The Proceedings of the 2018 Asia-Pacific International Symposium on Aerospace Technology (APISAT 2018)

Section 1. Icing does harm to the aerodynamic performance of aircraft and it is necessary to evaluate the aerodynamic performance of the aircraft in icing conditions. Icing wind tunnel test is the main method to study icing and anti/de-icing at present, but limited by icing wind tunnel test section size, sometimes it is impossible to carry out full scale model test in it. Instead of scaling the whole airfoil, a designed hybrid airfoil with a shorter chord length is usually used in icing wind tunnel which is composed of a leading edge geometry identical to that of the full scale leading edge and a shorter aft section. The aft section is usually designed to provide full scale flow field and droplet impingement on the leading edge. A newly developed method is to design the aft section to provide full scale airfoil surface pressure coefficient. This paper examined the “same surface pressure coefficient” design method based on icing similarity.Section 2. According to the icing scaling method, the icing process consists of six similarities: (1) Geometric similarity (2) Flow field similarity (3) Drop trajectory similarity (4) Water catch similarity (5) Energy balance similarity (6) Similarity of surface water dynamics. If all of the similarities are achieved the icing processes would be the same. This chapter examined each similarity and found main flow flied parameter to each similarity. Finally it found that, based on the basic hybrid airfoil design theory, if the velocity field near the nose section of the hybrid airfoil is the same as that near the full scale airfoil, the parameters such as β, freezing fraction and so on will be the same. And then all of the similarities would be the same, and so as the icing process.Section 3. A 1 m chord length NACA0012 airfoil model is chosen as a research model, which is also known as full scale airfoil model. The limits of water droplets impingement on the full scale airfoil leading edge will be predicted using numerical method. That part of the full scale airfoil is fixed for the hybrid airfoil, named nose section. The aft section of the hybrid airfoil is then designed to provide full scale airfoil surface pressure coefficient on the nose section of the hybrid airfoil. Flow field is calculated by solving Navier-Stokes equation and the catch efficiency distribution was acquired using an Eulerian method, and the prediction of the ice accretion shape is based on Messenger, multi-island genetic algorithm and gradient algorithm are used in designing. This paper assumes that if the percentage difference of the pressure coefficients between the two airfoils is within 10% (or within 0.05 in absolute value) at the grid points of the nose section, the design of the hybrid airfoil meets the requirements. Finally a 0.5 m chord length hybrid airfoil is acquired.Section 4. In order to compare the velocity fields between the hybrid airfoil and the full scale airfoil, the value of velocity and the angle of incidence of the velocity from 7 planes are compared respectively. The chosen planes are in front of the leading edge of airfoils. Results shows small differences of velocity and angle of incidence of the velocity between hybrid airfoil and full scale airfoil. Only when it is very close to the surface of the airfoils, the difference is only a little more. Results shows that the water droplet catch efficiency and the water droplet impinging limit trajectory of hybrid airfoil are almost the same with those of full scale airfoil. And comparing the ice shapes of hybrid airfoil and full scale airfoil, it shows that this hybrid design method is good for the ice accretion simulation. It can be concluded that the velocity field near the nose section is the same when the surface pressure coefficient is the same, and then the hybrid airfoil ice accretion would be the same with full scale airfoil.Section 5. Three conclusions are drawn from this study: (1) Ice accretion on the hybrid airfoil would be the same as that on the full scale airfoil if the velocity field near the nose section are the same. (2) Velocity field near the nose section would be the same between hybrid airfoil and full scale airfoil if the pressure coefficient on the nose section are the same. (3) Hybrid airfoil can be designed based on the same pressure coefficient on the nose section.

Based on the optimization software ISIGHT, a new optimization method is constructed that the arbitrary polynomial blade modeling program is combined with numerical simulation software NUMECA to reduce the process of parameterizing the blade. With this method, the design parameters for the blade modelling process are chosen as optimization variables. Adopting this method, a single stage transonic axial-flow compressor rotor is optimized by choosing multiple design parameters as optimization variables. Optimization variables are the aerofoil maximum thickness, the location of aerofoil maximum thickness, the design attack angle, the variation of deviation angle, the skew design and the swept design on six design stream surfaces. The mass flow and the pressure ratio are confined and the optimization objective is that the adiabatic efficiency reaches its maximum. At design rotation speed, the adiabatic efficiency of the single stage transonic compressor has increased by 2.94% in design controlling condition after optimizing. The result shows that the optimization method combining arbitrary polynomial blade modeling program and NUMECA is feasible and valid. And the method that multiple design parameters are chosen as optimization variables can effectively improve the performance of transonic compressor.

A general and flexible approximation model based on convolutional neural network (ConvNet) technique as well as a signed distance function (SDF) is proposed to predict aerodynamic coefficients of the airfoils in this paper. Traditional surrogate-based prediction methods are blamed for its limited dimensions of design variables and powerless for strong nonlinear engineering problems. Considering that ConvNets have been proven to be suitable for nonlinear and high-dimensional practical tasks in complex image identification and speech recognition, a two-layer ConvNet framework rather than conventional Kriging surrogate model is built to predict aerodynamic coefficients for large-scale nonlinear problems. In order to build the bridge between geometry information and the ConvNet, a new geometry representation method based on SDF is also applied. Furthermore, numerical studies are presented for wind turbine airfoils at a high angle of attack. Compared to ordinary Kriging model, the ConvNet-based method exhibits competitive prediction accuracy within the certain error margin. Moreover, the influence of the ConvNet’s nonlinear activation functions on the predictive effect is studied in both training and validation datasets.

This paper presents a multi-fidelity distributed aircraft design process using fully automated Computational Fluid Dynamic (CFD) analysis and optimization aiming to provide better prediction of aircraft characteristics in the early aircraft design stages for both conventional and unconventional aircraft configurations. Gradient based optimization algorithm in conjunction with adjoint sensitive analysis method is employed to tackle the more detailed shape arising with the increasing fidelity of the aerodynamic analysis tool. The design process is applied to a short-range transport aircraft. The design synthesis is applied to the original and redesigned configuration and the comparison of the synthesised aircraft characteristics highlight the effect of the shape design in the aircraft design process.

Blended-wing-body (BWB) aircraft concept coupled with distributed propulsion is proposed as a potential configuration to meet the N?+?3 goals. In this configuration, the boundary layer ingestion (BLI) effect resulting from the distributed propulsion enhances the aircraft?s aerodynamic performance significantly. The inlet&outlet boundary interacts with the upper airflow strongly, so that the pressure distribution differs from a clean airfoil. However, previous optimization design works mainly aimed at clean airfoil design at cruise conditions. Thus, in this paper, a two-dimensional section of this configuration considering the inlet&outlet boundary conditions of the propulsion system is designed through a multipoint optimization at cruise and climb conditions. We achieve a high lift to drag ratio 2D shape at cruise conditions while improving its climb performance. In addition, we use a weighted-integral method to improve the robustness of the optimal solution and enlarge the drag-divergence Mach number of the solution significantly. Our results may provide qualitative guidance on the future three-dimensional optimization design of the advanced aircraft aerodynamic shape.

In view of the head shape of a typical supersonic vehicle, the array pressure data of the supersonic vehicle is calculated by computational fluid dynamics numerical simulation method. The real-time solution scheme for algorithm model of the flush air data sensing system based on BP neural network technique and FPGA + DSP architecture digital signal processing is designed, which flight Ma number range is from 2.0 to 4.5. The experiment of real-time calculation of the principled sample machine is done in 1.2 m × 1.2 m supersonic wind tunnel. Test results are in good agreement with the measurement results of the wind tunnel system. The system is working without any fatal error in the whole test, which can reflect the flow parameters changes, and has good robustness and agility. Static pressure measurement relative error is ≤6.9%, the Mach number measurement error <0.1, the angle of attack and side slip angle measurement error <1o.

The calculation method of stability derivative which is based on CFD is employed to simulate the forced sinusoidal oscillating motion of low-aspect-ratio flying wing theory configuration and wind tunnel experiment configuration around body axis. Pitch\yaw\roll dynamic derivatives can be obtained with parameter identification technology from the time history of unsteady aerodynamic force and moment coefficient. The sting interference of static fluid characteristics and dynamic derivatives for low-aspect-ratio flying wing in transonic is studied quantitatively. The conclusions are: 1. The sting interference of pitch damping with angle of attack remain almost 5%–9%, except a rapid decrease to 2% at α = 15°; 2. The amount of interference of yaw damping is extremely large at small angle of attack (yaw damping of experimental model is almost a great many times of theory model), the amount of sting interference maintains in 10%–50% when the angle of attack is at α = 20°–25°; 3. The sting interference of roll damping maintain in 1%–2% at α = 0°–15°, and as the angle of attack increases, the interference rises gradually until 6% at α = 25°. The research can provide conference for the correction of sting interference of wind tunnel experiments for dynamic derivatives.

It is found Mars Pathfinder (MPF) experienced static instabilities along the trajectory when entering Martian atmosphere at 2 degrees’ angle of attack. The reason for the instability and the impact of the configuration of the sphere-cone probe shape have been investigated. Three-dimensional Navier-Stokes equations have been solved, where high temperature gas effects were considered by employing the chemical reaction models. The results of aerodynamic characteristics were compared with LAURA, which showed a good agreement. The distribution of the subsonic area at different trajectory point exhibits direct connection with the static instability. The sonic line movement leading to the pressure variation on the surface causes the instability. Different aeroshell configurations with various half cone angle have been simulated. The results show that the distribution of the subsonic zone changes as the aeroshell shape varies and the static instabilities are reduced. The static instability can be controlled by simply change the aeroshell configuration.

The icing with supercooled large droplet (SLD) can form the complex ice downstream of the deicing boots. Ice accretion due to the SLD may result in extremely severe performance degradation to hazard the flight safety. In this paper a numerical solver is developed to investigate the ice accretion, which includes the centered finite volume method for solving the N-S equations to get the air flow field, and the Lagrangian method for predicting the flow field of droplets, and a revised Messinger model for simulating the thermodynamic process of icing. Considering the SLD conditions, we investigate the droplet deformation and breakup using Taylor analogy breakup method. A splashing model is presented to analyze the splashing phenomenon and the droplet impact. And the Langmuir D distribution is studied to get the impact characteristics and icing result of multi-scale distribution of large droplets. Using the above methods, we complete the numerical simulation of ice accretion and icing effects at the supercooled large droplet conditions over the NACA 0012 airfoil. The calculation results are in good agreement with experimental data. The investigation has important engineering application value for the icing prediction of the SLD.

Effects of static aeroelasticity and Reynolds number on transonic pressure distribution are described. An engineering correction method of static aeroelasticity and Reynolds number effect on wind tunnel pressure distribution is constructed based on high-fidelity Fluid-Structure Interaction Simulation Method. And the fidelity of our engineering correction method has been verified with CRM tunnel pressure distribution from ETW tunnel.

This paper presents the effect different of slip boundary conditions on aerodynamics. The effect of continuum breakdown on aerodynamics of lift wing was shown. Present results keep an agreement with the results in reference about flow around the cylinder. The results predicted by Gokcen slip model is in excellent agreement with the DSMC results in higher Knudsen numbers comparing with other slip models. Peak transfer rate differences with different methods increase as the flow becomes more rarefied. Heat transfer rate is influenced much more by rarefaction effect than pressure coefficient by that.

Near-field signatures derived by equivalent area rule are traditionally used as inputs of sonic boom propagation calculation. However, CFD is more capable to present details of pressure perturbation distribution. Due to the local axisymmetric assumption in propagation calculation based on geometry acoustic theory and the asymmetry of rounding flow field, CFD solutions extracted on a cylindrical surface with radius of 1–2 times length of the projectile couldn’t be used as inputs directly. The multipole matching method proposed by Page and Plotkin was introduced to deal with this problem. At first, accuracy of CFD solutions was validated by two benchmarks, then the pressure perturbation extracted from the flow field around a 70 degrees–swept plate model was corrected before propagation calculation. Cylindrical radius and the order of expanding Fourier serials have been discussed. Ground sonic boom signatures showed that convergence was obtained with the correction.

The Thermal Protection System (TPS) in Space Shuttle is assembled by the ceramic tile, the strain isolation pad (SIP) and the substrate structure. The existent experiment data reveals that the pressure distribution in the gaps and in the SIP has important effect on the aerodynamic load of the tile. Restricted by the computational capacity decades ago, an approximated multi-tile flow model based on the porous media flow was proposed to analyze the aerodynamic load for Shuttle tiles. The present work utilizes computational fluid dynamic (CFD) technology to reexamine the effects of steady internal flow on the aerodynamic load of the tile. The two-dimensional TPS model is established by adding tile-SIP assembly on the surface of airfoil. The far-field flow speed is assumed at transonic regime to simulate the critical load condition with shock wave on the tile. The internal flow in the gaps and flow through the SIP are calculated by solving the Navier-Stokes equation. The Spalart-Allmaras turbulence model is adopted. The pressure distributions on the top of the tile, in the gaps and along the bond line between the tile and the SIP are obtained. Numerical results show: The strongest aerodynamic load is at Xs/L = 0.67; when the permeability of SIP is increasing, the moment in Z-direction and the force in Y-direction are decreasing and the force in X-direction is increasing.

Because of its low resistance and high stealth, the loading mode of the embedded weapon has become the most widely used weapon loading mode. In this article, the method of computational fluid dynamics combined with dynamic grid technique is used to simulate the extravehicular process of an embedded weapon under transonic conditions. The effects of 3 kinds of control measures on the trajectory and posture of the weapon are studied. The results show that the ejection force provides a certain speed and posture for the weapon to overcome the adverse effects of the shearing layer; inclined plate introduces the shear layer airflow outward, reduces the velocity gradient in the shearing layer, and facilitates the weapon out of the bay; the leading edge jet disturbs the shear layer by importing airflow, thereby reducing the impact of the shear layer on the weapon; all three control measures will have a favorable impact on the weapon exit process.

The objective of this research is to investigate the available approach for calculating critical loads of a maneuvering flexible aircraft in transonic flows. A time-domain state-space form aeroelastic model for maneuver loads analysis is developed firstly, in which unsteady aerodynamic influence coefficient matrices are generated based on the overset field-panel method, and rational aerodynamic approximation are constructed using Roger’s approximation. Then maneuver loads analysis for a transport aircraft in pitch are carried out, producing a continuous time history of the airframe states and loads acting on the components of the aircraft. Simultaneously, the effects of the structural deformation of the aircraft on maneuver loads are obtained. The results indicate that the peak loads acting on the wing and horizontal tail decrease 5.1% and 10.6% respectively, while those on the elevator increase 16.2%, due to the structural deformation.

The Blended-Wing-Body (BWB) airplane concept represents a potential revolution in subsonic transport efficiency for large airplanes. Except the Wing-Body blended way, the propulsion airframe integration of BWB airplane also takes a very important role for its good aerodynamic performances. A 300 seating civil BWB aircraft designed by the Airplane Concept Design Institute of Northwestern Polytechnical University, whose engine is podded on pylon located over the wing, after of the aircraft centre-body, is the original model for analysis. Because its body looks like a ship, we also call it Ship-Wing-Body (SWB). The propulsion airframe integration including researches in many areas, in this paper, only the engine’s position effect on the SWB’s aerodynamics is analyzed. The engine is simplified as a Flow-through Nacelle (FTN), and the pylon is deleted from the model to make clear the problem is only the engine position effect. First the effect of FTN at original position on SWB aerodynamic performance is analysed. Then the FTN Position Change Effects on SWB Aerodynamic Performance is analysed. Through the analysis, it could be seen that there are three main aspects of the effect: the high pressure region caused by nacelle’s leading edge stagnation point; the shock wave interface between nacelle and body; and the separation caused by shock wave. These three main aspects are more sensitive along Z-axis than X-axis. Under the influence of these aspects, the lift and drag coefficient of SWB are changed monotonous along X-axis; but along Z-axis, there is a critical position, at this position, the SWB has maximum lift coefficient and minimum drag coefficient.

In this paper, the commercial code FLUENT is employed to evaluate the aerodynamic characteristics of high-lift multi-element airfoils. A two-element airfoil consisting of a leading edge droop nose, a trailing edge dropped hinge flap, as well as a drooped spoiler is designed based on the SC(2)-0410 supercritical airfoil and the GA(W)-2 flap. Parametric study is conducted to analyze the aerodynamic influences of the deflections of the dropped hinge flap and the drooped spoiler. Then, comparison is made between the dropped hinge flap coupled with the drooped spoiler and a conventional Fowler flap without spoiler drooping, given the identical leading edge droop nose. Numerical simulation results show that: (1) With the same flap deflection, the lift coefficient slightly increases at small angles of attack and slightly decreases at large angles of attack, as the spoiler deflection increases; (2) With the same spoiler deflection, the lift coefficient remarkably increases at both small and large angles of attack, as the flap deflection increases; (3) With the same flap deflection of 30°, dropped hinge flap with drooped spoiler obtains a higher lift coefficient at small angles of attack and a lower lift coefficient at large angles of attack, compared to Fowler flap; (4) By increasing the dropped hinge flap deflection to 40°, dropped hinge flap with drooped spoiler obtains a higher lift coefficient at both small and large angles of attack, compared to the original Fowler flap.

In order to understand the effects of axial sweep and tangential lean of blades on aerodynamic performance of transonic axial compressor NASA ROTOR 37, a systematic study was carried out. The optimal design method was developed and applied. The blade stacking line was described by Bezier polynomials. The optimization problem was to achieve the maximum rotor isentropic efficiency and pressure ratio at the design point, with a constraint on the mass flow rate. Commercial software Design 3D was employed for aerodynamic evaluation based on Navier–Stokes code. Simulation results were validated with experimental data in the literature. The optimization technique used in this paper is based on the concept of function approximation. This approximation helps to realize a rapid method available to evaluate rotor aerodynamic performance. The method requires the existence of a database containing several blade geometries and their associated aerodynamic performances. These database samples are used to construct the Artificial Neural Network. Finally, the optimal rotor configurations were obtained and compared to the original design. The effect of swept and leaned blade on the rotor performance was analyzed.

Aerodynamic performance and surface flow of reentry capsule are important for altitude static pressure tapping location design and pressure coefficient calibration. A capsule CFD model is presented and aerodynamic performance in different Mach number, angle-of-attack (AOA) and angle-of-side-slip (AOSS) is investigated in this paper. Based on statistics principle of minimum variance and deviation, location optimization design of altitude static pressure tapping is also conducted in this paper. The result shows numerical accuracy of aerodynamic performance is validated through comparison with wind tunnel test results. Aerodynamic performance of the reentry capsule is in a steady level from −30° to 30° AOA and −10° to 10° AOSS. High absolute value of positive and negative AOA & AOSS can significantly induce deterioration of capsule aerodynamic performance. According to statistics principle of minimum variance and deviation, three appropriate tapping points are obtained. With pressure coefficient histogram analysis, ideal altitude static pressure tapping location is selected.

Gas particles produced by dissociation and ionization in the high temperature shock, have strong radiative heating on surface when spacecraft re-enter at high Mach number. Aiming at the radiation heating problem of the hypersonic reentry spacecraft, the finite volume method was used to calculate the radiation transfer equation, in which a ghost cell method is employed to processed the boundary condition. Using this method to calculate the typical cylindrical furnace. The results showed that the accuracy of the current methods was high, the grid convergence was good, and it satisfied the physical fact. The method was then used to calculate the radiative heating on the Fire II surface based on the calculation data of the chemical non-equilibrium flow field. The radiation heat flux on the surface of the spacecraft is calculated. The results were in good agreement with other researcher’s calculation value.

Distributed electric propulsion is a new concept in recent years. It makes full use of the scale-free nature of electric energy and distributes multiple propulsion devices on the wings or fuselage to improve the aerodynamic performance of aircraft. This study will analysis the effect of distributed propellers slipstream on aerodynamic characteristics of the wing. The study will analysis the propeller-wing system by means of momentum theory, actuator disk model and CFD method to explore the high-lift mechanism of the distributed propeller slipstream. Based on the momentum theory, a method to analyze the lift coefficient of the wing is proposed by using the velocity of the slipstream behind the propellers. Based on the actuator disk model, the influence of different number of propellers on the lift coefficient and lift coefficient distribution of the wing was investigated by OpenVSP. Based on CFD method, calculation and simulation were carried out through Fluent, and k-omega SST model was used to analyze the interaction of the propeller-wing system. The research results show that the distributed propeller slipstream will not only increase the airflow speed over the wing, the rotation of the slipstream also has a great influence on the wing. In addition, the smooth flow of wingtip vortex offset also has great effect on aerodynamic characteristics of the wing.

An analysis was conducted to investigate the flow and heat transfer of a turboramjet pre-cooler and a new concept pre-cooler used controllable octahedron porous media structure is developed. The pressure and temperature distribution of the pre-cooler is investigated with both numerical and experimental method. Additionally an optimization theory and evaluate factor of porous media structure was developed and applied into the pre-cooler optimization. At the present work the CFD results regarding the pressure drop and heat transfer performance were compared to available experimental data and were found to be in agreement. Thus, the porous media pre-cooler could be applicable to the turboramjet.

In order to simulate the influence of intake distortion on the performance and stability of multi-stage axial compressor accurately, an improved split actuator disk (SAD) model is developed based on actuator disk models. By comparing the simulation with the original model, it is verified that the split actuator disk model is more accurate. Further, the influence of the intake distortion condition on performance of multi-stage axial compressor is simulated using split actuator disk model. This paper shows that the split actuator disk can eliminate the error caused by the assumption of the original model effectively, and the simulation results are closer to the full three-dimensional simulation. Using the split actuator disk model can analyze the performance of the multi-stage compressor under the condition of total pressure distortion, and provide a feasible and accurate method for the joint simulation of the inlet-compressor system.

In this paper, aerodynamic and aeroelastic characteristics of flying wing with split drag-rudder are studied, using aeroelastic analysis method in time domain based on CFD/CSD (Computational Fluid Dynamics, CFD; Computational Structural Dynamics, CSD) coupling. Unsteady aerodynamic is obtained by solving Euler equations, and structural response is calculated by solving structural dynamic equations. CFD/CSD coupling uses predictive-multiple step loose coupling method, coupling interface data exchange uses constant volume transformation (CVT) method, and dynamic grid deforming uses transfinite interpolate (TFI) method. The flutter boundary of AGARD 445.6 wing is calculated with the coupled CFD/CSD method. Comparison between present results and experimental data is given, and there are good agreements in subsonic and transonic region, respectively, which illustrate the reliability of this method. Then, for the flying wing, the aerodynamic and aeroelastic characteristics are studied and revealed. Lift and drag coefficients with various cracking angles and angles of attack are calculated, and some change laws are concluded. Finally, flutter boundary of the flying wing is computed, and lift and drag coefficients in rigid and elastic conditions are given, after which several laws are summarized.

The thermal protection system plays a key role in atmospheric re-entry missions of aerospace vehicles. Generally, heat shield plates are assumed to have smooth surfaces. However, discontinuities or cavities exist due to sensor installation, fabrication tolerances and different expansion rates of non-similar materials. This paper tries to investigate low-density hypersonic flows at a group of altitudes, including 70 km, 75 km, 80 km and 90 km, past a flat plate with two- and three-dimensional cavities using the direct simulation Monte Carlo (DSMC) method, with an emphasis to analyze the influence of operation altitude on the flow-field patterns. This work shows that one primary vortex is formed as a result of flow separation and reattachment. In addition, rarefied gas effect plays an important role in flow-field structure and density distributions; as the flight altitude is increased, the primary vortex becomes slender, with its core moving up. Besides, the extra third-dimension precludes external flows from penetrating deep into the cavity, causing the primary vortex to rise, producing a three-dimensional blocking effect. Moreover, as operation altitudes grow, the three-dimensional blocking effect becomes more noticeable.

In this paper, the CFD method is used to carry out the numerical study on the influence of the flight trajectory of the helicopter in hovering and small speed forward flight, and the research selects the Caradonna-Tung rotor and the related fuselage and external store model after the proportional amplification, and uses dynamic patched and overset grid technique. The control equation of the flow field of the helicopter, which include the Reynolds Averaged Navier-Stokes (RANS) equations and the motion equations of the rigid body six of degree of freedom, are solved by coupling. According to the change of the centroid displacement trajectory and the Euler angle (pitching angle, roll angle and yaw angle), the flight trajectory of the external store in different conditions is analyzed, and the influence of the helicopter’s flow field on the flight trajectory of the external store is studied. Through the analysis of the influence of the external store separation trajectory, the research shows that the external separation is influenced by the free flow, the helicopter wash flow and the self-gravity and the posture presents the characteristics of low head, negative rolling and forward yaw. The research work provides valuable reference to the safety problem for the launch and separation of missiles, and the airdrop of personnel and materials from the helicopters.

