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Lv Z.,Nanjing University of Aeronautics and Astronautics | Xu J.,Nanjing University of Aeronautics and Astronautics | Mo J.,Xian Aerospace Propulsion Institute
Acta Astronautica | Year: 2017

The present study focuses on the unsteady mode transition process of an over-under TBCC exhaust system. The method of characteristics is applied to design the over-under TBCC exhaust system according to the entrance parameters of the turbine nozzle and ramjet nozzle at the design point. The dynamic mesh is adopted to adjust to the update of the computational domain, and the unsteady numerical method is employed to simulate the dynamic flowfield of the exhaust system during the mode transition process. The results show that the flowfield structure and the performance vary greatly during the mode transition. Owing to the interaction between the turbine exhaust jet and ramjet plume, the flowfiled in the turbine nozzle is affected by the ramjet exhaust jet considerably. The axial thrust of the turbine nozzle decreases, while that of the ramjet nozzle increases gradually during the mode transition, but the total axial thrust of the entire exhaust system varies smoothly. Both the axial thrust coefficient and pitching moment of the exhaust system increase along with the open of the ramjet nozzle, while the result for the lift is contrary. However, the axial thrust coefficient, lift and pitching moment all decrease rapidly with the shutdown of the turbine nozzle, and the decreases in axial thrust coefficient, lift and pitching moment are 1.04%, 67.15% and 80.92%, respectively. Besides, two sudden change of the axial thrust coefficient exist at the beginning and end of the motion of the splitter plate. © 2017 IAA

Zhou J.,Xian Aerospace Propulsion Institute | Yang Z.,Northwestern Polytechnical University
Zhendong yu Chongji/Journal of Vibration and Shock | Year: 2017

The equations of motion for nonlinear flutter of curved composite panels were developed with the finite element method. The reduced order modes constructed with the proper orthogonal decomposition (POD) method were used in reducing the order of these equations, and the equations of motion were transformed into a reduced nonlinear system under the POD modal coordinates, then the reduced equations were solved in time domain by using the numerical integration method. Compared with the results calculated using the traditional modal reduction method, the results using POD method based on reduced order models agreed well with the former, and also saved the computation time greatly. © 2017, Editorial Office of Journal of Vibration and Shock. All right reserved.

Lv Z.,Nanjing University of Aeronautics and Astronautics | Xu J.,Nanjing University of Aeronautics and Astronautics | Mo J.,Xian Aerospace Propulsion Institute
Acta Astronautica | Year: 2017

The performance of a single expansion ramp nozzle (SERN) is poor due to over-expansion at off-design conditions. The present study focuses on improving the SERN performance by secondary injection on the cowl and is carried out by using the k−ε RNG turbulence model. The incidence shock wave resulting from the secondary injection impinges on the expansion ramp, resulting in separation and the increase of the pressure distribution along the ramp. The performance of the SERN can be improved significantly, and the augmentation of the thrust coefficient, lift and pitch moment can be as high as 3.16%, 29.43% and 41.67%, respectively, when the nozzle pressure ratio (NPR) is 10. The location of the injection has a considerable effect on the lift and pitching moment, and the direction of the pitch moment can be changed from nose-up to nose-down when the injection is on the tail of the cowl. The effect of the injection on the axial thrust coefficient is much more apparent, if the operation NPR is far from the design point, and however, the results for the lift and pitching moment are opposite. The increases of injection total pressure and injection width have positive impacts on the SERN performance. And if the parameter φ maintains constant, the axial thrust coefficient would increase when the injection total pressure decreases, so low energy flow can also be used as the secondary injection without decreasing the lift and pitching moment. The mass flow rate of the injection can be decreased by applying the higher total temperature flow without reducing the performance of the SERN. © 2017 IAA

Yuanqi L.,Xian Aerospace Propulsion Institute | Hongjun L.,Xian Aerospace Propulsion Institute | Haohai X.,Xian Aerospace Propulsion Institute
Proceedings of the International Astronautical Congress, IAC | Year: 2016

