A NEW AEROMECHANICAL STABILITY ANALYSIS METHODOLOGY FOR COUPLED ROTOR/FUSELAGE SYSTEM OF HELICOPTERS

The aeromechanical st ability for the coupled rotor/fuselage system of helicopters in forward flight i s investigated. The periodic time-varying equations of motion are developed thr ough building a new 24DOF coupled rigid/elastic blended element based on the fle xible multibody system theory in this paper. It accounts for the effects of prec one, sweep, and the moderately large elastic deflections on the blade and elasti city of shaft and fuselage of the helicopter. The dynamic coupling between the r igid motion of blades about the flap, lag and pitch hinges of articulated rotor and moderately large elastic deflections are included. There is no restriction o n the rotation amplitudes of flap, lag and pitch in the formulation. The stabili ty of periodic solution is studied using the Floquet theory. The transition matr ix is calculated by the Newmark integration method. The aeromechanical stability of a new helicopter is studied. The results show that it is stable in the given forward flight. But the instability arises with the decrease of the bending and torsion stiffness of the shaft.

Prediction of Aerodynamic Interactions of Helicopter Rotor on its Fuselage

An iterative and full-coupled rotor/fuselage aerodynamic interaction analytical method is developed based upon the rotor free-wake model and the 3-D fuselage panel model. A close vortex/ surface interaction model using the Analytical/Numerical Matching (ANM) was adopted in the method in order to simulate effectively the unsteady close interaction between the rotor tip-vortex and fuselage surface. By the analytical method, the unsteady and steady pressure distribution on the fuselage surface, and the unsteady lift and pitching moment of the fuselage in a rotor interaction environment were calculated for different advance ratios. It is shown that the unsteady aerodynamic loads of the fuselage due to the rotor interaction have the same periodic characteristics as the rotor. The comparisons between the present close vortex/surface interaction model and a previous model, which simply excludes vortex filaments inside the fuselage, were also made and the advantages of the former over the latter were demonstrated in improving unsteady close interaction calculations.

Impact of Crash Environments on Crashworthiness of Fuselage Section

In order to study the crash resistance of the civil aircraft structure in different crash environments,two environmental models of soft soil and water are established to analyze the dynamic response of the fuselage section subjected to the vertical at the impact velocity of 7 m/s.Simulation results show that the soft crash environment can have a certain cushioning effect on the structure crash,but it will prolong the crash time and change the energy absorption mode.This work suggests that soft environment may not be suitable for forced landing.

A Realistic 3D Non-Stationary Channel Model for UAV-to-Vehicle Communications Incorporating Fuselage Posture

Considering the unmanned aerial vehicle(UAV) three-dimensional(3D) posture, a novel 3D non-stationary geometry-based stochastic model(GBSM) is proposed for multiple-input multipleoutput(MIMO) UAV-to-vehicle(U2V) channels. It consists of a line-of-sight(Lo S) and non-line-of-sight(NLo S) components. The factor of fuselage posture is considered by introducing a time-variant 3D posture matrix. Some important statistical properties, i.e.the temporal autocorrelation function(ACF) and spatial cross correlation function(CCF), are derived and investigated. Simulation results show that the fuselage posture has significant impact on the U2V channel characteristic and aggravate the non-stationarity. The agreements between analytical, simulated, and measured results verify the correctness of proposed model and derivations. Moreover, it is demonstrated that the proposed model is also compatible to the existing GBSM without considering fuselage posture.

EXPERIMENTAL INVESTIGATION OF AERODYNAMICINTERACTION EFFECT OF ROTOR WAKE ONFUSELAGE OF HELICOPTER

The interaction effect of rotor wake on fuselage of helicopter was investigated with experimental method. The results from experiment have proved that for the drag of fuselage the effect of rotor airflow is closely in connection with both the flight speed and the collective pitch of blades, while for the thrust and pitch moment of fuselage the collective pitch angle of blades plays more important role. A simple and effective computing method about aerodynamic interaction can be derived from the measured data. In order to implement the experiment, a fuselage model, a special sensor, the measurement and data acquisition and processing system were designed and manufactured according to the special requirements of this research project, thereby a good base was built up for carrying out experiments successfully with high quality.

Nonlinear aeroelastic coupled trim and stability analysis of rotor-fuselage

Based on the Hamilton principle and the moderate deflection beam theory, discretizing the helicopter blade into a number of beam elements with 15 degrees of freedora, and using a quasi-steady aero-model, a nonlinear coupled rotor/fuselage equation is established. A periodic solution of blades and fuselage is obtained through aeroelastic coupled trim using the temporal finite element method （TEM）. The Peters dynamic inflow model is used for vehicle stability. A program for computation is developed, which produces the blade responses, hub loads, and rotor pitch controls. The correlation between the analytical results and related literature is good. The converged solution simultaneously satisfies the blade and the vehicle equilibrium equations.

