The direct force and aerodynamic force compound control technology is proposed for achieving the rapid turning and precise guided strike in high-maneuver tactical missiles. To investigate the coupling interaction problems in the compound control method, both wind tunnel experiments and unsteady numerical simulations are conducted on a typical missile configuration with transverse jet nozzles. Specially, dynamic response to the direct force/aerodynamic force control are carried out experimentally in wind tunnel. The results shows that the rear transverse jets provide beneficial interaction for aerodynamic characteristics. As a consequence of the pitching motion, the upstream separation region is largely affected by the unsteady interaction flow. And the aerodynamic coefficients experience obvious dynamic hysteresis, which is enhanced by the pitching motion rate. The amplification factor of pitch moment is more seriously affected under large angle-of-attack in the pitching motion, which is deviated from the static counterparts and leads to unfavorable impact. The direct force/aerodynamic force compound control is proven to be capable of increasing the pitching-up velocity, and it is a reliable option for rapid turning and precise attitude control.

（School of Aerospace Engineering， Tsinghua University， Beijing 100084， China） Abstract: When the missile flies at high angle of attack, it will generate complex asymmetric vortex, resulting in large side force, which will have a serious impact on the flight control of the missile. It is often necessary to introduce artificial perturbation to describe the phenomenon of generating side force in the test by numerical method. In this paper, the improved delayed detached-eddy simulation （IDDES） method with shear layer adaptation is used to calculate the missile configuration with /without perturbation, and the calculated results are compared with the experimental results. Without adding perturbation and at 40 degree angle of attack, the fully symmetrical vortex is calculated, and there is almost no side force, which eliminates the uncertainty and non-physics of asymmetric effect caused by numerical error. Adding a suitable "strake" to the missile head to introduce the perturbation and simulate the asymmetric effect caused by the test can accurately predict the separation of the missile. The pressure distribution is in good agreement with the test results.

Aiming at quick over-the-shoulder launch implementation for air-to-air missiles, a novel launch strategy namely poststall re-orientation is proposed, which uses over stall maneuver of missiles at high angle of attack to re-orientate rapidly. The kinetic and aerodynamic characteristics of a slender missile with relaxed static stability design is calculated by solving unsteady RANS equations coupled with RBD equations, and unsteady RANS equations are discretized under the framework of finite-difference method on moving overset grids. The influence of preset deflection angle of rudders, ejection angular velocity and direct force control on the trajectory and attitude of the missile in overturning is investigated. Finally, a launching strategy combining preset rudder deflection and direct force control is proposed to realize re-orientation of missiles in 1.5 s, which presents valuable guidelines for fast backward-firing of air-to-air missiles.

The unsteady flow of slender body at Mach number 0.6~1.15 and angle of attack 0°~180° is calculated by using unsteady numerical simulation DDES method. The unsteady aerodynamic characteristics, frequency characteristics and vortex flow characteristics are analyzed. The main conclusions are as follows: aerodynamic force strong unsteadiness of slender body exists in the range of angle of attack from 45° to 165°, and the amplitude of unsteady fluctuation of lateral force is particularly strong, and its instantaneous magnitude is 1/4 to 1/2 of normal force coefficient. The lateral aerodynamic force has obvious dominant frequency while the longitudinal aerodynamic force/moment has no obvious dominant frequency in the range of angle of attack from 45° to 165°. The aerodynamic frequency is mainly brought by the slender body, and the wing contributes little. Complex vortex flow on the leeward side is the main cause of unsteady aerodynamic force, which is reflected in vortex generation, vortex switching and vortex shedding.

In this paper, the aerodynamic interference of jet plume to a slender body in subsonic and transonic flow is investigated. Simulation indicates that on the upstream side of the body, jet plume interacting with freestream will cause a more complex flow which induces low pressure area on the slender body downstream of the nozzle. At medium angle of attack or still less, tendency for the interacting effect of CN and mz on the slender body reveals differently. At large angle of attack, the jet nozzle on the upstream side of the slender body still keeps some maneuvering capability to pitching movement. The interference effect varies more violently when the upstream jet is working, while it’s more steady when the down stream jet is working. The intensity of the jet plume interference is greater while the slender body is placed in transonic flow than in subsonic flow.

In order to improve the test technology and improve the accuracy of the test data under the super large angle of attack, the influence analysis of the model support interference is carried out. Aiming at the support interference problem encountered by the slender configuration in the super large angle of attack test of the high-speed wind tunnel, a numerical simulation method is accepted to study the support interference problem of slender configuration at super large angles of attack in high-speed wind tunnel experiment. The interference of side and segmented support (rear / back support) is analyzed. Numerical results show that, in general, the interference of segmented support is less. Therefore, it's better to use segmented support for static tests. Since the side sting across flow field, additional shock and expansion waves are formed, which significantly change the pressure distribution about the model. When side support is accepted for continuous dynamic experiments at incidence ranging from 0 to 180°, it's necessary to analyze support interference and modify data combined with numerical method.

This article demonstrated the research work on the problem of canard-wing interference of rolling airframe missile via wind tunnel rotating force measuring experiment. According to the result of the wind tunnel experiment, the variation law of rolling rate with different canard-wing setting angle is obtained. The results obtained by four point average method or eight point average method show that the longitudinal aerodynamic problem of rolling airframe missile can be solved with quasi-steady method. The strong interference from down wash under small attack angle condition is one of the main causes which lead to large fluctuation of the longitudinal aerodynamics. The research concluded a rule of the Magnus effect caused by canard-wing interference refer to the rolling rate and change of attack angle. Magnus force and Magnus moment at medium angle of attack have a strong nonlinearity and large absolute value.