System static characteristic simulation is critical for the performance analysis of liquid rocket engine (LRE) and pre-design activities in particular. Analysis tool is a fundamental instrument along the entire design period from pre-design phase to parametric studies from tuning of the engine parameters. During the engine working process, interfering factors which will cause the engine parameters off-design conditions cannot be ignored. YF-115 engine, based on LOX/Kerosene staged combustion cycle, designed by Xi'an Aerospace Propulsion Institute is the upper stage engine of Chinese new generate launch vehicle Long-March 6 (CZ-6), which has been successfully launched in Sep. 2015. A new non-linear static characteristic simulation model for YF-115 is established in this paper. Compared with the old one, combustion chamber (CC) and gas generator (GG) performances are estimated using the NASA code Chemical Equilibrium with Applications (CEA) instead of looking up standard gas table. Broyden quasi Newton method is used to solve the non-linear algebraic equations. As a validation, comparison of the main engine parameters (such as chamber pressurethrust and rotation rate etc.) is carried out between the trial run measurement data and the simulation result. Results reveal that calculation stability and accuracy are greatly improved by the new algorithm. Furthermore, the influence of the interfering factors (instability of inlet pressure or pump efficiency for instance) on the engine parameters is analyzed. The results obtained can be used for the analysis of rocket engine test results, the reliability and faults analysis, and also for revealing the variation law of the engine parameters with various interfering factors. Copyright © 2016 by the International Astronautical Federation (IAF). All rights reserved.

Yuanqi L.,Xian Aerospace Propulsion Institute | Hongjun L.,Xian Aerospace Propulsion Institute | Xiaobo S.,Xian Aerospace Propulsion Institute
Proceedings of the International Astronautical Congress, IAC | Year: 2016

The LOX/Kerosene engine, YF-100, designed by Xi'an Aerospace Propulsion Institute, is the first stage engine of Chinese new generation launch vehicle Long-March 6 (CZ-6), which has been successfully launched in Sep. 2015. During the rocket launch or engine trial, launch pad or test bench will be heated intensely by high temperature gas exhausted from the engine nozzle. Because of the low oxygen-fuel equivalence ratio in thrust chamber of the engine, the combustion product is fuel-rich hot gas. Combustion occurs after the fuel-rich hot gas exhausting from engine nozzle and mixing with surrounding air. High temperature flame generates in the gas-air mixing layer. This phenomenon is called afterburning. Radiant heat holds a great quantity and convection holds a less, because the plume do not scour the launch pad directly. In this paper, researches on exhaust plume radiative heat transfer of LOX/Kerosene rocket engine is presented. Firstly, nozzle and plume flow field is obtained via solving Reynolds Averaged Navier-Stokes equations with k-e turbulent model. Finite-rate combustion model is employed to describe the equilibrium flow and afterburning. Secondly, the radiative heat transfer is investigated based on the temperature, pressure and species fraction of the plume flow field calculated previously. Discrete ordinates model (DOM) is used to simulate radiative heat transfer, and narrow band model is used to simulate absorption coefficient of gas. Furthermore, the accuracy and reliability of simulation model are verified through comparison with the radiation measurement in engine trial run. Analysis and simulation results show that afterburning effects lead to a much higher heat flux than expected. Thermal protection method of launch pad is suggested according to the results in the end of this paper. Copyright © 2016 by the International Astronautical Federation (IAF). All rights reserved.

Yang L.-J.,Beihang University | Fu Q.-F.,Beihang University | Qu Y.-Y.,Beihang University | Gu B.,Beihang University | Zhang M.-Z.,Xian Aerospace Propulsion Institute
International Journal of Multiphase Flow | Year: 2012

Gel propellant is promising for future aerospace application, but because it behaves as the non-Newtonian power-law liquid, it is difficult to atomize. Impinging jet injectors are often used for atomization of gelled propellant. To understand the atomization mechanism of gelled propellant, a linear instability analysis method was used to investigate the instability and breakup characteristics of the sheet formed by a gelled propellant impinging-jet injector. The maximum disturbance wave growth rate and dominant wave number were determined by solving the dispersion equation of a power-law liquid sheet. It was found that the maximum disturbances growth rate and the dominant wave number both increase as Weber number of the liquid sheet increases. Consistency coefficient and flow index were tested for their influence on the stability of the power-law liquid sheet. A modified model, to predict the breakup length and critical wavelength of the power-law liquid sheet, was adopted. To validate the power-law liquid sheet breakup model, experiments were performed with injectors of different configurations and a high speed camera was used to show detailed information of the liquid sheet breakup process. The rheology of the power-law fluid used in the present study was also investigated. Comparison between the theoretical and experimental results shows that the linear instability analysis method can be applied to predict breakup length and wavelength of the power-law liquid sheet. © 2011 Elsevier Ltd.

