Numerical and experimental analysis on the helicopter rotor dynamic load controlled by the actively trailing edge flap

Reducing the rotor dynamic load is an important issue to improve the performance and reliability of a helicopter. The control mechanism of the actively controlled flap on the rotor dynamic load is numerically and experimentally investigated by a 3-blade helicopter rotor in this paper. In the aero-elastic numerical approach, the complex motion of the rotor such as the stretching, bending, torsion and pitching of the blade including the deflection of the actively controlled flap (ACF) are all taken into consideration in the structural formulation. The aerodynamic solution adopted the vortex lattice method combining with the free wake model, in which the influence of ACF on the free wake and the aerodynamic load on the blade is taken into account as well. While the experimental method of measuring hub loads and acoustic was accomplished by a rotor rig in a wind tunnel. The result shows that the 3/rev ACF actuation can reduce the $3\omega$ hub load by more than 50\% at maximum, which is significantly better than the 4/rev control. While 4/rev has greater potential to reduce BVI loads than 3/rev with $\mu=0.15$. Further mechanistic analysis shows that by changing the phase difference between the dynamic load on the flap and the rest of the blade, the peak load on the whole blade can be improved, thus achieving effective control of the hub dynamic load, the flap reaches the minimum angle of attack at 90°-100° azimuth under best control condition; when the BVI load is perfectly controlled, the flap reaches the minimum angle of attack at 140° azimuth, and by changing the circulation of the wake, the intensity of blade vortex interaction in the advancing side is improved. Moreover, an interesting finding in the optimal control of noise and vibration is that an overlap point exist on the motion patterns of the flap with different frequencies.

1D modelling and PID control of helicopter Diesel engine rotational speed in torque changes

The article discusses the results of simulation tests concerning the operation of a diesel engine for a light helicopter. The tests were carried out in the AVL Boost software which is used to analyze dynamic phenomena in internal combustion engines. The research object was a newly designed diesel engine of a V8 structure and a power of 330 kW. This engine was designed to be used in the construction of a light class helicopter. The created one-dimensional simulation model included all the main engine components as well as the connection to the helicopter main transmission and the helicopter rotor. The tests consisted in selecting the P value in the PID controller used to control the amount of fuel injected into the engine. The change in the P value indirectly influenced the reaction of the engine to a change in power and torque during horizontal flight of a helicopter. These changes were introduced by changing thrust torque in the helicopter rotor. The fuel injection regulator was designed to maintain a constant engine rotational speed. The maximum speed deviations from the nominal speed of the engine operation due to both increasing and decreasing speed were analyzed. Additionally, the sum of the deviation values was analyzed until the rotational speed of the tested object stabilized. The results showed that the change of the P parameter affects all the analyzed parameters of the engine operation; however, the minimum deviation values for each parameter occur at non-equal PID settings, which makes it difficult to clearly indicate the appropriate value of the P element.

Hysteresis Compensation for a Piezoelectric Actuator of Active Helicopter Rotor Using Compound Control

Active rotor with trailing-edge flaps is a promising method to alleviate vibrations and noise level of helicopters. Hysteresis of the piezoelectric actuators used to drive the flaps can degrade the performance of an active rotor. In this study, bench-top tests are conducted to measure the nonlinear hysteresis of a double-acting piezoelectric actuator. Based on the experimental data, a rate-dependent hysteresis model is established by combining a Bouc–Wen model and a transfer function of a second order system. Good agreement is exhibited between the model outputs and the measured results for different frequencies. A compound control regime composed of a feedforward compensator and PID (Proportional–Integral–Derivative) feedback control is developed to suppress the hysteresis of this actuator. Bench-top test results demonstrate that this compound control regime is capable to suppress hysteresis at different frequencies from 10 Hz to 60 Hz, and errors between the desired actuator outputs and the measured outputs are reduced dramatically at different frequencies, revealing that this compound control regime has the potential to be implemented in an active helicopter rotor to suppress actuator hysteresis.

Review of Modern Helicopter Constructions and an Outline of Rotorcraft Design Parameters

This work contains the results of a modern helicopter construction analysis. It includes the comparison of almost seventy rotorcraft constructions in terms of size in line with EASA requirements – large and small helicopters. The helicopters are also divided because of a mission purpose. The proposed division for large aircrafts is: transport, multipurpose, attack and for small aircrafts: observation, training, and utility. The aircraft construction features are described. Average dimension values of airframes and rotors are shown. Helicopter rotor arrangements are presented in terms of an operational purpose. Next, the rotorcraft design inputs are described. The mathematical formulas for design inputs are given. The ratios are calculated and gathered for the compared aircrafts. Correlation between the analysed parameters is presented on charts. Design inputs are also presented in the paper as a function of MTOW. The function trends are determined to provide an evaluation tool for helicopter designers. In addition, the parameters are presented as possible optimisation variables.

Fracture analysis of AA6061-graphite composite for the application of helicopter rotor blade

The main objective of the work is to study the fracture behavior of AA6061-graphite material using both experimental technique and finite element simulation by considering helicopter rotor blade as a case study. From the case study, it has been observed that the helicopter rotor blade, made of AA6061, has been failed at the threaded portion of the hole. Experimental fracture toughness is carried out using the compact tension specimens as per ASTM standard testing procedure. Modeling of compact tension specimens and the threaded portion of the bolt hole was utilized to analyze the fracture toughness using a simulation tool. From the results and the comparison, it is recommended to use AA6061-9wt% graphite material as a replacement of AA6061 in the application of main rotor blades of the helicopter.