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The airfoil on modern fighters needs large positive moment to maintain longitudinal trim in addition to good aerodynamic and stealth performance. The novel design requirements would increase the likelihood of multimodality in design space. To address this issue, this paper uses parallel niching method to identify the existence and possible sources of multimodality in the aerodynamic and stealth shape optimizations based on a symmetrical airfoil. To isolate the instances of space multimodality, several subproblems are studied by varying the number and range of design variables, as well as positive moment constraints, respectively. The results demonstrate that the occurrences of multimodality in aerodynamic design space are highly related to positive moment constraint and ranges of design variables. Additionally, it’s found that some local minimums in aerodynamic space are very close to that of stealth space, indicating that the there’s not that large conflict between aerodynamic and stealth design under certain circumstances. This discovery may appeal a new approach of aerodynamic/stealth design of aircraft.

Taking the hard air refueling device as an object, the momentum distance method is used to model it, and the dynamic system model which can reflect its motion characteristics is obtained. The PID control law design and Simulation of the model are carried out. The simulation results show that the system has strong resistance to the vertical and horizontal coupling. The ground demonstration of automatic docking is completed on the ground test-bed of the boom aerial refueling device, which further verifies the feasibility of the control design and meets the needs of the engineering design.

The ever-increasing demands for risk-free, resource-efficient and environment-friendly air vehicles motivate the development of advanced design methodology. As a particularly promising design methodology considering uncertainties, robust aerodynamic design optimization (RADO) is capable of providing robust and reliable aerodynamic configuration and reducing cost under probable uncertainties in the flight envelop and all life cycle of air vehicle. However, the major challenges including high computational cost with increasing dimensionality of uncertainty and complex RADO procedure hinder the wider application of RADO. In this paper, the complete RADO procedure, i.e., uncertainty modeling, establishment of uncertainty quantification approach as well as robust optimization subject to reliability constraints under uncertainty, is elaborated. Systematic reviews of RADO methodology including uncertainty modeling methods, comprehensive uncertainty quantification approaches, and robust optimization methods are provided. Further, this paper presents a brief survey of the main applications of RADO in the aerodynamic design of transonic flow and natural-laminar-flow, and discusses the application prospects of RADO methodology for air vehicles. The detailed statement of the paper indicates the intention, i.e., to present the state of the art in RADO methodology, to highlight the key techniques and primary challenges in RADO, and to provide the beneficial directions for future researches.

The work develops the multi-dimensional dissipation strategy within advection upstream splitting methods (AUSM) in hypersonic flows based on the affordable shock-stable item by Chen et al. (J. Comput. Phys. 373 (2018) 662-672). This strategy is related to the sear velocity difference and the pressure-based sensing function. It strengthens the robustness of AUSM-family schemes against shock anomalies and meantime preserves shear layer. A series of numerical cases illuminate its potential capability for hypersonic flows, particularly in the extreme situations such as the large aspect ratio of cells, skewed non-orthogonal grid, and unstructured grid.

The geometric shape is not directly manipulated by free-form deformation (FFD) technique. A high level FFD control volume is always adopted in the refined aerodynamic design. The direct manipulated free-form deformation (DFFD) technique is developed in this paper to improve the shortcomings of the FFD technique. This method adopt a parameter identification technique which can identify the disturbance quantities and convert them into displacement values on the FFD control lattices. The DFFD technique can accomplish refined aerodynamic shape design using less design variables. The visual impression is much better than the traditional FFD technology, and the number of the design variables is not increased with using of the high level FFD control volume. The robust design system is established by DFFD technique, multi-groups collaboration particle swarm optimization (MCPSO) and adaptive sampling algorithm, combined with quaternion mesh deforming technique and the numerical simulation of boundary layer transition based on γ - Re θt transition model. A transonic laminar wing body is optimized by this system. The design results show that not only a wide range of laminar flow area is maintained, but also the development of shock is controlled.

The aerodynamic characteristics of elliptic airfoil are quite different from the case of conventional airfoil for Reynolds number varying from about 10 4 to 10 6 . In order to reveal the fundamental mechanism, the unsteady flow around a stationary two-dimensional elliptic airfoil with 16% relative thickness has been simulated using unsteady Reynolds-averaged Navier–Stokes equations and the γ - Re θ t ‾ transition turbulence model at different angles of attack for flow Reynolds number of 5 × 10 5 . The aerodynamic coefficients and the pressure distribution obtained by computation are in good agreement with experimental data, which indicates that the numerical method works well. Through this study, the mechanism of the unconventional aerodynamic characteristics of airfoil is analyzed and discussed based on the computational predictions coupled with the wind tunnel results. It is considered that the boundary layer transition at the leading edge and the unsteady flow separation vortices at the trailing edge are the causes of the case. Furthermore, a valuable insight into the physics of how the flow behavior affects the elliptic airfoil’s aerodynamics is provided.

