Feasibility study of a novel load alleviation system on the UH-60A Blackhawk helicopter

Though structural load alleviation (SLA) schemes for helicopters traditionally solely use rotor controls to reduce loads, Voskuijl [54] presents an SLA system for the Sikorsky UH-60A Blackhawk, making combined use of longitudinal cyclic and horizontal tailplane actuation. In high speed flight, when performing maneuvers around the longitudinal axis, utilization of these combined controls reduces pitch link loads and shaft bending moments, due to lower use of longitudinal cyclic. The maneuvre of choice is the 1.75g doublet at 130kts, mimicking the UTTAS maneuvre. This particular SLA system is of interest since its application to helicopters currently featuring a variable incidence stabilizer requires minimal hardware changes. The system is also practically implementable since use is made of two feedback signals that are measured aboard most helicopters. While results in the Flightlab simulation environment are promising when testing the system under design conditions, load reduction capabilities in off-design conditions are unknown. It is also unknown if this SLA system is safe if failure occurs. Furthermore, if loads of the regular UH-60A performing the 1.75g doublet are considered as a reference, the degree of component downsizing or gross weight increase is unknown. Within Flightlab, this research aims to assess off-design performance, safety after SLA system failure, as well as determining potential weight savings that the implementation of this system may allow. Four off-design conditions are considered, consisting of maximum forward or aft cg, and low or high speed flight. Due to adverse attitudes at airspeeds below 60kts, it is decided to phase in the SLA system at airspeeds above 60kts, being fully functional at 80kts. In all off-design conditions, shaft bending moment load reduction exceeds 10%. The third design variable used during SLA development, gross weight (GW), is not considered. Concerning safety, tailplane deflection limits are chosen based on longitudinal cyclic margin following a maximum tailplane deflection failure case. It is decided to limit the stabilator to incidence angles of -7deg (trailing edge up) and 10deg (trailing edge down). These limits also allow for a safe landing attitude at the maximum landing airspeed of 60kts. A tailplane actuation rate of 45deg/s is imposed, being the required amount for full load alleviation at 80kts forward speed. Tailplane deflection at this actuation rate, for all off-design conditions, yields maximum peak load factors of 1.54g. However, tailplane actuation limits may constrict the helicopter’s agility. This is assessed on the basis of five metrics: attitude quickness, agility quickness, flightpath agility quickness, flightpath bandwidth and pitch rate bandwidth. Of these, flightpath bandwidth is a novel approach developed here. These newly developed metrics are proposed as current metrics and norms do not paint a complete picture concerning agility. Agility assessment according to these five metrics indicate equal or superior performance of the SLA system for changes in pitch rate and flightpath trajectory. Three failure cases; -7deg, 10deg fixed tail incidence and 0.1deg/s rate limited tailplane demonstrate sufficient capability to change aircraft pitch attitude or flightpath trajectory. While ADS-33E solely requires the helicopter to remain within the OFE after FCS failure, it is felt that the proposed flightpath bandwidth metric would be a valuable addition to ADS-33E, more adequately ensuring safe operation after failure. If loads on the baseline UH-60A are considered as a reference for maximum tolerable loads, the design of the SLA equipped UH-60A can be altered, or GW can be increased, before similar shaft normal stresses are experienced during the 1.75g doublet at 130kts flight. It is found that the main rotor radius cannot be downsized based on this parameter. Rotor shaft downsizing results in a shaft weight reduction of 0.31kg, down from 11.11kg. When increasing payload instead of downsizing main rotor components, the mixed control SLA system allows the design gross weight to grow from 15283lbs to 16409lbs; an increase of 7.4%. Finally, the relevance of findings is determined by comparing the transients for variations in inflow and interference models. Differences in load predictions for similar pitch rate transients indicate valid results. Nonetheless, flight tests must be conducted due to the difficulty of proper tailplane interference modeling (Prouty [47]). Therefore, it can be concluded that the proposed SLA system provides significant load reductions in the considered off-design conditions, potentially allowing for a substantial GW increase. With the chosen tailplane actuator limits, the SLA system is considered safe for all failure scenarios, while allowing for similar or improved agility in terms of pitch attitude and flightpath trajectory.