Chinese Journal of Aeronautics

Published by Elsevier
Print ISSN: 1000-9361
In this paper, taking the relativistic effect of high-velocity moving target into account, the Doppler shift, polarization deflection, reflection coefficient and phase delay of reflected electric field are analyzed rigorously under the assumptions that the incident signal to the target is a plane wave and the target is a perfect conductor plane. The analytic expressions are obtained. The present results are of practical significance to some extent for the accurate expression of the wideband returned signal of high-velocity moving target in bistatic radar system and understanding of wide-band ambiguity functions
In order to improve the performances of an 11 cm-diameter turbine engine, this article suggests to substitute a new-style micro diffuser redesigned based on a new concept for the traditional diffuser having poor performances. The new diffuser comprises integral blades and splitters, which are taken for a series of ducts in designing. This article investigates the effects of the cross-section area distribution along the flow path on the redesigned diffuser's performances. Having furnished with the new diffuser in place of the original vaned one, the 11 cm-diameter prototype engine is tested on the rig for its performances. CFD and experiments have shown that the improved diffuser with the unchanged original size has gained excellent performance parameters of pressure coefficient over 0.65 and total pressure recovery coefficient over 0.9. Equipped with the redesigned micro diffuser, the engine increases the thrust by 11% and decreases the specific fuel consumption by 9%.
Designation of RRA treatments and those with an incorporated plastic deformation
TEM micrographs of 2A97 alloy after double-aging with retrogression treatment at 220 ×15 ℃ min following initial aging at 165 ×30 h (a), after triple aging T11 (b) and (c), and that process with an incorporated plastic deformation TP1 ℃ (d) and (e).
DSC analyses of the solution-treated and waterquenched alloy and naturally aged for two weeks.  
A new high strength 2A97 Al-Cu-Li-X alloy was subjected to triple-aging of retrogression and re-aging treatments (RRA). Transmission electron microscopy (TEM), differential scanning calorimetry (DSC), and tensile tests were used to investigate the effects of RRA treatment on the microstructures and properties. DSC test reveals the reversion temperature range of the strengthening δ' (Al3Li) phase. The results show that the microstructure consists of δ' (Al3Li) phase, T1 (Al2CuLi) phase and θ″/θ'(Al2Cu) phase for 2A97 alloy treated by a triple-aging of a retrogression and re-aging treatment in the following order: (1) at 165 °C× 30 min, (2) at 220 °C or 240 ° × 15 min, (3) at 165 °C× 24 h. The plastic deformation, incorporated into the treatment after secondary high temperature aging, promotes the T1 precipitation during final re-aging. The tensile properties of the alloy treated by the retrogression and re-aging treatment reach the peak level of alloy single-aged at 165 °C in T6 temper.
Comparison of performance between four target extraction algorithms mm
Implicit image correction coefficients
Performance comparison of with and without image correction mm
Accuracy evaluation in real application
Accurate three-dimensional (3D) target positioning is of great importance in many industrial applications. Although various methods for reconstructing 3D information from a set of images have been available in the literature, few of them pay enough attention to the indispensable procedures, such as target extraction from images and image correction having strong influences upon the 3D positioning accuracy. This article puts forward a high-precision ellipse center (target point) extraction method and a new image correction approach which has been integrated into the 3D reconstruction pipeline with a concise implicit model to accurately compensates for the image distortion. The methods are applied to a copyright-reserved close range photogrammetric system. Real measuring experiments and industrial applications have evidenced the proposed methods, which can significantly improve the 3D positioning accuracy.
Space object recognition plays an important role in spatial exploitation and surveillance, followed by two main problems: lacking of data and drastic changes in viewpoints. In this article, firstly, we build a three-dimensional (3D) satellites dataset named BUAA Satellite Image Dataset (BUAA-SID 1.0) to supply data for 3D space object research. Then, based on the dataset, we propose to recognize full-viewpoint 3D space objects based on kernel locality preserving projections (KLPP). To obtain more accurate and separable description of the objects, firstly, we build feature vectors employing moment invariants, Fourier descriptors, region covariance and histogram of oriented gradients. Then, we map the features into kernel space followed by dimensionality reduction using KLPP to obtain the submanifold of the features. At last, k-nearest neighbor (kNN) is used to accomplish the classification. Experimental results show that the proposed approach is more appropriate for space object recognition mainly considering changes of viewpoints. Encouraging recognition rate could be obtained based on images in BUAA-SID 1.0, and the highest recognition result could achieve 95.87%.
Upper stage solid rocket motors (SRMS) for launch vehicles require a highly efficient propulsion system. Grain design proves to be vital in terms of minimizing inert mass by adopting a high volumetric efficiency with minimum possible sliver. In this article, a methodology has been presented for designing three-dimensional (3D) grain configuration of radial slot for upper stage solid rocket motors. The design process involves parametric modeling of the geometry in computer aided design (CAD) software through dynamic variables that define the complex configuration. Grain burn back is achieved by making new surfaces at each web increment and calculating geometrical properties at each step. Geometrical calculations are based on volume and change-in-volume calculations. Equilibrium pressure method is used to calculate the internal ballistics. Genetic algorithm (GA) has been used as the optimizer because of its robustness and efficient capacity to explore the design space for global optimum solution and eliminate the requirement of an initial guess. Average thrust maximization under design constraints is the objective function.
