采用粒子图像测速(Particle image velocimetry,PIV)技术对一缩比共轴双旋翼模型在悬停和以不同前进比前飞时的流场进行水洞实验.测量得到了旋翼流场的瞬时涡量的速度分布,桨尖涡的脱落轨迹,悬停时的尾迹边界和前飞时的尾迹边界等流场特性参数分布.研究了不同状态下共轴双旋翼流场的气动干扰特性.在悬停时,下旋翼的桨尖外侧有上洗流现象,而下旋翼则没有.与共轴双旋翼性能试验数据比较得出,在悬停时共轴双旋翼形式存在有利的相互气动干扰现象.实验还得出了悬停和不同前进比前飞时桨尖涡的脱落轨迹.

According to the characteristics of the compound coaxial helicopter, the general aerodynamic forces and moments of rotors, wing, fuselage, stabilator, vertical stabilizers (containing the rudder) and propeller were modeled and a trimming equation set was established accordingly. In view of the controlling redundancy problem, a control strategy was studied for the compound coaxial helicopter under conversion mode, using the linear transition method and minimal power optimization transition method. A sample compound coaxial helicopter was taken as an example to demonstrate the effectiveness of the method. Then the trimming values, the attitude angles and the power of these two methods were compared. The result shows that both the two methods can solve the controlling redundancy problem, while the trimming values and the attitude angles are reasonable and the linear transition method is better in the smooth change of the control values, the difficulty level of the controlling system and the flight quality.

A calculation model of the control stick vibration characteristics was set up.Based on the research of helicopter hinge moment,a calculation model of the swashplate control stick load was established.With reference to conventional airplane measurement cells,the measurement cell with enough strength and sensitivity was developed,solving the contradiction between measurement accuracy and flight safety by ground tests.The flight test was successfully conducted,and the important parameter of swashplate control stick load was obtained.The calculation results are basically consistent with experimental data,and the excitation force frequency is close to 1 and 2 times of the main rotor speed.Engineering computation requirement can be achieved,providing a reference for the design and test of helicopter control system.

According to the characteristics of the compound coaxial helicopter, general aerodynamic forces and moments of rotors, wing, fuselage, horizontal stabilizer (contains the elevator), vertical stabilizers (contains the rudder) and auxiliary propeller are modeled. A control strategy and its corresponding trimming algorithm are studied for the compound coaxial helicopter under various steady-state flight conditions. The conditions include the hover and low speed flight in helicopter mode, high speed flight in airplane mode, especially the conversion mode from helicopter to airplane with the linear transition method. Together with the control strategy, the trimming method is investigated to obtain the trimming values. A small-scale compound coaxial helicopter is taken as an example to demonstrate the effectiveness of the method. The result shows that the trimming values, the lift distribution of rotors and the wing under the whole steady-state flight conditions are reasonable.

Based on the engineering reality, a calculation model of the rotor unsteady aerodynamics of the small-scale coaxial helicopter had been set up. An airfoil unsteady aero- dynamic model was established by introducing a set of Leishman-Beddoes semi-empirical indi- cial response formula. The induced velocity of the aerodynamic interaction between upper and lower rotors of a coaxial helicopter was set up based on the rigid wake model, free wake model and the dynamic inflow model with interference factors. Considering the teetering rotor＇s particular flapping dynamic formula, the Runge-Kutta method was used to obtain the numerical solution of the rigid blade flapping angle. According to the calculation and analy- sis, the thrust response to a rapid collective pitch increase in both the hovering and forward flight states were obtained, and the blade＇s hinge moment of the upper and lower rotors were compared.

The yaw control of coaxial helicopters is gained from the differential control of collective pitches of up and down rotors. Hence, a coupling phenomenon can be found. Employing the free wake arithmetic, the aerodynamic analysis for the small coaxial helicopter was made, mainly focusing on the coupling between yaw and collective pitch controls in hover flight. The computation showed the relationship between collective pitches of two rotors when the yaw control or collective pitch control was given. From those results, the characteristics of decoupling control could be identified. And those data was transformed into several simple functions using the curve fitting method. These results will be significant for designing the flight control system.

Based on the engineering reality,a calculation model of the rotor unsteady aerodynamics of the small coaxial helicopter in hovering state was set up.An airfoil unsteady aerodynamic model was established by introducing a set of Leishman-Beddoes semi-empirical indicial response formula.The induced velocity of the aerodynamic interaction between upper and lower rotors of a coaxial helicopter was set up by bringing in an interferential factor in the dynamic inflow model,the Runge-Kutta method was used to obtain the numerical solution of the rigid blade flapping angle.According to the calculation and analysis,the thrust and torque response to a rapid collective pitch increase in hovering state was gotten,and the course response differences between the semi-differential and full-differential direction control modes was compared.The results would be a necessary preparation for the research of coxial helicopter＇s maneuvering flight.

Based on the engineering reality, a calculation model of the rotor unsteady aerodynamics of a small coaxial helicopter in hovering state had been set up. An airfoil unsteady aerodynamic model was established by introducing a set of Leishman-Beddoes semi-empirical indicial response formulas; the aerodynamic interaction between upper and lower rotors of the coaxial helicopter was set up by bringing an interferential factor in the dynamic inflow model, and the Runge-Kutta method was used to obtain the numerical solution of the rigid flapping angle. According to the calculation and analysis, the thrust and torque responses to a rapid collective pitch increase in hovering state were obtained, and the direction response's differences between the semi-differential and full-differential direction control modes were compared.

Based on the engineering reality,a calculation model of the rotor blades hinge moment about coaxial helicopter was set up.An airfoil unsteady aerodynamic model was established by introducing a set of Leishman-Beddoes semi-empirical indicial response formula.The induced velocity model of the aerodynamic interaction between upper and lower rotors of a coaxial helicopter was created by introducing an interferential factor into the Pitt-Peters dynamic inflow model;the Runge-Kutta method was used to obtain the numerical solution of the rigid blade flapping angle,and a modal superposition method used to calculate the distortion of the flapping direction.According to analysis of the coaxial helicopter,the total hinge moment is basically the same,although the partial hinge moments of upper and lower rotors are different.In medium and high speed,the hinge moments of upper and lower rotors are accumulated at the lower swashplate to form a moment making the lower swashplate lean forward.

由于直升机的速度较低,一般最大速度不超过350km/h,机身的气动外形对飞行性能的影响相对固定翼飞机来说较弱。因此,有人说直升机气动特性主要是旋翼气动特性。就直升机本体技术而言,传动系统和旋翼系统是直升机最重要的关键部件,反映了直升机技术的本质和特征。