Based on the Favre averaged three dimensional compressible Navier Stokes equations and Baldwin Lomax algebraic turbulence model, the Jameson scheme and matrix artificial dissipation was used for simulation of the integrated flow field around a missile and mixed compression intakes in "X" array Boundary layer suction and starting of mixed compression intake were first proposed in numerical simulation and the impact of boundary layer suction on the flow field was investigated The characteristics of the integrated flow field were discussed and the contour maps of Mach number and the velocity distribution were described The impact of the location of intakes, boundary layer suction and attack angle on the performance of intakes and the impact of the outer shape of intakes on the lift of the missile were also discussed The numerical results show that boundary layer suction is necessary to mixed compression intake It also shows the influence of the location and the outer shape of intakes The result can be referenced in the design of missile and intakes

In order to further uncover the interaction of tip clearance flow with three-dimensional separation in the corner region of a compressor cascade, detailed experimental investigations were carried out. The outlet aerodynamic parameters were measured by five hole pneumatic probe, and the distribution of the static pressure on the endwall was obtained. The effects of tip clearance size and inlet flow angle on the interaction of tip clearance flow with three-dimensional separation in a compressor cascade were analyzed in detail. The results show that the three-dimensional separation on the blade suction surface is largely removed by the clearance flow for a proper tip clearance size and the aerodynamic performance of the cascade is improved. The removal mechanism is associated with preventing the endwall fluid from interacting with the suction surface boundary layer by the clearance flow. The interaction between the leakage vortex and the tip side passage vortex turns stronger with increasing the clearance size and the flow turning. 

Calculating performance of PDTE can provide important data for design and further experiment. Based on the theory of ZND detonation wave and cycle analysis,a model (ZND cycle analysis method, model-1) to calculate the performance of PDTE was established.The model and former thrust wall model (model-2) with PDC characteristic chart method (model-3) was compared. And using the three models the design point performance of a PDTE was calculated. The result shows that, compared with model-2 the error of specific thrust (FS) by using model-1 reduces 1.75%. And the error of specific fuel consumption (sfc) reduces 4.04%. Thus model-1 is more precise than model-2. Compared with model-3 which need more experiment data the method-1 is more convenient. This model can be used to fast evaluate the performance of PDTE.The change of performance gain of PDTE was calculated with the change of flight state. 

The feasibility of γ-Reθ transition model on hypersonic flows was studied by an in-house CFD software. The model was verified by turbulent flows over the T3flat plate for low speed flows. Numerical simulations were conducted for hypersonic flows,taking the around compressible corners and double cone with laminar flow as the examples, with traditional turbulence models and also transition model in order to assess the capability of these models for high speed flows. The computed surface pressure and heat transfer rate were compared with the experimental data. The results indicate that although this transition model is correlated for low speed flows, it could be used to predict the flow transition for hypersonic flows .

In order to study the effect of kerosene distribution uniformity on combustion performance, experiments were carried out by changing the number of injector holes, the direction of injection (one-row and two-row on a strut) and the distribution of kerosene injector holes (3 holes or 4 holes in one-row) in one strut, with the free stream Mach 6 and total temperature 1650K. Some conclusions can be drawn from the analysis of the experimental data in different conditions such as the integrated balance data and the wall pressure distributions:compared to the other two struts, the top holes of the 1×4 strut are nearer to the wall and its bottom holes are closer to the cavity, so the kerosene enters the low-speed and high-temperature zones such as the boundary layer and the circumfluence zone of the cavity more easily, which is helpful for the ignition and steady combustion of the kerosene; for the opposite injection scheme, most kerosene and its combustion are in the core area of the combustor, so the wall heat load is reduced, but too rich fuel has bad influence on the total combustion performance; for the both-side injection scheme, some kerosene is injected into the corner of the combustor wall, which results in more well-proportioned distribution in the cross section of the combustor, so ignition becomes much easier, and better combustion effect and greater net thrust are obtained. 

The history of electric propulsion technology development was summarized. The features and application status of different types of electric thruster were analyzed. Some suggestions for speeding up the development of Chinese electric propulsion technology were made, based upon the analysis for the technology trends and the comparison between China and abroad.