Wang X.M.,Northwestern Polytechnical University | Wang Y.F.,AVIC The First Aircraft Institute | Lu Z.Z.,Northwestern Polytechnical University | Deng C.H.,Xian Aerospace Propulsion Institute | Yue Z.F.,Northwestern Polytechnical University
Mechanics of Materials | Year: 2010

This paper presents the results of a series of experiments to investigate the superelastic cyclic stress-strain responses of NiTi shape memory alloys under tension-torsion biaxial loading conditions. The uniaxial tension and torsion experiments were also conducted to make comparisons. Experiments were controlled by axial displacement and torsional angle in sine wave form with fixed maximum values. Saturation is reached after 30 cycles. The evolutions of equivalent stress-strain curves as well as the separated tensile and shear stress-strain curves during cycling are analyzed. Dissipated energy density, characteristic stresses and strains as functions of deformation cycles are paid attention. Results show that the mechanical responses are significantly affected by the loading path. The stress-strain behaviors under proportional loading are similar to those under uniaxial loading. With the increase of the out-of-phase angle during the non-proportional loading, the special phase transformation exhibition totally disappears in the equivalent stress-strain curves. © 2009 Elsevier Ltd. All rights reserved.

Li P.,Northwestern Polytechnical University | Li W.-L.,Xian Aerospace Propulsion Institute | He G.-Q.,Northwestern Polytechnical University
Guti Huojian Jishu/Journal of Solid Rocket Technology | Year: 2012

A numerical calculation on the flow field of mixing combustion chamber without chemical reaction was carried out for hydrazine monopropellant air-turbo-rocket under uniform inlet flows. The induction and decay evolution of streamwise and normal vortices were obtained, and an analysis was made on the influence of this course on turbulent mixing efficiency of the bypass and core flows. A quantitative analysis revealed that streamwise vortices induced by a spanwise array of large scale secondary flows at trailing edge play a dominant role in the downstream mixing process of the bypass and core flows. It is preferable to employ scarfed lobed mixers with large lobe penetration rate. A preliminary analysis of turbulent mixing efficiency was also made between computational and experimental data for two mixing schemes including three kinds of lobed mixer, and results show that nonuniform inlet flow conditions have a significant impact upon turbulent mixing and combustion efficiency for small dimension air-turbo-rocket.

Ma A.-J.,Xi'an Technological University | Li H.,Xian Aerospace Propulsion Institute | Chen W.,Xi'an Technological University | Hou Y.,Xi'an Technological University
Polymer - Plastics Technology and Engineering | Year: 2013

The surface functionalized silicon carbide/carbon fiber (SiC/CF) hybrid fillers were introduced to improve the thermal conductivities and mechanical properties of the epoxy resin composites. Results revealed that, the thermal conductivities of epoxy resin composites were increased with the increasing volume fraction of SiC, and a higher thermal conductivity of 1.226 W/mK could be achieved with 28 vol% treated SiC/CF hybrid fillers (25 vol% treated SiC +3 vol% treated CF), about 6 times higher than that of native epoxy resin (0.202 W/mK). Both the flexural strength and impact strength of the epoxy resin composites increased up to 5 vol% incorporation, but decreased with the further addition of SiC. However, the electrical conductivity was decreased with the increasing volume fraction of SiC. For a given SiC/CF hybrid fillers loading, the surface functionalized SiC/CF hybrid fillers exhibited a positive effect on the thermal conductivities and mechanical properties of the epoxy resin composites. © 2013 Copyright Taylor and Francis Group, LLC.

Zhou J.,Xian Aerospace Propulsion Institute
Advances in the Astronautical Sciences | Year: 2010

As a significant sub-system for most space vehicles, the space propulsion system was used to provide the power for vehicle attitude control, orbit maneuver, and so on. Under the circumstance to develop the micro-propulsion system for micro space shuttle, micro-satellite, and micro deep space exploration vehicle (with a mass between 1∼100kg), a 1 Newton bi-propellant thruster using the propellants of H2O2/CnH 2n+1OH and stainless steel material was designed. Using the method of numerical simulation, contours of temperature, pressure, Mach number and velocity in rectangular combustion chamber and nozzle were achieved, and the influence of near-wall layer was analyzed. Moreover, the fire test was conducted and fire properties of the engine were obtained.

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