Nominal composition of Inconel 600 alloy 
n values and parabolic oxidation rate constants k P for cyclic oxidation of coated systems in CO 2 gas for 100 h at 10 h cycle 
IN 600 alloy was coated with two different types of coatings, Cr-modified aluminide coating this is called aluminizing-chromizing and Y-doped chromium modified aluminide coating this is called aluminizing-chromizing-yttriumizing. Diffusion coating was carried at 1 050 °C for 8 h under Ar atmosphere by simultaneous aluminizing-chromizing process and by simultaneous aluminizing-chromizing-yttriumizing. Cyclic oxidation tests were conducted on the uncoated and on the coated Inconel 600 alloy in the temperature range 800–1 000 °C in CO2 for 100 h at 10 h cycle. The results showed that the oxidation kinetics for uncoated Inconel 600 alloy in CO2 is parabolic and the phases present are NiO, (Fe, Cr)2O3, NiFe2O4 and NiCrO4. The oxidation kinetics for both coated systems in CO2 was found to be parabolic and the value of kP for both coated systems were found to be lower than that for uncoated Inconel 600 alloy. Oxide phases that formed on coated systems are Al2O3 and NiCrO4. The role of yttrium can be attributed to its ability to improve the adherence of the oxide scale.
Low bulk density expanding vermiculite is prepared, and the surface modification of hollow Al2O3-SiO2 microspheres and the composition of the low density ablative coating are studied. Organic silicon epoxy resin and phenolic aldehyde resin are applied as film forming matters to get a series of ablative coating having a density of 0.4-0.6 g/cm3. The performance of the low density ablative coating is evaluated by mechanical, thermodynamic and oxygen acetylene ablation tests, and the results are as follows, adhesion is in the range of 2.97-4.63 MPa, conductivity is no more than 0. 1 kcal/(m·h· °C), line ablation rate is no more than 0.30 mm/s, mass ablation rate is in the range of 0.11-0.18 mm/s.
A theory of composite material patch winding is proposed to determine the winding trajectory with a meshed data model. Two different conditions are considered in this study. One is Bridge condition on the concave surface and the other is Slip line condition in the process of patch winding. This paper presents the judgment principles and corresponding solutions by applying differential geometry theory and space geometry theory. To verify the feasibility of the patch winding method, the winding control code is programmed. Furthermore, the winding experiments on an airplane inlet and a vane are performed. From the experiments, it shows that the patch winding theory has the advantages of flexibility, easy design and application.
The technique of creep feed grinding is most suitable for geometrical shaping, and therefore has been expected to improve effectively material removal rate and surface quality of components with complex profile. This article studies experimentally the effects of process parameters (i.e. wheel speed, workpiece speed and depth of cut) on the grindability and surface integrity of cast nickel-based superalloys, i.e. K424, during creep feed grinding with brazed cubic boron nitride (CBN) abrasive wheels. Some important factors, such as grinding force and temperature, specific grinding energy, size stability, surface topography, microhardness and microstructure alteration of the sub-surface, residual stresses, are investigated in detail. The results show that during creep feed grinding with brazed CBN wheels, low grinding temperature at about 100 °C is obtained though the specific grinding energy of nickel-based superalloys is high up to 200-300 J/mm3. A combination of wheel speed 22.5 m/s, workpiece speed 0.1 m/min, depth of cut 0.2 mm accomplishes the straight grooves with the expected dimensional accuracy. Moreover, the compressive residual stresses are formed in the burn-free and crack-free ground surface.
The paper presents a finite volume numerical method universally applicable for solving both linear and nonlinear aeroacoustics problems on arbitrary unstructured meshes. It is based on the vertex centered multi-parameter scheme offering up to the 6th accuracy order achieved on the Cartesian meshes. An adaptive dissipation is added for the numerical treatment of possible discontinuities. The scheme properties are studied on a series of test cases, its efficiency is demonstrated at simulating the noise suppression in resonance-type liners.
Experimental study of synthetic jet produced by pulsed direct current (DC) discharge is presented. High velocity jet is activated electro-thermally by high frequency pulsed DC discharge in small cavity. A cavity of 2.38 mm diameter cylinder bounded by circular electrode is made in a ceramic plate and a small orifice of 1.78 mm diameter is drilled in the middle of cavity. High frequency pulsed DC discharge instantaneously heats air in the cavity and produces high velocity jet at the exit of the orifice. Schlieren imaging at high framing rate of 100 kHz reveals the presence of supersonic precursor shock followed by the jet emerging from the orifice. The jet velocity reaches as high as about 300 m/s. Jet with smaller cavity volume produces lesser effect and jet velocity reaches maximum at certain cavity volume with given discharge current and orifice size. As duty time of pulse increases from 5 to 20 μs at fixed frequency of 5 kHz, the jet velocity also increases and becomes nearly constant with further increase in duty time. At fixed duty time of 20 μs, higher frequency pulsing of 10 kHz produces degradation of the jet as the discharge pulse continues. The jet developed in this study is demonstrated to be strong enough to penetrate deep into supersonic boundary layer and to produce a bow shock when the jet is issued into Mach 3 supersonic flow.
Variation of emission intensity at 337.1 nm in x direction.
Vibrational temperature vs applied voltage at different driving frequencies.
Electron temperature and density vs applied voltage.
This article carries out synthetic measurements and analysis of the characteristics of the asymmetric surface dielectric barrier discharge plasma aerodynamic actuation. The rotational and vibrational temperatures of an N2(C3IIu) molecule are measured in terms of the optical emission spectra from the N2 second positive system. A simplified collision-radiation model for N2(C) and N2+ (B) is established on the basis of the ratio of emission intensity at 391.4 nm to that at 380.5 nm and the ratio of emission intensity at 371.1 nm to that at 380.5 nm for calculating temporal and spatial averaged electron temperatures and densities. Under one atmosphere pressure, the electron temperature and density are on the order of 1. 6 eV and 1011 cm−3 respectively. The body force induced by the plasma aerodynamic actuation is on the order of tens of mN while the induced flow velocity is around 1. 3 m/s. Starting vortex is firstly induced by the actuation; then it develops into a near-wall jet, about 70 mm downstream of the actuator. Unsteady plasma aerodynamic actuation might stimulate more vortexes in the flow field. The induced flow direction by nanosecond discharge plasma aerodynamic actuation is not parallel, but vertical to the dielectric layer surface.