The purpose of this paper was designing a uniform aero-heating blunt body, and a sphero-cylinder with radius 25 mm was chosen as the benchmark. Axial symmetry parametric shape was designed by a quadratic spline curve profile. An automated aerothermodynamic optimization process was constructed to search objective shape. For the axial symmetry flow field, aerodynamic heating calculations was estimated by solving the Reynolds Averaged Navier-Stokes (RANS) equations in conjunction with the Spalart-Allmaras (SA) turbulence model. The certain design flow conditions corresponded to Mach number 6 flight at an altitude of 25 km and with no angle of attack. The grid of each shape for aerodynamic heating calculation was adaptively refine to catch the shockwave based on the fluid field, and all the grids were generated automatically through Gridgen-based script. For optimization method, the Multi-Island Genetic Algorithm (MIGA) was employed. The optimization strategy was to minimize the heat-flux difference value in a define region of blunt body. As a result, the optimized shape would perform with uniform aerodynamic heating distribution in the region. When compared with the sphero-cylinder benchmark, the final objective body performed a nearly uniform aero-heating distribution, furthermore, it decreased the peak heat-flux by about 27%. Robustness analysis indicated that peak heat flux of objective body performed lower than the baseline of sphero-cylinder with angle of attack no more than 23°. Moreover, validation calculations indicated objective body also performed approximately uniform distribution at higher Mach number with lower altitude and lower Mach number with higher altitude. Heat conduction simulations showed the objective body have a significant decrease of temperature when compared with the original sphero-cylinder body.

The flying-wing configuration has a ideal/unsatisfactory vertical stability since the lack of vertical stabilization or control surfaces. The evaluation of dynamic derivative is critical for the flying-wing study and design. To improve the accuracy of forced oscillation test, the support interference was analyzed by computational method. The simulation of clean model without control surfaces deflection was validated with the results obtained from AVIC ARI FL-51 wind tunnel test by a 2DOF apparatus. Then, cases of model with/without stings and apparatus were simulated. Apparent interference mainly appeared at angle of attack of 35° and 40° where strong vortex flow happened. Both a bent sting and a rear sting were compared for the interference. The model with bent sting was more sensitive to the effect of apparatus than the model with rear sting, but the total interference was still lower than the other. And finally, some discussions of the flow field structure during dynamic motion were given based on the spatial streamlines result.

The free navigation of Fridsma planing hull model is studied using OpenFOAM with different inflow velocities. Numerical calculation method with 6 degree of freedom motion of rigid body and unsteady two phase flow coupled is adopted for the study. The finite volume method is used to solve the Reynolds-averaged Navier-Stokes (RANS) equation and the VOF method is used to capture the free surface of air-water two phase flow. The calculation results show that the hydrodynamic lift of the planing hull increases while the wetted area decreases with increasing inflow velocity. As a result, the hydrodynamic drag is effectively controlled in this way. Furthermore, the error at low inflow velocity between the calculated drag and the corresponding experimental value is less than 10%, while the error at high inflow velocity is increased evidently, which is thought to be relevant to the complexity of the computation of unsteady 6 degree of freedom motions.

Under the constraint of pilot perspective angle, fuselage loading space, and the tail-down angle of after body, to change the parameters of the nose camber, the middle fuselage section’s flatness, roundness, and after-body contract ratio. Use CFD numerical simulation to assess the fuselage’s viscous drag, pressure drag, and longitudinal moment, attain the fuselage shape with low drag and acceptable longitudinal moment. The analysis and study show that the cross section profile of fuselage have greater impact on the fuselage pressure drag and longitudinal moment.

Ice ridges deteriorate aerodynamics performance and endanger flight safety. With structural grid and central finite volume method the airflow field is obtained by solving N-S equations. Lagrange method is employed to get the trajectory of water droplets. Based on Messinger icing thermodynamics model, icing accretion on the leading edge of NACA 0012 is simulated and the results agree well with the experiment data, which verifies the feasibility and accuracy of the method. Based on this, on the conditions of small and large water droplets, the effects of ambient temperature and liquid water content on ice ridges are studied respectively. Besides, with the plasma phenomenological model the effects of nanosecond-pulse dielectric barrier discharge plasma actuator on ice ridges are studied. The results show that: whether small or large water droplets, the decrease of ambient temperature will cause the ice ridge height to increase and the increase of liquid water content will lead both the height and the range to increase; in the case of the small water droplets ambient temperature has little influence on the range; however, in the case of large water droplets the increase of ambient temperature will lead the range to increase; nanosecond-pulse dielectric barrier discharge plasma actuator can remove the ice ridges caused by the small water droplets and can postpone the position of ice ridges caused by the large water droplets.

The aerodynamic drag reduction characteristics of superhydrophobic structures were numerically simulated using Reynolds averaged N-S equations and SST k-ω turbulence model. The drag reduction mechanism, flow field and main factors affecting drag reduction characteristics are studied. The results show that the vortex in superhydrophobic structure reduces viscous drag and produces pressure differential drag. The drag reduction characteristics of superhydrophobic structure is the result of the interaction of these two forces. The free shear area ratio and the free stream Mach number are the main factors affecting the drag reduction characteristics of super hydrophobic structures. The drag reduction rate of superhydrophobic structure increase with the increase of free shear area ratio, decrease with the increase of Mach number in the given Mach number range.

The present study is concerned with the effect of Görtler vortices on hypersonic boundary-layer transition over concave surfaces. Numerical simulations and linear stability theory are performed to investigate the nonlinear development of Görtler vortices and interaction with Mack modes. The result indicates that the transition location from the linear growth to the nonlinear growth is affected by introducing Görtler mode at the inlet and different Mack modes, but the value of the linear growth rate and nonlinear rate differ a litter by different Mack modes. Transition location moves ahead when Görtler mode interacts with Mack modes. Whereas, the forward distance induced by Görtler mode differ very few to Mack modes with different spanwise wavelength. Thus transition advance by introducing Görtler mode has no obvious wavelength selection mechanism for different Mack modes under the scope of calculated spanwise wavelength.

This primary objective of this paper is to study the unsteady characteristics in the reflected shock wave/boundary layer interaction in an end-wall shock tube by a numerical simulation. It is discovered that the shock bifurcation mainly consists of a leading bifurcated shock, a recirculation region and a tail shock. Especially the shape of the entire recirculation zone seems like a crooked earthworm and the vortices in the rear of recirculation region continuously drop off from the walls when the shock bifurcation moves forwards. By analyzing the flow turbulent properties in the bifurcation, the turbulent intensity in the stream-wise direction is dominant in both value and direction, and the distribution of Reynolds shear stress becomes obviously asymmetric, which indicates an underlying instability of the shock bifurcation.

An intake is a device for supplying external air into the internal engine of an aircraft. In order to reduce the radar reflection on the engine face for reducing the observability, intake has an S-shaped duct. When applied to s-duct of the fuselage, the uniformity of the flow at the engine face varies depending on the state of flow at the inlet of the duct. In this study, the effect of active flow control at the inlet of S-duct was analyzed using a commercial computational fluid dynamics tool. The purpose of this study is to investigate the effect of the active flow control factors for an RAE M 2129 S-duct. The study was conducted by changing the position, length, and mass flow of the active flow control, and the mass flow was within 1% of the s-duct internal mass flow. The OLHD method was applied to 15 models for three factors and the k-ω SST turbulence model was used to predict the flow separation and secondary flow by adverse pressure gradient. In conclusion, the performance of the S-duct was found to have the greatest influence on the position of the flow control, and it was confirmed that the performance was the worst when the position was 0 in the duct standard.

The laminar interaction hypersonic flowfield induced by the sweep blunt fins, which are mounted on the flat plate, is numerical studied. Finite volume method and multi-block patched meshes are used to solve three dimensional compressible Navier-Stokes equations. The influence of grid density, numerical schemes and limiter functions are studied. Spatial structure and surface flow separation pattern are descried and the influence of different sweep angles is also given. The results show that grid numbers, numerical schemes and limiter functions have huge impact on the computation results. On the basis of the computational stability, high density grid numbers, low dissipation numerical schemes and limiter functions are preferred to obtain a proper result. For the interaction flowfield, the extent of the separation region and the peak value of the heat flux are decreased as the sweep angles increase.

The turbulence-transition model developed from the RANS methods has become the most used model in studying flow transition of hypersonic flight vehicles since it overcomes the flaw of stability analysis method that large-scale distributed parallel computing is not applicable. The current turbulence-transition model makes a numerical correction on compressibility in the prediction of transition within the hypersonic boundary layer, resulting in poor generality, so it needs to be further improved and perfected. Targeting this, the equations for the dimensionless boundary layer of compressible flow with pressure gradient are established, with the velocity profile and temperature profile similar outcome of the laminar plate obtained using the local similarity method. The influence of Mach number and pressure gradient on key parameters such as momentum loss thickness and shape factor are explored. The simulation results show that pressure gradient is the key factor for judging transition of the boundary layer. A new fitting equation for determining hypersonic transition model is established.

In this work we investigate evolution process of local perturbations in Mach 2.25 flat plate laminar boundary layer by a high-order discontinuous Galerkin spectral element simulation. The perturbations are produced by nanosecond pulsed dielectric barrier discharge (NSDBD). An empirical phenomenological model is applied to define the fast gas heating which is the dominant effect of NSDBD to surrounding air. The simulations indicate the fast heating produced local high frequency fluctuations of pressure and velocity which are distributed and propagated almost along boundary layers. These are different to the half circular cylinder compression wave with a tail in almost quiescence free stream.

In this paper, a parametric design method of point-line-arc was introduced firstly, which could analyze the influence of design parameters on the aerodynamic performance of high bypass ratio separate exhaust system. And the sensitivity analysis is carried out by perturbing different parameters, such as the angle of the cowling, the length of the cowling, the outlet position of the external bypass, the turning position of the external bypass and the length of the tail cone. The numerical simulation method was used to compare the influence of the five parameters on aerodynamic performance. Then, a rapid calculation method was developed for obtain the engine performance parameters in a very short time. Those five parameters and the corresponding large number of aerodynamic performance results were taken as input data, and those data were trained by artificial neural network. After that, a new combination of geometric parameters could be input and the corresponding aerodynamic parameters could be output through the artificial neural network in a few seconds and a good fitting result is obtained. Finally, a set of original geometric parameters were input to calculate the fitting data, the result shows that the largest fitting error is only 0.51%, which indicates that the error of fitting method is very small.

The numerical investigation has been conducted to analyze characteristic of RCS jet interaction on a reusable hypersonic vehicle. The numerical simulation method suitable for jet interaction flow fields is established based on finite volume method and the multi-block patched grid technique. The typical hypersonic wind tunnel test results are selected to verify the reliability of the numerical method applied in this study. The complex jet flow field structure, including spatial and surface interaction range, produced by a single RCS engine in yaw channel are visualized in detail. The variation of the lateral force and the yaw moment amplification factors with the angle of attack, height and Mach number are analyzed. The results show that the RCS jet interaction of the yaw channel has a remarkable enhancement effect on its control ability, the greater the dynamic pressure, the bigger the lateral force and the yaw moment magnification factor.

High-temperature effects need to be considered for a better design of hypersonic and reentry vehicles. They affect both the boundary layer flow and its flow transition, whose primary stages can be investigated through modal stability analysis. In this work, physical and numerical tools for high-temperature flows are presented and the efficiency of the new developed in-house boundary layer and stability solvers is tested. Specially, we focus on the stability of a flat plate flow in thermochemical non-equilibrium through an investigation of growth rates under the influence of various flow parameters.

A numerical simulation was carried out in order to investigate the effects of Reynolds number on performance in centrifugal compressor stage. The simulation results show that total-to-total pressure ratio, efficiency decrease 0.18% and 4.3% respectively as the Reynolds number drops. The separation flow under low Reynolds number is more severe. The interaction in impeller among blade surface secondary flow, main tip leakage flows from main blade and splitter blade transfers more circumferentially, leading to higher value and larger region area of entropy. The flow inhomogeneity along the blade height is enhanced at diffuser inlet, and the interaction between the passage wake from impeller and the separation flow from pressure surface secondary flow leads to the high entropy region grows larger and forms the backflow at the rear part of the blade pressure side. The effect of Reynolds number mainly focused on tip leakage flow and passage wake in impeller and wake in vaned diffuser.

Numerical investigation of dielectric barrier discharge (DBD) plasma actuation has been conducted to quantify the momentum transfer without and with the incoming flow over a flat plate. A comparative analysis between numerical investigations and experiments is presented. A phenomenological model is used to simplify the effect of plasma on air flow based on the momentum balance equation. For steady actuation, the plasma will produce a steady jet close to the wall, the effect region of plasma is gradually getting larger and the maximum velocity of the jet is reducing along the chord of the actuator. The main way of flow control is to inject momentum into incoming flow, and the induced jet is continuously transferred with nearby fluid. For unsteady actuation, the plasma will produce a series of vortex structures and the intensity of them decreases gradually along the chord of plasma actuator. The mechanism of unsteady plasma actuation is both the momentum and vorticity injection.

Aircraft icing damages the original aerodynamic shape, especially the ice accretion on the leading edge of the wing and tail fins, which may have serious impact on the flow field and aerodynamic characteristics. For the numerical simulation of flow field and aerodynamic characteristics around the ice configuration, based on the basic LBM model, constructing LBM algorithm based on multi-layer partitioned rectangular gird technology, establishing boundary condition processing formats and calculation methods for different areas such as object, far field and layered, reducing the number of grids effectively and improving computation efficiency. Simultaneously, a Smagorinsky eddy viscosity model combining multiple relaxation time method and large eddy simulation has accomplished the numerical simulation of flow field around airfoil and multi-element airfoils with ice at a high Reynolds number. By comparing the calculation results with the experimental data, the feasibility and accuracy of the calculation method is verified. The results of the study show that the iced airfoil increases the drag, reduces the lift and angle of stall. The flow field has significant unsteady flow characteristics. High efficiency and accurate forecasting of the unsteady dynamic characteristics in the ice configuration flow field is the key to study the influence law after ice accretion.

In the preliminary design of the wing ice protection system, the external thermal load estimation of the wing is studied. Based on the principle of heat transfer and mass transfer and some simplified assumptions, the theoretical model of thermal load of electrothermal ice protection system was established. Based on the secondary development of FENSAP software, the skin surface thermal and mass balance simulation was carried out according to the external environment and existing conditions. A method of determining the ice - resistant thermal load by boundary condition iteration is proposed. According to the design conditions, the thermal load calculation results of the two models are compared, and the error is as low as 11%. Then, in the flight envelope, 20 typical working conditions were selected to verify the theoretical model. Through the analysis of the heat flow items, the trend of heat load is consistent with the theory of heat and mass transfer, which proves that the calculation model is suitable for the whole flight envelope. When the wing ice protection system works at the specified power, the temperature of the skin surface of the aircraft in the flight envelope varies between 1.7 °C and 15.4 °C, which ensures the safety and energy conservation.

This paper presents a code to simulate rime and glaze ice accretion. A Shear Stress Transport (SST) $$ k - \omega $$ model with roughness extension is developed to directly calculate the convective heat transfer coefficient. The airflow field is solved based on Reynolds-averaged Navier-Stokes (RANS) equations and the roughness extension of SST k– $$ \omega $$ model. The droplet flow field is solved based on Eulerian method. Ice accretion model based on Stefan problem was solved to simulate the film flow on the icing surface. This roughness extension model is applied to predict the skin friction coefficient and Stanton number of the MSU plate. The results are in good agreement with the experimental data. Furthermore, rime and glaze ice accretion simulations over NACA0012 airfoil are accomplished. Results show that the ice shapes of both rime ice and glaze ice agree well with the result of the IRT experiments.

A 3D glaze icing numerical method for aircraft and aero-engine entry Components was presented. The Eulerian - lagrangian model was employed to simulate the air- water two phase, and the distribution of impingement droplet was obtained. Mass and energy balance were established in a control volume, then ice accretion code was completed using UDF (User Defined Function), which was provided by FLUENT. This program can calculate glaze icing, Rime icing and mixed icing. The calculation of icing on NACA0012 3D airfoil is carried out, and compared with the experimental results in literature. The results show that the maximum icing thickness is in good agreement with the experimental results, and the predicted ice shape is consistent with the experimental ice shape development trend. The calculation results show that the proposed icing calculation method is effective and reliable. On the basis of this, centre conical body is calculated, the qualitative analysis of results. It can be seen from the calculation results that three typical ice types are formed near the stationary point of the centre conical body under different inlet temperature conditions. It is also found that the temperature has not a great effect on the amount of icing. The research results provide an alternative way for the calculation of three-dimensional icing formation for aircraft and aero-engine, and can be used to improve the design technology of anti icing system.

In this paper a numerical solution method for compressible Euler equations on arbitrary grids with high order boundary elements, based on discontinuous Galerkin space discretization, is presented. Unstructured grids with both triangular and quadrilateral elements are utilized, which contain high order curved edge elements near wall boundary and low order straight edge elements elsewhere. In this study, the polynomial basis functions up to third order are utilized, and the inter cell fluxes are calculated with HLLC approximate Riemann solver. This paper includes steady state results for a circular cylinder flow in two dimensions, the results are compared with different grids and different boundary discretization orders. The accuracy of this method as well as the needs of high order boundary discretization are demonstrated.

Tri-linear interpolation method is the most popular approach in data transferring between overset grids due to its briefness and robustness, however it is not sufficient to guarantee the conservation of flow variable. An improved tri-linear interpolation method for overset mesh is presented. A cell-cut algorithm is applied to construct a local supermesh in the overlapping grid area, which could be used to improve the conservation of flow variable. The number of donor cells is expanded reasonably to improve the accuracy of interpolation, and a volume weighted method is applied to eliminate interpolation error coming from the grids mismatch in the overlapping grid areas. Numerical results show that compared with original tri-linear interpolation method, the improved method could reduce the numerical errors, interpolate the flow variables more accurately. For linearly and non-linearly distributed flow variables, the improved tri-linear method is more conservative because it has inherited the conservation of supermesh, and it is robust and engineering-practical for complex engineering applications.

A numerical simulation with RANS method in riblets drag reduction is performed in this paper. Inspired by the roughness model, modeling of the riblets is developed about ω in k-ω SST turbulence model. Based on current experimental data, a simplified method to reproduce an up-shift of the logarithmic region is proposed to represent the drag-reduction behavior of riblets in the wall region. Pressure gradients and geometry curvature are modeled in the establishment for realistic configurations by introducing the Clauser parameter β. The results show that the modeling method could simulate the drag and flow mechanism. By comparing with the experiment database on airfoils with riblets, it’s proved that the modeling can predict the drag reduction in complicated wall boundary conditions and adverse pressure gradient. Cases are provided for the estimate of riblets drag reduction in engineering applications.

Based on Maximum Principle analysis, a unification of multi-dimensional limiter construction for unstructured finite volume discretizations was presented. To enhance flow resolution, a new type of multidimensional limiter was developed with special limiting strategies, and was compared with various conventional limiters. Numerical experiments showed that due to less restriction in gradient reconstruction, the new limiter can preserve designed accuracy for smooth flow. For complex flow with shock and contact discontinuity, it is less dissipative and has better flow resolution.

Low diffusion shock-capturing schemes are known to produce numerical shock anomalies in hypersonic flow computations. In the current study, we conduct a stability analysis of the Roe approach. It has been found that the shock instability can be effectively suppressed if enough entropy production can be guaranteed inside the numerical shock structure. To improve the robustness of the Roe solver, a modification is proposed to enforce enough entropy production of the numerical flux function. Numerical results show that the proposed scheme is quite robust against shock instabilities.

Although shock-capturing methods have been well developed in the past few decades, there are still some challenging issues that need to be addressed with caution, especially in hypersonic flows. It is well known that Godunov-type schemes will suffer from shock anomalies such as carbuncle phenomenon and post shock oscillations for hypersonic flow computations. What is worse is that numerical schemes which have minimal dissipation on linear waves are more prone to the shock instability. To circumvent the instability problem, a rotated methodology is proposed to combine two interlinked Godunov-type schemes, i.e. a low dissipative HLLEM scheme and a shear dissipative HLLEC scheme. It is demonstrated that the new scheme can be implemented in a simple and economic manner in the form of the Roe solver, which provides comparative resolution to Roe-type and Osher-type schemes while being free from the shock instability. A stability analysis is conducted to verify its robustness against strong shocks. Numerical results of several carefully chosen strong shock wave problems are investigated to demonstrate the robustness of the proposed method.

An improved unsteady preconditioning boundary condition is developed for solving the tiltrotor aircraft’s hovering flow problems of different speed scales. On this basis, the set of numerical methods are validated, and the principle of preconditioning parameter selection is determined. Then through it, the wake vortex structure evolution of a 0.25-scale tiltrotor model is carefully analyzed, and the effects of “fountain effect” and “ground effect” on the aerodynamic performance of tiltrotor are revealed and compared. The results show that “ground effect”, coupled with “fountain effect”, inhibits the evolvement of rotor wake. The “ground effect” affects the whole azimuth region, making the thrust coefficient of the installed rotor slightly increased by 5.26% compared with the isolated rotor, but the total thrust coefficient of the tiltrotor aircraft is only 82.46% of the isolated rotor’s. The conclusion shows that the novel method proposed here is effective to simulate the complicated flow field in the hovering configuration, and the flow mechanism found here is of referential. Aerodynamic interaction has a significant impact on the overall aerodynamic performance of the tiltrotor aircraft.

Bypass and laminar-separation-bubble induced transition phenomena are very common in turbine blade cascade flow with high turbulence environment. In order to predict the transition position and aerodynamic characteristics accurately in high turbulence environment, the similarity velocity profiles at various shape factors are obtained by solving the Falkner-Skan boundary layer similarity equations. Then, the local formulas of non-local variables are fitted based on the similarity velocity profiles and combined with a new indicator factor to form transition criterion, so as to construct the turbulence transition model. The transition model was used to predict and simulate the flows of zero pressure gradient plates, T3C series pressure gradient plates, and PAK-B turbine blade cascade. The predicted results under various turbulence and pressure gradients are all according with the experimental data. It proves to be that the model proposed in this paper is reasonable and feasible.

A novel scheme is proposed to reduce numerical dissipation error based on the third-order weighted essentially non-oscillatory (WENO) scheme. The improvements include two parts: firstly, an additional candidate stencil is added downwind to the original WENO scheme, which aims at decreasing the dissipation of the linear part of the scheme; secondly, a new nonlinear weighting procedure, which is similar to the third-order WENO-Z scheme with a new fourth-order global smoothness indicator, is conducted. Numerical results on four cases in hypersonic flows demonstrate that the novel scheme has better performance on resolving complex flow structures and predicting heat transfer while maintaining the numerical stability at discontinuities compared with original WENO scheme.

Based on the structural overlapped grid technology, the aerodynamic characteristics of DLR-F11 model, which is issued by the Second High Lift Prediction Workshop (HiLiftPW), are predicted by self-developed PMB3D solver. PMB3D is a RANS-equation solver, employing finite-volume discrete and implicit method. The computational accuracy of PMB3D solver with structural overlapped grid is verified by comparing the simulation results with some other CFD solvers, such as FUN3D, CFL3D and OverFlow et al. The simulation results demonstrate that (1) PMB3D could rival the other CFD solvers, since it can precisely predict the aerodynamic characteristics, the pressure distribution and the flow characteristics of the model; (2) the use of the structural overlapped grid technology could largely reduce the difficulty of grid generation for complex geometry, without affecting the accuracy of the flow simulation; (3) PMB3D with structural overlapped grid could quickly and accurately predict the aerodynamic characteristics, and it is a reliable analysis tool for aerodynamic design.

Taylor-Culick flow is of importance on the interior ballistics of solid rocket motor (SRM). The combustion of solid fuel results in the motion of the wall boundary of the flow. Therefore, the understanding of Culick-like flow with a regressing wall plays a significant role in the study of SRM. However, the complex geometry of real fuel surface of SRM leads to considerable difficulty on the analysis and numerical simulation. In this paper, Navier Stokes equations coupled with Level-set equation are resolved to mimic the flows inside the combustion chamber with grain regression of the propellant. The four typical geometrical structure which corresponds to classical propellant grain shapes are simulated. The results demonstrate that the different geometrical structure in associate with flow direction could significantly affect the flow inside the combustion chamber. Besides flow turning could be substantially changed based on different geometrical configuration, isolated vortex could also be introduced under the influence of different injection schemes. Considering both flow turning and vortex play an important role in the combustion instability of the SRM, the geometrical design of grain should be taken additional care except for the traditional requirement of thrust.

The laminar separation phenomenon on SD7003 airfoil under the low Reynolds numbers condition was investigated by large eddy simulation. Based on the structured patched mesh, the numerical method adopted the implicit sub-grid-scale model, and utilized the AUSM scheme for spatial discretization and dual-time-step method for time marching. The numerical simulations captured the instantaneous flow structures clearly. Besides, the averaged and pulsed flow field was obtained by statistics. The computed results agree well with the computational and experimental data in literatures. The comparison with the transition model further reveals that, the large eddy simulation method is capable of simulating complex flow phenomena of separation, transition and reattachment accurately. And the mechanism of separation induced transition is discussed.