FINITE ELEMENT ANALYS IS AND OPTIMUM DESIGN FOR THE FRONT－FUSELAGE WITH SANDWICH CONSTRUCTIONS

AbstractFinite element analysis and optimization subject to stress, displacement and side con-straints for composite sandwich structures are mainly treated. The square isoparametricsandwich plate / shell elements are used to perform structural analysis. The thickness ofthe faceplates and the depth of the core are taken as design variables in optimization pro-cess. The number of layers for each laminate is also taken as design variables if the compo-site faceplates are used. A few widely applied approximation concepts, such as design vari-able linking, regionalization method and temporary deletion technique of passive con-straints are employed to reduce the number of both design variables and constraints. Theadvanced hybrid approximation techniques combining with dual solutions are cmployed inoptimization. The corresponding software is applied to the analysis of experimental modeland to the optimum design for the composite sandwich front-fuselage, and satisfactory re-sults are obtained.

Effect of Fuselage Cross Section on Aerodynamic Characteristics of Reusable Launch Vehicles

An experimental study on examining aerodynamic characteristics of fuselage cross sections for RLVs (Reusable Launch Vehicles) was conducted at Mach number 0.3, 0.9 and 4.0 in the wind tunnel of ISAS (Institute of Space and Astronautical Science), JAXA (Japan Aerospace Exploration Agency). Three bodies, having the same projected area and length, with and without a set of fins, were tested. Their cross sections are a circle, a square and a triangle with rounded corners. The results showed that the fuselage cross sections had large effects on aerodynamic characteristics in subsonic and transonic flow. The lift coefficient of the model having the triangular cross section with a set of the fins was larger than that of the others in high angles of attack region due to contributions of the separation vortices generated from the fuselage expanding to the wing surface.

Drop test and crash simulation of a civil airplane fuselage section

Crashworthiness of a civil airplane fuselage section was studied in this paper. Firstly, the failure criterion of a rivet was studied by test, showing that the ultimate tension and shear failure loads were obviously affected by the loading speed. The relations between the loading speed and the average ultimate shear, tension loads were expressed by two logarithmic functions, Then, a vertical drop test of a civil airplane fuselage section was conducted with an actual impact velocity of 6.85 m/s, meanwhile the deformation of cabin frame and the accelerations at typical locations were measured. The finite element model of a main fuselage structure was developed and validated by modal test, and the error between the calculated frequencies and the test ones of the first four modes were less than 5%. Numerical simulation of the drop test was performed by using the LS-DYNA code and the simulation results show a good agreement with that of drop test. Deforming mode of the analysis was the same as the drop test; the maximum average rigid acceleration in test was 8.8 l g while the calculated one was 9.17g, with an error of 4.1%; average maximum test deformation at four points on the front cabin floor was 420 mm, while the calculated one was 406 mm, with an error of 3.2%; the peak value of the calculated acceleration at a typical location was 14.72g, which is lower than the test result by 5.46%; the calculated rebound velocity result was greater than the test result 17.8% and energy absorption duration was longer than the test result by 5.73%.

Study on flow behavior and structure over chined fuselage at high angle of attack

A study of leeward vortex structure over chined fuselage and the effects of micro tip perturbation on its vortex flow have been carried out in wind tunnel experiments at Reynolds numbers from 1.26×105 to 5.04×105 with PIV and pressure measurement techniques.Firstly,the experiment results have proved that micro tip perturbation has no effects on the vortex flow and its aerodynamic characteristics over chined fuselage at high angle of attack,in which there are not any non-deterministic flow behaviors.Secondly,the evolution of leeward vortex structure over chined fuselage along the axis of model can be divided into four flow regimes:linear conical developed regime,decay regime of leeward vortex intensity,asymmetric leeward vortex break down regime and completely break down regime.And a correlation between leeward vortex structure and sectional aerodynamic force was also revealed in the present paper.Thirdly,the experiment results show the behavior of leeward vortex core trajectories and zonal characteristics of leeward vortex structure with angles of attack.Finally,the experiment results of Reynolds number effect on the leeward vortex flow have further confirmed research conclusions from previous studies:the flows over chined fuselage at high angles of attack are insensitive to variation of Reynolds number,and there is a little effect on the secondary boundary layer separation and the suction peak induced by leeward vortex.

