旋翼/机身/平尾/尾桨非定常气动干扰是导致众多新型直升机试飞出现机体振动载荷增加、低速纵向“抬头”、航向操纵性变差等问题的直接原因，因此直升机设计阶段准确分析旋翼/机身/平尾/尾桨气动干扰至关重要。目前全机气动干扰研究主要依赖于实验，而大部分直升机综合分析软件仍然难于快速、准确分析全机非定常气动干扰。本文基于Lagrangian体系的粘性涡粒子法描述旋翼尾迹，通过增加旋翼尾迹非定常项构建非定常面元法以计算桨叶、机身、平尾的气动载荷，并根据涡量等效和Neumann物面边界条件建立一套适用于旋翼/机身/平尾/尾桨非定常气动干扰分析的非定常面元/粘性涡粒子混合法。通过悬停和前飞状态的Caradonna-Tung、ROBIN、AH-1G、Maryland等实验算例验证方法的准确性。基于本文方法研究旋翼/机身气动干扰下的机身非定常压力和旋翼拉力变化特性，分析表明旋翼/机身气动干扰增加旋翼拉力峰峰值，机身气动载荷呈现桨叶片数的倍频，且机身非定常压力主要受桨叶通过性影响，尾梁主要受桨尖涡/尾梁干扰影响。随后分析旋翼与机身距离、桨叶片数、旋翼轴前倾角、桨尖下反角及后掠角等旋翼参数对旋翼/机身气动干扰的影响规律。基于本文方法分析旋翼/平尾气动干扰对平尾气动载荷的影响规律，耦合直升机飞行动力学模型，再现和揭示飞行测试中出现的纵向“抬头”现象及原因。随后分析平尾构型对平尾气动载荷和直升机操纵特性的影响，并根据平尾等效迎角提出消除“抬头”现象的可动平尾安装角变化规律设计方法。基于本文方法研究悬停、各风向、右侧滑状态下旋翼/尾桨气动干扰对尾桨性能的影响，分析表明旋翼/尾桨气动干扰降低尾桨性能，且60°右侧滑状态最为显著，并出现拉力“迅速恢复”现象。随后分析尾桨旋转方向、高度和纵向位置对尾桨性能的影响，揭示了60°右侧滑状态尾桨性能下降特性主要由旋翼桨尖涡/卷起桨尖涡与尾桨干扰形式的转变和位置决定。研究表明底向前尾桨有利于减小性能损失；低尾桨悬停性能损失较小，但60°低速右侧滑状态存在“迅速恢复”现象；高尾桨悬停性能损失略大，但避免了60°右侧滑状态的“迅速恢复”现象，更适合于侧滑要求较高的军用直升机。
Unsteady aerodynamic interaction of rotor/fuselage/empennage/tail rotor leads to airframe vibration load increasing, low-speed “pitch-up”, and poor yaw controllability which were found in flight test of many helicopters, therefore, accurate analysis of aerodynamic interaction of rotor/fuselage/empennage/tail rotor is essential in helicopter design phase. Most of the current comprehensive rotorcraft analysis codes are still difficult to fast and accurately analyze the unsteady aerodynamic interaction which is mainly depended on experiments. Therefore, base on the Lagrangian formulation, an unsteady panel/viscous vortex particle hybrid method is established through Neumann boundary condition and by converting panels to wake vorticities in which the rotor wake is described by the viscous vortex particle method, and the unsteady aerodynamic loads of rotor blade, fuselage, empennage are represented by the unsteady panel method with unsteady term of rotor wake. The accuracy of the present approach is validated through experiments of Caradonna-Tung, ROBIN, AH-1G, and Maryland rotors in hover and forward flight condition.The unsteady pressure distribution of fuselage and rotor performance under rotor/fuselage interaction is analyzed through the present method. It is shown that the peak-to-peak of rotor thrust coefficient is increased due to the rotor/fuselage interaction, and the frequency of fuselage’s aerodynamic loads is proportional to the number of blades. The unsteady pressure of fuselage is mainly affected by blade passage effects, the tail boom is mainly influenced by the interaction between rotor tip vortex and the tail boom. The influence of distance between rotor and fuselage, number of rotor blades, rotor axis tilt angle, anhedral and sweep angle of rotor blade on rotor/fuselage interaction is then discussed.The effect of rotor/empennage aerodynamic interaction on empennage’s loads is studied through the present method. The phenomenon of “pitch-up” is simulated by coupling with helicopter flight dynamic model, and the reasons of “pitch-up” are explained. The influence of empennage’s type on aerodynamic loads of empennage and handling characteristics of helicopter is analyzed, and a design method of moveable empennage which is used to eliminate “pitch-up” is then proposed through effective angle of empennage.The tail rotor performance under main rotor/tail rotor aerodynamic interaction in hover, different wind direction, and right sideslip condtion is analyzed through the present approach. It is shown that the main rotor/tail rotor interaction decreases the tail rotor performance, the 60° right sideslip is the most significant condition, and there is “sudden recovery” phenomenon in right sideslip condition. The influence of rotation direction, height and longitudinal position on tail rotor performance is studied. The decrease of tail rotor performance is due to the transformation of rotor tip vortex/tail rotor interaction into roll-up tip vortex/tail rotor interaction and the position of roll-up tip vortex. The performance of tail rotor with bottom forward is better than that of bottom backward, the decrease of performance for low-set tail rotor is less than high-set, however, there is significant “sudden recovery” in 60° right sideslip condition. Although the decrease of perfomance is larger in hover condition, the “sudden recovery” in 60° right sideslip condition is avoided for high-set tail rotor, therefore, the high-set tail rotor is more suitable for military helicopter with higher sideslip requirement.