This paper introduces the design performance parameters, flow field calibration results and standard model force measurements of Φ120 hypersonic wind tunnel of Zhejiang University. This wind tunnel adopts conventional "blow-suck" layout. The diameter of the nozzle outlet is 120 mm; the design operation Mach number is 5.0, 6.0 and 7.0; the total pressure of the incoming flow is 0.2~2.0 MPa; the total temperature of the incoming flow is 400~700 K; and the running time is not less than 10 seconds. The flow field calibration results show that the uniform region diameters of the 5.0/6.0/7.0 nozzles are more than 90mm; the average Mach numbers of the uniform region are 5.07, 6.05 and 6.94; the rms deviations are 0.018, 0.015 and 0.023; and the axial Mach gradients of the uniform region are 0.021, 0.016 and 0.031 respectively. All the above key parameters reach the GJB1179A-2012 qualified indices, and some parameters reach the GJB1179A-2012 advanced indices. The aerodynamic force measurement results of the AGARD HB-2 standard model in the Φ120 hypersonic wind tunnel are consistent with the simulation results of in-house code GRAND and the reference experimental results in GJB4399-2002. In summary, Φ120 hypersonic wind tunnel has a wide range of parameters, which is applicable for hypersonic aerodynamic teaching experiments and advanced scientific research subjects such as hypersonic complex flow mechanism and flow control for heat and drag reduction.

Subjected to the accuracy of surrogate model and premature convergence sometimes occurs, the solution of classical efficient global optimization (EGO) usually can be improved. Regarding this point, this paper presents a thorough study of the EGO with high accuracy optimal solution. The algorithm is based on the Kriging surrogate model, in which the Kriging believe strategy based Expected Improvement (EI) function is adopted, it can help to lead the optimal solution to local optimum in the late period of iteration. Besides, in order to improving the accuracy, cooperating with quasi-Newton method and Powell method are also incorporated. Several representative numerical examples are selected to test the algorithms above, the result shows that the algorithms in this paper can reach more accurate global optimal solution than classical EGO, with less additional cost. At last, an aerodynamic optimization problem is developed, it shows that the drag coefficient decreased 1.11% further than the former EGO, shows its practicability in an engineering environment.

In order to quickly and accurately estimate the aerodynamic characteristics of the aircraft in the early stage of aircraft development, an empirical engineering correction algorithm that integrates physical laws is proposed. The important parameters of the empirical formula are corrected and fitted by a small amount of CFD component aerodynamic coefficient data, so as to improve the calculation accuracy of the engineering algorithm. Firstly, the applicability of the data is verified, and the applicability of the formula source Datcom calculation data and the CFD data is verified; then, the normal force coefficient is analyzed from the perspective of components, and a small number of CFD components are used to correct the aerodynamic coefficient data. The empirical formulas of aerodynamic coefficients of aircraft components are fitted, and the calculation equations of aerodynamic coefficients of the algorithm is confirmed by comparing the calculation results of the revised formulas with the CFD prediction set.

In this paper, the conical shape dispenser and submunition are used, and the backward separation simulation is carried out by numerical simulation. The interference effect of dispenser on submunition separation under the conditions of variable separation speed and variable angle of attack are studied. The research shows that, during separation process of submunition, the trajectory and attitudes of submunition change greatly, which caused by the interference of dispenser. It is also found that the separation speed mainly affects the relative displacement between dispenser and submunition, and higher speed will accelerate the submunition keeps away from dispenser. The separation angle of attack mainly affects the altitudes of submunition, the divergence of roll angle will be faster and oscillation amplitudes of pitch and yaw angle will be larger with higher separation angle of attack, which is negative for separation.

The heat distribution on the surface of the blunt body is calculated using CFD method and compared with the experimental data to verify the reliability of the CFD methods. The temperature distribution inside the structure at different times is calculated using iterative coupled calculation and one-way solving methods respectively. The iterative coupled calculation completes the integrated calculation through the mutual data exchange between the flow field and the structure in single time iteration. The one-way solution obtains the aerodynamic heating of the surface through CFD method, and then obtains the internal temperature distribution through solving the Structure Heat Transfer Equation. The calculation results show that the stagnation point of the blunt body is the most seriously affected by the aerodynamic heating; the structural temperature calculated by iterative coupling calculation method is much closer to the experimental results than the one-way solution method.

To alleviate the requirement of mesh resolution in the near-wall region, an immersed boundary method, called the local domain-free discretization (DFD) method, is extended to large eddy simulation (LES) of turbulent flows. The discrete form of partial differential equations at an interior node may involve some nodes outside the solution domain. The flow variables at these exterior dependent nodes are evaluated via linear extrapolation along the direction normal to the wall. A wall model based on the turbulence boundary layer equations is introduced. The wall shear stress yielded by the wall model and the no-penetration condition are enforced at the immersed boundary to evaluate the velocity components at an exterior dependent node. For turbulence closure, a dynamic subgrid scale (SGS) model is adopted and the Lagrangian averaging procedure is used to compute the model coefficient. The SGS eddy viscosity at an exterior dependent node is set to be equal to that at the outer layer. Numerical experiments of backward-facing step flow on relatively coarse meshes have been conducted to verify the ability of the present LES-DFD method. The predicted results agree well with the published experimental or numerical data.