Numerical simulations using ILES, DDES and URANS, respectively, have been carried out to study the wake flow of a circular cylinder at Reynolds number of 3900. The three-dimensional compressible Favre-filtered/averaged Navier-Stokes equations are solved using Roe scheme for flux difference splitting and WENO scheme for reconstruction. The Sparlart-Allmaras one-equation turbulence model is adopted for the DDES and URANS simulations. Comparative studies are carried out based on mean flow quantities, mean integral quantities and turbulence statistics. Based on comparison between the predicted results and experimental measurements, the ILES/LES, DDES using fifth-order WENO scheme with minimum dissipation are found to be adequate to predict the complex flow field. However, the URANS method is found to be not able to reproduce such complex flow. The influences of the SGS model, parameter ε and order of WENO scheme have also been investigated. It is found that the SGS model has a small impact on the LES simulation, and ILES could give results closer to the experiment than LES with a SGS model. The parameter ε and the order of WENO scheme are found to have a strong impact on the simulations. A larger parameter ε and higher order are preferable to obtain a better prediction.

The porous wall can delay hypersonic boundary layer transition by suppressing the second mode instability, which helps in reducing the weight, cost and complexity of thermal protection systems of hypersonic vehicle. A linear stability theory model is developed to investigate the performance of suppressing second mode instability. Based on the fourth Runge-Kutta scheme, the basic flow of hypersonic boundary layer is obtained using a shooting technique. The admittance of porous wall is considered in the boundary condition of stability equations. The second mode growth rate in temporal stability problem is solved by the Chebyshev spectral method. The prediction results agree well with the existing data. The porous walls with high porosity and relatively high aspect ratio provide maximal damping of second mode wave. Specifically, high porosity is associated with strong attenuation effect. Relatively high aspect ratio corresponds to relatively shallow pores, which results in a cancellation effect.

Accurate simulation of high Reynolds number flow around airfoils beyond stall is a challenging computational fluid dynamics (CFD) problem of significant importance in aerospace industry. In this study, detached eddy simulations (DES) have been performed on the flow around NACA0021 airfoil beyond stall at 60° angle of attack. Detailed investigations are carried on the effect of time step, the effect of span-wise length and the effect of grid resolution, respectively. The present DES results show good agreements with the experimental data and some other published pioneering work. The DES model in OpenFOAM shows a good capacity in the prediction of unsteady flow with separation and vortex shedding. Compared to the unsteady Reynolds-averaged Navier-Stokes (URANS) simulation, DES simulations show visible improvements on the prediction of force coefficients and the flow structures.

Many fluid flows of engineering interest, though very complex in appearance, can be approximated by low-order models governed by a few modes, and Dynamic Mode Decomposition (DMD) has been proved effective in analyzing the coherent structures of complex flows. In this article, we present the formulation and design progress of a parallel dynamic mode decomposition program, especially the parallel I/O strategy, as a significant supplementation of parallel dynamic mode decomposition algorithm presented in others literatures. Parallel I/O performance with different data block size and processor number is demonstrated with a 6.9 GB file generated by 1001 snapshots of Rayleigh-Taylor instability flow. Analysis of flow structure and spatio-temporal coherent structure are performed by Fast Fourier Transformation (FFT) and dynamic mode decomposition for flow field resulted from high order weighted essentially non-oscillatory (WENO) schemes. For test case of Rayleigh-Taylor instability flow with Atwood number A = 0.5, we find a significant phenomenon that WENO9 with very fine grid (h = 1/1920) exhibits the characteristic of large unsymmetrical bubble-like plumes, but the others take the form of symmetric bubble-like plumes.

The fixed geometric inlet only has better performance near the design Mach number, and it is difficult to meet the demand of the combined cycle engine. A multistage adjustable variable geometry inlet with rotating cowl and ramp is proposed. The numerical simulation of variable geometry inlet with rotating cowl and ramp is performed by solving Reynolds Averaged Navier-Stokes (RANS) equations with dynamic mesh technique. The cowl and ramp rotating angular velocity have been modified using a parametric user defined function (UDF) that covers the whole range of operating points. Two methods (Spring-based Smoothing and Local Re-meshing) have been used to achieve the mesh deformation and re-meshing. The general performance of rotating cowl and ramp are investigated, respectively. The variable geometry scheme for the hypersonic inlet under different flow conditions has been studied by two-dimensional numerical simulation. Results indicate that both the cowl and the ramp rotating with constant angular velocity can help to enhance the self-starting ability. The variable geometry inlet with rotating cowl and ramp has better performance in the entire operating range. Especially, it can effectively reduce the inlet start Mach number and the inlet can self-start at Ma2. The variable geometry inlet has better flow capture ability. The mass flow rate coefficient comes to 0.53 at Ma2. The multistage adjustable variable geometry inlet scheme effectively solves the contradiction between the start and flow capture capability at low Mach number. Moreover, the variable geometry scheme for hypersonic inlet is simple and feasible in engineering application.

Morphing vehicle would change aerodynamic configuration in real time with the flight condition changing to keep the best performance in the whole flight envelope, which would result in strong flow unsteadiness and the response problem between flow and the airframe. Aiming at these problem, multiple methods were applied to study how the flow would change during the morphing. Unsteadiness brought by the supercritical airfoil thickness and camber continuous deformation was studied with computational fluid dynamics (CFD). The results show that deformation of airfoil thickness and camber would change the pressure distribution obviously. And the unsteady effect is strong with distinguishable lift and drag hysteresis loops. As the morphing frequency and amplitude increase, the unsteadiness would be strengthened. The unsteady effect resulted from the camber deformation is relatively stronger than that from the thickness deformation. The research show that the unsteadiness come from the hysteresis between the flow structure (location/strength of shock and the boundary separation during) and the airfoil morphing.

Differing from Large scale rotor, nano rotor always rotates at an ultra-low Reynolds number, which leads to the rapid deterioration of aerodynamic performance. The traditional aerodynamic shape optimization method fails to improve the rotor propulsion performance at such a low Reynolds number. To improve the rotor propulsive performance of the nano rotor, an investigation on the unsteady flow mechanism of a bio-inspired nano rotor is carried out. A pitching-blade motion is introduced to the rotating rotor based on the delayed stall mechanism inspired from the flapping insect. Aerodynamic model of the pitching-blade rigid nano rotor is established using the Multiple Reference Frame (MRF) technique. Then the propulsive performance of the nano rotor with and without bio-inspired motion is studied with computational fluid dynamics solver based on k-epsilon turbulence model, respectively. The influence of pitching frequency on the rotor propulsive performance is investigated as well. Results indicate that the pitching-blade motion of the nano rotor improves the aerodynamic performance of nano rotor and the improvement augments with the increasing of the pitch frequency. Flow field results show that the evolvement of the LEV is the main unsteady flow mechanism for the improvement. This research will profit for the design of small rotor.

The aerodynamic behaviour of transonic flow around a supercritical aerofoil is strongly influenced by shock-wave/boundary-layer interaction (SBLI) due to compressible and viscous effects. SBLI causes undesirable effects in various manners including flow instability, drag rise, and buffet, which crucially limit the flight envelop hence operation. In this paper, a numerical investigation is conducted for an OAT15A supercritical aerofoil under a typical buffet onset condition. Unsteady Reynolds-Averaged Navier-Stokes (URANS) equation is used to simulate the compressible, viscous flowfield. A two-dimensional (2D) surface bump based on preceding research on SBLI control is employed as a flow control device. It is placed on the suction side of the aerofoil relative to the shock position, with a fixed location of 27% of the chord length. A freestream condition of Mach 0.73 and a 3.5° angle of attack have been considered for the unsteady flowfield. It has been found that the trailing edge vortices within the separation bubble have considerable influence on self-sustained shock oscillation by scrutinising the flowfields in the presence/absence of bump control. The establishment of a λ-shock structure effectively restricts the motion of the front shock leg without incurring significant re-expansion generated by the moving rear shock leg. This subsequently suppresses flow separation at the trailing edge within an acceptable range, and attenuates the periodic lift fluctuation associated with the oscillating shock movement.

For a closed-circuit wind tunnel driven by a multiple-stage compressor, continuous mode test can achieve higher efficiency and better data density in comparison of pitch/pause mode test, but fulfilling flow accuracy requirement using the rotation speed of compressor as the only actuator during pitch variation is a challenging task because of the system lag caused by the fluid transport delay. A new method which integrates feed forward and feedback algorithms to accomplish continuous mode test in a 0.6 m × 0.6 m transonic wind tunnel is introduced in this paper. Feed forward parameters is deduced from ‘compressor speed-pitch angle’ curve based on either calculation or empirical data. In the experiments using the new method under the required conditions (Mach number range: 0.6–0.9, speed of pitch angle: 0.2°/s or 0.1°/s, pitch range: −4°–10°), Mach number is well controlled within the error band of ±0.001 in the continuous mode test, and the duration of continuous mode test is much less in comparison of pitch/pause mode test.

Symmetrical aerospace configuration design has greatly increased currently. Compare with conventional aerospace vehicles, those aerospace vehicles’ aerodynamic forces have the characteristics of strong unsteady, nonlinear and longitudinal/directional/lateral coupling. To study those characteristics, accurate prediction and accurate measurement the dynamic stability derivatives of the vehicles is strongly needed. One important tool is dynamic derivative wind tunnel test. With these issues in method, the difficulties of dynamic derivative wind tunnel test in hypersonic wind tunnel will be solved, and a forced oscillation test rig for measure dynamic derivative of hypersonic vehicle will be built here. The test rig has been completed with a symmetrical hypersonic aerospace model under Mach number 5.0 to 8.0, and the direct dynamic derivative and coupling dynamic derivative has been obtained. The test result shows that: comparing the direct dynamic derivative results of forced oscillation test with free oscillation test, the relative deviation is less than 10%. The new test technique can meet the requirement of the dynamic derivative test of hypersonic vehicles.

Multistage compressor’s experiment performance often differs from its numerical simulation, because the real flow field is much more complicated than the CFD’s assumption of steady and uniform cascade inlet flow. While CFD cannot provide a reliable prediction for multistage compressor, analyzing experiment’s data could possibly be the only way to directly diagnose the compressor’s flow field and performance. Using the testing data between rotors and stators, this program can solve the S2 stream surface field which is hard to directly observe in the compressor test rig. By calculating meridian field with experimental and formulated model, the solution reveals the essential performance parameter of each stage and each element such as flow coefficient, loading coefficient, efficient, etc. These solutions are the base of figuring out the weakest design of the compressor flow field, and the result reveals the subsequent modification quantitatively. Compressors performance can hardly reach the design goal without several such trial and error iterations.

Near-space hypersonic vehicles use a sharp leading edge and a large lift-to-drag ratio profile, and the aerodynamic heating environment is extremely harsh. At present, in order to improve its performance and reduce costs, hypersonic vehicles are in urgent need of low-redundancy thermal-protection structure design, and require ground arc heating assessment tests to have better accuracy and repeatability. Because the stagnation radius of the leading edge is too small, the traditional heat flux measurement technology cannot measure the heat flux density at the leading edge stagnation point. How to determine the tip front stagnation point heat flow in the arc wind tunnel condition is a technical problem to be solved urgently. In this paper, based on the experimental measurement and numerical analysis techniques, a method for identification and measurement of heat flux at the sharp leading edge stagnation point is proposed and verified by experiments. The results show that the method can obtain the heat flux distribution in the stagnation area of the leading edge with small leading edge radius, which has high credibility.

Advancement on the improvement of transonic-supersonic flow field quality and compatibility for complicated testing facility was conducted in FD-12 Tri-sonic wind tunnel of China Academy of Aerospace Aerodynamics (CAAA). A newly-built transonic test section was supposed to be versatile and will promote the flight testing ability while obtaining testing data with higher accuracy. This versatile platform based on a modular structure could be used for different complicated flight testing by replacing the inner core while the shell and sustentation are shared. Furthermore, new supersonic test sections incorporating nozzle were built to eliminate the assembly step difference. New nozzle curves were designed and manufactured with test section wall divergence angle. The flow field quality will be improved by eliminating the residual shock waves which were caused by nozzle curves design methods and the steps from assembling for nozzle and test section. Flow field calibration test had been conducted to check the flow field uniformity in these new transonic and supersonic test sections. The calibration result showed that the flow field quality had achieved the classic requirement absolutely or proximally. Both the versatile section and incorporated supersonic section had been used in flight testing and more and more test technologies were developed.

The spoiler of civil aircraft plays an important role in flight as one of the important control surfaces. Different from the high lift devices such as flap, slat and so on, the spoiler is used for reducing the speed or the altitude in flights, making roll maneuver with aileron, and breaking lift on the ground. Moreover, the deflection of spoiler is usually reached a large value, the aerodynamic load acting on the spoiler can be remarkable due to the large designed deflection angle. Thus, it is necessary to research the load acting on the spoiler in the design of civil aircraft load. This paper introduces the principle of the spoiler load of civil aircraft, and gives an engineering method for the spoiler load design of civil aircraft.

An experimental study of interaction flow induced by a perpendicular blunt fin was presented. The tests were carried out in a hypersonic gun tunnel with freestream Mach number of 6. Fluctuating heat-transfer rate was measured along model centerline and high speed schlieren photos were also taken to show complex interaction flow characteristics. Steady and unsteady features of turbulent interactive flowfield are presented. The results show that the fin induced separated flow is very complex. The distributions of the heat transfer rate are changed obviously due to the fin interaction. In the streamwise direction on the plate centerline two peaks are observed. Unsteady characteristics is very obvious. In the incipient separation region ahead of the fin the instantaneous heat transfer is highly intermittent. In the separation region, amplitude of the fluctuations of the heat transfer rate becomes extensive.

The hypersonic separated flowfield features induced by cylindrical protuberance which was mounted on the flat plate were studied experimentally. The double-color oil flow method with real-time oil supplying is developed to clearly visualize the surface flow features of complex separated flow at hypersonic speed. The tests were performed in FD-03 hypersonic blow down wind tunnel at Mach number 5 by using the color oil flow and schlieren technology. The surface and symmetry plane flow characteristics are studied. The results show color oil flow technology can provide a clear visualization of separated flow induced by cylinders. The color oil flow technology distinguishes the incoming flow and separated flow more clearly than the conventional single white oil flow technology. The information of separated flow characteristics was obtained in the present study and the effect of different protuberance height was presented. The height of the cylindrical protuberance has huge impact on the range of separation region.

This paper introduces the experimental study of the thrust vector nozzle performance. On the basis of the calibration tank and original equipment, a nozzle entrance airflow measurement device, portable critical flow venturi-nozzle flowmeter and a reference nozzle are newly designed and processed, and a nozzle performance testing system in calibration tank is established. Using these devices, the thrust coefficient, the rudder deflection angle and nozzle flow coefficient of a stealth thrust vector nozzle are tested. The experiment results show that thrust characteristic measurement with the separate nozzle in calibration tank is consistent with the law theory. As a consequence, this set of calibration tank nozzle performance testing system can be used to test the thrust coefficient, effective vector angle and nozzle flow coefficient of vector nozzle. This study is significant to rapid test of preliminary design scheme of vector nozzle and its performance analysis.

As the angle of attack increases to a certain size, the induced flow field of a slender conical forebody will become asymmetric even the freestream is symmetrical. This paper is aim to study the characteristics of asymmetric vortex pair under steady and unsteady plasma actuation when the vortex flow field is non-bistable. The asymmetric vortex is controlled by alternating current dielectric barrier discharge (AC-DBD) plasma actuators. The effect of the different duty-cycle frequencies is studied by using particle image velocity (PIV) and pressure distribution measurements. Reynolds number based on the base diameter of the cone forebody is 40,000. The angle of attack is 28°. The duty-cycle frequency includes 5 Hz, 100 Hz, and 1000 Hz. The duty-cycle ratio changes from 0.01 to 0.99. Static pressure measurement results show that the local side force changes linearly with the duty-cycle ratio when the duty-cycle frequency is relatively low; it still represents almost linearly when the duty-cycle frequency is in a medium value; while it becomes clear non-linear but continuous when the duty-cycle frequency is relatively high. At last, the mechanism of control is analysed with the results of PIV.

A semi span structural dynamic FEM model and flutter model were built for a typical flying wing, natural modes and body freedom flutter characteristics were investigated. In order to model the typical support conditions in wind tunnel test, two vertical springs were attached to the model and the natural frequencies and flutter characteristics were calculated. To keep the model stable in the flow field during test, a winglet on the wingtip was built. It is found that the impact of suspension stiffness and wingtip on the flutter characteristics is so significant that even change the flutter type. The spring added on the plane can obviously increase pitching and plunging mode frequencies, and finally encountered plunging/pitching coupled flutter when pitching frequency is lower than plunging. The winglet could be regarded as a additional mass on the wingtip that caused the increasement of inertia, as a result, the pitching frequency became lower. This research is relevant to the suspension design of flying wing’s body freedom flutter wind tunnel test model.

An experimental study on the circulation control around a modified NCCR1510-7067N airfoil is presented. The plasma-induced wall jet is achieved by the use of Dielectric Barrier Discharge plasma actuators, which are placed around the curved circular trailing edge. A wind tunnel test is performed in order to quantify the effects of applied voltage and actuator position on the lift enhancement. The lift and pressure measurements are taken using micro pressure electronic scanning valve. Additionally a smoke flow visualization test is performed in order to elucidate the fundamental working principles of the concept. Preliminary results indicate that the variation of the lift coefficient with the applied voltage was nonlinear, the short laminar separation bubble appeared at the leading edge in advance with applied voltages above 8 kV; consequently, the optimal applied voltage was 8 kV. The lift coefficient was increased by 0.154 and the lift augmentation efficiency reached 108.7 with a single actuator on the lower surface. A smoke flow visualization showed the delay of separation position and the formation of virtual flap effect.

An experimental study is performed to characterize the static hysteresis of the lift and drag on an FX 63-137 airfoil at the chord Reynolds number of $$ 1.8 \times 10^{5} $$ . A long strip of plasma actuator is installed on the upper surface near the leading edge of the airfoil. The purpose of this work is to study the effect of plasma flow control on the static hysteresis of lift and drag at low Reynolds numbers. Pressure measurements are used to determine the surface pressure distribution around the airfoil. Lift and drag forces acting on the airfoil are calculated from the measured pressures and normalized by the chord. The detailed surface pressure distributions over the baseline airfoil reveal that it is the laminar separation bubble (LSB) on the upper surface delays the stall for the forward process of the angles of attack (AOA); while for the backward process of AOA, the flow cannot establish the same laminar separation bubble as the forward process, resulting in the static hysteresis of the lift and drag. The plasma flow control results show that static hysteresis of lift and drag can be reduced by affecting the LSB at low Reynolds numbers.

An anti-icing application of AC-SDBD plasma actuator has been explored through experiments. The purpose of this study is to study both the thermal and aerodynamic effects on plasma anti-icing. Two types of multi-SDBD actuators were designed with different induced flow directions. One type (Type-1) is designed to generate an induced flow with direction same to the incoming flow, while the other (Type-2) is used to generate an induced flow with an opposite direction to the incoming flow. The anti-icing tests were carried out in an icing-wind-tunnel. The ice accretion and corresponding surface temperature have been recorded using a high-speed digital camera and infrared imaging system respectively. The results show that the Type-1 actuator can ensure that the leading edge of the airfoil remained smooth and did not have the ice accretion, while the Type-2 actuator had ice accumulation on the leading edge, but it can effectively postpone the ice location on the upper surface of the airfoil. Such results show that the anti-icing effects are directly related to both the thermal and aerodynamic effects of the DBD plasma actuator.

Mitigation of sonic boom to an acceptable stage is a key point for the next generation of supersonic transports. Meanwhile, designing a supersonic aircraft with an ideal ground signature is always the focus of research on sonic boom reduction. This paper presents an inverse design approach to optimize the near-field signature of an aircraft making it close to the shaped ideal ground signature after the propagation in the atmosphere. Using the proper orthogonal decomposition (POD) method, a guessed input of augmented Burgers equation is inversely achieved. By multiple POD iterations, the guessed ground signatures successively approach the target ground signature until the convergence criteria is reached. Finally, the corresponding equivalent area distribution is calculated from the optimal near-field signature through the classical Whitham F-function theory. To validate this method, an optimization example of Lockheed Martin 1021 is demonstrated. The modified configuration has a fully shaped ground signature and achieves a drop of perceived loudness by 7.94 PLdB. Finally, a non-physical ground signature is set as the target to test the robustness of this inverse design method.

The coupled aero-structural adjoint-based approach for large civil aircraft multi-disciplinary optimization is studied in this paper. Firstly, computational techniques are introduced. Then the coupled aero-structural adjoint system (CASA) is developed and constructed, solving the structural adjoint equations derived from the structural static equations, with the help of the parallel flow adjoint code PADJ3D and lagged coupled adjoint method. Afterwards, the multi-disciplinary optimization model for the large civil aircraft is established, while the sequential quadratic programming algorithm is used for optimization, freeform deformation method is employed for geometric parameterization and parallel RBF-TFI grid reconstruction technology is used for mesh perturbation. On accounting of the coupled aerodynamic and structural disciplines, the multi-disciplinary optimization for large civil aircraft wing is carried out based on the developed CASA system. The simulation results demonstrate the effectiveness of the CASA system. Under the stress constraints, the aerodynamic drag and the structural weight can be effectively optimized.

Formation flight is usually seen in birds’ migration and military operations, and will be more often in unmanned aircrafts applications in future. It is known that drag reduction and fuel saving can be achieved in formation flight and positional parameters are important to this effect. But few literatures about efficiency of formation flight types on this effect are published. Wake vortexes of the leading aircraft in formation flight are assumed to be a pair of reversed vortexes. Using potential flow techniques, models of the trailing aircraft’s lift and drag coefficients are constructed which are functions of lateral spacing and vertical spacing. An estimation of energy saving is conducted on three formation flight types-V mode, Λ mode and echelon mode. The formation studied consists of three the same aircrafts. The total drag of three aircrafts in a formation is regarded as an index for energy saving efficiency under a precondition that all aircraft’s lift equals to their own weights. Each type is under its best status. Results show that echelon mode has the highest energy saving efficiency, Λ mode is a little low and V mode is markedly lower than the two others. This result accords well with formation flight types used in birds’ migration.

Preliminary design of aircraft wings requires multiple cycles of optimization to compromise the influences from different disciplines and constraints. Surrogate-based optimization is a popular choice in this circumstance because surrogate models can be reused in different cycles once constructed, while gradient-based optimization might be less efficient despite its fast convergence in each cycle. However, surrogate-based optimization suffers from the curse of dimension. The selection of design variables has a big influence on the optimization efficiency and performance. We investigate two different approaches to drag minimization in the design of transonic wings: (1) optimize all the section shapes and their twists independently (2) scale all the wing sections together and optimize them with their twists. We find that surrogate-based optimization by the former approach cannot ensure a better solution in spite of its larger design space. The latter approach is generally more efficient, which is more practical in preliminary design of wings.

The paper present a supersonic aero-structural optimization method based on the gradient information obtained by adjoint approach. The adjoint approach works on the acquisition of the partial derivative matrix and the iterative operation of the matrix. A swept wing model is studied to illustrate this methodology, in which the objective function is the lift-to-drag ratio considering the elastic deformation of wing structure, and independent variables are the thickness of the skin and the size of the spar. The results show that the higher lift-to-drag ratio can be achieved, while the weight of wing structure keeps light.

The future development trend of helicopter will possess the function of high speed, far distance and high ceiling. Now, the main development trend of compound helicopters at home and abroad are the conventional configuration with auxiliary propulsion unit and the wing. But for this configuration, the force on the parts are very different from the ordinary. To solve this problem, it need to ensure the allocation of the lift between rotor and wing in the stage of general design. In this paper, for Rotor/wing compound helicopter, build a calculation model of required power and put up a kind of lift distribution strategy between the rotor and the wing and analyze it in different flight status. Through the research, the project of lift distribution has been confirmed. At the same time, to improve the working efficiency of the propellers, build an optimization model to modify propellers geometric parameters in I-sight. According to the optimized parameters, the working efficiency of the propellers have been verified to promote a lot.

Improve the efficiency of propulsion system is an effective mean to archive flight performance improvement of electric-powered tilt-rotor. In this paper, the motor is constructed by combining the motor equivalent circuit model and the positive polynomial loss model. The proprotor is modeled by Goldstein vortex theory and the validity is verified. Based on this, aimed at an effective optimization goal, a comprehensive optimization method for the propulsion system is proposed using optimization algorithm. Then the optimization of the propulsion system was carried out and an effective optimization program was obtained for a small electric-powered tilt quad rotor. The research results show that this optimization method can effectively solve the problem of motor and proprotor matching. By optimizing the design of the propulsion system, the power demand can be effectively reduced and the flight performance can be improved.

Airfoil is very important in the aircraft performance. To figure out the conflicts between the requirements of the aerodynamic and stealthy design of supersonic fighter airfoil, we established a multi-objective optimization platform, which is based on high fidelity method, Parsec parameterization method, artificial neural network and Pareto genetic algorithm, to reduce the supersonic drag and radar cross section (RCS) of airfoils. Three optimized airfoils are obtained based on the optimization design. Compared with the initial airfoil, the geometric shapes of optimized airfoils have smaller leading-edge radius, later maximal thickness location and diamond-shaped thickness distribution. Reductions of the supersonic drag coefficients, the values of pitch moment and RCSs in the key azimuth of optimized airfoils are included in the design condition. Meanwhile, the lift coefficients are improved. Results also indicate that the multi-objective optimization framework could be implemented.