A model-based prognostics method for fatigue crack growth in fuselage panels

This paper proposes a model-based prognostics method that couples the Extended Kalman Filter(EKF) and a new developed linearization method. The proposed prognostics method is developed in the context of fatigue crack propagation in fuselage panels where the model parameters are unknown and the crack propagation is affected by different types of uncertainties. The coupled method is composed of two steps. The first step employs EKF to estimate the unknown model parameters and the current damage state. In the second step, the proposed efficient linearization method is applied to compute analytically the statistical distribution of the damage evolution path in some future time. A numerical case study is implemented to evaluate the performance of the proposed method. The results show that the coupled EKF-linearization method provides satisfactory results: the EKF algorithm well identifies the model parameters, and the linearization method gives comparable prediction results to Monte Carlo(MC) method while leading to very significant computational cost saving. The proposed prognostics method for fatigue crack growth can be used for developing predictive maintenance strategy for an aircraft fleet, in which case, the computational cost saving is significantly meaningful.

Influence of artificial tip perturbation on asymmetric vortices flow over a chined fuselage

An experimental study was conducted with the aim of understanding behavior of asymmetric vortices flow over a chined fuselage.The tests were carried out in a wind tunnel at Reynolds number of 1.87 · 105 under the conditions of high angles of attack and zero angle of sideslip.The results show that leeward vortices flow becomes asymmetric vortices flow when angle of attack increases over 20.The asymmetric vortices flow is asymmetry of two forebody vortices owing to the increase of angle of attack but not asymmetry of vortex breakdown which appears when angle of attack is above 35.Asymmetric vortices flow is sensitive to tip perturbation and is nondeterministic due to randomly distributed natural minute geometrical irregularities on the nose tip within machining tolerance.Deterministic asymmetric vortices flow can be obtained by attaching artificial tip perturbation which can trigger asymmetric vortices flow and decide asymmetric vortices flow pattern.Triggered by artificial tip perturbation, the vortex on the same side with perturbation is in a higher position, and the other vortex on the opposite side is in a lower position.Vortex suction on the lower vortex side is larger, which corresponds to a side force pointing to the lower vortex side.

Developing an engineering-statistical model for estimating aerodynamic coefficients of helicopter fuselage

The design of the geometric shape of a helicopter fuselage poses a serious challenge for designers. The most important parameter in determining the shape of the helicopter fuselage is its aerodynamic coefficients. These coefficients are determined using two methods： wind tunnel test and computational fluid dynamics（CFD） simulation. The first method is expensive, time-consuming and limited. In addition, estimates in regions away from data can be poor. The second method,due to the limitations of numerical solution, the number of nodes and the used solution, is often inaccurate. In this paper, with the aim of accelerating the design process and achieving results with reasonable engineering accuracy, an engineering-statistical model which is useful for estimating the aerodynamic coefficients was developed, which mitigated the drawbacks of these two methods.First, by combining CFD simulation and regression techniques, an engineering model was presented for the estimation of aerodynamic coefficients. Then, by using the data from a wind tunnel test and implementation of statistical adjustment, the engineering model was modified and an engineering-statistical model was obtained. By spending less time and cost, the final model provided the aerodynamic coefficients of a helicopter fuselage at the desired angles of attack with reasonable accuracy. Finally, three numerical examples were provided to illustrate the application of the proposed model. Comparative results demonstrate the effectiveness of the engineering-statistical model in estimating the aerodynamic coefficients of a helicopter fuselage.

一种多段机翼水面起降地效无人机气动特性

为改善水面起降性能和气动性能,基于船身式机身、多段机翼和T尾融合设计思路,提出并设计了一种新型仿生式多段机翼地效无人机方案,采用N-S方程和K-Ω-SST湍流模型,详细研究了该型无人机在不同状态下的压力云图、压力系数及升阻特性。仿真结果表明,地效作用随离水高度的增加而减小,离水高度与平均几何弦长之比(高度弦长比H/c)接近1时,地效作用较为显著,无人机在离水高度0.2 m时,升力系数、升阻比分别提升21.91%和40.37%,阻力系数降低15.22%;对提出的多段机翼布局,地效飞行主要影响下表面压力系数和压力云图,下表面压力系数展向上由内向外正压增幅逐渐减小,弦向上前后缘附近压力系数较小,结合压力云图分析,地效对升力增幅的影响主要集中在中段和内段机翼下方区域;地效飞行可明显提高升力线斜率(H/c为1时提高了8.89%),迎角增加时升力系数增幅和阻力系数降幅均逐渐变大,升阻比增幅(H/c为1)在迎角2°后均达到26%以上;通过验证机的多轮水面起降和有、无地效飞行试验,证明设计方案具有优秀的气动性能和飞行性能,可为水质检测、水面地效运输、搜救侦察等提供应用平台。