A fault tolerant control (FTC) design technique against actuator stuck faults is investigated using integral-type sliding mode control (ISMC) with application to spacecraft attitude maneuvering control system. The principle of the proposed FTC scheme is to design an integral-type sliding mode attitude controller using on-line parameter adaptive updating law to compensate for the effects of stuck actuators. This adaptive law also provides both the estimates of the system parameters and external disturbances such that a prior knowledge of the spacecraft inertia or boundedness of disturbances is not required. Moreover, by including the integral feedback term, the designed controller can not only tolerate actuator stuck faults, but also compensate the disturbances with constant components. For the synthesis of controller, the fault time, patterns and values are unknown in advance, as motivated from a practical spacecraft control application. Complete stability and performance analysis are presented and illustrative simulation results of application to a spacecraft show that high precise attitude control with zero steady-error is successfully achieved using various scenarios of stuck failures in actuators.
Coefficients of turbulence models
The flows behind the base of a generic rocket, at both hypersonic and subsonic flow conditions, are numerically studied in this paper. The main concerns are addressed to the evaluation of turbulence models and the using of grid adaptation technique. The investigation focuses on two configurations, related to hypersonic and to subsonic experiments respectively. The applicability tests of different turbulence models are started on the level of two-equation models calculating the steady state solution of the Reynolds averaged Navier-Stokes equations. All used models - the original Wilcox k-ω, the Menter SST and the EARSM formulation - predict an asymmetric base flow in both cases caused by the support of the models. A comparison with preliminary experimental results indicates a preference for the SST and EARSM results over the results from the older k-ω model. Some further calculation of the flow fields around these simplified configurations are suggested to reach a better agreement with experiment results.
This article provides a flexible-joint-manipulator, which incorporates with three means to make its mechanical arm come into compliant contact with the objects with a force kept within an acceptable range. At first, the Cartesian impedance control law is introduced on the basis of virtual decomposition to realize the compliance control. Then, adaptive dynamic joint compensators on all joints are used to achieve more precise control. Finally, a Cartesian force-feedback path generation is developed for collision detection and force control. Experiments are performed on a 4-degree of freedom (DOF) satellite on-orbit self-servicing (SOOSS) manipulator. The results of the trajectory tracking and collision experiments demonstrate the effectiveness and feasibility of the proposed method.
An adaptive approximation-based optimization (AABO) procedure is developed for the optimum design of a composite advanced grid-stiffened (AGS) cylinder subject to post-buckling. The design taking account of post-buckling under ultimate load will be able to promote the structural efficiency compared to the conventional design in which only the linear buckling is allowed. The beam-shell offsets technique is utilized for modeling the stiffener-skin connection, and the Newton-Raphson method is employed for the post-buckling analysis. A few structural analysis efforts are carried out for establishing the Kriging model of the collapse load of the AGS cylinder for optimization to significantly increase the optimization efficiency. The multi-island genetic algorithm (MIGA) is utilized for global optimum search. An adaptive approximation framework is proposed to resolve the computational burden caused by the large domain of design variables, and it is demonstrated that much less computational expense than that of the traditional approximation-based optimization method can be achieved. The utility of making use of commercial optimization package iSIGHT in conjunction with the finite element (FE) code MSC.MARC to develop the preliminary design tool of the composite AGS cylinder is evaluated as well.
The methodology for adaptive control of helicopter ground resonance with magnetorheological (MR) damper is presented. The adaptive inverse control method is used to control the output damping force of MR damper and the range of the damping force is given. Through the adaptive inverse control, the damping force of MR damper is fit to a desired damping force. With the background of applying MR damper to control of helicopter ground resonance, a model of loss force and an adaptive arithmetic for stabilization of the coupled rotor/fuselage system are presented. The simulation shows that the controller presented in this paper can stabilize the rotor/fuselage coupling system quickly and control the helicopter ground resonance effectively.
Mounting capacity of F-22.
Weight data of mainstream combat aircraft
Typical engine parameters
The nature and characteristics of attack unmanned combat aerial vehicle (UCAV) are analyzed. The principles of selecting takeoff thrust-weight ratio and takeoff weight of attack UCAV are presented by analyzing the statistical data of weights for various main combat aircraft. The UCAV airborne weapons are analyzed, followed by the preliminary estimation of the payload weight. Various typical engines are analyzed and one of them is selected. Then the takeoff weight of the UCAV is determined. Based on some basic parameters and assumptions, the qualitative decomposition calculation for takeoff weight is completed. The key factors for obtaining longer endurance of aircraft with small aspect ratio configuration are found to be high lift-drag ratio and internal space. On the basis of the conclusions mentioned above, a highly blended flying-wing plus lifting body concept is proposed. According to this concept, the UCAV configuration is designed and optimized. Finally, the UCAV configuration with small aspect ratio, high lift-drag ratio, and high stealth characteristic is obtained.
This article is aimed to experimentally validate the beneficial effects of boundary layer suction on improving the aerodynamic performance of a compressor cascade with a large camber angle. The flow field of the cascade is measured and the ink-trace flow visualization is also presented. The experimental results show that the boundary layer suction reduces losses near the area of midspan in the cascade most effectively for all suction cases under test. Losses of the endwall could remarkably decrease only when the suction is at the position where the boundary layer has separated but still not departed far away from the blade surface. It is evidenced that the higher suction flow rate and the suction position closer to the trailing edge result in greater reduction in losses and the maximum reduction in the total pressure loss accounts to 16.5% for all cases. The suction position plays a greater role in affecting the total pressure loss than the suction flow rate does.
There is introduced a new low-reaction, highly-loaded axial compressor design concept which is coupled with boundary layer suction method. The characteristic features of the concept are made clear through its comparison with the MIT boundary layer suction compressor. Also are pointed out the potential applications of this concept as well as its key technological problems. Based on this concept, a single-stage, low-reaction and low-speed axial compressor is constructed in association with analysis and computation of boundary layer suction on vanes with the aid of a three-dimensional numerical approach. The results attest to the effectiveness of this way to control separation in blade cascades by the boundary layer suction and the feasibility of this proposed design concept.