Aircraft Aerodynamic Shape Optimization is a complex, large-scale and expensive optimization problem. Adjoint-based gradient optimization method plays a significant role in aerodynamic shape design. In this paper, we mainly focus on the optimization algorithm applied to aircraft aerodynamic design field. Sequential Quadratic Programming (SQP) is employed here, and some experience-based improvements are made in line search to accelerate the convergence of the algorithm. After improvements, the physical change of variables in each iteration can be evaluated, set and control. The optimization model with the drag coefficient minimization objective and wing thickness constraints for Wing-Body-Tail Common Research Model (CRM) is established. The optimization strategy and the improved SQP algorithm are detailed and verified afterwards. Aerodynamic shape design for CRM with Wing-Body-Tail Configuration is carried out. The optimization results are compared and discussed. Our optimization procedure reduced the drag from 167.9 counts to 149.4 counts (an 11% reduction) within 40 iterations, calling CFD solver 49 times. The optimization results demonstrate the effectiveness of the improved SQP method proposed in this paper.

To analyze the influence of different methods on the productivity of aircraft assembly line more quickly and flexibly, this paper proposes a virtual simulation method combined with industrial engineering theory instead of traditional empirical formula and Gantt chart method. On the basis of the process flow, key data and layout of the aircraft assembly line, the simulation model is built. The productivity, productive tempo and utilization of equipment of the assembly line are studied, and the bottleneck of the assembly line is analyzed. At the end of this paper, the method of productivity upgrading is studied. The optimization of process flow, equipment, processing technology, and processing time are summarized. Taking the cost as the constraint, the simulation experiment method is proposed to analyze different schemes, and the optimal scheme to meet the capability requirements is obtained.

Gaps between feathers can be observed when some birds are flying. They can control the opening and closing of gaps to achieve the goal of good aerodynamic performance. The effects of the wing gaps on aerodynamic performance is investigated by experimental study. Firstly, a kind of flapping wing with one-way holes is designed and tested. The holes are connected by hinge, which is open during the up-stroke process and close during the down-stroke process. One-way holes can increase the lift significantly. However, the hinge connection makes the movement process less continuous. And there’s more resistance because the windward area is increased by hinged holes. Further on, the method was improved by using one-way deforming characteristics of bird feathers. The goose feathers are used as a covering skin of flapping wing, which can have continuous opening and closing gaps, making the flapping process more smoothly and getting smooth aerodynamic characteristics. The results show that the one-way holes/gaps flapping wing can effectively increase the lift and reduce the power consumption. The using of natural feathers as flapping wing covering skin is a conducive way to improve the aerodynamic performance.

Although wind tunnels are the most popular aerodynamic measurement tool of today, whirling arms are another type of tool which is especially useful for the measurement at very low airspeed, including zero. The authors developed a modern whirling arm facility for the measurement of the characteristics of small-scale propellers. In this work, an experiment to measure the characteristics of APC SF 8x6 propeller at a negative advance ratio (from 0.0 to –0.8) is conducted. The rotation of the arm is controlled by a servo motor to maintain the steady rotational speed (i.e. axial airspeed of the propeller) against the thrust fluctuation of the propeller attached at the end of the arm. The very small standard error and standard deviation of the thrust and torque measurement demonstrate the developed system’s ability for precise aerodynamic measurement. In a certain range of advance ratio (from –0.4 to –0.8), remarkable fluctuation of thrust and torque was observed, which suggests the propeller was in a non-steady working condition such as vortex ring state.

An analysis scheme and a physics-based mission system model are applied to evaluate the parameter influence on combat effectiveness of anti-ship missiles in cooperative operations. The cooperative operation is divided into 3 steps: cooperative penetration, cooperative detection and cooperative strike. The technical models which include radar detection model, hit probability model and vulnerability model are adapted to a suitable level of detail to combat effectiveness evaluation. The synthesis evaluation method of cooperative combat effectiveness is analyzed according to operation steps. The case studies are specifically applied to the problem of linking the design parameters to mission effectiveness. The simulation results indicate that increasing the number of anti-ship missiles is the major way of improving combat effectiveness especially in cooperative operations. Besides, increasing of speed, armor thickness and maneuver overload as well as decreasing the RCS of anti-ship missiles are also effective approaches to improve combat effectiveness. Although the models are at basic level, the complexity of the cooperative operation is demonstrated. The mission system model applied is feasible for assessing the combat effectiveness of cooperative operations.

In order to study the rotor/wing compound helicopter characteristic of flight dynamics, this paper establishes the helicopter flight dynamics model, including rotor, propeller, wing with aileron, horizontal tail with elevator, vertical tail, fuselage and other aerodynamic models; and then the CFD and momentum source method are used to analysis the aerodynamic interference of the fuselage/wing/flat tail to rotor, the rotor to the fuselage/wing/flat tail, the propeller to the wing/flat tail, and corrected the aerodynamic calculation model of each component in the form of interference coefficient. The target pitch angle is introduced as a constraint to solving the problem of lift distribution between rotor and wing in the configuration aircraft. The manipulation redundancy problem is solved by setting the three stages as hovering low speed, pre-transition flying, and high-speed front flying to assign the manipulations. The helicopter is mainly controlled by collective pitch, horizontal/longitudinal period variable pitch and the propeller pitch in the hovering low-speed stage; In the high speed stage it is controlled by aileron, elevator and propeller pitch; It is controlled by all the manipulations through introducing the weighting factor in the pre-transition flight stage. Finally, get the trimming strategy of the rotor/wing compound helicopter through the trimming analysis of the example.

This paper carries out a modeling method of ducted propulsion system of vertical take-off and landing aircraft. On axial flow condition, the model of ducted propulsion system, based momentum theory, is established and verified. Based on the non-dimensional definition of open propeller, the non-dimensional method, applied to the ducted blades, is proposed. The research shows that the ducted propulsion system model and the ducted blade non-dimensional method are reasonable and effective, for high performance aircraft configuration design with tilting duct-fan propulsion system.

Following the commercial trend of space sector and taking advantage of smaller satellites to be launched especially in LEO; OrbitPlus Technology Group began the feasibility study and preliminary system design for a 3U CubeSat which can support a general designed payload in one unit and two other units which filled by essential subsystems. This paper contains the main subsystems initial conditions to support longer orbital lifetime with altitudes higher than 500 (Km) for different types of payload. To have a clear idea on required subsystems, Structural, Command and Data Handling, Telecommunication, Power Generation, Attitude Control, Propulsion and End-of-Life Disposal subsystems have been investigated. This study can be first step of initial detail system design of OrbitPlus CubeSat open platform with respect to specific requirements of different payloads.

Planform leading edge curve (PLEC) is one of the waverider’s characteristic curves. As sweep angles of the waveriders designed for given the curve can be controlled directly, the design method for given PLEC is worth being studied. In this paper, a parameterization design method for a more practical waverider given PLEC by use of the osculating cone theory was introduced in detail, and the key of the design process is to parameterize the PLEC and compute the leading points and the flowfield on osculating plane. Constraints of the method were also given in this study, when the leading points are existed. As the cubic spline curve could be controlled parametrically more conveniently, this kind of curves were used to describe PLEC. A series of waveriders were generated for given different PLECs. Furthermore, the numerical simulation method was used to verify the proposed design method and to analyze the effects of the key parameters of PLEC on the waverider configurations. Conclusions were got: different special parts of waveriders could be controlled. In the conditions this paper set up, waveriders with double body could be generated when controlling two design points, and waveriders with finlets and with chined back were generated when controlling three design points, waveriders with wings were got when controlling four design points. Some aerodynamic performances of the typical waveriders were also calculated.

A hybrid-electric propulsion system combines the advantages of fuel-based systems and battery powered systems and offers new design freedom. To take full advantage of this technology, aircraft designers must be aware of its key differences, compared to conventional, carbon-fuel based, propulsion systems. This paper gives an overview of the challenges and potential benefits associated with the design of aircraft that use hybrid-electric propulsion systems. It offers an introduction of the most popular hybrid-electric propulsion architectures and critically assess them against the conventional and fully electric propulsion configurations. The effects on operational aspects and design aspects are covered. Special consideration is given to the application of hybrid-electric propulsion technology to both unmanned and vertical take-off and landing aircraft. The authors conclude that electric propulsion technology has the potential to revolutionize aircraft design. However, new and innovative methods must be researched, to realize the full benefit of the technology.

Civil aircraft is a complex system. With the development of aviation industry, scale and complexity of the systems are getting larger and larger. Hence conventional design process can no longer satisfy the following requirements: integrality and consistence of information, capability of describing different activities and flexibility of requirements changes. However, MBSE (Model based System Engineering) has shown its potential of handling the challenges. Instead of natural language, MBSE adopts different models as the basic elements to storage and transfer data. Hence the relation between requirements of different design levels will be more intuitive and a faster response to requirements modification become possible. In this paper, from the top requirements of civil aircraft, we introduce a V&V activity model to the existing Harmony-SE to construct a both efficient and effective design framework. Comparing with conventional V design process model, our method enables the incremental and iterative developing method as well as a validation step after each design stage. These will produce better-quality aircraft within shorter development period.

Design of wave rider is attracting more and more attention right now and is of special interest for hypersonic applications. The article imports Shock Fitting Methods to calculate the three dimensional hypersonic flow to capture shock wave precisely. The choice of shock generating body is very important, because it affects the shock flow field. It can be single-stage or multiply-stage. In this paper, the double-cone shock generating body is introduced in order to modify lift/drag and the center of pressure of wave rider. According to the restriction of ratio of length to width, the shock surface is cut and then the outer line of wave-riding surface is decided. Relying on engineering application, the leeward surface is designed. The results show that lift/drag of engineering wave rider is relatively higher than ordinary vehicle.

The cabin is a part of a civil commercial aircraft that directly provides service to the end user, including necessary safety facilities for passengers and cabin crews, the physical environment such as sound, light and temperature, essential service facilities and service facilities which enrich the travel experience. Specifically, it includes cabin lighting system, environment control system, passenger oxygen system, passenger address and cabin interphone system, in-flight entertainment system, cabin external communication system, cabin decoration and necessary cabin service equipment such as the seat, the kitchen and the bathroom. Ultimately, the cabin management system implements integrated control and display operation. The cabin of civil aircraft is not only a platform which provides the most direct service to passengers and cabin crews, but also an embodiment of the civil aircraft’s brand value. All systems and components in the cabin need to be considered as a whole in order to ensure their consistent design style and the unique design concept. The application of system engineering thought to carry out the Top-down integrated design guarantees the high value of cabin brand, also it reduces the project risk and avoid the cabin potential safety hazard. This paper will study and summarize the method of cabin design based on system engineering thought.

A flying wing layout is an aircraft layout with no tail only huge wings. It has the advantages of lightweight, low flight resistance and good stealth, but it also has shortcomings such as poor maneuverability. The use of the active flow control technology to replace or enhance the control surfaces, therefore improve the lateral maneuverability of flying wings has attracted the interest of various research teams. The plasma flow control technology can change flow near the actuator to achieve the effect of controlling the local pressure of wing, thereby completing the lateral manipulation of the unmanned aerial vehicle (UAV). Many teams have also carried out relevant experiments, which laid the theoretical foundation for this paper. This paper mainly designs a sensor board system of flying wing UAV, which is used to collect the local pressure on the wing of UAV and control the switch of actuator. An experiment was carried out on the ground. First, this board can collect and monitor the pressure data at a certain point. Second, the circuit system can also react to specific situations (such as local pressure values) and automatically control the actuator’s switches. Third, the plasma actuator can also be actively controlled by remote control. This experiment lays the theory and practice foundation for the UAV flight experiments in the future.

In this paper, the method of converting time-domain data into frequency data processing is studied. Flight test data is recorded according to the time interval of time domain data, test data such as frequency, power spectrum analysis, frequency domain identification characteristics shall be carried out in the frequency domain, this needs to be time domain data transformation processing, converted into frequency domain data.

This paper describes the method of creating an aircraft level system integration test facility by connecting distributed individual system test integration benches using a fiber–optic network and signal acquisition and reconstruction. The facility has been used for x type aircraft integration testing to support first flight. The experience proved this aircraft level system integration test facility can satisfy the aircraft function integration test requirements. Connecting existing system test facilities to reduce the integration tasks on the aircraft reduced the time taken to get to first flight. The cost of a new aircraft level system test facility was saved by reusing the distributed test resources.

Flapping Wing Micro Air Vehicles (FWMAVs), which are inspired by nature’s flyers and mimic their flight, have numerous advantages compared with conventional fixed wing and rotating wing aircrafts at small scale and low Reynolds numbers, such as ability of hovering and anti-disturbance, high lift aerodynamic performance, good maneuverability. Due to the dimensional constrains and demands of compact design and low power consumption, suitable angles of attack and size of wings have to be tested to reach high aerodynamic efficiency. This paper presents results of experimental investigation of angle of attack (changed by a root deflection angle) and wingspan on the aerodynamics and power of a FWMAV. Based on the measured results, a quasi-steady model for lift and power estimation is suggested. The difference between the estimated and measured lifts and power is less than 10%, which reveals that the quasi-steady model is reasonable for a preliminary design. The results indicate that the wing with a root deflection angle around 15° shows the highest aerodynamic efficiency, and larger wings are preferred to reach a higher ratio of lift to power, implying that larger wing could give a higher lift under certain power input. Finally, we have obtained a best performance wing, which can generate about 40 g of lift at a power input of about 8 W when flapping at 22 Hz.

A new concept of compound helicopter to achieve high speed of two times of a conventional helicopter is described. This configuration consists of a single main rotor and a main fixed wing. Antitorque is performed through two electric driven propellers installed at the wing tips. Main propeller to achieve high speed is aft-mounted at the tail of the fuselage. Design of a flyable model of the compound helicopter proposed by JAXA is described. Following the flight test of the 1st concept demonstrator, it is found that the compound helicopter with single main rotor combined with a set of wing-tip propellers worked as anti-torque device is stable and can be controlled satisfactorily through the main rotor controls. However, it was difficult to simulate the high advance ratio flight conditions of the main rotor with the 1st concept demonstrator because the tail propeller was connected to the main rotor and driven by a same motor, thus decreasing the main rotor rotating speed also resulted in decrease of the tail propeller thrust. A new design which aims to achieve high speed flight and demonstrate the controllability of the aircraft even in high advance ratio is illustrated. Flight test results of the 2nd generation scaled-down model is reported. Advance ratio as high as more than 0.8 is achieved during test flight without any controls on the fixed wings.

The morphing aircraft can change the shape of the vehicle by local or whole to improve the aerodynamic efficiency of the aircraft in a wide speed range, so as to achieve multi-task functions or multiple control purposes. Based on the self-developed numerical simulation platform ARI_CFD, a three-dimensional numerical simulation method is used to study the aerodynamic influence of the variable forward sweeping schemes and the main parameters of the forward-swept wing on the aircraft. The results show that, the two morphing schemes including “The forward-swept wing embedding in the main wing” and “the main wing rotation” both have advantages and disadvantages in the design aspects of the forward-swept wing and its morphing mechanism, in addition, the aerodynamic evaluation shows that the aerodynamic characteristics of the two morphing schemes are not very different and can be selected according to the specific condition as appropriate. With the increase of the exposed area of the forward-swept wing, the maximum lift-drag ratio of the whole aircraft increases first and then decreases with a non-monotonic changing law. Backward shifting of the forward-swept wing position is beneficial to improving the lift and lift-drag ratio of the whole aircraft, and is beneficial to increase the variable forward-swept wing size. By changing relative thickness of the forward-swept wing airfoil from 0.04 to 0.06, the lift-drag ratio of the forward-swept wing itself increases from 33.55 to 37.72, making transonic lift-drag ratio of the whole aircraft increases from 11.29 to 11.60. The use of a high-lift airfoil with bigger camber on the forward-swept wing can effectively increase the lift-drag ratio of the layout, and the aerodynamic efficiency gain of the whole aircraft can reach up to 13.2%. The optimization of the wing tip planform can enhance the aerodynamic efficiency near the wing tip, but the effect is very limited. The results show that a well-designed morphing aircraft with forward-swept wing can significantly improve transonic aerodynamic performance. It is verified that forward-swept wing morphing aircraft has much development potential as a multi-task vehicle within a wide speed range. This study can provide references for the design of the morphing aircraft.

The results of a statistical investigation of 42 fixed-wing, small to medium sized (20 kg−1000 kg) reconnaissance unmanned air vehicles (UAVs) are presented. Regression analyses are used to identify correlations of the most relevant geometry dimensions with the UAV’s maximum take-off mass. The findings allow an empirical based geometry-build up for a complete unmanned aircraft by referring to its take-off mass only. This provides a bridge between very early design stages (initial sizing) and the later determination of shapes and dimensions. The correlations might be integrated into a UAV sizing environment and allow designers to implement more sophisticated drag and weight estimation methods in this process. Additional information on correlation factors for a rough drag estimation methodology indicate how this technique can significantly enhance the accuracy of early design iterations.

A reasonable and precise degradation model plays a vital role during the process of structure degrade research. Enlightened with the practical engineering examples, we systematically review the consideration and classification about the contemporary research of degradation model, which in terms of data type and its ample extent. First of all, this paper expound the model set process, their suitable occasion and research status, then, some relate questions and application comparison of these models were made to explain their characters during the practical engineering. Finally, some feasible research directions and challenges were emphasized in the conclusion.

According to radial basis functions, the greedy method and the subspace method are used to develop a deformation solver for hybrid unstructured mesh. The solver is constructed with massive parallel processors to improve the deformation efficiency of complex boundary. ONERA M6 wing of million mesh magnitudes and X48B flying wing of ten million mesh magnitudes are selected as test cases. The parallel acceleration, robustness and efficiency of the solver are validated and compared with different CPU cores and basis functions. The results indicate that the mesh quality can be guaranteed after deformation, and the deformation efficiency can be increased more than 80 times with massive parallel system.

Inverse airfoil design algorithm can obtain the airfoil geometry according to the target pressure coefficient (Cp) distribution. Recently, the rapid development of machine learning method provides new idea to solve engineering problem. Multi-output least-squares support vector regression machines (MLS-SVR) is a multi-output regression machine learning method which can make prediction for several outputs simultaneously through learning a mapping from a multivariate input feature space to a multivariate output space. In this paper, MLS-SVR is used to learn the mapping from the Cp distribution to the geometry, which can be seen as a multi-output regression problem. Through iteratively adding the predicted airfoil geometry and its pressure coefficient distribution into the sample database, the precision of MLS-SVR to predict the right airfoil geometry corresponding to the target Cp distribution is improved. A low speed, transonic and supersonic airfoil inverse design problem are used to validate the efficiency of the proposed algorithm, and the experimental results show that the proposed algorithm can save 34.1% and 58.6% CFD evaluations for low speed and transonic cases respectively to obtain satisfactory airfoil.

A radar cross section gradient calculation method based on adjoint equation of method of moment (MoM) is proposed. The adjoint equation of method of moment as well as the variation of radar cross section is derived. Backscattering RCS of cylinder and missile model are firstly calculated and compared with measured data or FEKO MoM result to verify the reliability of the program adopted. Then, the cylinder and missile are parameterized and disturbed with domain element method. The gradient of design variables are computed with both adjoint method and finite-difference method. Numerical results of both methods are in good agreement proving that adjoint method has high reliability and precision. RCS gradient calculation based on adjoint equation has high efficiency, accuracy and can be applied in the calculation of complex geometry. It forms the basis of the gradient-based aerodynamic-stealth optimization platform.

A three-dimensional supersonic flow at Mach 1.4 over a rectangular open cavity of length-to-depth ratio L/D = 6, width-to-depth ratio W/D = 2 is studied with large eddy simulation. The numerical results are validated against the experimental data provided in the literature of Dudley and Ukeiley [1] through comparing time-averaged flow patterns and turbulent velocity in the cavity flow. The agreement is reasonable. Furthermore, we analyze the essential dynamics of the cavity flow, e.g. the resonant vortex-acoustic interactions, the dominant flow oscillation. The compressibility effects on the base flow and the pressure and velocity fluctuations are studied subsequently. Additionally, a two-dimensional open-cavity flow at the identical Mach number, Reynolds number and length-to-depth ratio is simulated with direct numerical simulation for reference. The two-dimensional result shows the vorticities are accumulated gradually in the initial stage and then switch the flow pattern eventually. Conversely, the flow in the three-dimensional open cavity could achieve a “stable” turbulence instantly and then maintain the flow pattern.

Whether war wounded aircraft can be repaired quickly and effectively determines the operational capability of aircraft in the war environment. In order to ensure the combat readiness and continuous strength of the aircraft, it is necessary to scientifically determine the spare parts requirement of the war wounded aircraft. In this paper, a ranking method for the importance of combat aircraft damage spare parts is established. Firstly, the model of aircraft combat spare parts is established to determine the type of combat spare parts to be ranked. Secondly, the kill probability and survival probability of aircraft components in one mission are analyzed and determined. Thirdly, the independent existence state of aircraft and the occurrence probability and system performance of each independent state are determined. Finally, Griffith degree of importance of aircraft residual lethal parts was calculated and the order of war damage spare parts was conducted.

In this paper, a force-measuring platform which is used to test the heading load of landing gear is analyzed, as well as the influences of its inertial forces on the accuracy of the test results. Two sets of drop test simulation models, considering or not considering the dynamometer platform, are built based on the dynamic simulation platform named LMS Virtual.lab motion. The installation gap between the heading force sensor and the dynamometer platform is simulated through the form of spherical contact. The gap size was changed at two subsidence speeds of 1 m/s and 1.5 m/s, and the relationship between the gap size and the heading inertia force of the force-measuring platform was analyzed. According to the design requirements of the drop test in Book 14 of the Aircraft Design Manual, a drop test bench was built. The landing gear load of a high-speed UAV landing gear was measured and compared with the simulation results.

On the basis of summarizing the research results of CPS and CPS architecture for NC machining process, the application of the edge computing architecture and its characteristics, the relationship between cloud computing and fog computing, and the application value are discussed. Based on the application scenarios of the construction technology and the edge computing technology, an edge computing architecture based on the intelligent monitoring and control CPS built in the NC machining process is proposed, and its function positioning, function model and key technologies are discussed.

We developed an optimization design approach which can be used to the high subsonic Natural Laminar Flow (NLF) airfoil design at low Reynolds number. The aerodynamic characteristics are evaluated by the γ-Reθt transition model, and a modified CST method based on the disturbing CST basis function is introduced in the airfoil parameterization. After that, we build up the optimization system which employs the MCPSO (Multi-Groups Cooperative Particle Swarm Algorithm) as the search algorithm, and then apply the optimization methodology to the design of a propeller tip airfoil. More specifically, the laminar separation bubble and transition location are controlled in the optimization process. Our simulation results indicate that the optimized distribution shows better ability to control separation position and reattachment position, and the pressure drag can be greatly reduced when the laminar separation bubble is weakened. We demonstrate that it is a reasonable design idea that the friction drag increment is proposed as a constraint condition. These design experiences can provide valuable reference data for the design of the high subsonic NLF airfoil at low Reynolds number.

When reducing the drag of the aircraft using winglet, the size of the winglet is limited mainly under the wing root bending moment increment (ΔBM) due to the winglet. The existence of the maximum possible drag reduction (ΔCD) at certain ΔBM is not touched by the majority of public resources. 24 winglet plans are evaluated in Computation Fluid Dynamic (CFD) to study the impact of 5 major geometric parameters (span, cant angle, incident angle, twist angle, sweptback angle) on the ΔCD and ΔCm. It is found that for a fixed span, a linear trend line exists between ΔCD and ΔCm for various winglet plans that the ΔCD/ΔBM is the highest. This trend line is defined as the maximum ΔCD achievable at certain ΔBM at a fixed span. With increasing span, the obtainable ΔCD is higher but the ΔCD/ΔBM decreases. The concept of the ΔCD/ΔBM trend line as the drag reduction limit will greatly simplify the optimization process of winglet design.

The improvement of the lift to drag ratio (L/D), and the decrease of the required power for flight, are important measures to enhance the range and endurance of propeller aircrafts. The research investigates the effect of variable camber wing on the flight performance of a transport aircraft based on wind tunnel tests. The wing camber is altered by deflect the trailing edge flap in small angles. The tests reveal that the maximum L/D reaches its peak when the flap deflects 6°, and is 4.0% higher than the configuration with flap retracted. With flap deflecting 6°, the lift coefficient for the maximum L/D is 16.1% higher. The flight performance analysis shows that the range and the endurance increase by 4.0% and 13.7% respectively at the selected flap angle. By flying at larger lift coefficient and larger L/D, the required power for endurance flight decreases significantly, which is the reason for the improvement of the range and endurance performance.