Modeling of small perturbation State Variable Model (SVM) for the design of aeroengines’ multivariable robust control system was studied. The small perturbation SVM of an aeroengine was derived directly by fitting its nonlinear simulation data under small perturbation about the steady operating point. The SVM established by this approach has high accuracy since the modeling error is minimized under the meaning of least square. Moreover the approach is not limited by the plant order. The application of the approach to establishing the small perturbation SVM for a turbofan engine shows its effectiveness.

The effects of nanometer PbCO 3 on combustion behavior of NEPE solid propellant have been studied The compatibility of nanometer PbCO 3 and nitrate ester , as well as the effects of nanometer PbCO 3 on the curing reaction and on the thermal decomposition behaviors of ammonium perchlorate and nitroamino compounds were studied by Differential Scanning Calorimetry (DSC) It was found that nanometer PbCO 3 exhibited good compatibility with nitrate ester which had observable acceleration on the thermal decomposition of nitroamino compounds and on the curing reaction but little effect on ammonium perchlorate When nanometer PbCO 3 was applied to NEPE solid propellant, the pressure exponent were effectively brought down to 0 54 and 0 52 with which content was 1% and 2%respectively

In order to assess the effects of stator-rotor cavity rim sealing flow on the performance of a turbine rotor, a detailed 3D unsteady numerical investigation is presented. Compared with no cavity model, the time-solved results show that the rotating ingress and egress structures change the incidence angle before the rotor blades and stretch the horseshoe vortex in the hub region, thereby resulting in large changes in turbine exit conditions. The existence of stator-rotor cavity with no sealing flow can ingest the vane wake fluid and has positive effects on turbine performance. The interaction of sealing flow and upstream vane weak gives rise to the entropy in the blade passages. The turbine efficiency decreases by 2.1% when the sealing mass flow rate is 1.37% of the mainstream mass flow rate. This proves that the effects of sealing flow on turbine rotor performance are very large and should be considered in turbine design. 

Nanometer-CuO powder was synthesized by solid-state reaction using CuCl 2·2H 2O with NaOH at room temperature. Average particle sizes of nanometer CuO were about 10 nm. The effect of nanometer CuO on thermal decomposition characteristics of RDX was investigated by DSC. The results show that the peak temperature of thermal decomposition of RDX shifts 12 ℃ downward due to the effect of nanometer CuO. The effect of nanometer CuO on the thermal decomposition characteristics of RDX is different from that of normal CuO.

Aimed at investigating the effects of cowl lip shape on 2-D inlet, Fluent was adopted for numerical simulation. A conclusion is drawn that the reduction of area of the cowl lip can cut down the starting Mach number of the inlet effectively at price of sacrificing some mass flow. Meanwhile, the total pressure recovery is improved. Based on that, an optimized design method was put forward finally.

The thermal decomposition characteristics of HMX influenced by the addition of aluminum,nickel,copper with different particle size(general and nano)were studied by PDSC and TG The result shows that the nano copper have the greatest influence on the properties of condensed phase decomposition of HMX among the metal powders Such catalysis effect of nano copper will be weakened by the decrease of the content of naon copper or the increase of the system pressure Based on the kinetics result inferred from the isothermal DSC data,this catalysis mechanics are described to the efficacy of first catalysis,secondary catalysis and reaction site effect of nano copper on HMX

In order to explore the thrust performance of microwave thruster，the thrust produced by microwave thruster system was measured with three-wire torsion pendulum thrust measurement system and the measurement uncertainty was also studied，thereby judging the credibility of the experimental measurements. The results show that three-wire torsion pendulum thrust measurement system can measure thrust not less than 3mN under the existing experimental conditions with the relative uncertainty of 14%. Within the measuring range of three-wire torsion pendulum thrust measurement system，the independent microwave thruster propulsion device did not detect significant thrust. Measurement results fluctuate within ± 0.7mN range under the conditions 230W microwave power output，and the relative uncertainty is greater than 80%.