Heights of separation lines 
Increments of energy loss coefficient 
This article experimentally studies the effects of air injection near the blade trailing edge on flow separation and losses in a highly loaded linear compressor cascade. Aerodynamic parameters of eight cascades with different air injection slot configurations are measured by using a five-hole probe at the cascade outlets. Ink-trace flow visualization is performed to obtain the flow details around the air injection slots. The static pressure distribution is clarified with pressure taps on the endwalls. The results indicate that air injection has little effect on the static pressure distribution on the endwalls, but improves the flow behavior at the corners between the suction surfaces and the endwalls with the decrease in losses at midspan. Slot positions have great effect on the compressor cascade performances. The optimal slot location is 25% of the blade span. The energy loss coefficient is reduced by 5.5% at most.
Robust optimization approach for aerodynamic design has been developed and applied to supercritical wing aerodynamic design. The aerodynamic robust optimization design system consists of genetic optimization algorithm, improved back propagation (BP) neural network and deformation grid technology. In this article, the BP neural network has been improved in two major aspects to enhance the training speed and precision. Uniformity sampling is adopted to generate samples which will be used to establish surrogate model. The testing results show that the prediction precision of the improved BP neural network is reliable. On the assumption that the law of Mach number obeys normal distribution, supercritical wing configuration considering fuselage interfering of a certain aerobus has been taken as a typical example, and five design sections and twist angles have been optimized. The results show that the optimized wing, which considers robust design, has better aerodynamic characteristics. What's more, the intensity of shock wave has been reduced.
An optimization strategy is proposed to deal with the aerodynamic/stealthy/structural multidisciplinary design optimization (MDO) issue of unmanned combat air vehicle (UCAV). In applying the strategy, the MDO process is divided into two levels, i.e. system level optimization and subsystem level optimization. The system level optimization is to achieve optimized system objective (or multi-objective) through the adjustment of global external configuration design variables. The subsystem level optimization consists of the aerodynamic/stealthy integrated design and the structural optimization. The aerodynamic/stealthy integrated design aims at achieving the minimum aerodynamic drag coefficient under the constraint of stealthy requirement through the adjustment of local external configuration design variables. The structural optimization is to minimize the structural weight by adjusting the dimensions of structural components. A flowchart to implement this strategy is presented. The MDO for a flying-wing configuration of UCAV is employed to illustrate the detailed process of the optimization. The results indicate that the overall process of the surrogate-based two-level optimization strategy can be implemented automatically, and quite reasonable results are obtained.
Some years ago the national CFD project MEGAFLOW was initiated in Germany to combine many of the CFD development activities from DLR, universities and aircraft industry. Its goal was the development and validation of a dependable and efficient numerical tool for the aerodynamic simulation of complete aircraft which met the requirements of industrial implementations. The MEGAFLOW software system includes the block-structured Navier-Stokes code FLOWer and the unstructured Navier-Stokes code TAU. Both codes have reached a high level of maturity and they are intensively used by DLR and the German aerospace industry in the design process of new aircraft. Recently, the follow-on project MEGADESIGN and MEGAOPT were set up which focus on the development and enhancement of efficient numerical methods for shape design and optimization. This article highlights recent improvements of the software and its capability to predict viscous flows for complex industrial aircraft applications.
Sketch of a large aspect ratio wing with two spars. First, the wing with half amount of fuel is defined as the nominal model, in which the lumped masses of fuel tanks numbered 1-8 are 2, 3, 2, 3, 2, 3, 2, 3 kg, respectively. Considering the first six normal modes, the nominal p-k flutter analysis of the wing with half amount of fuel gives a flutter speed of V f = 93.0 m/s at a frequency of 11.2 Hz. Then, the three cases of perturbations of struc
Robust analysis-curves for Case I.
Robust analysis-curves for Case II.
Robust analysis-curves for Case III.
Air vehicles undergo variations in structural mass and stiffness because of fuel consumption and the failure of structural components, which might lead to serious influences on the aeroelastic characteristics. An approach for aeroelastic robust stability analysis taking into account the perturbations of structural mass and stiffness is developed. Applying the perturbation method and harmonic unsteady aerodynamic forces, the frequency-domain linear fractal transformation (LFT) representation of perturbed aeroelastic system is modeled. Then, the robust stability is analyzed by using the structured singular value μ-method. The numerical results of a bi-spar wing show its effectiveness and low computational time in dealing with the robust problems with mass and stiffness perturbations. In engineering analysis for solving aeroelastic problems, the robust approach can be applied to flutter analysis for airplane with the fuel load variation and taking the damage conditions into consideration.
High-aspect-ratio flexible wing data
Effects of drag on LCO at flight speed of 60.0 m/s without gravity loads considered.
Effects of drag on LCO at flight speed of 61.0 m/s without gravity loads considered.
The aeroelastic analysis of high-altitude, long-endurance (HALE) aircraft that features high-aspect-ratio flexible wings needs take into account structural geometrical nonlinearities and dynamic stall. For a generic nonlinear aeroelastic system, besides the stability boundary, the characteristics of the limit-cycle oscillation (LCO) should also be accurately predicted. In order to conduct nonlinear aeroelastic analysis of high-aspect-ratio flexible wings, a first-order, state-space model is developed by combining a geometrically exact, nonlinear anisotropic beam model with nonlinear ONERA (Edlin) dynamic stall model. The present investigations focus on the initiation and sustaining mechanism of the LCO and the effects of flight speed and drag on aeroelastic behaviors. Numerical results indicate that structural geometrical nonlinearities could lead to the LCO without stall occurring. As flight speed increases, dynamic stall becomes dominant and the LCO increasingly complicated. Drag could be negligible for LCO type, but should be considered to exactly predict the onset speed of flutter or LCO of high-aspect-ratio flexible wings.