The aerodynamic characteristics of the upswept tail of a transport aircraft are studied and the optimum arrangements of ventral fins to reduce the drag are evaluated. Two symmetric vortices emerge under the fuselage afterbody, which are the major causal factor of the additional pressure drag. A pair of fins installed under the fuselage, extruding the core of the vortices effectively damp the vortex. Parametric study shows that the length, height, location and yaw angle of the fins are the sensitive factors of drag reduction. Drag reduction of 0.0021 is achieved in wind tunnel test for typical cruise angle of attack (AoA). The pitching moment has nose down tendency and the longitudinal stability is reduced. The reason is that the pressure recovery of the bottom surface of the tail is improved by adding the fins, which is dependent on the AoA.

JAXA is proposing a concept of a high speed EMS compound helicopter with the maximum cruise speed 500 km/h which is twice as fast as conventional helicopters. In order to realize this high speed within an envelope of the existing propulsion system, it is essential to develop a very low drag airframe. Furthermore, the compound helicopter has several rotor/propellers and lifting surfaces in the vicinity, which may cause the aerodynamic interference problems. This paper describes the review to the mechanism of the high drag of helicopter airframe followed by the resolution proposals for the airframe with the low drag and the less severe aerodynamic interference.

The distributed propulsion (DP) with the effect of boundary layer ingestion (BLI) can improve the flight performance of unmanned aerial vehicles (UAV), and the mounting parameters of DP can further affects the propulsive efficiency and the aerodynamic performance of airfoil. To explore the effect of DP mounting parameters, computational fluid dynamics (CFD), sensitivity analysis and Kriging surrogate models methods were used, results indicate that in a wide range of design space, the effect of BLI significantly improves the aerodynamic performance and propulsive efficiency at medium/small angle of attack ( $$ 2^{ \circ } $$ and $$ 6^{ \circ } $$ ), but the undesirable configuration will exert adverse effects on the propulsive efficiency and pitching moment, lower mounting height or larger setting angle of DP will produce the aerodynamic effect of the reflex airfoil, which is suitable for flying-wing UAV.

Two alternative designs of a flying wing configuration unmanned aerial vehicle (UAV) with different aspect ratios were investigated by the numerical calculation. With the comparisons of lift-drag characteristics, lateral-directional stability and efficiency of control surfaces, the scheme of the smaller aspect ratio was adopted. Then the wind tunnel tests of the selected configuration were conducted to investigate the aerodynamic characteristics. In order to obtain more dynamic parameters desired in the flight simulation, a simple and useful method for the engineering estimation of dynamic derivatives was proposed. The longitudinal and lateral-directional dynamic derivatives of the UAV respectively calculated by the estimation method and CFD method were then compared. The comparison results showed that the estimation method can accurately reveal the dynamic aerodynamic characteristics of the aircraft.

Vertical take-off and landing aircraft (VTOL) has been a hot research topic in the aeronautics. This paper discusses the feasibility of a new concept of rotor/fixed-wing which can switch the flying mode during flying. In this design, wings on both sides of the UAV could rotate respectively around the pitch axis to switch the flying mode, so as to achieve the purpose of VTOL or the cruising flight. The paper made an analysis on the aerodynamic characteristics and the mechanical characteristics of this kind of UAV. Based on these study, a physical prototype were made and conducted a flying test, which the recorded date was basically in line with the design goals, the power to weight ratio of the rotor/fixed-wing is significantly greater than the VTOL aircraft which use traditional propeller. It is proved that the analysis on the mechanical characteristics and the aerodynamic characteristics is valid. It illustrates that this aircraft configuration had certain feasibility and paved the way for potential application prospect in the future.

The high probability of the occurrence of separation bubbles or shocks and early transition to turbulence on surfaces of airfoil makes it very difficult to design high lift and high-speed Natural-Laminar-Flow (NLF) airfoil for high altitude long endurance unmanned air vehicles. To resolve this issue, a framework of uncertainty-based design optimization (UBDO) is developed based on the polynomial chaos expansion method. The $$ \gamma { - }\overline{\text{Re}}_{\theta t} $$ transition model combined with the shear stress transport $$ k - \omega $$ turbulence model is used to predict the laminar-turbulent transition. The particle swarm optimization algorithm and surrogate model are integrated to search for the optimal NLF airfoil. Using proposed UBDO framework, the aforementioned problem has been regularized to achieve the optimal airfoil with a tradeoff of aerodynamic performances under fully-turbulent and free transition conditions. The tradeoff is to make sure its good performance when early transition to turbulence on surfaces of NLF airfoil happens. The results indicate UBDO of NLF airfoil considering Mach number and lift coefficient uncertainty under free transition condition shows a significant deterioration when complicated flight conditions lead to early transition to turbulence. Meanwhile, UBDO of NLF airfoil with a tradeoff of performances under fully-turbulent and free transition conditions holds robust and reliable aerodynamic performance under complicated flight conditions.

The scheme of maglev-rotor unmanned aerial vehicle (UAV) is presented based on research of the existing multi-rotor UAV with vertical take-off and landing (VTOL) capability in this paper. Besides calculation and CFD analysis of aerodynamic characters of the maglev-rotor system and the transition between the VTOL mode and the forward flight mode of the maglev-rotor UAV, test flight are carried out to verify the calculation method and the feasibility of the scheme. The prototype of conventional electric motors is specially designed and built on this purpose. The attack angle of fuselage, the tilt angle of rear duct, the power of lift duct and the power of rear tilt duct are measured at different wind speeds. It is confident that the transition mode is available since the changes of the fundamental parameters, such as the attack angle of fuselage, the power of the lift duct fan and the rear tilt duct fan, the tilt angle of the rear duct fan, go on smoothly and continuously.

Unlike traditional helicopters, ground effect has great influences on the performance of vertical take-off and landing (VTOL) ducted fan unmanned aerial vehicles (UAVs) during take-off and landing. A testing rig including a scaling ducted fan system and an imitation floor is designed and manufactured to investigate the influence of ground distance on the thrust and required power of the ducted fan system. The testing results show that the thrust is decreased and the driving power is increased because of the ground effect, and the loss of thrust and the increase of power become larger when the ground distance decreases. Finally, an acceptable distance is provided according to the testing results, and the ground effect should be specially considered during the design of ducted fan VTOV UAVs.

The aircraft family is a set of aircraft products that share common components but vary in configurations based on different mission performance and requirements. The reconfigurable UAV family using modular design can improve the efficiency of executing multiple missions and enable the acquisition cost benefits, which is one of key competitive edges in military applications. This paper aims to study an effective optimization method for conceptual design of the reconfigurable UAV family with interchangeable components that can be reconfigured between missions. First, a Flying-wing UAV family defined for combat and reconnaissance missions is used as an example for demonstration of the method, and an appropriate comprehensive analysis model is developed. Next, the optimization formulation for the conceptual design of UAV family is presented in detail, including design variables, design constraints, and objectives. A hybrid optimization strategy is then applied to solve the optimization problem of UAV family conceptual design. The results after optimization indicate that the mission performance and requirements for each configuration are satisfied, and costs are reasonably compromised for the UAV family.

The article reviews the necessity of simulation in laminarization as an important optimization technique in aircraft design. Transition models e.g. k-ω SST with γ-Reθt model and k-kl-ω model are popular in industry and embedded in many optimization framework. The article aims to evaluate their accuracy in laminar transition predicting as well as in evaluating the effect of laminarization, i.e. the extended length of laminar brought by optimization in shape. The object for numerical simulation and experiment validation for laminar transition position predicting are one original airfoil and one laminarized airfoil under Re~2.62 × 106, Ma~0.785. The mesh with boundary conditions, the criterion for judging laminar transition, and the method to monitor laminar/turbulence distribution over model surface in wind tunnel experiment are introduced. The result shows that due to many factors, the two models cannot very accurately predict the laminar transition position, but the effect of laminarization can be evaluated quite satisfactorily by the k-ω SST with γ-Reθt model. Therefore, the laminarization applied on three-dimensional aircraft part, e.g. wing, nacelle can be constructed using this model in the future optimization framework of the authors.

The market of light hybrid and all-electric aircraft is developing rapidly, many start-up companies have introduced their concepts and designs which are to enter to service in the next few years. At the same time, university and government research programs on hybrid and all-electric aircraft have focused more on regional, medium and long-haul aircraft. This article will emphasize the application of electric propulsion for eco-friendly light general aircraft. The goal of this paper is to introduce and develop aircraft and propulsor concepts, create methodology and a tool to calculate and make analysis of propulsor and electric motors that they contain. The Electric Ducted Fan (EDF) was chosen as a propulsor, four different types of EDF cores were introduced and developed in this article. For EDF parameters calculation analytical methods were chosen, additionally guidelines for creating a MATLAB code for making calculating easier will be presented. The design point was to achieve the same thrust for all types of EDF by changing design parameters to be able compare all other EDF’s parameters. The results represent calculations for four core types of EDF and their analysis. Electric motors are developing rapidly, so research was held in regard to technological development. Based on the analysis guidelines of usage of different types EDF for role aircraft were presented.

The blade is a key part of aeroengines. It is necessary to investigate the modal parameters of blades for guaranteeing structural integrity. Operational deflection shape (ODS) and operational modal analysis (OMA) are all modal test methods for aeroengine blades. First of all, the theory of ODS and OMA is introduced separately. Then, the ODS and OMA test systems are established with the high frequency vibration table and scanning laser Doppler vibrometer, by which the ODS and OMA of aeroengine blade are researched within the range of 10 kHz frequency respectively. The natural frequencies and mode shapes of the blade are educed through ODS data, while the natural frequencies, mode shapes and damping ratios are acquired through OMA data. Finally, the two test results obtained by ODS and OMA are contrasted, which indicated the two results are consistent. It is proved that the two test systems of ODS and OMA in this study are designed reasonably, and the test results are analyzed validly.

Combining iterative physical optic method (IPO) with equivalent edge currents method (EEC), a code for calculating the radar scattering characteristics of the nozzle was developed. The reliability of it has been validated by calculating a model introduced from a reference. Ray Tracing method was proposed to improve the efficiency of geometric blanking judgment. By comparing with traditional method, RTM can at least make computational efficiency tenfold in geometric blanking calculation. The aerodynamic and electromagnetic scattering characteristics of axially symmetrical nozzles S-shaped nozzle and double S-shaped nozzle were studied. The results show that the code developed is reliable. The performance of the nozzle was evaluated from the thrust, flow coefficient and total pressure recovery coefficient, and then the flow fields of the nozzles were analyzed. The S-shaped nozzle can effectively shorten the high temperature tail flame area, but it will bring about 2% of thrust loss. The electromagnetic scattering characteristics of three nozzles show that the S-bend structure can effectively reduce the RCS of the nozzle. Compared with the axisymmetric nozzle, double S-shaped nozzle can reduce RCS by at least 74.4%

This paper adopts the method of injection via an array of holes in three-dimensional numerical simulations of the rotating detonation engines (RDE) with hollow combustor using the premixed stoichiometric hydrogen-air mixture. The calculation is based on the Euler equations coupled with a one-step Arrhenius chemistry model. The array hole injection method is more practical than previous conventional simulations where ideal full area injection method is used. The wave structure of the flow field is composed of obverse-rotating waves (ORWs) propagating clockwise and reverse-rotating waves (RRWs) propagating counterclockwise. ORW is detonation wave (DW) near the outer solid wall while degenerate to shock wave (SW) near the nominally inner wall. This phenomenon is never found in the previse numerical studies.

Variable cycle engine with FLADE, abbreviation for Fan on Blade, is one of the research focus for future military and civilian aircraft power plants, showing outstanding performance advantages. A FLADE calculation method was established by calculating bypass flow and inner flow independently, to developing a steady-state performance simulation model for double bypass variable cycle engine with FLADE. The dynamic equations that can reflect the rotor inertia effect and component volume effect were added, to developing a transient performance simulation model. The transient characteristics of mode transition were analysed during opening/closing FLADE duct, putting emphasis upon the influence of geometric parameters adjustment and its different combinations. The results indicate that the mode transient characteristics of FLADE variable cycle engine are only influenced by FLADE vane angle and FLADE nozzle area, which should be increased or decreased synchronously. Both of the engine bypass flow and core flow are affected by opening or closing FLADE duct slightly.

Bird strike is a great threat to the safety of aircraft. The bird’s collision with the aero engine may lead to engine power loss, fires. Large blade debris in high-speed produced by the collision could even cause the disaster of aircraft crush. In this paper, the simulations of the bird striking a S-shaped inlet and blades were carried out to study the influences of the inlet to the blade’s capability of crashworthiness. Besides, the influences of bird’s incidence pitch angles to the bird strike inlet was studied. In the simulation, the SPH method and the linear Mie–Grüneisen equation of state were used in ABAQUS/Explicit to describe the bird’s fluid-like behavior during the bird strike events. The SPH method and the material model were verified by simulating the bird strike on rigid plate. Compared the simulation of the bird strike on fan blades with and without a S-shaped inlet, which could not only weaken the bird’s after-impact kinetic energy but also make the bird crush into pieces. Both of them can enable the engine fan blades withstand a heavier bird impact.

Particles depositing is a severe damage to turbine vanes. The depositing on upper surface for the plate models with film cooling configuration is experimentally studied for the attack angle −5º with a kind of wax, which is melted and atomized to generate particles. Some models are trenched along the row of film cooling holes to study the effects of different trenches with different depths on the depositing. All the trenches could distribute the depositing on the near downstream of the row of film cooling holes more uniformly and the deeper trench could make the depositing on this area more uniform. When the blowing ratio is 0.98, all the trenches could decrease the depositing on the upper surface and the depositing mass decreases with the trench depth increasing. When the blowing ratio is 1.47, all the trenches increase the depositing on the upper surface although the increment is not vary large and the depositing mass increases with the trench depth increasing. However, even if the blowing ratio is 1.47, the trench may reduce the negative effect of the depositing on turbine vanes since it could delay the depositing to the area far from the row of film cooling holes.

In this paper, the influence of overflow outlet at the trailing edge of the turbine blade on the flow and heat transfer in the trailing edge cooling passage is studied by numerical simulation method. The influence of the number of overflow holes, overflow hole radius and overflow contact area of the overflow hole (change the number of holes and hole radius, and unchange the flow cross-sectional area) on the distribution of flow coefficient of the trailing edge overflow hole were compared. And the influence of the flow distribution from the trailing edge and the lower edge plate, the influence on the heat transfer of the inner cooling wall ware compared. The results show that the number, radius, and flow contact area of the trailing edge overflow hole have large influence on the distribution of the flow coefficient of the trailing edge overflow hole. The decrease of the number of holes, the increase of the hole radius, and the reduction of the flow contact area all make the flow coefficient of the edge overflow hole rises; Increase of the number of hole and hole radius both increase the outflow ratio at the trailing edge and increase the average heat transfer coefficient of the inner cooling passage. The overflow contact area of the overflow hole has little influence on the overflow outflow ratio and the average heat transfer coefficient.

In view of the problems of the traditional aero engine development, such as the design elements are difficult to reuse, the traceability of requirements is not good and lack of top-level logic verification, the model based systems engineering (MBSE) method is introduced. Apply the Harmony for Systems Engineering to the top-down design flow of aero-engine turbine system. The system requirement model is established through the creation and classification of turbine system requirements and the definition of system use cases. To establish the function model, the functional, expressive, interface, and other requirements are transformed into a clear description of the system function. The system architecture is analyzed and designed on the basis of the system function, then the function of each use case is decomposed and allocated into the subsystems and components of turbine system. Based on the system model, the requirements of the system are refined, traced and verified. The results show that the MBSE method can perfect the requirement definition, complete the mapping of the requirements to the system elements, realize the function logic verification, and support requirements verification at all stages, which provides an effective practical approach for the development of aero-engine.

Fun surge often occurred for an augmented turbofan engine, which is related to the mechanical hydraulic nozzle control system (NCS). In order to fully realize the potential of the engine, it is now hoped that the mechanical hydraulic controller will be changed to a digital electronic controller. First of all, the control law of the engine is introduced. Secondly, modification scheme of a digital electronic NCS is proposed and a digital electronic controller is designed. Then the model of engine and NCS was established in AMESim and Simulink, and the closed-loop simulation of the NCS was realized. Finally, it is verified that the Digital Electronic NCS can ensure the normal operation of the engine in different inlet conditions and different dynamic processes, and its behavior is better than the mechanical hydraulic NCS. The digital electronic controller helps to reduce the occurrence of faults.

In this paper, the reduction method using the gas dynamic similarity is analyzed, and an integrated reduction model using the gas dynamic similarity theory and the linear parameter-varying (LPV) strategy is proposed, which can describe the JT9D turbofan engine dynamic behavior at different powers around the full defined flight envelope. Compared with the general reduction method using the gas dynamic similarity, the proposed model is more accurate in the full envelope. To study the selection of scheduling parameters, several reduction LPV models are built by different parameters. The comparisons between these models and the reference model indicate that the model, which is constructed by the parameters consisting of polynomials of the corrected shaft speeds and the total temperature at the engine input, performs the similar accuracy with the models using more complicated parameters. This set of parameters take advantages considering both accuracy and complexity during the reduction LPV model constructing.

Due to the complicated non-linearity of aircraft engine, it is very hard to design the control system directly. Now, linear model which could take place of real engine is necessary during the design process of aircraft engine control system. Normally, partial derivative method [1] and fitting technology [2, 3] usually are used to the linear model establishment of aircraft engine. Partial derivative method could be used for the steady state point identification, and fitting could extend the state variable model to transient process. The accuracy of fitting generally better than partial derivative method. However, when the noise exists, the accuracy of previous two method is not satisfactory.In this paper, the statistical characteristic of input, output and noise signal is analyzed, and a maximum likelihood identification method in frequency domain is proposed. This method establishes an optimized parameter estimation criterion function in frequency domain. By designing a multi-sinusoidal excitation signal, the Levenberg-Marquardt algorithm is applied to solve the criterion function to establish the control system model. The proposed method is applied to identify the shaft speed system of aircraft engine and the linear models at different points are established. From the simulation results, it can be seen that the aircraft engine models established have high accuracy.

An Extended Kalman Filter (EKF) was proposed for fan speed signal infusion of the turbo fan engine. Firstly, a nonlinear discrete time analytical engine model was identified using a general nonlinear engine mathematical model based on least square method. Afterwards, fan speed signal infusion EKF algorithm was designed based on optimal filtering theory for nonlinear multistage dynamic process. Then, Kalman gains of EKF algorithm were offline tuned, and fan speed signal was synthesized by using 6 different combinations of 4 sensed parameters, T25, T3, Ps3 and EGT as input of the infusion EKF algorithm, and the infused fan speed signal under different infusion combinations were compared with the actual fan speed signal sensed directly from nonlinear engine model using small step test cases. After that, fan speed closed loop control simulations were conducted using the infused fan speed signal as feedback, and algorithms with 6 different infusion combinations were analyzed with respects to control performance and stability using small step command test case under ideal environment. Finally, closed loop control simulation was conducted with the fan speed EKF infusion signal from selected 3 parameter infusion algorithm as feedback using both small and big step command test case under stochastic environment. The results show that, under both ideal (without consideration of noise) and stochastic conditions, the proposed EKF fan speed signal infusion algorithm has a good closed loop control performance

To investigate the effect of fuel type on gas turbine combustor performance, a reacting solver using flamelet combustion model is developed on open source CFD code OpenFOAM. The flow field of an industrial gas turbine combustor under three different power setting conditions including 20%, 58% and 100% is simulated with the new solver. The major features of the flow field with two different fuels are analyzed. Furthermore, the main performance parameters of the combustor are discussed. The results indicate that the major features of velocity and temperature distribution using nature gas remains unchanged, in comparison with using the diesel oil. The natural gas fueled case produces more water in the whole combustion zone than the diesel oil case at the same power setting. However the combustion efficiency and total pressure recovery coefficient are both improved. Thus, the outlet temperature distributes more uniformly using the natural gas.

Aimed at providing a good thermal protection for a pintle injector from being overheated, a three-dimensional combustion and heat transfer coupled simulation is conducted based on a 500 N GOX/GCH4 pintle engine model. Further for such a pintle injector with double rows of injection holes, the effects of the ‘skip distance’ (Ls) and the interval between two rows of injection holes (Li) on the injector’s characteristics of combustion and heat transfer are studied. The results indicate that given the diameter of pintle injector body (Dp), the maximum temperature on the outer surface of the pintle injector decreases from 1756 K to 1404 K when Ls/Dp increases from 0.64 to 1.5, meanwhile the combustion efficiency has a change less than 0.2% with an average value of 0.9510. But when it changes Li, the lowest maximum temperature on the pintle surface of 1326 K and the biggest combustion efficiency of 0.9525 occurs at a middle value of Li/Dp = 0.18 when Li/Dp increases from 0 to 0.38. This paper tries to explain the influences of Ls and Li from the view of the changes of flowfield structures, which show an intensive sensitivity with different pintle injector configurations. The conclusions this paper gets would be greatly helpful for the engineering design of a pintle injector with double rows of injection holes.

A dual orifice gas generator project based on aero-engine combustor was proposed. A set of mono-injector gas generator was designed and manufactured. Hot firing tests were conducted under different conditions to verify the performance of the gas generator, and researches on the working pattern of ethanol supply system, lean blowout limits tests and system sequence optimization were carried out. The test results show that the technical design scheme of the generator is feasible, the mass flow rate, temperature and pressure all meet the design target. The smooth ignition curve indicates a reliable ignition of the generator and steady performance of combustion, and the gas generator is characterized by advantages of step-start; through optimizing the ethanol supply sequence, the problem about impact effect of supply pipeline caused by valve work sequence is resolved, and combing with air pressure regulation optimization, the start-up time of gas generator is greatly shortened; the gas generator has a large lean blowout limits, and is able to work steadily under a wide working range.

In this paper, based on the scramjet, by the mode of characteristic calculation of one-dimensional scramjet engine, the effect of three and four-section combustion chamber geometry, fuel distribution ratio on scramjet performance is studied. The results of the comparison and evaluation of engine performance provide a certain reference for the geometric configuration of the combustion chamber and fuel injection parameters of the design point. It is mainly to determine the appropriate combustion chamber geometry and fuel injection method to obtain better combustion chamber performance and overall engine performance. The results show that increasing the thermal throat area of the combustion chamber can effectively increase the thrust of the aircraft during low-speed flight condition, and increase the fuel distribution ratio and the length of the expansion section in the front section of the combustion chamber without thermal choke and without affecting the inlet start, can improve the specific impulse of the engine at high speed.

A numerical simulation and experimental measurements are performed to characterize the flowfield of oxygen/oil gas-jet facility. Chemical non-equilibrium Navier-Stokes equations have been solved to acquire the axisymmetric combustion-heated stream condition. The numerical results indicate that chemical reactions have an impact on flow oscillation, which alleviate the rate of decay of flow and make shock region move downstream. Flow diagnostics has been conducted to measure flow properties. Distributions of the pressure, temperature and Mach number along the centerline of the exit plane are experimentally obtained. The comparison has been presented between simulation and measurement. Static pressure distribution measurements show a good agreement with the numerical results. The computed temperatures are higher than measured ones, but the trends of oscillation and declension are nearly the same. Possible reasons of the difference are mainly because of uncertainly in chamber pressure and combustion efficiency.

The numerical simulation of combustion is one of the most important research tools for developing alternative fuels and high-performance combustion engines, which is widely used in the field of aerospace. However, the broad range of time scales and complex chemical kinetics bring great challenge to combustion simulation. With traditional implicit methods, most of the CPU time will be spent on solving the stiff ODEs caused by detailed chemistry mechanism. So, the object of this study is to use an accurate and efficient explicit method to solve the stiff ODEs, by which the computational efficiency can be greatly improved. Developed by Gear and co-workers, the projective method is utilized to solve stiff ODEs with a broad range of time scales, which is suitable for combustion problems. In this study, the homogeneous ignition process and spherical flame propagation process of methane are both investigated. Compared to the results from a traditional implicit method, VODE, projective method gives almost the same results, while the calculation speed is one-order faster than that of the VODE method. Due to its high accuracy and efficiency, the projective method will have great potentials in combustion simulation.

Avionics cooling is quickly becoming the limiting factor of aircraft/spacecraft performance and reliability, particularly with the rapidly increasing power density, and decreasing module size. This paper looks at the high-efficiency heat removal using loop heat pipe technology, as an advanced two-phase thermal control method by reducing the thermal resistance between the heat sources and heat sinks. Two high performance loop heat pipes (LHPs) are developed and experimented. The test results show that LHPs can well work at the heat load up to 663 W, with low thermal resistance of 0.042°C/W.

