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.

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.

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.

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.

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.

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.

STUDY OF WING SHIELDING EFFECT OF PROPELLER AIRCRAFT

The calculation of wing shielding effect starts from solving Ffowcs Williams and Hawkings equation without quadrupole source in time domain. The sound scattering of the wing and fuselage which are surrounded by a multi propeller sound field is modeled as a second sound source. A program is developed to calculate the acoustical effects of the rigid fuselage as well as wings with arbitrary shape in motion at low Mach number. As an example, the numerical calculation of the wing shielding of Y12 aircraft with an approximate shape is presented. The result manifests clearly the shielding effect of the wing on the fuselage and the approach is more efficient than that published before.

Some improvements to the model of discrete sound field nearby multi-propeller aircraft

Improvements to the mathematical-physical model of discrete sound field of multi-propeller aircraft have been made by using aeroacoustic analogy method, which considers the effects of fuselage boundary as well as the interference process of the multiple propeller sound field. The calculated results illustrate the effects of fuselage on the propeller sound field, the ’beating noise’ phenomenon and the principle of noise control of synchrophaser system. The model for boundaries with arbitrary shapes can also be used to calculate the effects of rigid boundaries in other harmonic sound fields. Results for sound scattering of a rigid sphere in a planar harmonic wave calculated by using this model are satisfactorily coincident with those by the analytical method.

Application of supersonic particle deposition to enhance the structural integrity of aircraft structures

Aircraft metal components and structures are susceptible to environmental degradation throughout their original design life and in many cases their extended lives.This paper summarizes the results of an experimental program to evaluate the ability of Supersonic Particle Deposition(SPD),also known as cold spray,to extend the limit of validity(LOV)of aircraft structural components and to restore the structural integrity of corroded panels.In this study the potential for the SPD to seal the mechanically fastened joints and for this seal to remain intact even in the presence of multi-site damage(MSD)has been evaluated.By sealing the joint the onset of corrosion damage in the joint can be significantly retarded,possibly even eliminated,thereby dramatically extending the LOV of mechanically fastened joints.The study also shows that SPD can dramatically increase the damage tolerance of badly corroded wing skins.

Applied parametrized and automated airframe modeling methods in the preliminary design phase

The design process of new air-and rotorcraft is commonly separated into three different consecutive phases.In the conceptual design phase,the viability of different designs is investigated with respect to customer requirements and/or the market situation.It usually ends with the identification of a basic aircraft lay-out.In the subsequent preliminary design stage the various disciplines are introduced,thus redefining the design process as a multidisciplinary optimization(MDO)task.The objective of this design stage is to enhance the initial aircraft configuration by establishing an advanced design comprising a loft provided with primary structure.This updated aircraft configuration represents a global optimum solution for the specified requirements which will then be optimized on a local level in the concluding detailed design phase with particular regard to manufacturing aspects.The investigations in the preliminary design phase comprise the generation of numerous similar but still different analytical and finite element(FE)models.Even though computational power is constantly increasing the model generation process is still a time-consuming task.Moreover,it is also a potential source of errors which—in the worst case—may lead to time-and cost-intensive redesign activities during the detailed design.As the preliminary design stage,therefore,is of particular importance during the overall design process the model generation process benefits from parametric models and automated process chains.The presented paper overviews the tools used for the automated generation of FE models developed and used at the Institute of Structures and Design(BT)of the German Aerospace Center(DLR)for the subsequent use in numerical simulations.Furthermore,basic requirements for the effective use of parametrization and automation like a common data format and infrastructure will be introduced.Exemplary models and applications will be presented to illustrate the positive impact on efficiency in aircraft design.Concluding,future development steps and possible applications will be discussed.

VISCOUS FLOWS AROUND BODY-WING JUNCTURES

When an incoming boundary layer flow encounters the large pressure gradient in front of a protruded bluff body such as that of a wing/flat-plate juncture, a horseshoe vortex or a system of horseshoe vortices is generated. These horseshoe vortices travel downstream and form streamwise vortices in the wake. Such a flow phenomenon occurs in many practical applications such as the flow around wing/fuselage juncture of an aircraft, around appendage/hull junctures of a ships, around high-rise building, around bridge piers in rivers, and around blade passages in turbomachines. Due to the vortex flow generated at the juncture, the drag is in general increased and the heat transfer capacity is also degraded. For marine applications, there is an additional and perhaps more significant undesirable feature. That is the streamwise vortices in the wake cause the inflow to the propulsor system to become highly non-uniform. As a consequence, significant unsteady forces may result as the propeller is in operation. The purpose of this paper is to develop an integrated approach of experimentation, numerical evaluation and computer flow simulation for studying and control of the complex three-dimensional vortex flow in this matter.

WORLD

RUSSIA The burnt fuselage of an Aeroflot SSJ-100 passenger plane is seen on the tarmac at Sheremetyevo International Airport in Moscow on May 6. Forty-one deaths were reported after the plane caught fire before making an emergency landing.