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YUH-61A

The Boeing-Vertol 179 / 237 / YUH-61A was a competitor in the UTTAS program to choose a utility helicopter for the US Army (won by the Sikorsky UH-60). Of the three prototypes built (1974), one was modified for the LAMPS III programme. It had a four-blade rotor of composite material.

The YUH-61A helicopter transmission has a rated power level of 2080 kW (2792 hp) at an output rotor shaft speed of 286 rpm. The transmission had two twin inputs. Input spiral bevel gears mesh with a combining spiral bevel gear for the first reduction stage. The second reduction stage consists of a four planet, fixed-ring gear planetary. The overall transmission reduction ratio was 25.096:l.

The helicopter transmission test stand at NASA Lewis was initially designed for testing the U.S. Army's 3000-hp Utility Tactical Transport Aircraft System (UTTAS) helicopter main rotor transmission. The 3000-hp test stand was designed and built by the Boeing Vertol Company, refurbished by NASA, and put on line at NASA Lewis in March 1981. The test stand operates on a torque-regenerative principle. Power to the test transmission flows through two inputs (simulating two engines) and two outputs (main rotor and tail drive). Power i s provided by a constant speed 600-kW (800-hp) induction motor and speed is controlled by an eddy current clutch. Torque is induced independently i n each loop by planetary torque units. The stand is also capable of applying lift loads, moment loads, and drag loads on the transmission output shaft. Efficiency and vibration data were taken in the 3000-hp stand. The mechanical efficiency measured 98.7 percent at full power.

The mounting of an engine in a helicopter requires a relatively rigid system to prevent excessive shaft misalignment and at the same time, limit the velocities and accelerations of the engine in response to rotor and shaft excited frequencies. The YUH-61A mount study began in 1972. The problem was to select an engine mounting system which provided the lowest vibration environment for the engine which entailed avoidance of any resonance of the engine on its mounts in any mode at the exciting frequencies of the rotor(s), shafting, etc.

The YUH-61A engines are mounted to the aircraft at four locations. The mounts were provided with the spring rates as shown to detune the coupled engine/airframe natural frequencies from the important potential rotor forcing frequencies of the YUH-61A. These frequencies for the YUH-61A are 4/rev of the main rotor at 19.1 Hz, 8/rev at 38.2 Hz, and 1/rev at 4.77 Hz with NR at 286 rpm.

Early in the testing of the ground test vehicle (GTV), exhaust ducts cracked. In addition, the vibzation level at the exhaust frame of the engine exceeded the engine manufacturer's limit of 2.5 in/sec at high frequencies (above 50 Hz). In February and March 1974, a set of then-current design airframe engine nacelle components were installed on an engine at the engine manufacturer's facility and a vibration test was conducted. The test consisted of a static shake test and a powered abusive vibration test. The static shake test used a small shaker attached to the center of the shake table to which the powerplant package was mounted through the airframe/engine mount isolators.

For the abusive vibration test, the engine was supported by YUH-61A mounts on a shake table. The shake table was mounted on fou2 air bags. Vibratory excitation of the shake table was provided by means of three vertical electrohydraulic actuators, one beneath the aft mount plane in the center of the table, the other two at the forward end of the shake table, on each side of the table.

In November 1974, during the first 50 hours of testing the ground test vehicle (GTV), the exhaust ducts on each engine cracked. From January 1975 through 1975, both Boeing Vertol's and the engine manufacturer's exhaust ducts were vibrated at the engine manufacturer's facility by mounting the exhaust duct under study on an engine exhaust frame which was mounted to a shake table and vibrated at 4gs.

The interim solution was to coat the splines of existing quill shafts with nylon to permit a custom fit of the quill shaft splines to the mating splines. This pinpointed the source of the vibration and a method of suppression. The nylon coating was not a satisfactory production solution, since it does wear. However, it permitted the completion of the aircraft test program without incurring the lengthy delay that would have been required if new quill shafts with close outer diametral fits were required for piloting. The transmission splined component would also have to be replaced since this internal spline h~as a full radius which is unsuitable for use as a pilot.

There was no significant difference between the MTBF estimates of the UTTAS candidates in DT II. Considering the last 200 hours of OT II, the Boeing MTBF was significantly higher than the Sikorsky MTBF at the a=.15 level. Both UTTAS candidates demonstrated a higher system MTBF in OT II as compared to DT II, However, the Boeing UTTAS achieved a larger MTBF increase than Sikorsky. The Boeing increase was significant at the a=.Ol level. The Sikorsky increase was significant at the cz=.15 level. Engineering modifications in DT II and differences in the DT and OT flight profiles may be contributing factors to the reliability improvement of both UTTAS candidates in OT II.

During both DT II and OT II, the Boeing V56 prototype consistently demonstrated a higher system MTBF than tiat of any other Boeing or Sikorsky UTTAS prototype. It should be noted that the V56 aircraft was extensively refurbished after an accident which occurred during contractor flight testing in November 1975. At the time of the accident, the MTBF of the Boeing V56 prototype was significantly lower than the MTBF of the V57 prototype. The DT/OT I1 results, which now shnw the V56 aircraft with a significantly hiqher MTBF than "57, suggest that the V56 aircraft may have derived considerable benefit from the rebuilding after its accident.

An in-flight artificial icing evaluation was conducted of the YUH-61A helicopter equipped with a prototype deice system. During the test program 3.2 hours were flown in the artificial icing environment. Of this time, 2.8 hours were flow with the deice system functioning, and 0.4 hour was flown with the system not functioning. Anti-ice systems for the engines, engine air induction systems, pitot tubes, and windshields were used during all flights and functioned satisfactorily. With the deice system not functioning, ice accretion on the airframe and flight control surfaces caused significant increases in power required for level flight and significant decreases in autorotational rotor speed with collective full-down. Also noted were increased airframe vibration levels caused by random asymmetrical shedding of ice from the main rotor blades, and damage to the tail rotor blades and transmission fairing caused by ice impact. These adverse results preclude safe operation of the YUH-61A in an icing environment without a main rotor deice system. With the deice system functioning, the YUH-61A successfully flew in artificial icing conditions simulating moderate icing. Three deficiencies were noted which should be corrected prior to flight in icing conditions. These deficiencies are the inability to activate the deice system following an ice detector malfunction; the lack of a system to monitor the integral particle separator turbine operation; and the erratic and unreliable pitot-static indications in level and climbing flight caused by the irregular ice accretion patterns on the lower fuselage nose area.