In order to improve its performance and reduce the cost, hypersonic vehicle urgently needs to design the thermal structure with low redundancy. This requires the ground electric arc heating test to be more closer to the real aerodynamic thermal environment of the aircraft with better accuracy and repeatability. The surface heat flux density of the model is the important basis for determining the surface ablation state and total heat addition in the test parameters. In particular, the accuracy and repeatability of the heat flux density on the model surface directly affect the ablation and thermal insulation performance of the models. Therefore, it is necessary to accurately monitor and measure the heat flux density of model surface during long time ground assessment test in arc-heated wind tunnel. In this paper, based on the conventional water calorimeter measurement principle, combining with theoretical analysis, we design a new type spherical water calorimeter, which can measure the sum of spherical heat addition and stationary point heat flux density of the model. The internal cooling structure of spherical water calorimeter is optimized by means of three-dimensional heat transfer structure optimization, and the experimental verification is carried out. The results show that the quantity of the calorimeter with good accuracy and repeatability. This calorimeter can get the stationary point heat flux of the model during the long time test, and can diagnose the parameter fluctuation of the flow field in arc-heated wind tunnel.

As the turbine front temperature of the modern aero-engine gets higher and higher, the problem of gas ingestion and the related rim sealing is becoming more and more serious. In this paper, the unsteady numerical simulation of the rim seal model with cavity is carried out, based on the gas ingestion experiment rig of Bath University. Firstly, the CFD results is compared with the experiments to valid the accuracy. The comparison for the pressure distribution shows that the simulation agrees well with the experiments. But the results for the sealing efficiency have some relatively errors. Secondly, the change process for some parameters about the ingestion is discussed. The simulation shows the pressure distribution of main flow is the decisive factor, and the interaction between the blades and the vanes affects the pressure distribution. Thirdly, the effect of the interaction between blade and vanes is provided a close analysis. Interestingly, the periodicity of the sealing efficiency is not consistent with that of the rotating blades. Also, the effect of the axial spacing is obtained through reducing the spacing by twenty percent. The decrease of axial spacing between vane and blade will cause the increase of the peak value of the pressure in the high pressure region.

As the Mach number and range of aerial vehicles increases, the heat protection of the combustion chamber has been a more and more important issue. Besides developing heat-tolerant materials and new combustion chamber cooling techniques, the development of advanced cooling technology is also a significant way to the thermal protection of combustion chambers, such as film cooling, intersecting-grids cooling configuration with internal network channels, laminate cooling and the combination of thermal barrier and film cooling. The internal flow condition and cooling features of the intersecting-grids cooling structure that can be applied in the combustion chambers and turbine vanes and blades of modern gas turbine engines is investigated numerically in the paper, to get knowledge of the cooling features of such structure and discuss the feasibility of applying it in cooling systems of aero-engines.

With planar laser-induced fluorescence (PLIF) technique and multi-points injection mode, the quantitative relationship between the equivalence ratio and fluorescence intensity of kerosene is studied. Using the research results, the quantitative distribution of local equivalence ratio of kerosene in the crossflow is analyzed. The operating conditions are as follows: pressure ranging from 0.15 to 0.25 MPa, temperature ranging from 580 to 700 K, and the equivalence ratio ranging from 0.3 to 0.8. The results show that the equivalence ratio and fluorescence intensity of kerosene satisfy linear relationship; and the local equivalence ratio of kerosene in the crossflow shows an obvious delamination phenomenon; the spray combustible area decreases gradually with the increase of pressure and equivalence ratio.

In the aero engine, the nozzle is one of the most important parts. Nozzles are very small and difficult to be measured. The cone hole in the nozzle is hard to be machined, the quality of the cone hole seriously affects the atomization performance of the nozzle. Different from conventional measuring methods, this paper uses image processing to measure the angle of the cone hole. A method based on Depth From Focus was proposed, which makes use of a series of images to reconstruct the 3D model of the nozzle. This paper uses telecentric lenses and CCD to get images from different distance, then some experiments were made to prove this method can goodly measure the angle of the cone hole. Also this paper finds out that the smaller the moving interval is, the more accuracy there will be, but the more time will be cost.

Fuel nozzle is an important component of aerospace engine, with the typical features are swirling slots, micro-holes and micro-cones. The consistency in structure and size has a major impact on the performance of the engine’s combustion chamber. Aiming at the requirement of the nozzle swirling slot measurement, a four-axis non-contact measuring system based on the laser probe with conoscopic holography is built. According to the spatial geometric distribution characteristics of the swirling slot, the inspection path planning of laser scanning based on the probe spot is designed to measure the spiral line, then the point on the spiral of the swirling slot can be obtained. The coordinate values of points are quickly picked up, and then the fitting analysis is performed to get the relevant characteristic parameter values of swirling slot. The experimental results show that the measurement method can obtain the structure size of the nozzle swirling slot quantitatively, meeting the measurement accuracy requirements, and realize the judgment of the consistency of the size and structure of the swirling slot.

For airplane deicing problem, two dimensional electrothermal deicing model is established. The numerical study on heat transfer characteristics during electrothermal deicing process is presented. The Enthalpy-Porous Medium method is applied to describe the phase change process for the numerical simulation, which is a kind of enthalpy model. The computational domain is treated as a porous medium including solid ice and water liquid and mushy zone. The structured mesh topology is used to distribute the computational domain. The finite volume method is adopted to discretize the governing equations. The temperature is obtained by iteration of the energy equation coupled with the liquid volume fraction formula. The properties such as thermal conductivity of the mushy zone can be obtained by linear interpolation. The heat transfer characteristics are studied systematically for deicer pad including phase change process. The numerical study emphasizes on the effects of heating mode, cooling time, heater power and heater gap on phase change heat transfer characteristics. It shows that periodic heating mode for high heater power is superior to continuous heating mode for low heater power if reasonable combination of cooling time and heater power are adopted to obtain better deicing performance and less energy consumption. The existence of heater gap will decrease energy consumption less and make the temperature distribution more reasonable. Therefore, the reasonable distribution of heater gap can largely improve the deicing efficiency which is beneficial for the design of electrothermal deicer.

A steady, three-dimensional numerical model is developed for the low power arcjet thrusters. Numerical investigation was carried out to determine the effects of swirl on the flowfield and heat transfer characteristics inside the arcjet nozzle. Two different three-dimensional arcjet nozzle configurations with axial injectors and radial injectors are used in this numerical study in order to reveal the effects of swirl on the flowfield and heat transfer characteristics inside the arcjet nozzle. Comparisons of the flowfield and heat transfer characteristics in the two different three-dimensional configurations, i.e., with axial injectors and radial injectors respectively, are performed to reveal the effects of the swirl flow. The distributions of pressure, temperature, axial velocity and tangential velocity in the two different three-dimensional configurations are presented. The modeling results show that there exist evident three-dimensional effects of the flowfield in both the configuration with axial injectors and the nozzle with radial injectors. The swirl in the configuration with axial injectors is dissipated with small magnitude while it is inerratic for the case with radial injectors. The tangential velocity magnitude in the nozzle with radial injectors is big enough to persist downstream and is enhanced at the constrictor entrance. It is found that the swirl flow has evident influences on the flowfield, heat transfer characteristics inside the arcjet nozzle and little effect on electric field and thrust of the thruster when the arcjet thruster operated under the operation condition used in the study.

The cusped field thruster (CFT) is a class of advanced electric propulsion (EP) technology for satellite and space missions, offering advantages over other types of EP including enhanced electron confinement owing to the magnetic mirror and reduced particle loss effects at the dielectric wall. The increasing demand for downscaling for micro-satellite class platforms while keeping performance at similar level has led to considerable efforts dedicated to physical modeling and performance characterization of downsized CFT. Multi-objective design optimization is conducted in this study by employing performance parameters of downscaled CFT, namely, thrust, total efficiency, and specific impulse as the objective functions to maximize and design parameters including anode voltage and current, mass flow rate, and inner and outer magnet radii as the decision variables. Two geometric configurations are considered, i.e., those comprising three magnets with fixed thickness and four magnets with variable thickness to gain insights into the influence of magnet thickness on the performance. Considerable effects of magnet thickness on the performance have been found, including thrust increase of up to approximately 20% and increase in specific impulse by up to approximately 10%, as compared to the configuration with fixed thickness magnets.

The development of new propulsion system for exploring the moon and other planets in the solar system or deep space is necessary to achieve short mission term and large payload ratio. In recent years, magnetic nozzle is focused on as the candidate system to attain above objectives. Magnetic nozzle is the system which generates the thrust by converting the thermal energy of the plasma injected in the nozzle magnetic field formed by radial magnetic field into the directed kinetic energy. The objective of the present study is to clarify the performance of the new proposed plasma source which consists of LaB6 cathode. Furthermore, it is also the objective to measure the thrust and evaluate the performance of magnetic nozzle with LaB6 plasma source. The performance evaluation of LaB6 plasma source is conducted by measuring the plasma temperature and density by applying the double probe method. The thrust evaluation is conducted in the vacuum chamber. In the present experiment, we succeeded in measuring the thrust of thermal plasma and magnetic nozzle. As a result, the thrust of thermal plasma is 4.7 mN, and the thrust of magnetic nozzle is 16 mN. Therefore, the thrust of magnetic nozzle is 3.3 times larger than that of thermal plasma. Furthermore, thrust-power-ratio is 8.99 mN/kW, specific impulse is 4437.8 s and propulsive efficiency is 19.5%. Every performance of magnetic nozzle is also improved than those of thermal plasma.

Flight safety, economics and environmental protection are important factors to consider when designing an aircraft. The severity of the noise and emission output of the aircraft are heavily influenced by the manner in which the aircraft takes off. Aircraft manufacturers have developed and distributed their own software to calculate take-off parameters, departure paths, and noise data. However, this software does not have the capability to accurately estimate pollution output. Therefore, it is necessary to develop accurate mathematical models and software to calculate emission based on aircraft departure flight paths and take-off parameters. This paper presents aircraft pollution emission models based on the International Civil Aviation Organization (ICAO) database and Boeing Methods 2 (BM2) data. Methods for converting (Boeing Climb-Out Program) BCOP’s flight data to meet the calculation of emission are constructed in this paper. After then, the Aircraft Pollution Emission Analysis Tool (APEAT) was developed to calculate pollution emissions for various take-off conditions and parameters. This paper also proposes equivalent CO2 and weighted equivalent CO2 metrics to analyze the overall effect of take-off parameters on airport regional air quantity. Finally, some suggestions were made that can be used to reduce the aircraft pollution emissions in take-off phase.

Beidou satellite navigation system (BDS) is a global satellite navigation system built by China. It will provide global coverage of all-weather, all-time, high accuracy, high reliability positioning, navigation and timing services around 2020. With the development of BDS application, users’ requirements for accuracy, integrity, availability, continuity have been gradually improved. In BDS service performance specification, only signal-in-space continuity is defined, but no positioning service continuity is defined. In this paper, we first put forward the definition of the positioning service continuity standards, and evaluate the availability and continuity of BDS positioning service by using the records from several receiver sites in China. The results show that the availability of BDS can reach 99.47% and the continuity can reach 0.9991/h. The conclusion of this paper can provide theoretical reference for the civil aviation application of BDS.

Due to the recent increase in air transportation demands, safe and smooth traffic flow formation in a congested airspace is regarded as one of the key issues in need of attention. With this in mind, we propose a distributed and real-time trajectory optimization method for traffic flow formation. The proposed method adopts an MIQP (Mixed-Integer Quadratic Programming)-based MPC (Model Predictive Control) to calculate the trajectory of each aircraft merging into the traffic flow. To construct the model, non-convex conditions in the state equations, constraints, and conditions on merging traffic flow are represented as linear or convex quadratic constraints that include some binary variables. We confirm the reasonableness of the computed trajectories and the near real-time applicability of the method by numerical simulations.

Jeffrey’s rule of conditioning is an effective tool to update the current information under the given information. However, traditional Jeffrey’s rule can only process information under the framework of probability theory. So we generalize it based on Dempster-Shafer evidence theory which was seen as a generalization of probability in this paper. In this generalization, interval-valued prior and conditional probability which satisfies weaker conditions is joined with the basis of original Jeffrey’s rule and Dempster-Shafer evidence theory. And we then achieve the update of information by building an optimization model under interval prior and conditional probability. We achieve comparison of interval-valued belief degree with the basis of TOPSIS. One of the main advantages of this generalization is its ability to handle information with wider imperfections. Finally, we demonstrate the application of our generalization on an example of multi-criteria decision-making.

The ordered weighted averaging (OWA) operator is a very useful tool in dealing with the decision-making of multi-dimensional information especially in multi-criteria decision-making. In this paper, we consider measure representation of the uncertain information as a trapezoid fuzzy number, and there is various types of uncertainty in the criteria satisfactions, a new method OWA aggregation of multi-criteria on the framework of Z-valuation is proposed. In our method, firstly, the framework of Z-valuation is built up for the multi-criteria decision problem. Then, OWA operator is used to calculate a final trapezoid fuzzy number to make decision. An illustrative example of the whole procedure is provided.

Two dependent evidences can be viewed as resulted from orthogonal sum of one dependent original evidence and two independent original evidences, respectively. The original method has so many iterations and big calculation, based on these disadvantages, a fast combination method for dependent evidences in the framework of hyper-power sets is proposed in this paper. Equipollent classic Dezert-Smarandache (DSm) rule of combination can be got through importing the commonality function, according to the results of model analysis, theorem proving and example comparison show the feasibility and effectiveness of the proposed method.

This paper describes an effort to develop a realistic and flexible simulation platform that will simulate not only the current or future flight technologies but also the impacts to the whole dynamic airspace network evolution. Serving as a core role for airspace operation, special attention is paid to the flight of aircraft by integrating multiple real-time flight simulation clients into the system, enabling to take human factors into account and improving the degrees of fidelity. A large number of flight simulation clients, air traffic control constraints and meteorological units are connected by a distributed and cluster-type network. In order to meet the extensively requirements for the exploration of intelligent flight and collaborative decision-making research, the design of environment perception functionality for aircraft will be leveraged. Then a test case of 4-Dimensional trajectory is presented. The goal of research is to support the evaluation and promotion of new flight technologies while reducing risk and cost.

Blended wing body (BWB) aircraft usually equips with multi-group elevons for longitudinal and lateral control, those elevons have non-ignorable dynamic characteristics and tend to degrade the transient response to pilot command. However most control allocation methods do not account for effector dynamics, which is suitable for conventional aircraft but not for BWB. In order to handle the effector dynamics properly, four control allocation methods including explicit ganging, daisy chain, generalized inverse and dynamic control allocation are selected based on the guideline of choosing control allocation methods proposed by Boeing Company, then control allocators are designed based on those methods. The four control allocators are evaluated through command tracking and gust alleviation simulation, the simulation results imply that the dynamic characteristic of control surfaces have obvious effect on aircraft initial command response. Besides, the dynamic control allocation method have more potential than three other control allocation methods, but this method still needs further study to improve its performance.

The objective of this study is the development of a helicopter flight dynamic analysis framework with state-of-the-art modelling and simulation techniques. The helicopter model consists of fully coupled continues-time unsteady vortex lattice method (UVLM)-free wake aerodynamic model with multibody dynamics, and is expressed in a set of first order Differential-Algebraic Equations (DAEs). The wake dynamics and multibody dynamics are solved simultaneously by one single general purpose DAE solver. An efficient trim methodology is proposed by constrained motion of the helicopter and Jacobian method, without introducing simplified model or numeric transient. The present method is validated against Caradonna-Tung test, Carpenter & Friedovich experiment and UH-60A maneuver flight test. The predicted airloads, wake geometry and aircraft motion response agree well with experiment data. The computational efficiency of present method is higher than that of time-marching free-wake while keeping good fidelity.

This paper discusses the design and application method of flight control actuation system models based Modelica. According to the physical topology layout of aircraft flight control system, the model development of flight control system is decomposed from top to bottom, and all component parts of flight control system are developed by using development tool based on Modelica language, and the tree structure model library of flight control system is set up. Based on the established model library resources, the flying control system of specific aircraft can be integrated and encapsulated in accordance with the requirements. The model verification work is made through several levels of model verification to ensure that the function of the model has reached the development requirement. The confirmed flight control system models, which can be used to analyze the influence of flight control system on different user cases of aircraft ‘flight scenes, the models also can be used in real-time simulation of human in the loop with visual virtual scene. In the upgrading plan of an engineering simulator, the flight control system models are transformed into FMU models of FMI interface standard and loading into the engineering simulator to run successfully.

The dynamics of spacecraft relative motion greatly counts in the planning, guidance and control of many space missions, such as on-orbit servicing and formation flight. This paper develops a Hamiltonian model for spacecraft relative motion in the framework of analytical mechanics. Considering the structure-conserved property of Hamiltonian systems and deficiency of the classical explicit fourth-order Runge-Kutta method, the symplectic Runge-Kutta method is utilized to propagate the Hamiltonian model. Additionally, the applicability of the Hamiltonian model is numerically analyzed via a comparison with the well-known Clohessy-Wiltshire equations and Tschauner-Hempel equations.

Traditionally iron bird integration rig is a key test platform for testing and integration of mechanical systems for an aircraft under development. To fully verify the aircraft-level functions and requirements associated with mechanical system, avionics and electrical systems, a new method of testing is proposed based on a new pilot-in-the-loop integrated platform in the lab. The platform couples existing iron bird with copper bird and avionics integration rigs using state-of-the-art technology and provides good emulation of the physical and functional status for systems on-board. Examples of test cases performed on it are carefully studied in order to cover the aircraft’s safety and functional aspects before and after its first flight. The practices of the tests so far are also demonstrated in this report.

The cockpit validation and optimization is the significant phase during the civil aircraft design process. However, the aircraft flight deck and the corresponding sub-systems design cannot be manifested concurrently in process of the development life cycle. As a result, the flight deck modification in the later phase of product design process brings much more difficulty and amount of cost to the OEM. This paper conducts a basic research on the interactive simulation design, especially for the purpose of civil aircraft cockpit assessment and optimization. The methodologies, for example Model-View-Controller (MVC) design pattern, modulation design, and encapsulation and integration have been illustrated detailed and demonstrated with design prototype in the following chapters. This proposed design schema offers the aerospace engineer a preliminary reference or guidance documentation for the cockpit simulation design. Finally, the advantages and disadvantages of the simulation design methodologies for civil aircraft cockpit assessment and evaluation have been discussed and summarized in the last section.

Combining the extreme value theory and the multi-factor coupling system modelling, and considering the randomness of icing and pilot operation, the pilot-aircraft-environment simulation system was built. Monte Carlo simulation experiments were used to extract the extreme values of flight risk key parameters. The thick-tail characteristics of the parameter extreme and the generalized extreme value (GEV) distribution law were verified. The flight risk criterion was defined and the flight risk probability was calculated. This method can provide theoretical support for the prediction and prevention of flight accidents.

The increasing complexity of the aircraft is forcing the aerospace industry to develop innovative techniques for the coming challenges. Simulation has been proven to be a reliable tool that can be used to give benefit to designers, testers, aircraft vendors and crew training. However, the potential of simulation has not fully been explored in some phases of the aircraft development process, especially in the field of aircraft flight deck human machine interaction (HMI) design. System engineering concentrates on the design and application of the whole aircraft as distinct from parts, which looks at a problem in its entirety. This paper documents the implementation of Model Based System Engineering (MBSE) concepts in support of civil aircraft flight deck HMI simulation design, their benefits, and lessons learned that can be applied to the use of MBSE in support of future aeronautics projects. Finally, a Civil Aircraft Research Project: civil aircraft cockpit HMI simulation bench has been already studied for the purpose of flight deck assessment and evaluation.

Air transportation has played an important role in the economic development of the Asia/Pacific region. To accommodate increasing air traffic demand and reduce its environmental footprint, trajectory based and seamless air traffic management should be introduced. As a step towards this, we propose to develop a free route airspace concept for the Asia/Pacific region and assess its feasibility, benefits and problems, and propose solutions. The paper introduces the current airspace situation in northeast Asia, and overviews the European Free Route Airspace initiative that gives greater route efficiency and includes areas with seamless cross-border air traffic management. We then suggest initial ideas about airspace and route design that take into account differences between the European and the Asia/Pacific environments, and the research efforts that will be required to develop a free route airspace concept, and outline our planned study approach and assumptions.

As air traffic management moves towards trajectory-based operations, route flexibility will increase and aircraft will be less constrained by the current fixed air traffic services route network and airspace structures. To gain maximum benefit from this increased flexibility, it will be necessary to plan so-called “wind-optimal” tracks that take account of wind, especially for medium- and long-haul flights. The Electronic Navigation Research Institute (ENRI) has developed a dynamic programming-based trajectory optimiser that can generate ideal wind-optimal trajectories for research purposes, but its high computational burden and the difficulty of applying operational constraints mean that it is not suitable for creating practical flight trajectories. We are therefore exploring alternative means of creating practical wind-optimal trajectories. In this paper, we describe a proof-of-concept program that represents possible aircraft lateral routes as a graph and calculates wind-optimal trajectories by a shortest-path search using fuel consumption or flight time as a metric instead of distance. We present preliminary results of track computation assuming constant altitude and airspeed in the north Pacific free route area, and show a sample qualitative comparison with Pacific Organised Track System tracks and an actual flight track that demonstrates the potential feasibility of our approach.

Path planning is one of the key aspects of autonomous unmanned aerial vehicles (UAVs). In this paper, an effective path planning approach based on a hybrid ant colony optimizations (ACO) algorithm for UAV patrolling in forest fire prevention missions is proposed. The proposed approach takes two steps, namely local path planning and global path planning, to find the shortest feasible path flying through multiple target points with obstacle avoidance. In local planning phase, a dubins-path based A* algorithm is applied to find the optimal path between every two target points, the resulting path would be flyable and safe. Later, the visiting order of each target point would be determined by an improved ACO algorithm in order to minimize the length of the final path. Simulation result shows the proposed algorithm can efficiently find a shortest flight path that fulfills the requirements of UAV based patrolling task in forest fire prevention mission.

For launch vehicles with constant thrust and without auxiliary propulsion system, this article proposes a simultaneous method to solve dynamic online trajectory planning problems with terminal attitude constraints. According to strict engineering requirements on real-time and reliability, advances with a trajectory online planning and closed-loop guidance strategy based on trajectory rapid optimization algorithm. Real-time-ability and reliability of the algorithm can be enhanced by assigning proper guidance cycle, initial value and route constraint control. Based on simulation result, simultaneous method based closed-loop guidance method achieves high precision autonomous guidance and reliable convergence, fulfilled requirements on real-time and route constraints. Furthermore, the method can operate trajectory online re-planning in malfunction situation, can serve as foundation to solve more complicated guidance mission in the future.

For the simulation of aero-optical effects in engineering application, the influence of coherent structure on the optical transmission cannot be considered because RANS method ignores the details of the flow fluctuation. A random phase screen method based on turbulence sphere vortex model is improved to predict the optical distortion caused by instantaneous coherent vortex structure. The main idea of improvement is to relate the sphere vortex parameters to the local dynamic characteristic of the flow field. The simulation of optical distortion in compressible mixing layers shows that the improved random screen method has good agreement with the results from the direct integration of the intensity and the accuracy is great higher than the original method.

Situation estimation and threat assessment in air combat are crucial issues. Due to the limitations of the sensor’s precision, terrible battlefield environment and the deliberately interfere of the enemy, the information of threat assessment is often ambiguous and uncertain. Belief function theory has unique and significant advantages in uncertain information processing and fusion, so it is suitable for threat assessment. Our proposed approach handles information begins with feature extraction using fuzzy sets, then interspecific fusion and time domain fusion are carried out by belief function theory and evidence discount. Finally, the data is processed into a definite number to measure the threat index of the target. In the fusion process, evidence discount improves the quality and accuracy of convergence. The simulation example verifies that our proposed method is not only able to accurately analyze and convey the sensor information, but also has better real-time performance and the ability to process uncertain information.

A new algorithm based on expectation maximization (EM) is presented for identifying the parameters and noise covariance matrices in an aircraft dynamic system. The proposed algorithm contains two steps. The first step is to estimate the state of the system using the Kalman filtering (KF) and the current estimator of these unknows. In the second step, the parameters as well as the noise covariance matrices are recursively updated by using the online EM algorithm and the multidimensional stochastic approximation strategy. In order to make a comprehensive comparison of the proposed algorithm and the traditional algorithm, the proposed algorithm is tested by using simulation data and shows desirable estimation accuracy.

In this paper, cooperative and non-cooperative interception of a highly manoeuvrable aircraft target is studied. Each interceptor has a radio frequency seeker and they can cooperate by sharing relative position data of the target between each other. Based on federal filter, the filtering information of multiple interceptors is fused to provide high performance tracking of the target and solve the problem that radar is vulnerable to deception jamming in guidance phase. An augmented proportional navigation law (APN) is designed to overcome the influence of target acceleration on guidance accuracy. The tracking error and miss distance of the cooperative and non-cooperative interception scenarios are evaluated via Monte Carlo simulations for increasing number of missiles, and the miss distance of cooperative interception with and without deception jamming is compared. The simulation demonstrate the advantage of detecting and dealing with the deception jamming of the suggested method, and resulting a small tracking error and miss distance. In addition, results also show that the average miss distance decreases with the number of cooperative missiles fired at the target.

