Development of a Helicopter Airframe Flight-by-Flight Fatigue Load Spectrum

To predict airframe service fatigue cracking, it is necessary to develop a service fatigue stress spectrum at certain critical locations on the airframe. Fatigue stress/strain spectra consist of maximum and minimum stress pairs with their frequency of occurrence in each regime of a flight. Flight-by-flight usage of 1,838 flight hours consisting of 2,886 flights was recorded with the usage monitoring system installed on 8 Navy and 11 Marine helicopters. This information was utilized to obtain a flight sequence and regime sequence within a flight and the associated number of cycles in each regime to which the airframe was subjected during each flight. The measured rotor shaft loads were available for a limited number of flight hours. A harmonic analysis of the measured rotating hub loads was performed to obtain the amplitude and frequency content. The amplitude and frequency content of rotating frame loads was utilized to construct fixed system loads using Fourier Coordinate Transformation (FCT). The front and aft rotor hub loads thus determined were balanced by inertia loads at the aircraft center of gravity (c.g.) using the equation of motions. The aircraft linear and angular accelerations required to achieve equilibrium were evaluated and correlated with measured flight test data to verify the mathematical model. In the next step, local linear and angular accelerations were calculated for the discrete mass model of the aircraft. Local linear and angular accelerations and mass distributions were used to determine distributed inertia loads. The distributed inertia loads and front and aft rotor hub external loads were transferred to the airframe FEM loading nodes to construct the so-called external load vector. The load vectors were created for all the regimes in the usage spectrum and used later to compute maximum and minimum stresses near critical locations by performing airframe FEM analysis. The stresses thus calculated form the global finite element model were compared with the measured stresses recorded during the stress survey for selected locations. A local model will be developed to obtain the close correlation. The number of cycles in each regime was then combined with the maximum and minimum stresses obtained from the FEM analysis to develop a stress spectrum. The stress spectrum thus developed will be utilized to compute fatigue crack initiation life with the local stress-strain method and crack growth life. This approach enables the Navy to plan airframe changes and achieve flight safety. It also provides tools to assess individual airframe fatigue damage in the Navy’s aging fleet without strain gage instrumentation on each aircraft.