In this paper, the operating conditions, technical requirements, performance characteristics, design ideas, application experiences and development trends of aerospace engine bearings, including material technology, integration design and reliability, are reviewed. The development history of aerospace engine bearing is recalled briefly at first. Then today's material technologies and the high bearing performances of the bearings obtained through the new materials are introduced, which play important rolls in the aeroengine bearing developments. The integration design ideas and practices are explained to indicate its significant advantages and importance to the aerospace engine bearings. And the reliability of the shaft-bearing system is pointed out and treated as the key requirement with goals for both engine and bearing. Finally, as it is believed that the correct design comes from practice, the pre-qualification rig testing conducted by FAG Aerospace GmbH & Co. KG is briefly illustrated as an example. All these lead to the development trends of aerospace engine bearings from different aspects.
Thanks to recent advances in manufacturing technology, aerospace system designers have many more options to fabricate high-quality, low-weight, high-capacity, cost-effective filters. Aside from traditional methods such as stamping, drilling and milling, many new approaches have been widely used in filter-manufacturing practices on account of their increased processing abilities. However, the restrictions on costs, the need for studying under stricter conditions such as in aggressive fluids, the complicity in design, the workability of materials, and others have made it difficult to choose a satisfactory method from the newly developed processes, such as, photochemical machining (PCM), photo electroforming (PEF) and laser beam machining (LBM) to produce small, inexpensive, light-weight aerospace filters. This article appraises the technical and economical viability of PCM, PEF, and LBM to help engineers choose the fittest approach to turn out aerospace filters.
Baseline parameters of 2-DOF plunging-pitching airfoil 
Reduction in torsional stiffness for solid doublewedge wing due to aerodynamic heating. Effects of thickness ratio and altitude.
Pitch LCO amplitude versus flight Mach number for a 2-DOF system with all nonlinearities. Time histories and phase portraits represent the uncontrolled and controlled system, respectively.  
Effects of nonlinear active control on system encompassing all nonlinearities.  
Designing re-entry space vehicles and high-speed aircraft requires special attention to the nonlinear thermoelastic and aerodynamic instability of their structural components. The thermal effects are important since temperature environment brings dramatic influences on the static and dynamic behaviors of flight structures in supersonic/hypersonic regimes and is likely to cause instability, catastrophic failure and oscillations resulting in structural failure due to fatigue. In order to understand the dynamic behaviors of these “hot” structures, a double-wedge lifting surface with combining freeplay and cubic structural nonlinearities in both plunging and pitching degrees-of-freedom operating in supersonic/hypersonic flight speed regimes has been analyzed. A third order piston theory aerodynamic is used to estimate the applied nonlinear unsteady aerodynamic loads. Also considered is the loss of torsional stiffness that may be incurred by lifting surfaces subject to axial stresses induced by aerodynamic heating. The aerodynamic heating effects are estimated based on the adiabatic wall temperature due to high speed airstreams. As a recently emerging technology, the active aerothermoelastic control is aimed at providing solutions to a large number of problems involving the aeronautical/aerospace flight vehicle structures. To prevent such damaging phenomena from occurring, an application of linear and nonlinear active control methods on both flutter boundary and post-flutter behavior has been fulfilled. In this paper, modeling issues as well as numerical simulation have been presented and pertinent conclusions outlined. It is evidenced that a serious loss of torsional stiffness may induce the dynamic instability; however active control can be used to expand the flutter boundary and convert unstable limit cycle oscillations (LCO) into the stable LCO and/or to shift the transition between these two states toward higher flight Mach numbers.
With the aid of multi-agent based modeling approach to complex systems, the hierarchy simulation models of carrier-based aircraft catapult launch are developed. Ocean, carrier, aircraft, and atmosphere are treated as aggregation agents, the detailed components like catapult, landing gears, and disturbances are considered as meta-agents, which belong to their aggregation agent. Thus, the model with two layers is formed i.e. the aggregation agent layer and the meta-agent layer. The information communication among all agents is described. The meta-agents within one aggregation agent communicate with each other directly by information sharing, but the meta-agents, which belong to different aggregation agents exchange their information through the aggregation layer first, and then perceive it from the sharing environment, that is the aggregation agent. Thus, not only the hierarchy model is built, but also the environment perceived by each agent is specified. Meanwhile, the problem of balancing the independency of agent and the resource consumption brought by real-time communication within multi-agent system (MAS) is resolved. Each agent involved in carrier-based aircraft catapult launch is depicted, with considering the interaction within disturbed atmospheric environment and multiple motion bodies including carrier, aircraft, and landing gears. The models of reactive agents among them are derived based on tensors, and the perceived messages and inner frameworks of each agent are characterized. Finally, some results of a simulation instance are given. The simulation and modeling of dynamic system based on multi-agent system is of benefit to express physical concepts and logical hierarchy clearly and precisely. The system model can easily draw in kinds of other agents to achieve a precise simulation of more complex system. This modeling technique makes the complex integral dynamic equations of multibodies decompose into parallel operations of single agent, and it is convenient to expand, maintain, and reuse the program codes.
In order to improve the drift precision of air supported gyroscope, effects of surface roughness magnitude and direction on vortex torque of air supported gyroscope are studied. Based on Christensen's rough surface stochastic model and consistency transformation method, Reynolds equation of air supported gyroscope containing surface roughness information is established. Also effects of mathematical models of main machining errors on vortex torque are established. By using finite element method, the Reynolds equation is solved numerically and the vortex torque in the presence of machining errors and surface roughness is calculated. The results show that surface roughness of slit has a significant effect on vortex torque. Transverse surface roughness makes vortex torque greater, while longitudinal surface roughness makes vortex torque smaller. The maximal difference approaches 11.4% during the range analyzed in this article. However surface roughness of journal influences vortex torque insignificantly. The research is of great significance for designing and manufacturing air supported gyroscope and predicting its performance.