Both the air traffic demand and the use of drones have continued to expand recently. As a result, the number of collision accidents between aircraft and drone has increased. We propose a stochastic model predictive control (SMPC) system for collision avoidance and landing which takes uncertain information of wind and obstacle positions into consideration. We carried out vertical and lateral simulations with static and moving obstacles using linear aircraft model. The simulation result showed that the controller was able to maintain the reference trajectory. Our proposed SMPC system for collision avoidance also proved to be effective for avoiding static obstacles with constant linear motion. Further improvements are needed to avoid obstacles with more complex, random movement.

For answering the question about whether an approach power compensation system (APCS), using techniques to keep angle of attack (AOA) of an aircraft invariable during carrier landing phase, can hold the airspeed, a basic control structure of automatic carrier landing system (ACLS) with a vertical rate reference (H-dot) was focused to analyze three typical maneuvering scenarios. Then, for AOA-maintained APCS, the effect of its major commands on the longitudinal control ability of H-dot ACLS was simulated by F/A-18A fighter’s control structure. The results indicate that, the primary flight state variables were affected from different levels by the command signals of an AOA-maintained APCS, and such APCS can also hold the airspeed as long as the flight track angle is constant in the steady flight.

Aircraft loss of control (LOC) inflight is studied as a similar reason for accident. The study for the LOC includes the envelope determination and the LOC protection. And the envelope, which can forecast the advance signal of LOC, is needed to be studied first because of the precondition for protecting the LOC. According to the characteristics of LOC, the dynamic envelope based on the region of attraction theory is proposed to be the envelope for analyzing the problems of LOC. And the reason that the dynamic envelope can be regarded as the LOC envelope represents first. And primary reason is that the stability boundary has the similar performance of the LOC envelope. In addition, the relationship between the dynamic envelope and the scheduled control law are established to illustrate that there is a possibility for the dynamic envelope proposed in an online flight control system. And the characteristics of the dynamic envelope under the scheduled control law are studied in details. And the results show that the influence on the dynamic envelope of the flight altitude is the same as that of the velocity, and they can be combined as the dynamic pressure, which is also important for the scheduled control law. In addition, the dynamic envelope determination for a flight control system can be done at the same time with the control coefficient of the scheduled control law.

In this paper, a new optimal linear-quadratic guidance law with terminal acceleration constraints is proposed for intercepting maneuvering target. Because the angle of attack is approximately proportional to the normal acceleration for aerodynamic control missile, not only miss distance, but also terminal acceleration is required to converge to zero for increasing effectiveness of the warhead for destroying armored target in the proposed guidance law. The gain characteristic of proposed guidance law is studied by comparing with traditional optimal guidance law with considering autopilot dynamics lags. Numerical simulations are performed to examine performance of the proposed optimal guidance law, and its robustness with respect to the time constant of autopilot dynamic lags is also studied by considering time constant typical errors.

PIO (Pilot-Induced-Oscillation) is due to the mutual influence between abnormal aircraft response and pilot’s dynamic characteristic which causes the phenomenon that aircraft-pilot closed loop is instable. Therefore, airworthiness regulations (FAA25 CS25 CCAR25) 143 (a) (b) have given specific requirements for these. At the preliminary design stage of control law design, PIO tendency shall be analyzed by engineering prediction methods which will confirm whether the design of the FBW control law is suitable. In the detailed design stage, PIO tendency will be analyzed by means of simulator test with pilot in loop. During verification of worthiness compliance, evaluation criteria of PIO shall be determined according to recommended methods of AC25-7C and verify airworthiness compliance through MOC6 and MOC8. Finally, the data of flight test and/or simulator test with pilot comments show that there is no PIO tendency which meets the requirements of airworthiness regulations.

This paper focuses on the cooperative formation of manned aerial vehicles (MAVs) and unmanned aerial vehicles (UAVs). Firstly, the characteristics of the MAV/UAVs and the mission assignment of MAV and UAVs in the formation is analyzed, and the mathematical model of MAV and UAV are built. At the same time, the relative kinematics equations of the MAV/UAVs formation based on Leader-Follower mode are also built. These work lays foundation for the later research on cooperative operation of MAV/UAVs formation. Secondly, new formation shapes are designed referring the operation experience of MAVs formation and considering the unique features of MAV/UAVs formation. And these formation shapes are encapsulated into a formation library with which the fleet can expand or adjust the formation flexibly according to real battle demands. Moreover, UAV’s maneuvers are designed and these maneuvers are also encapsulated into a maneuver library. Lastly, the flight control system is designed for UAVs in order to enable them to fly stably following expected flight state. Furthermore, the management strategies of MAV/UAVs formation including formation organization, formation maintain and formation reconstruction are analyzed based on the formation library and maneuver library. The result of this paper is meaningful for both theoretical research and practical application of MAV/UAVs formation.

This paper describes the establishment and analysis of a robust attitude control system for a distributed tilt-wing UAV in the flight state transition procedure. Firstly, a complete nonlinear dynamic model of the target UAV was developed. Secondly, a selection of the transition strategy and a unified nonlinear control allocation model were determined, based on which, an optimal linearized control allocation method towards four deflect angles was developed. Thirdly, robust attitude controllers towards four deflect angles were designed through $$ \mu $$ synthesis method and D-K iteration. In the end, a scheduling model is involved to simulate the whole procedure of the flight state transition and the results of the simulation reach the requirement of performance.

The flight safety envelope can be viewed as a safe set which could ensure safety as long as the flight state keep within it. Once aircrafts are impaired, such as engine loss or a jammed elevator, dynamic characteristics of the aircrafts will be damaged and the flight safety envelope will shrink deeply. Unaware of these upset conditions, the flight system may operate under a conflicting envelope. Especially in landing phase, it is difficult to keep the flight state within the safety envelope and the accidents are generally associated with it. Therefore, study on estimation of the safety envelope under upset conditions, is critical to improve flight safety. In order to estimate the safety envelope, the reachability analysis based on the level set method is presented. The reachable set is obtained via computing the Hamilton-Jacobi partial differential equation (HJPDE), which is based on optimal control theory. Examples are provided using NASA’s generic transport model (GTM), and the results can be applied to flight safety risk assessment, providing theoretical guidance for the design of envelope protection systems.

For the integrated control of the missile and the compound adjustable ducted rocket, the Integrated Flight/Propulsion Control (IFPC) system was constructed by hierarchical dispersion method based on the mathematical model of the system and the coupling characteristics analysis between the subsystems. According to the work principle of the integrated system, the integration and the coordination of the subsystems were achieved by the motion control commands and the coordination control commands respectively. The IFPC method was validated by simulation based on typical mission. The results show that the IFPC system can control the motion of the missile and coordinate the operating status of the subsystems effectively.

In recent years, the development of navigation technology improves the reliability in flight of Micro Aerial Vehicles (MAVs), and cooperative operation using multiple MAVs attracts a great deal of attention. The purpose of this research is to construct a system that efficiently explores unknown indoor environments using multiple MAVs when a disaster occurs. In order to achieve the purpose, this paper proposes a cooperative exploration method using a new heuristic function. The heuristic function is evaluated with connectivity constraints between all MAVs and the base station. Therefore, the proposed method is effective when the method is applied to actual MAVs. We implement the proposed method and executed two types of simulations in three different environments. The results show that the proposed method can realize more efficient exploration than the existing method in many cases while the connectivity constraints are satisfied. Finally, in order to explore unknown environments more efficiently, future works are indicated.

Kyushu Institute of Technology has been developing fully reusable space transportation system since 2005. WIRES#015, the subscale flight model of suborbital vehicle will be launched from Mojave Desert, California in 2020. The engine employs thrust vector control (TVC) system to improve the control performance during the ascent phase where dynamic pressure is low for instance at the launcher exit. However, natural frequency of the vehicle attitude motion may become higher than the cut-off frequency of the actuator, when the vehicle experiences maximum dynamic pressure, and the actuator performance becomes saturated. In this research, the attitude control law on the longitudinal and lateral motion using thrust vector control and aerodynamic control surfaces in powered flight was designed based on eigenvalue analysis, and the response of pitch angle and TVC actuator performance was evaluated based on non-linear flight simulation. In conclusion, two findings are obtained. At first, the designed control system has appropriate damping ratio on short period mode and enough stability margin of open loop transfer function, which meet the control design requirement. Second, flight simulation shows that the control law stabilizes vehicle attitude with the influence of gimbaling reaction torque, while TVC actuator is not saturated. In future, possible disturbances such as steady wind and combined use TVC and aerodynamic control surfaces for fault tolerance and high angle of attack maneuverer in abort flight will be installed into this flight simulation.

The application of touch screen in the cockpit of civil aircraft is a trend for the development of wide-body aircraft. So this paper analyses those elements that should be considered when the touch screen is introduced into the civil aircraft cockpit for application. The paper firstly combs the requirements of the current aerospace standard or specifications for touch screen design, then points out the design requirements of the characteristics of touch screen itself, the human-machine interface, system integration, and environment factors which should be considered when the touch screen is applied to the cockpit. Finally, the ergonomics requirements caused by the touch screen in cockpit application are also analysed based on the human factor handbook, human factor reports and the characteristics of the touch screen. The paper aims to provide a design reference for the application of the touch screen in civil aircraft cockpit.

In order to realize reusable winged sub-orbital transportation system, the Space Systems Laboratory, Kyushu Institute of Technology has worked in research, development, and flight test for winged rocket which is called “WIRES” (WInged REusable Sounding rocket) since 2005. Autonomous flight system is one of the most important research fields for space transportation system.As space transportation vehicles have important nonlinear dynamics, control system design tends to be complex and difficult. Dynamic inversion (DI) is known as one of effective nonlinear control theories, and authors try to apply DI theory to attitude control for WIRES.DI controller is concerned about increment computational costs, because it is necessary to calculate the inverse dynamics. In this research, authors evaluate the calculation cost of DI controller by implementing to WIRES#015 which is subscale test vehicle of suborbital flight.As a result, the cost is small enough compared to control cycle. It can be said that the microcomputer has enough performance to realize longitudinal attitude control.

A method of controlling flight-path angle via adaptive backstepping terminal sliding mode is presented. At the first two steps of the design process of controller, Radial Basis Function (RBF) neural networks are employed to approximate the unknown parameters uncertainty online and the dynamic surface control is combined with backstepping design technique to design the virtual controller, so that the explosion of complexity in traditional backstepping design is avoided perfectly. In the last step, the high-order sliding mode control law is designed by the non-singular terminal sliding mode to eliminate the chattering and make the system robust to uncertainties. It is proved by Lyapunov method that all signals in the closed-loop system are semi-global uniform ultimately bounded, and the tracking error can be adjusted by adjusting the controller parameters to converge into a small neighborhood of zero. Finally, the simulation results demonstrate the validity of the proposed method.

We studied the basic UAV technology necessary for an arresting gear method under freezing conditions or unraveling of the runway surface occurring in the severe weather typical of Mongolia: A large number of tests were conducted in both winter and summer, and further analysis was conducted using a numerical model. We concluded that the wire model using multiple weights with free endpoints is optimal for the UAV arresting gear system. Application of this model resulted in safe stopping on a freezing runway after a straight run. Fixing the end points or fixing only one side created a danger that the aircraft would change direction if the UAV’s hook hit the wire at a point other than the midpoint. Using numerical analysis, a simulation that included large friction and dampers caused the UAV speed to decrease linearly, so the UAV took time to stop. Conversely, the collision model with multiple weights showed that no matter where it was caught by the wire, the UAV did not change its direction, ran straight on the runway, and showed a parabolic deceleration during the experiments. The method resulted in safe landings. We recommend using this method to reduce the UAV operation and development costs because the crashes of UAVs mostly occur on the runway during landing.

Ground-based augmentation system (GBAS) of the Global Navigation Satellite System (GNSS) can use ranging error correction to achieve satellite integrity monitoring. For purposes of integrity monitoring, GBAS uses error overbounding method to calculate protection level (PL). The general error overbounding modeling method is to establish a zero-mean Gaussian model for pseudorange error. However, there are problems such as “thick tail”, “asymmetric” and “zero mean” in practical applications. In this paper, we use the properties of stable distribution to propose a four-parameter stable distribution modeling method. Based on the estimation of the characteristic parameter and the dispersion coefficient, the stable distribution of symmetry parameter and position parameter are estimated at the same time. Simulation analysis shows that the four-parameter stable distribution method improves the accuracy of the protection level calculation.

In this paper, a multiple model Kalman filter (MMKF) is proposed for attitude/position estimation of monocular vision, in which the noise disturbance and data loss during camera imaging process is considered. Firstly, by establishing an acceleration model of artificial markers in the camera image plane, a Kalman filtering (KF) strategy is adopted to deal with the disturbance and data loss problem of markers. Especially for the data loss situation, an extended Kalman filter (EKF) with equality constraints, taking into account the prior information between the artificial markers, is designed. Furthermore, after comprehensive consideration of the computational complexity and accuracy of the above two strategies, we present a novel MMKF estimation scheme to ensure the speed and accuracy of attitude/position estimation. The effectiveness of the proposed MMKF estimation strategy is evaluated by numerical simulation experiments.

The global navigation can be provided by GNSS satellite constellations such as GPS, GLONASS, and the recently emerging BEIDOU and GALILEO in medium and higher earth orbits. Generally speaking, those GNSS satellites, that are in the altitudes of about 20,000 km, move around the Earth with an orbital speed of approximately 3.88 km/s relative to the Earth. According to the Doppler Effect, the relative movement of GNSS satellites and GNSS receivers lead to a Doppler shift in the GNSS signals frequency on the receiver. It would be even worse if the GNSS receiver also has a high velocity relative to the GNSS satellites. A GNSS receiver in space, especially in Low Earth Orbit (LEO), suffers from the high dynamics. A satellite in the altitude of 550 km travels in an orbital speed of approximately 7.58 km/s which leads to a high Doppler frequency shift of about ±60 kHz. That is why tracking the GNSS signals for LEO satellites is always of a great challenge. In this paper, it is attempted to suggest and analyze a suitable method which can surmount the difficulty of high dynamic GNSS signal tracking in LEO. Vector tracking is the method which is selected for this purpose and its performance is analyzed regarding the requirements. It is shown that compared with FLL tracking loop, vector tracking increase the dynamics by at least 200%.

Design of the Korea Lunar Pathfinder Orbiter (KPLO) attitude control subsystem is a unique challenge that requires refined models that simulate the cis-lunar and lunar orbital environments including coordinate frame changes, and a new sun-pointing control mode that satisfies Sun and communication pointing constraints. This paper introduces LUNASIM, a simulator for verifying the attitude control performance of KPLO. The simulation architecture is described, and attitude control performance verification results of the sun-pointing control sub-mode are presented.

On-orbit target observation, fault diagnosis, and robotic on-orbit service (OOS) for malfunctioned satellites has attracted people’s extensive attention. One of the key technologies for OOS is the relative attitude determination between the target and chaser. This paper propose a new method to determine the target attitude using point cloud data. First, the feature planes are extracted from the point cloud dataset. Secondly, the rectangular planes are recognized, and then the side length parameters are determined to identify which plane of the satellite is being observed. Thirdly, the normal vectors of the feature plane are utilized to determine the target attitude. Experiment is performed to evaluate the propose algorithm by using a TOF camera.

In order to meet the needs of aircraft multi-electric development, an electric drive emergency release system was designed for the upper gear lock mechanism of a certain type of aircraft landing gear. Firstly, the mechanical equilibrium equation of the upper lock mechanism was established and theoretical analysis was carried out. Based on LMS Virtual. Lab/Motion and AMESim software, the dynamic models of the emergency lock release system of the landing gear and the electric drive system under different braking modes were established respectively, and the joint simulation was further realized. Finally, the effect of different braking modes on electric drive emergency release system was discussed in detail. The results show that the electric drive actuator of the short-circuit brake system, where the brushless DC motor drives the ball screw to unlock the upper lock and realizes the emergency unlock function, possesses the best driving effect. The analysis results provide theoretical support for the design of electric drive actuators of aircraft landing gear emergency system.

This paper is aimed at a design of a local electro-hydraulic generation system following the trend of more electric aircraft. This design integrating a motor and a pump is called electric hydraulic pump and it is applied as the power unit of the whole system. Modeling and sub-models based on AMESim are built as well as the primary experiment validation. To enhance the precision of the model in AMESim and verify the pressure fluctuation as well as the pulsation when the system is working, the model of pump designed as a super component with nine pistons is present. Furthermore the leakage from the swash plate and the slipper, pistons and the valve plate are taken into consideration. The preliminary experiments show that the performance of the local electro-hydraulic generation system is also acceptable. Electric hydraulic pump that brings the weight and noise benefits can be the substitute of traditional pumps and motors.

Thrust vector control (TVC) actuation system is an important part for swing angle control of the launch vehicle nozzle. In this paper, the control methods of electro-mechanical hydrostatic actuator (EMHA), a new type of actuators, are studied. The block diagram of the system is firstly introduced based on the mathematical relationships of key components. Then the PID control with multiple loops of the system is given. The simulation analysis shows that the notch filter can prove the dynamic performance of the system. Finally, the dynamic performance analysis of the system adding the saturation limits is also done. The paper shows the control algorithms of EMHA and provides theoretical support for the subsequent controller design.

Aiming at the current popular fault methods’ disadvantages on diagnosing sensor faults existing in electromechanical actuator, this paper presents a fault diagnosis approach based on dynamic parity space. Firstly, we derive the mathematical model of EMA and establish the state space equation. Secondly, a sensor fault diagnosis approach based on structural residual parity space is introduced. Lastly, the sensor fault such as current sensor fault, motor rotary encoder fault and LVDT fault is detected and isolated based on structural residual, which prove that the proposed fault diagnosis approach is effective in diagnosing the sensor fault of electromechanical actuator.

Starting system is an important part of aero-engine. The power output characteristic of starter determines the starting reliability and starting quality of the engine. Direct-current (DC) starting motor is one of the aero-engine starter, aiming at the parameter selection and power matching in starting system design first, the characteristics of the DC starting motor four stages starting process are analyzed. Based on the voltage balance and torque balance, the mathematical model of the starting motor system is established, then the differential equation model is transformed into state space model and a simulation model is built in the Simulink/Matlab environment. In the case of given aero-engine starting load, the dynamic response of the starting motor system is simulated, and the simulation results show that the parameter s are coincide with the theoretical trend and the model can meet engineering requirements. The model can provide a basis for parameter selection and starting power matching of engine DC starting motor.

This research is based on the scientific special project “research and application validation of key technologies for intelligent production line of aircraft ducts”. This project number is MJ-2016-G-64. This paper is aimed at the small holes on the flange face of the aviation conduit, two different measuring methods are discussed for measuring the aperture and center coordinates, that is, the accuracy and repeatability of projection method and cylindrical fitting method. The triangular laser scanning probe used in the system is more convenient for data acquisition and less affected by light occlusion. With this system, high precision, high efficiency and high repeatability can be selected.

A new method for calculating flutter characteristics is presented using Arbitrary basis mode shapes and reduced-order modeling (ROM) techniques. It can be applied to aeroelastic uncertainty analysis efficiently and accurately, without recalculation of normal modes and aerodynamic force though structural parameters vary. First, a number of normal mode shapes of different structure samples is calculated by changing the baseline structural parameters, such as mass or stiffness variations. Then the reduced arbitrary basis mode shapes are extracted from the above mode shape samples using the principal component analysis (PCA) method. Therefore, the physical mode shapes of varied structures can be regarded as linear combination of the reduced arbitrary basis mode shapes. Hence, under the coordinate system of reduced arbitrary basis mode shapes, there is no need to recalculate the physical normal mode shapes when structure parameters vary. The Modal Assurance Criterial (MAC) is used to evaluate the accuracy of normal mode shapes represented by the reduced arbitrary basis modes. Afterwards, the generalized aerodynamic force coefficients under the dynamic motion of arbitrary basis mode is calculated by the CFD technique, which is identified to a reduced-order model by the observer method. Finally, in order to perform flutter analysis with aerodynamic reduced-order model under the coordinates of arbitrary basis mode, the uncertain aeroelastic equation is deduced under the new coordinate system, to consider the structural variations. Flutter analysis and uncertainty quantification is conducted on a flat plate by the polynomial chaos expansion (PCE). Compared with the Monte Carlo Simulation, results indicate that this method with ROM and PCE is accurate and much more efficient.

A nonlinear flutter wind tunnel test of a very flexible wing model has been conducted at the FD-09 Wind Tunnel of China Academy of Aerospace Aerodynamics (CAAA) to acquire experimental data suitable to correlate with and validate nonlinear aeroelastic analysis methods. Nonlinear aeroelastic analysis methods are necessary to accurately predict the aeroelastic response and critical velocity of high altitude long endurance (HALE) aircrafts, such as Helios or Zephyr. However, very little test data is available for validating such methods up to date. A 1.5 meter-long 0.05 meter-width flexible straight wing model was designed, whose aspect ratio is 30. The model was vertically downwards mounted on the rotating mechanism of the wind tunnel for adjusting root angle of attack (AOA). The model was fully instrumented to collect structural response data, such as strain and acceleration response data. Two different test cases were performed, with root AOA was 1° and 2°, respectively. During the test process, a low speed instability region was discovered and limit cycle oscillation (LCO) behavior was observed. As increasing in the flow speed, the aeroelastic response of the flexible wing model experienced the damping vibration, simple harmonic LCO and periodic LCO phase, but no divergent behavior occurred. A maximum deflection of the wing tip was about 30% of the model length.

This paper analyses the reasons for the serious drop in lift after the deformation of the large sweep angle aircraft. It is determined that the lift loss is mainly related to the sweep angle and the bending deformation amount of the wing, and also determined that the inevitability of the lift loss according to the traditional composite layering method. Drawn on the successful experience of aviation aircraft design, the optimization method of stiffness tailoring for aircraft wing and the rounding treatment method for optimized composite layer are proposed. With a composite aircraft as the object, the optimized design, processing and ground test verification of the wing are completed. Studies have shown that, after the stiffness tailoring optimization design, the maximum bending deformation is almost constant (2% increase), and the deformation difference between the front and the rear edge is reduced by 80%. After optimization, it not only satisfies the design requirements of pure bending deformation, reduces the lift loss from 32% to about 4%, but also reduces the structural weight by 12%. The research results have a certain guiding significance for the subsequent aircraft designing.

The recent computational progress in multidisciplinary time-domain analysis calls for an efficient grid deformation method which can produce the deformed grid with high quality and can consider the rigid movement and elastic deformation meanwhile. To meet these demands, a highly efficient grid deformation strategy based on the proper orthogonal decomposition (POD) method is developed. In this strategy, the accuracy, efficiency and quality preserving which may contradict with each other in a single grid deformation method are separated and implemented in different methods, such as finite element interpolation (FEI) for accuracy, POD for efficiency and radial basis functions (RBFs) for quality preserving. There are two stages in this strategy: (1) In the pre-processing stage, snapshots, POD basis vectors and RBF coefficients of POD modal weights are calculated in sequence and the focus is quality preserving in this stage; (2) In the simulation stage, the deformed grid is calculated by the weighted summation of POD basis vectors, which is followed by shape preserving process based on the FEI and transfinite interpolation (TFI). The second stage focuses on the efficiency and accuracy. The strategy is then applied to a 2-dimensional biconvex airfoil and a F6 wing-fuselage conjunction. The deformed grid quality is shown to be preserved as well as the RBF’s, while the grid orthogonality is also maintained well, especially at and near the moving surface. As for the efficiency, this strategy is much better than the RBF: 24.3 times for the 2-dimensional biconvex airfoil and 1810 times for the F6 wing-fuselage conjunction.

In this paper, a wind tunnel experiment model for elastic aircraft with high-aspect-ratio is designed. The numerical simulation method is used to predict the exact state of the wind tunnel experiment, and the low-speed wind tunnel experiment is carried out. The purpose of this study is to verify the effect of geometric nonlinearity on the aerodynamic force of aircraft with high-aspect-ratio and to obtain the aeroelastic trim parameters and influence of elastic deformation on aerodynamic forces. Some parts of the model are fabricated by 3-D print using nylon to reduce weight. The aircraft aeroelastic deformation is measured using the principle of binocular vision measurement. Aerodynamic is measured using balance and angle of attack is measured using attack angle sensor. Research show that the wind tunnel experiment device and supporting measurement device designed for high aspect ratio aircraft can effectively acquire the influence of geometric nonlinearity on aircraft aerodynamic force.

The frequency of the structure is the important dynamic characteristic of the structure, for aircraft, when the wing and other large scale structures deform under the action of unsteady aerodynamic forces, its frequency will also change. For large flexible structural aircraft represented by high altitude long endurance UAVs, it is very important to explore the variation rules of the structural dynamic characteristics of the aircraft in order to develop the design technology of the aircraft and improve the design method of the aircraft. This paper introduced a test method of the ground electromagnetic dry wind tunnel. A principle experiment method of ground electromagnetic dry wind tunnel is established. The electromagnetic field and Ampere force are used to realize the simulation of aerodynamics force, and the model is statically and dynamically loaded without touching the model. Based on this new test technique, a quasi modal test of a large flexible structure is carried out to study the dynamic characteristics of the structure under the condition of large deformation.The results show that the bending deformation of the structure affects the bending and torsion frequencies of the structure, and affects the dynamic characteristics of the structure. This new test method is conducive to the study of aeroelastic characteristics of large flexible structures, and is of great significance for the design and development of new aircraft.