飞机复合材料机身壁板装配技术分析与展望

复合材料以其优异的综合性能在航空航天领域得到了大量应用,其应用范围逐渐从次承力结构向主承力结构扩展,包括在机身结构的组成方面,传统的金属组装壁板逐渐被复合材料整体壁板所取代。由于复合材料壁板具有不同于传统金属材料壁板的装配工艺特点,因此对其装配方法和装配工艺提出了新的要求。针对圆筒状机身复合材料机身壁板的装配过程,分别介绍了复合材料机身壁板大尺寸测量技术、复合材料机身壁板装配定位调姿技术和复合材料机身壁板的先进制孔连接技术,系统总结了近年来国内外相关研究进展和应用情况,指出了飞机大尺寸复合材料结构装配技术未来的研究与应用方向。

数据驱动的机身对接过程仿真技术研究

针对传统机身对接仿真中理论模型偏差较大、有限元模型消耗时间长等问题,开展数据驱动的机身对接过程仿真技术研究。首先通过虚实融合技术,构建六自由度平台运动学模型和机身–调姿平台位姿转换模型,通过数据交互实现仿真数据实时传输从而驱动虚拟模型,利用测量数据和虚拟模型进行了机身对接仿真,确定了对接过程位姿参数,为调姿平台参数修正提供数据基础。之后为提升数据求解效率,根据求解机身变形的有限元模型,计算出变形量并转化成机身位置变化量,将机身姿态角和位置变化量作为输入输出值构建代理模型,并验证方法有效性。最后开发了机身对接过程仿真系统,以机身试验件的对接过程为例,验证了仿真系统的可行性。

电动飞机蒙皮结构的冲击损伤试验与优化设计

飞机蒙皮结构常受到冰雹、维修碰撞等低能冲击,机身结构易出现损伤,引发飞机性能减退。为了提高蒙皮结构的抗冲击性能,结合Hashin失效准则建立了一种复合材料泡沫夹层结构低速冲击有限元等效模型,并利用超声C扫描对冲击后的复合材料泡沫夹层进行无损检测,试验结果表明,与无损检测结果相比较,模拟结果的误差低于10%,证明了该冲击等效模型的合理性。最后利用该有限元等效模型对某型电动飞机机身复合材料泡沫夹层蒙皮结构进行优化设计,以提高抗冲击能力、减小吸收的破坏能量、降低结构损伤程度为目标,在相同铺层数量下,得到了最优的复合材料泡沫夹层结构铺层设计方案。

局部外形参数对高速飞行器的RCS影响

为了降低高速飞行器的雷达散射截面(radar cross section,RCS),采用多层快速多极子算法和物理光学法研究局部外形参数对飞行器雷达散射特性的影响。在此基础上,提出一种变半径弧形的边缘钝化形式,以提高飞行器仰视对抗能力。所提出的变半径弧形钝化外形具有良好的RCS减缩能力,与传统圆弧钝化外形相比,在前向重点角域内,RCS对数均值降低了22.88%;随着仰角的增大,变半径弧形钝化形式还具备全向RCS减缩能力,在全向角域内RCS对数均值降低了13.37%。最后,研究所提出的变半径弧形钝化方式在不同雷达波频段下的RCS特性,结果表明这种钝化方式对于频率较高时仰视对抗能力更好。

基于时间动态规划的机身冗余控制对接算法

为了提高中机身的对接精度,保证对接效率,减小支撑底板受力变形,提出了一种基于调姿时间动态规划的机身冗余控制对接算法。根据机身初始位姿以及机身对接精度要求计算机身目标位姿,以姿态调整量和调姿速度为约束条件计算调姿时间,规划调姿轨迹。通过调姿轨迹计算定位器各轴驱动量,控制电机联动实现中机身调姿对接。仿真实验结果表明,经过基于调姿时间动态规划的冗余控制对接,飞机机身调姿定位过程中支撑底板无受力变形,减小调姿时间的同时对接精度相对较高,满足机身对接同轴度要求。

声发射技术在复合材料壁板压-剪复合载荷疲劳试验损伤扩展监测中的应用

利用声发射技术对某飞机复合材料机身壁板压-剪复合载荷疲劳试验中位于壁板长桁上的两处初始损伤开展了连续跟踪监测,试验过程中对各监测通道声发射信号的幅值、能量、计数等特征参数进行实时统计分析,在试验中成功监测到其中一处损伤的扩展,为试验结论分析提供了重要数据支撑。