Modes and frequencies of free vibration 
Trim variables in level flight condition 
shows the frequency response of pit ch rate to the deflection of the elevat or in the typical condit ion. F rom the figure it is show n that t he re sponse curves of t he rigid airplane and t he flex ible airplane are basely match together at low frequen cies, but have large diff erence at elast ic vibrat ion frequencies. T he curves of t he flexible airplane show a peak in gain and 180% decrease in phase.  
Based on the equations of motion of flexible air vehicles including rigid-body modes and elastic structural modes, and applying influence coefficients of linear aerodynamics, a set of equations are derived and a method is presented for analysis of flight loads and dynamic characteristics. The problems in the fields of flight mechanics and aeroelasticity such as static aeroelastic divergence, trim and deformation, aerodynamic loads distribution, flutter and flight dynamics can be solved by the procedure. An airplane with high aspect ratio wings is analyzed, and the results show that the coupling between rigid-body modes and elastic modes is distinct and should not be overlooked.
This article presents a parameterized configuration modeling approach to develop a 6 degrees of freedom (DOF) rigid-body model for air-breathing hypersonic vehicle (AHV). The modeling process involves the parameterized configuration design, inviscous hypersonic aerodynamic force calculation and scramjet engine modeling. The parameters are designed for airframe-propulsion integration configuration, the aerodynamic force calculation is based on engineering experimental methods, and the engine model is acquired from gas dynamics and quasi-one dimensional combustor calculations. Multivariate fitting is used to obtain analytical equations for aerodynamic force and thrust. Furthermore, the fitting accuracy is evaluated by relative error (RE). Trim results show that the model can be applied to the investigation of control method for AHV during the cruise phase. The modeling process integrates several disciplines such as configuration design, aerodynamic calculation, scramjet modeling and control method. Therefore the modeling method makes it possible to conduct AHV aerodynamics/propulsion/control integration design.
Altitude characteristic is of great importance for studying when an air-breathing pulsed laser thruster works in the dense atmosphere condition of 0-30 km altitude. The experimental findings all over the world show that the similar relationship between impulse coupling coefficient and altitude. According to strong explosion theory and an ideal gas model, a dimensionless factor indicating energy law of similitude is introduced, and formula of impulse coupling coefficient is deducted. Then theoretical study of altitude characteristic is carried out and mechanism of altitude characteristic is further explained. The results indicate that there is a maximum value of impulse coupling coefficient if the dimensionless factor equals to 0. 41 in theory, and whether the phenomena of maximum appear or not depends on the range of the dimensionless factor related to altitude. As to a conical nozzle with the fixed length of 120 mm, the relationship between the sonic velocity and the dimensionless factor causes the maximum phenomenon at the altitude of about 12. 5 km, and maximum theoretical impulse coupling coefficient is also found in the experimental investigations. The mechanism of altitude characteristic for air-breathing pulsed laser thruster is discovered in this article, which will provide reference for further research on altitude characteristic.
AbstractWhen the stagnation temperature of a perfect gas increases, the specific heat ratio does not remain constant any more, and start to vary with this temperature. The gas remains perfect, its state equation remains always valid, except it will name in more calorically imperfect gas or gas at High Temperature. The goal of this work is to trace the profiles of the supersonic Minimum Length Nozzle with centered expansion when the stagnation temperature is taken into account, lower than the threshold of dissociation of the molecules and to have for each exit Mach number several nozzles shapes by changing the value of the temperature. The method of characteristics is used with a new form of the Prandtl Meyer function at high temperature. The resolution of the obtained equations is done by the second order of finite differences method by using the predictor corrector algorithm. A study on the error given by the perfect gas model compared to our model is presented. The comparison is made with a calorically perfect gas for goal to give a limit of application of this model. The application is for the air.
As an important parameter in the single airborne passive locating system, the rate of phase difference change contains range information of the radio emitter. Taking single carrier sine pulse signals as an example, this article illustrates the principle of passive location through measurement of rates of phase difference change and analyzes the structure of measurement errors. On the basis of the Cramér-Rao lower bound (CRLB), an algorithm associated with time-chips is proposed to determine the rates of phase difference change. In the measurement of the rates of phase difference change, phase discrimination in the frequency domain outperforms that in the time domain when signal noise rate (SNR) is lower. Multi-chip processing can significantly reduce variance of the measurement of rates of phase difference change. Simulations demonstrate the validity and accuracy of the proposed algorithm. The simulations carried out on the typical single airborne passive location have proved its adaptability to dynamic measurements. The proposed algorithm to determine the rates of phase difference change proves simple and easy to implement with less computation workload.
Bird strike sites displacement. 
Central point displacement at impacted point B. 
Camber line displacement at 3 ms. 
Damage-modified nonlinear viscoelastic constitutive equation and failure criterion are introduced and the three-dimensional incremental forms are deduced based on the updated Lagrangian approach. A simple tensile test model and a split Hopkinson pressure bar model are built to verify the accuracy of the subroutine implemented within the non-linear finite element program LS-DYNA. A numerical model of bird strike on windshield is established to study the responses of windshield under three different bird velocities at three sites. The bird is represented by a cylinder with a hemisphere at each end and the contact-impact coupling algorithm is used in this study. It is found that the implemented subroutine can properly describe the mechanical behavior of polymethyl methacrylate under low and high strain rates and large deformation, and can be used validly.
Modeling with vortex ring.
Induced velocity at central axis.
Velocity in vortex core.
Four assumed points.