The flying wing has an unconventional aerodynamic configuration with many advantages over traditional aerodynamic configurations. However, due to the light weight and flexibility of the flying wing structure, the aeroelastic problem is more significant. In the flutter analysis, it is found that the short-period mode of the flying wing is coupled with the bending mode of the wing that leads to flutter instability, which is called body freedom flutter. In order to study the body freedom flutter, This paper introduces two typical test cases. We use MSC.Patran to establish the finite element structural dynamic model of the flying wing and use MSC.Nastran for modal analysis. The aerodynamic lifting surface model of the flying wing is built in MSC.Flightloads. The flutter characteristics of the model are calculated by MSC.Nastran, and the flutter speed and frequency under the different stiffness are obtained. Finally, comparative analysis was performed using ZAERO software.

Compared with traditional composites, interpenetrating phase composites (IPCs) consist of two phases which are each interconnected in three dimensions. The mechanical properties of IPCs are seriously determined by their various micro-structures. Therefore, it is very important to obtain a reasonable mechanical model to characterize IPCs based on their preparation technologies and realistic micro-structures. A routine is compiled to describe the spatial distributions of interpenetrating phases by solving the phase field equation. And a 3D random finite element (FE) model based on the phase field method is presented which can characterize the realistic microstructure of IPC. The main content of this paper is vibration damping properties of polymer-metal IPC cantilever beam, combined with theoretical analysis and FE analysis. Based on viscoelastic cantilever beam vibration theory, the theoretical prediction formula of loss factor and the natural frequency are deduced. The vibration behaviors of polymer-metal IPC are simulated, and the predictions accord well with experimental data.

Experiments enforcing bearing load were occupied on GLARE 2A 6/5 laminates in this paper. The damage progression and failure modes of specimens were observed using C-scan, macro photography and scanning electron microscope (SEM). Test results indicate that the damage of those laminates initially occurred as plastic deformation of metal layers, while it progressed, delamination appeared and developed due to plastic deformation becoming larger and bearing loads mainly carried by the matrix of 0° ply. As load increased, matrix in 0° ply cracked, followed by delamination progressed rapidly and few fibers buckled in the region near the hole, triggering final shear-out failure of the laminates. To predict final failure modes and failure loads, a finite element model was developed in which shear-out failure in GLARE laminates are classified as in-ply failure and inter-ply failure, modelled with Hashin criteria in strain form, cohesive element approach and alloy’s plasticity. This model is also capable of predicting the locations where in-ply damage and delamination occur firstly as well as simulating the progression of damage. The calculated results are in good agreement with test results, meaning that the model is able to simulate the behavior of GLARE laminates under bearing load effectively.

Brushless DC motor (BLDCM) is increasingly used in industrial and military applications. This paper presents the design of electromagnetic performance and structure for a small BLDCM used in aircraft fuel pump, then test the motor for the design. The result shows the effectiveness of the small-size technology and the design of the sealing structure, and can be used as a reference for similar motors.

The performance of the metering device of aeronautical vehicle can not be carried out without the feedback of displacement sensor. In order to control the engine accurately, the metering device for controlling the fuel oil quantity in the aeronautical vehicle needs accurate displacement feedback. In order to realize more lightweight and integrated aviation metering device, the measuring principle of capacitive grating displacement sensor is introduced. According to the principle of capacitive grating displacement sensor, an integrated displacement sensing structure is proposed, and the displacement sensing layer is integrated on the surface of the structure. Realize the integration of sensing and structure. The structure and material of integrated sensing layer are studied, and the influence of material on sensing precision is found. The integrated displacement sensing metering device is designed, and the method for acquiring the stable signal is obtained. The testing platform of micrometer-level displacement feedback is built, the resolution of the integrated displacement sensing device is 1 µm, the measuring range is 15 mm, and the integrated displacement sensing structure replaces the original separated displacement sensor, thereby reducing the structure volume and the weight of the aviation metering device.

Considering of local energy loss of section that the fuel flows through, the equation of fuel flow rate flowing through metering needle was derived, based on which, the function, structure and working principle of constant pressure difference valve were analyzed detailed. According to continuity equation and force equilibrium equation, by using linearization method, the model of constant pressure difference valve was established, and the structure diagram was drawn. In terms of the structure diagram, the transfer function of constant pressure difference valve system was attained, and then, the stability was analyzed. The steady-state error of step input was calculated. The results show that the modeling method which is based on continuity equation and equilibrium equation is feasible, the condition of stability and the main influence factors of the steady-state error are attained. The model of constant pressure difference valve is the base for constructing fuel regulator model and simulation calculation.

To ensure the relative consistency of cumulative direction and size for the accumulated error between two assemblies, controlling method for assembly coordination accuracy was studied under digital assembly environment. Based on the qualitative analysis on the main factors that would cause the consistency problems, and the key technical problems to be solved on interchangeability and coordination for actual production, the coordination accuracy guarantee method was developed. It can be decomposed by four detailed controlling schemes, i.e. (1) identification and controlling of coordination elements for aircraft products, (2) analysis on the mapping relations between the coordination elements of aircraft products and the manufacturing process coordination elements, (3) analysis on assembly coordination relationship, and (4) analysis on assembly error propagation and coordination error chain construction between different assemblies. Then a whole closed cycle process, namely “4 + 1”, was proposed to accomplish the feedback adjustment process for coordination dimension. With the adjustment on manufacturing process and the modification on coordination error links, the research on the controlling method for coordination accuracy can improve the quality and efficiency of the dimension and shape transfer process.

With the fast development of the society economy in recent years, the Chinese aviation network is playing a more and more critical and irreplaceable developing role in today’s integrated transportation system. In this paper, an empirical structural analysis of Chinese airline network is made utilizing complex network theory. The results show that the presented network is a small network with two regime pow-law degree distribution. Meanwhile, the betweenness-degree relevance takes a positive correlated relationship, while its node average degree-degree relevance is obviously negative assortative reflecting high degree nodes preferring to relate to small degree nodes, and the clustering coefficient distribution between degree follows negative correlated relationship reflecting node with small degree tends to cluster into group.

Traditional empirical formula of stress intensity factor (SIF)/energy release rate (ERR) by double cantilever beam (DCB) specimen ignores the effect of the deformable uncracked segment of DCB specimen. Based on plane theory of elasticity, this paper presents a general solution of SIF/ERR of DCB accommodating the strain energy of deformed uncracked segment. This improvement provides a theoretical solution which is consistent with numerical results of finite element method, and effective, accurate, and valuable in engineering designs.

A new kind of metal spring tire is put forward and the variation range of the vertical displacement for inner and outer springs are solved by establishing the mechanical models. The inner spring is the main load-bearing part of the tire in the vertical direction. So the inner spring’ vertical stiffness is simulated and its theoretical model is validated. In order to get the exact solution of the stiffness values of the spring tire, a stiffness test system is built. Finally, the three-direction stiffness values and grounding characteristics of the tire are obtained, and the corresponding theory and simulation model are further validated.

Combining with the effect of root deformations of the double cantilever beam (DCB) specimens, this paper presents two kinds of variational-analytical solutions of stress intensity factor (SIF)/energy release rate (ERR) for Mode II crack through force method and displacement method. The plane theory of elasticity is implemented for deformation and stress analysis of DCB in these two kinds of solutions. Comparison is made between the present solutions and finite element solutions which clearly substantiates the former accurate, applicable and effective in engineering designs.

After the conceptual design of a six-rotor UAV, a detailed 3D model was built in CATIA, in which complete details of the structure was designed. Then static and dynamic structural analysis was carried out by using ANSYS. By applying the loads and the other forces which will effect on the structure, analysis was done on the complete UAV. Also, the structure analysis is carried out on the landing gear by applying all the forces which are acting on it. Through the analysis of the stress on the six-rotor UAV under different working conditions, the results show that the UAV is strong enough under the loading condition of 5 kg.

In this paper, a method for fatigue life prediction of the actuator structure is proposed. The size and direction of the first principal stress of the part are obtained by finite element software. Considering the size coefficient, surface processing coefficient and heat treatment coefficient, stress concentration factor is introduced. Based on the stress concentration factor, the standard S-N curve of the material manual is corrected by modifying the first principal stress. Finally, the fatigue damage of the structure under the design fatigue spectrum is estimated with the linear cumulative damage theory, thus providing a feasible way for the actuator fatigue analysis. Compared with the method of linear interpolation for fatigue life, this method is more secure and more in line with high safety and reliability requirements of civil aircraft.

Composite material tail fin of helicopter has complex structure and lay-up, whose strength and stiffness are verified by experiment before, and whose strength analysis is not carried out. In this paper, MSC serial software is applied to create the finite element mode of composite material structure for helicopter tail fin so as to verify the structure strength under certain critical load cases by Tsai-Wu failure criterion. The calculation results are given to determine the load-carrying characteristic of structure, further to verify the experimental results of the structure. The structure can be evaluated according to the analysis and the experimental results.

In this paper, taking the planetary roller screw for aircraft electro-mechanical actuator as the research object. The planetary roller screw exist four different designs which are widely used for different applications: these four designs are respectively known as standard planetary roller screw, inverted planetary roller screw, recirculating planetary roller screw and differential planetary roller screw. The composition, thread form, working characteristics and applicable occasions of these four different designs are analyzed and compared. Considering the possible relative slide between components, the angular and axial motion relationships are demonstrated in detail, and the relationships of structural parameters between components were obtained. The force analysis was performed at the contact point to analyze the transmission efficiency. The relation between the transmission efficiency and contact angle, helix angle was also deduced in consideration of the rolling friction and the relative sliding friction of the component. The results show that the transmission efficiency of the planetary roller screw rapidly increases with the increase of the helix angle and then gradually decreases. In addition, the efficiency gradually increases with the increase of the contact angle, and finally tends to be stable.

Civil aircraft load design regulations generally require load calculations for various combinations of weight, center of gravity, moment of inertia, speed, and altitude etc. Considering different maneuvers and gusts, the number of load conditions will reach tens of thousands. If we consider the load of each point in the time history of the simulation, the number of load points will be unacceptable. It is very difficult to filter out the valid data from a large number of databases and get the critical load of the aircraft. On the one hand, if the selecting method is not precise enough, it may cause some critical load conditions to be missed, which will affect the structural safety of the aircraft. On the other hand, if the selecting method is too complicated, it may lead to too many critical load conditions and some invalid data, which will influence the efficiency of the aircraft design. Based on years of engineering practice, this paper proposes a kind of efficient and reasonable selecting methods for civil aircraft load calculations, which has been verified in several types of aircraft design.

The objective of this study was to develop a weight design platform for hybrid wing body (HWB) sizing and analysis with an emphasis on Vehicle Sketch Pad (VSP). A new method for HWB platform discretization has been introduced and studied, which ensures the platform is smooth and continuous. With VSP model, multi-fidelity analysis tools have been linked, and this integration facilitates the configuration design and accelerates the design iteration. The wing-body is taken as a major component of HWB in the example, and the model is generated in VSP. The example shows a good capability for a rapid sizing and analysis of the HWB configuration.

Compressor component work at high temperature, high pressure and high centrifugal force, the shape at cold state is different form hot state. The aim of a component cold-hot state conversion is to obtain the original size and shape, by the condition of knowing the size after deformation and the load at the hot state, which will be subtracting by the size difference caused by temperature, pressure and centrifugal force. Conventionally, the Cold-Hot State conversion carried out by compressor engineer is more like a “tip clearance conversion”, people pay more attention to the clearance between blade tip and casing, after gaining the deformation of blade and casing, they just translate the blade tip or casing to get an appropriate clearance to make sure the compressor work safety. For a compressor with small size which the blade transformation is tiny, this traditional method may be appropriate, but while the size of compressor is larger, the blade deformation itself cannot be neglected, so, in this paper, a finite-element thermal-structure based iteration modify method was established, which can achieve the goal of both cold to hot and hot to cold bidirectional conversion.

The rapid development of the additive manufacturing technology (3D printing) has brought much greater optimization space for the structural innovation design. The combination of topology optimization design and additive manufacturing technology can further reduce the structural weight, thus improving the bearing efficiency of the structure. The main purpose of this paper is to explore the feasibility of reducing weight for civil aircraft structure by means of optimization design. The structural topology optimization design of typical hinge structure for civil aircraft is carried out. Compared with the traditional structure, the topology optimization structure achieves a weight reduction of 25% and a structural stiffness increasing of 75.7% under the condition that the strength requirement is satisfied, which reaches the expected goals.

A geometric modeling method for cold blade is proposed based on hot-to-cold finite element (FE) transformation calculation. The main idea is to interpolate the deformation result calculated from hot-to-cold transformation to the control points of the hot blade profile. Subtract the interpolated displacement values from the coordinates of the hot control points then the coordinates of the cold control points are obtained. The resulting coordinates could be directly used for solid modeling in commercial CAD programs. Both the hot-to-cold FE transformation calculation part and the output part of the cold blade profile’s control point coordinates could be accomplished in a whole process automatically in the commercial FE program ANSYS15.0. The flow chart of the entire process is given attached with ANSYS APDL commands. Finally, a numerical example based on a blisk sector model is provided. The hot-to-cold FE transformation calculation is performed, the finite element model (FEM) is obtained and solid model is built. Additionally, the effect of mesh density on the modeling accuracy is investigated.

The helicopter rotor lead-lag damper dynamic characteristic modeling has the important role and significance for helicopter rotor system dynamic analysis in all three major research fields that are structural loads, motion response and system stability. The difficult part of this work is the variety of dampers means different working principle and the nonlinearity of the stress-strain relationship. The existing model can’t cover all those problems above and the lacking of a unified form both has brought inconvenience to rotor system dynamic modeling. The approach of helicopter lead-lag damper dynamic characteristic modeling based on the fractional calculus theory is attempted to cover elastomeric damper, fluid-elastomeric damper and MRFE damper. By adding the initial order variable function, the influence of current amplitude on MRFE damper dynamic characteristic is considerate. And the possibility of using fractional derivative damper model in research on fluid-elastomeric damper and MRFE damper dynamic characteristic is validated.

A time-varying mesh stiffness model of spur gear pair, considering the influence of friction, is established based on potential energy method. The time-variant characteristic of the mesh stiffness can be represented more intuitively, by expressing the model with angle variables. And it is feasible to gain the friction coefficient of any meshing point, by establishing the time-varying friction coefficient model under mixed elastohydrodynamic lubrication. And in the model, the viscous-pressure and viscosity-temperature effects are taken into account. The results show that the friction coefficient, which has obvious time variability, is mainly affected by tooth surface roughness and sliding speed. Meanwhile, for gear mesh stiffness, friction can directly influence its value which changes more significantly in single tooth meshing area. And it can be decreased by lubrication. In addition, the model built in this paper is relatively close to the actual running state, so that it can provide a new time-varying mesh stiffness model for subsequent gear dynamics analysis.

Aiming at the problem that modal damping of mechanical system is difficult to be identified accurately, and then a method called phase tangent slope method was presented. The method is utilized amplitude-frequency and phase-frequency curves of vibration response data, and then modal damping ratio is indirectly gained through calculating the phase tangent slope at the resonance point. The method was applied in simulation analysis and experimental verification and compared with the half-power bandwidth method, and the results demonstrated that the calculated modal damping ratio in the simulation analysis is very close to the theoretical value, and the maximum error of the two method is less than 0.06%. In the modal damping identification of the real rotor, the mean values of the identified modal damping ratios in the two directions have the same order of magnitudes, and their relative error is 7.18%, which further validates the correctness of the method.

The control surface in flight suffers severe working environment and usually a kind of colored noise excitation. Sometimes the modal parameters of control surface will be missed when using NExT-ERA identification method. By analyzing the characteristics of colored noise, it is pointed out that the uneven distribution of excitation energy is the main reason for the modal omission. Combined with stability diagram and fuzzy clustering, an automatic identification method based on NExT-ERA is put forward in this paper to solve the problem of modal omission. At the same time, the problem of NExT-ERA which needs manual assisted ordering is solved. Finally, the method is validated by an experiment of control surface excited by shaker table with two typical colored noises.

Taking the high bypass ratio turbofan engines as the research object for the dual-rotor system under the load of blade loss, a simplified four disk dual rotor system is established. The motion differential equation of the dual rotor system is established by the finite element method, and the transient response of the dual rotor system under the blade loss load is simulated and calculated with the Newmark-β integral method, considering the effects of rotor speed, eccentricity, supporting stiffness and damping. The results show that the transient response of the system has obvious impact characteristics under the excitation of the blade loss load, and the stiffness and damping of the bearing at the rear of the fan have a significant influence on the transient response.

In order to measure the radar cross section (RCS) characteristics of inner surface of Spherical Convergent Flap Nozzle (SCFN), low electromagnetic scattering carrier is designed, and its RCS is measured in the anechoic chamber. In addition, we use the inverse synthetic aperture radar (ISAR) imaging method to obtain the scattering center distribution of the carrier and the carrier with the SCFN. The results show that the RCS of the carrier is much smaller than the RCS of the SCFN in the range of ±50°, and the average of the difference between the two first increases and then decreases with the rise of the frequency, the maximum value is −19.8079 dB. Experiments have shown that this low electromagnetic scattering carrier can be used for RCS measure of cavity targets.

At present, the study of Computerized Numerical Control (CNC) machine tool state monitoring is still in the primary stage, and the mechanism of fault early warming has not industry norms yet. Most of the internal information collection of CNC system are used for the display of machining state, and the research of external information collection and processing mostly installs the sensor on the specific parts, which is lack of engineering feasibility. This paper puts forward a method of CNC machine tool state monitoring based on two-dimension feature. The internal data of CNC system and the data of external sensor are integrally studied. Firstly, the internal collection method of CNC system based on Dynamic Data Exchange (DDE) is proposed. Through running the standard program, this paper carries on the automatic collection, processing and analysis to the key indexes, such as power, current, temperature and so on. Secondly, by calculating and analyzing the characteristic frequency of the key components, the integral installation method of sensor is put forward to improve the feasibility of engineering application. Finally, by combining with the spot practical situation, the feasibility of the above research is verified. This study preliminarily realizes the engineering application of CNC machine tool state monitoring, and lays the foundation for the norm of following fault early warning of CNC machine tool.

This paper discusses the use of both extrinsic Fabry-Perot interferometry and digital image correlation method to measure strain at elevated temperatures. Strain measurements of TC4 Titanium Alloy under 500 °C have been performed based on extrinsic Fabry-Perot interferometry approach. The results show that the maximum absolute error was 0.052 × 10−3 compared with theoretical strain. When the load was higher than 0.2 × 10−3, the performance of EFPI sensor was better than that of the conventional resistance strain gauge, and the relative error was less than 14%. Digital image correlation method was also applied to a 310S stainless steel subjected to a uniaxial tensile test at 400 °C. Noticeable in-plane strain heterogeneities were highlighted throughout the test. These spatial heterogeneities were clearly visible from the beginning of the test and markedly intensified at the last stages before the macroscopic fracture. Comments on the investigated techniques and suggestions for future work are finally discussed.

By modeling an aero-refueling hose as a flexible pipe, an experimental research is performed on measuring the sectional bending stiffness of a flexible pipe withstanding a internal-pressurized fluid. A Revised Midpoint-Concentrate-Force-Loading Methodology (RMCFLM) with its valid range is developed to measure the sectional bending stiffness of a flexible pipe. The total system measurement error would be controlled within 7%. Three test specimens of a silica gel pipe, a polyvinyl chloride pipe and a rubber hose with fabric insert are conducted to measure their sectional-bending-stiffness in variable internal-pressurized fluid. The sectional-bending-stiffness increases with an incremental fluid pressure in a flexible pipe. Furthermore, the equivalent elastic modulus of a pipe is smaller and a incremental rate of the sectional-bending-stiffness is larger.

In this paper, bulk notched sample was designed to introduce crack and shear band interaction in bulk metallic glasses (BMGs). Deformation morphologies on the polished surface demonstrate that crack in BMGs might be deflected or arrested by surrounding shear bands. Distinct fracture morphologies could be observed in the interaction-induced soften region, indicating a transition of the mechanism dominating crack propagation. A hyperelastic model was used to discuss crack and shear band interaction. It’s proved that crack propagation is dominated by local elastic properties rather than global linear elastic properties due to shear induced softening and multiple shear bands. Our study suggests that multiple shear bands with a proper spacing are helpful to inhibit catastrophic crack propagation and to improve the plasticity of bulk metallic glasses.

For planning the flight load survey, several key points are stated in this paper by following sequence: 1. Stage selection for flight load survey as a constituent part of the whole process of flight tests- conducting as earlier as possible, as well as some relationship with other test subjects are accounted for; 2. Choosing proper test methods, such as measurements by strain gages or by pressure distribution, depending on characteristics of different structures, and then conducting the modifications; 3. Ground calibration tests to give the load equations revealing the relationship between strain gage bridges and net loads; 4. Choosing proper flight matrix and confirming the validity criteria for different kind of maneuver, with typical time history diagrams cited to show some flight maneuvers such as roller coaster, horizontal tail-elevator trade, yaw and rolling; 5. Arrangements of flight test details such as choosing proper time and airspace for flight, real-time monitoring of key parameters, communicating with pilots during the flight, and so on; 6. Analysis of test data, and validations of design methods to show compliance with relevant airworthiness regulations, or further correction of the methods.

The corrosion damage is easy to be formed on the surface of a metal structure in the corrosion environment, then the structural mechanical properties are degraded. For this purpose, the cellular automata method was adopted to simulate the uniform corrosion damage behavior of aluminium alloy in this paper. The uniform corrosion of aluminium alloy surfaces is simulated in complex corrosive environments and different concentration of corrosion solution. The forming processes of corrosion and the changed rules under different concentration of corrosion solution c and ambient temperature T are all obtained. It was shown that with the increasing concentration of corrosion c and the ambient temperature T, the number of corrosion cells increased, and the number of corrosion cells presents an approximate power function of etching time t.

The functions and requirements play an extremely important role in developing complex systems, providing a framework so that all engineers involved in a project have a common understanding of the scope and characterization of the system. Meanwhile, the functions and requirements are used to specify the system itself, as well as relevant auxiliary products and services which the stakeholders expect and desire. In this paper, a functional requirements development and management approach applied in a civil aircraft program is introduced. The operational scenario/use case definition, functional analysis, functional requirements and interfaces capture processes for a civil aircraft are presented in detail. In particular, the model-based functional analysis process and method are presented. The aircraft functional requirements hierarchy is studied and defined, taking account of the aircraft product breakdown, organization, and statement of work in order to define the requirements topology and traceability relationships. A typical aircraft functional requirements development case is taken to illustrate the functional requirements development method. Furthermore, a requirements management process is depicted throughout the aircraft program life cycle including requirements identification, validation, allocation, verification, and change control activities. A requirements management platform is developed according to the requirements management process. It shows that it is effective to perform functional requirements development based on aircraft operational scenarios (Use Cases) and aircraft program stakeholders’ needs, to improve the completeness of requirements and to ensure engineers’ common awareness of the aircraft to be developed. Furthermore, it’s necessary to perform a common requirements management process throughout the whole aircraft life cycle to ensure that the requirements can be properly validated, satisfied, implemented and verified so that stakeholders’ needs and expectations are met and that aircraft development, operation, and commercial success can be achieved.

The objective of this paper is to evaluate the effects of the stress from actual flight environment on the pilot workload, with the comparison analysis of the pilot workload under the simulated and actual flight environment. Twelve pilots participated in three simulation flight tasks with different workload levels and three pilots in the actual flight task. ECG indicator parameters were recorded with the biological feedback system in flight experiments. The workload was evaluated with four subjective evaluation rating scale after each flight task. There are strong correlations among workload from four subjective evaluation scales but strong correlation among few ECG parameters under three simulated flight tasks, strong correlations among workload value from subjective evaluation scales and ECG indicators in actual flight task. There are significant differences on pilot workload in the different phases of all the flight tasks from subjective evaluation scales, ECG indicators existed significant differences among all the flight phases just in actual flight task. Subjective evaluation methods can be used to measure the workload under simulated and actual flight tasks, but ECG only can be used well in the flight environment with higher stress. There are huge effects of stress from actual flight environment on pilot workload.

The three-dimensional measuring software is one of the key technologies of coordinate measurement machine (CMM). There are no enough technical support and competitive technical achievements in the development of CMM software in China, the reason is that there are many key technologies involved in developing CMM software. Based on the research platform provided by the AVIC Beijing Precision Engineering Institute Aircraft Industry (AVIC303), this paper has made some significant breakthroughs in the field of CMM, and then developed CMM software based on MVC software architecture. It can not only complete the whole measurement process, but also guarantee the measurement accuracy of the measured part. What’s more, the CMM software has laid solid foundation on the homemade of the CMM software and information security in the aerospace industry.