feed, c.g. management, thermal management and provides failure information of the fuel components to the warning and caution display. The B-2 system contains three computers and uses an active-standby architecture. The primary computer controls the fuel subsystem. In the event of a failure of the primary computer, the bus controller turns off the primary computer and activates the secondary computer. The secondary computer removes control of the system from the primary computer and takes over operation of the system. This removing of control of the primary computer is necessary in the event that the failure mode of the primary computer would not allow the primary computer to turn off. In the event of failure of the secondary computer, control is transferred to a manual panel. The third computer provides only fuel quantity information to the manual panel and critical warnings and cautions.

The B-2 fuel subsystem design started with an arbitrary requirement for ten seconds of negative "g". The fuel subsystem contractor proposed gravity activated, double ended inlets to meet the requirement. A mission analysis was conducted to define the requirement and it was determined that it was not possible for the B-2 to be subjected to a single duration of ten seconds of negative "g". The requirement was conservatively estimated to be three seconds. The boost pump inlet design was modified to meet the three second requirement.

One common instance is when manufacturing oversight leaves debris such as cleaning rags in the tank. During the early operation of the B-2 it was found that there was an unusually high level of cotton linters in the fuel tanks which collected on the fins of the heat exchangers during engine operation. These fibers were thought to come from the cotton coverall and cleaning cloth of the workers during installation of the fuel components and sealing of the tanks. The roughness of the graphite composite tank material could account for the increase of cotton fiber as compared to that found in metal tanks. Improved flushing procedures were incorporated to insure removal of the contamination prior to engine operation and flight.

In air vehicle number 1 and air vehicle number 2 the flush cleaning process did not effectively remove this contamination. After flushing, sump samples were taken and the samples examined visually as suggested by T.O.42B-1-1. The procedure allows 10 or less fibers to be in a 1 quart sample. The samples appeared acceptable. After taxi tests and early flights, cotton fibers were found on the face of a heat exchanger in the engine feed line. More extensive flushing procedures and fuel quality inspection procedures were incorporated. The defuel line used to remove the fuel after flushing contained a screen. This screen was inspected for fibers after each flush defueling. When the screens appeared clean, sump samples were taken. The sump samples were filtered to measure solid contaminant, and then the filter paper was inspected under a microscope for fibers. It became obvious that short fibers which could not be detected visually could be counted under the microscope. The inspection procedures were modified to require that only fibers longer than 1500 microns in length be counted. For air vehicle number 3, eight flushes were required to pass the less than ten fiber count.




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