Different from the traditional helicopter, there would be a severe interaction between rotor and fuselage with canard and horizontal tail during hovering for canard rotor wing (CRW) aircraft. The un-steady RANS (URANS) equations discretized on overset grid are solved in this paper to simulate this kind of aerodynamic interaction. A UH60-A rotor hovering case is used to validate the numerical method. Subsequently, the numerical method is employed on two configurations, i.e., Isolated rotor configuration and Full CRW configuration under numerous collective pitch angle. The force and moment coefficients for different parts with respect to different collective pitch angles are obtained. The result shows that downwash of rotor has a stronger impact on fuselage than canard wing and horizontal tail. Compared with isolated rotor, fuselage with canard wing and horizontal tail has a slight impact on the trust and torque of the rotor in full CRW aircraft configuration.

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.

In the work, it is shown the numerical investigations about the unsteady inviscid results obtained for the pitching oscillating wings at different angles of attack. The results are obtained by solving the unsteady Euler equations in a bodyfitted coordinate system. It is based on the four-stage runge-Kutta time stepping scheme. Meanwhile to increase the time step that is limited by Courant limit (CFL), the implicit residual smoothing with local variable parameters is used. As a result, the unsteady aerodynamics about a rectangular wing and a delta wing, which are oscillated in pitching with different frequencies, are shown in this paper. The properties of the unsteady aerodynamics in these cases are researched here.

The work presented here shows the unsteady inviscid results obtained for the two-and three-dimensional wings which are in rigid and flexible oscillations. The results are generated by a finite volume Euler method. It is based on the Runge-Kutta time stepping scheme developed by Jameson et al.. To increase the time step which is limited by the stability of Runge-Kutta scheme, the implicit residual smoothing which is modified by using variable coefficients to prevent the loss of flow physics for the unsteady flows is engaged in the calculations. With this unconditional stable solver the unsteady flows about the wings in arbitrary motion can be received efficiently. The two- and three-dimensional rectangular wings which are in rigid and flexible pitching oscillations in the transonic flow are investigated here, some of the computational results are compared with the experimental data. The influence of the reduced frequency for the two kinds of the wings are researched. All the results given in this work are reasonable.

An implicit upwind finite volume solver for the Euler equations using the improved flux-splitting method is established and used to calculate the transonic flow past the airfoils with heaving, pitching oscillations and the control surface. Results are given for the NACA64A-10 airfoil which is in harmonic heaving and pitching oscillation and with the control surface in the transonic flow field. Some computational results are compared with the experiment data and the good agreements are shown in the paper.

[Abstract] This paper describes the modeling of aircraft’s maneuver to support pilot-inloop flight simulation. The aircraft’s maneuver is modeled on base of analyzing the kinematical features of aircraft in different flight modes, related with airborne sensor working state and load release. According to the kinematical features, the relation between aircraft’s maneuver and airborne sensor’s working state and load release path/envelope is modeled. Through the model, the Pilot Control Instructions (PCI) can be simulated. The model can reflect not only the kinematical features of an aircraft but also the relation between aircraft motion and airborne sensor’s working state and load release in modern flight. Resulting from the model, the simulated PCI can satisfy a pilot’s needs in the flight simulation. The model is estimated through pilot-in-loop interception flight simulation. The results indicate the rationality of the model. The model can also be used in other states of aircraft’s pilot-in-loop simulation system.

This paper presents the computer simulation methodology for complication system's conceptual design. The conceptual design of complication system, such as a fighter, is a multi-system and multi-discipline synthesis course. And it includes a lot of factors. Finding the main factor and modeling of the relation between the main factor and the other factors should be focused on in the design. Computer simulation is a good method to visualize these relations. Through the method, conceptual design of a complication system can be effectively done. As an example, the relation between the fighter's maneuver and radar working state and missile launch is modeled. Resulting from the model, a pilot-in-loop air combat simulation can be accomplished. Through the simulation, the factors affects fighter performance can be analyzed. The method has been successfully used in fighter conceptual design. So, it can also be used in other complication system's conceptual design

Taking into account the influences of scatterer geometrical shapes on induced currents, an algorithm, termed the sparse-matrix method (SMM), is proposed to calculate radar cross section (RCS) of aircraft configuration. Based on the geometrical characteristics and the method of moment (MOM), the SMM points out that the strong current coupling zone could be predefined according to the shape of scatterers. Two geometrical parameters, the surface curvature and the electrical space between the field position and source position, are deducted to distinguish the dominant current coupling. Then the strong current coupling is computed to construct an impedance matrix having sparse nature, which is solved to compute RCS. The efficiency and feasibility of the SMM are demonstrated by computing elec-tromagnetic scattering of some kinds of shapes such as a cone-sphere with a gap, a bi-arc column and a stealth aircraft configuration. The numerical results show that: (1) the accuracy of SMM is satisfied, as compared with MOM, and the computational time it spends is only about 8% of the MOM; (2) with the electrical space considered, making another allowance for the surface curvature can reduce the computation time by 9.5%.