In order to investigate the effects of air throttling on scramjet combustor, air throttling numerical simulation and experimental measurements were performed.The parameters(position,flux,air throttling off time) of air throttling were analyzed. Results show that under the inflow air condition,Ma=2.0,T _e=548.8K,p _e=101.6kPa,the stabilization time of shock train is shorter with the location of air throttling at 745mm from the combustor entrance and flame stabilization is failed. But the flame is successfully stabilized with the location of air throttling at 875mm. When the flux of air throttling and air throttling off time increase, the wall pressure of scramjet combustor increases, which may influence the inlet starting. 30% flux of inflow air and air throttling off time under 440ms for air throttling are fit for this inflow air condition. 

The basic theory and technical state of all kinds of electric rocket engine were discussed According to its operation features and application examples, the application situation of all its kinds were analyzed It is forecasted that the electric propulsion has great prospect in aerospace applications

On the basis of the design concept of the traditional three-dimensional internal waverider inlet， an aerodynamic combination design concept of the forebody and three-dimensional inward turning inlet with the characteristic of waverider theory is presented. In combination with the flow field of two incident shock waves and the oblique shock wave theory， a flow field to give consideration to the two-dimensional flow in upstream and three-dimensional flow in the downstream is obtained. Afterwards， the aerodynamic combined surface of the inward turning inlet and waverider forebody could be designed by using the streamline traced method. A forebody/inlet configuration is subsequently derived from this concept and numerically studied. The results show that， at the design point (Ma6.0)， the mass flow rate of the inward turning inlet is 0.96， and the total pressure recovery is 0.53. In addition， at off-design points (Ma4.0)， the mass flow rate and the total pressure recovery are 0.71 and 0.70， respectively. Compared with the two-dimensional mixed compression inlet， the performance of aerodynamic combination configuration significantly improves. Especially， the mass flow rate at the design point increases by 4.1%.

The characteristics, development history and current situation of hybrid rocket motor were presented. The potential application of hybrid rocket motor was analyzed based on the astronautic development in China. The hybrid rocket motor can be used in sounding rockets, low cost target drones and missiles, suborbital vehicles, large launch boosters, advanced upper stages and orbital transfer systems.Thus,the application prospect of hybrid rocket motor is extensive. The design method for hybrid sounding rocket in Beijing University of Aeronautics and Astronautics was summarized. The key technologies which affect the developments and applications of hybrid rocket motors were analyzed. 

A new fast and efficient algorithm was introduced to solve the nonlinear, hyperbolic partial differential equations governing the unsteady flow of the propellant lines by Chebyshev super spectral viscosity (SSV) method. Compared with conventional spectral method, the method can stabilize the numerical oscillation, accelerate the convergence and improve the computational efficiency. The details of the method are presented with an illustration of the water hammer problem in a simple piping system (consised of a tank, a pipe and a valve). The results obtained from the Chebyshev super spectral viscosity method exhibit greater consistency with conventional water hammer calculations. In addition, Chebyshev super spectral viscosity method requires less computer calculation time with less numerical error, indicating that the method is more suitable to solve liquid transients in propellant lines. 

In order to further reveal the effects and action mechanics of the blended winglet on the tip leakage flow of compressor rotor，a numerical simulation has been carried out to investigate the effects of blade tip winglet on controlling tip clearance flow in an axial compressor rotor. Emphasis was put on the analysis of effectiveness of blade tip winglet with different winglet geometries，suction side winglet and pressure side winglet. The simulation results show that rotor tip leakage vortex trajectories are changed by the blended tip winglets. Thus the degree of suction surface boundary layer separation is also changed. The specific pressure side winglet is found highly effective to reduce the compressor rotor stalling flow coefficient by 8.20% with a slight reduction of efficiency.

In purpose of making a study on ablation morphology and performance for C/C material under the direct impact from particles and enriching C/C material ablation study, particle erosion experiment was undertaken. The study has been made for microcosmic ablation morphology by resorting to such measuring approaches as scanning electron microscope, 3D microscopic imaging and etc. The results show that because of direct impact from particle erosion, small dimples are appeared in erosion areas, while with protruded axial rod fibers, on the surfaces of these rods, pits in different sizes are observed with partial monofilaments breaking from matrixes.Therefore, the microcosmic cone-shape holed morphology is formed. It shows that holes and notches are occurred in partial areas of radial fiber bundles, and the whole morphology is no longer in consistency.While the microcosmic morphology for the non-erosion part is in accordance with that morphology at throat with bamboo-shoot monofilaments.