This article deals with real-time hi-fi simulation of large aircraft flying in turbulent wind in a simulator to study its takeoff and landing behavior in microburst wind shear. A parameterized three-dimensional (3D) microburst model is built up on the basis of vortex ring and Rankine vortex principle. Complicated microburst wind fields are simulated by means of vortex ring declination and multi-vortex superposition. Based on the modeling data of Boeing 747-100, a dynamic model with wind shear effects considered is established and a general method to modify the aerodynamic model is proposed. A controller for longitudinal and lateral escapes is designed and verified in simulated microburst wind field. Results indicate that, with high extensibility, reasonability and effectiveness, the 3D microburst model with wind shear effects considered is fit to simulate real wind fields. Different escape schemes can be adopted to fly through a wind field from different locations. The model can be used for real-time flight simulation in a flight simulator.
A tilt rotor is an aircraft of a special kind, which possesses the characteristics of a helicopter and a fixed-wing airplane. However, there are a great number of important technical problems waiting for settlements. Of them, the flight control system might be a critical one. This article presents the progresses of the research work on the design of flight control system at Nanjing University of Aeronautics and Astronautics (NUAA). The flight control law of the tilt rotor aircraft is designed with the help of an inner/outer loop control structure and an eigenstructure assignment algorithm on the basis of a proper mathematical model already verified by the wind tunnel tests. The proposed control law has been born out through the construction of the flight control system and the flight tests. Now, the flight tests are still underway on a prototype of small unmanned tilt rotor aircraft. The results have evidenced the credibility of the aircraft design and the effectiveness of the flight control system for the tilt rotor working in the helicopter mode. A full envelope flight test is planned to carry out further researches on the flight control law.
RCS curves corresponding to two types of stealth aircraft (under the S wave band).
Relevant radar detecting data of series models (δ ave = −10 dBsm, K A =7/72)
Taking into account the limitations of existing stealth performance analysis methods, a method termed as the integrated stealth performance analysis method is proposed for evaluating the stealth ability of the penetration aircraft. Based on various target radar cross section (RCS) scattering characters, this article integrates the relevant parameters needed for building up target circumferential RCS scattering model and proposes the RCS scattering controlling parameters to control the changing trends of the relevant model RCS scattering characters. According to the radar dynamic detecting characters during the whole penetration course, a dynamic stealth performance evaluating model is proposed accompanied by a series of stealth ability estimation rules. This new analysis method can enhance the integrality and dependability of the stealth analysis conclusions and summarize the relationship between the target RCS scattering characters and their effects on stealth performance. The rules indicated by this relationship can be used as the reference for designing new type of stealth aircraft and setting up specific penetration tactics.
The previous study on modeling of the tilt rotor aircraft used to put a premium on the complicated aerodynamic computation, and the research on the motion equations is often constrained to frequently use the oversimplified 6-degree of freedom (DOF) rigid body equations. However, the transfiguration of aircraft during transition stage, is complicated due to the aerodynamic interference and the change of center of gravity (CG). Moreover, the gyroscopic moment caused by tilting the high-speed revolving rotors seriously interferes with the aircraft attitude. The above-cited 6-DOF single rigid body equations do not take the inertia coupling effects into account during transition. For this sake, the article, reckoning the body, the nacelles and the rotors to be independent entities, establishes a realistic model in the form of multi-body motion equations. First, by applying Newton's laws and angular momentum theorem to a mass of elements of the aircraft, the multi-body motion equations in inertial frame as well as in body frame are obtained by integrating over all elements. As the equations are of implicit nonlinear differential type, the consistent initial value problem should be solved. Then, a numerical simulation of the differential equations is conducted by means of the Runge-Kutta-Felhberg integral algorithm. The modeling and the simulation algorithm are verified against the data of XV-15 as an example. The model can be used in the area of flight dynamics, flight control and flight safety of tilt rotor aircraft.
In order to find out the optimal press bend forming path in fabricating aircraft integral panels, this article proposes a new method on the basis of the authors' previous work. It is composed of the finite element method (FEM) equivalent model, the surface curvature analysis, the artificial neural network response surface and the genetic algorithm. The method begins with analyzing the objective's shape curvature to determine the bending position. Then it optimizes the punch travel at each bending position by the following steps: (1) Establish a multi-step press bend forming FEM equivalent model, with which the FEM experiments designed with the Taguchi method are performed. (2) Construct a back-propagation (BP) neural network response surface with the data from the FEM experiments. (3) Use the genetic algorithm to optimize the neural network response surface as the objective function. Finally, this method is verified by press bending a complicated double-curvature grid-type stiffened panel and bears out its effectiveness and intrinsic worth in designing the press bend forming path.
R eference fr ames of the carrierair craft multibody system wit h flexible links 1. 1 Inertial frame Ixi y iz i
Simplification of t he aircr aftlandingg ear system and the outer for ce analysis
T he time history of the carrier pitch attitude
T ime histo ry of t he elevator ang le ( command si gnal and actual deflection ang le)
T ime histor y of the AOA
A general mathematical model of carrier based aircraft ski jump take-off is derived based on tensor. The carrier, the aircraft body and the movable parts of the landing gears are treated as independent entities. These entities are assembled into a multi-rigid-body system with flexible links. Dynamical equations of each entity are derived on the basis of the Newton law and the Euler transformation. Using the invariance property of the tensor, the dynamical and kinematical equations are converted to tensor forms which are invariant under time dependent coordinate transformations. Then the tensor formed equations are expressed by the matrix operation. Differential equation group of the matrix form is formur lated for the programming. The closure of the model is discussed, and the simulation results are given.
Process of "searching from different directions".
Structure of PMOTS algorithm.
Running results of HPMOTS algorithm.
For dealing with the multi-objective optimization problems of parametric design for aircraft, a novel hybrid parallel multi-objective tabu search (HPMOTS) algorithm is used. First, a new multi-objective tabu search (MOTS) algorithm is proposed. Comparing with the traditional MOTS algorithm, this proposed algorithm adds some new methods such as the combination of MOTS algorithm and “Pareto solution”, the strategy of “searching from many directions” and the reservation of good solutions. Second, this article also proposes the improved parallel multi-objective tabu search (PMOTS) algorithm. Finally, a new hybrid algorithm—HPMOTS algorithm which combines the PMOTS algorithm with the non-dominated sorting-based multi-objective genetic algorithm (NSGA) is presented. The computing results of these algorithms are compared with each other and it is shown that the optimal result can be obtained by the HPMOTS algorithm and the computing result of the PMOTS algorithm is better than that of MOTS algorithm.
The coupled numerical simulation of flow field, solid temperature field, species concentration field and gas radiation transfer/ energy field based on statistical narrow-band correlated-k (SNBCK) model, is employed to accurately predict aerothermodynamic characteristic of aircraft exhaust system. A series of methods to increase computational efficiency and descend computational resources make it possible to finish the calculation in PC. The parameters of narrow-band model are evaluated by HITEMP line-by-line database. Three examples have proved the accuracy of using these methods to solve flow heat transfer coupled problem and radiation transfer/energy equation, which are the calculation of temperature distribution of water-cooling nozzle in rocket engine, the calculation of carbon dioxide absorptivity at 4.3 micron band, and the gas radiation heat transfer evaluation of the cylindrical furnace. Finally, the inner flaps temperature distribution of ejecting nozzle with floating outer flaps is computed, under high-altitude, high-speed and afterburning conditions. Two completely different air-inlet schemes of ejecting channel almost achieve the same effect in cooling inner flaps.
Model-based control and diagnostics logic.
Sensor fault detection isolation using bank of Kalman filters.
Simulation of sensor FDI using bank of Kalman filters.
A duty in development of an on-line fault detection algorithm is to make it associate with estimation of engine's health degradation. For this purpose, an on-line diagnostic algorithm is put forward. Using a tracking filter to estimate the engine's health condition over its lifetime, can be reconstructed an onboard model, which is then made to match a real aircraft gas turbine engine. Finally, a bank of Kalman filters is applied in fault detection and isolation (FDI) of sensors for the engine. Through the bank, the real faults that have occurred can be detected and isolated. The on-line fault detection algorithm has the ability of maintaining the effectiveness over the engine's lifetime and is verified by simulation using a nonlinear engine model.
Parameters of HAVE PIO configurations
Requirements of bandwidth in each flight phase
Results of predicting aircraft FQ
Boundary of MAI criterion and improved criterion.
During the process of aircraft design, the mathematical model of pilot control behavior characteristics is always used to predict aircraft flying qualities (FQ). This is one of the important methods to avoid pilot-aircraft adverse coupling. In order to study the FQ criterion based on closed-loop pilot-aircraft systems, first, an experimental database is built, which includes 40 aircraft dynamics configurations and the corresponding flight simulation results. Second, the mathematical pilot models with a set of different aircraft configurations are obtained by this experimental database. Then, two FQ criteria, Neal-Smith criterion and Moscow Aviation Institute (MAI) criterion, are analyzed. And the relationship between the FQ level evaluated by actual pilot and the parameters of closed-loop pilot-aircraft systems is studied. Finally, an improved criterion of aircraft FQ is built based on the above two criteria. This new criterion is further used to predict FQ for four new aircraft dynamics configurations, and the prediction results verify its accuracy and practicability.
Recently, frequency-based least-squares (LS) estimators have found wide application in identifying aircraft flutter parameters. However, the frequency methods are often known to suffer from numerical difficulties when identifying a continuous-time model, especially, of broader frequency or higher order. In this article, a numerically robust LS estimator based on vector orthogonal polynomial is proposed to solve the numerical problem of multivariable systems and applied to the flutter testing. The key idea of this method is to represent the frequency response function (FRF) matrix by a right matrix fraction description (RMFD) model, and expand the numerator and denominator polynomial matrices on a vector orthogonal basis. As a result, a perfect numerical condition (numerical condition equals 1) can be obtained for linear LS estimator. Finally, this method is verified by flutter test of a wing model in a wind tunnel and real flight flutter test of an aircraft. The results are compared to those with notably LMS PolyMAX, which is not troubled by the numerical problem as it is established in z domain (e.g. derived from a discrete-time model). The verification has evidenced that this method, apart from overcoming the numerical problem, yields the results comparable to those acquired with LMS PolyMAX, or even considerably better at some frequency bands.
tSN cur ves and SN curve slope m ( t ) degenerat es into a constant mt , and that t he log arithmic int ercept is linearly related t o precorrosion t im e, i e. , lg C ( t ) = a+ bt . T he former characterist ic can be called t he constant slope, and the lat ter is named the loglinear inter cept here. T hat m t = 3202 98 and C ( 0) = 10 11644 92 are respectively approximate to m 0 = 3325 8 and C 0 = 10 11 852 5 of t he SN curve obt ained in general environment [ 5] indicates t he f easibilit y of t he new equation, w hich is show n in F ig1. ( 2) T he geometrical meaning of the constant slope is clearly illust rat ed by the parallel straig ht lines in F ig. 1. In addit ion, if a preliminary life predict ion of an aircraft operat ing in w eak corrosive environment is demanded, mt may t ake m 0 value. Finally, this charact eristic facilit at es parameter es t imat e and helps reduce bot h corrosion and fatigue test s.  
To quantitatively evaluate the effects of corrosion during grounding on fatigue life of aircraft structures, a new power equation is proposed using two variable linear regression method. That the slope is a constant and the logarithmic intercept is a linear function of pre corrosion time makes this equation advantageous: it has a simple form, its parameters have unambiguous technical and geometrical meanings, and it facilitates engineering applications. Three parameter equations after pre corrosion are obtained from back-calculation of fatigue limits, which have been successfully used to predict safe life of aircraft structures in corrosive environment.
Top-cited authors
Shao-Ping Wang
Wenfeng Ding
  • Nanjing University of Aeronautics & Astronautics
Chang He Li
  • Qingdao University of Technology
Y B Zhang
  • Qingdao University of Technology
Jiuhua Xu
  • Nanjing University of Aeronautics & Astronautics