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Commercial Aerospace: The Next Launchpad For the Chinese Economy

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China Report ASEAN 2024年第2期9卷 22-25页

作者： Zhang Lijuan 不详

On December 29,2023,the first launch service tower was completed at the Hainan International Commercial Aerospace Launch Center,marking a key step forward in building the launch capacity of China’s first commercial aerospace launch site.On December 10,a Hyperbola-2 methane-liquid oxygen reusable verification rocket was launched successfully,marking the first recovery of reusable carrier rockets in China.

关键词： Launch launch rocket

Overview of Al-based nanoenergetic ingredients for solid rocket propulsion

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Defence Technology（防务技术） 2018年第5期14卷 357-365页

作者： Luigi T.DeLuca Politecnico di Milano MilanItaly Nanjing University of Science and Technology NanjingChina

The introduction of nano-sized energetic ingredients first occurred in Russia about 60 years ago and arose great expectations in the rocket propulsion community, thanks to the higher energy densities and faster energy release rates exhibited with respect to conventional ingredients. But, despite intense worldwide research programs, still today mostly laboratory level applications are reported and often for scientific purposes only. A number of practical reasons prevent the applications at industrial level: inert native coating of the energetic particles, nonuniform dispersion, aging, excessive viscosity of the slurry propellant, possible limitations in mechanical properties, more demanding safety issues, cost, and so on.This paper describes the main features in terms of performance of solid rocket propellants loaded with nanometals and intends to emphasize the unique properties or operating conditions made possible by the addition of the nano-sized energetic ingredients. Steady and unsteady combustion regimes are examined.

关键词： Nanoaluminum Solid rocket propellant Burning rate Combustion Propulsion Performance

Numerical investigation of transcritical liquid film cooling in a methane/oxygen rocket engine

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航空动力学报 2011年第4期26卷 903-916页

作者： YANG Wei SUN Bing School of Astronautics Beijing University of Aeronautics and Astronautic Beijing 100191 China

Transcritical film cooling was investigated by numerical study in a methane cooled methane/oxygen rocket engine.The respective time-averaged Navier-Stokes equations have been solved for the compressible steady three-dimensional(3-D) flow.The flow field computations were performed using the semi-implicit method for pressure linked equation(SIMPLE) algorithm on several blocks of nonuniform collocated grid.The calculation was conducted over a pressure range of 202 650.0 Pa to 1.2×107 Pa and a temperature range of 120.0 K to 3 568.0 K.Twenty-nine different cases were simulated to calculate the impact of different factors.The results show that mass flow rate,length,diameter,number and diffused or convergence of film jet channel,injection angle and jet array arrangements have great impact on transcritical film cooling effectiveness.Furthermore,shape of the jet holes and jet and crossflow turbulence also affect the wall temperature distribution.Two rows of film arranged in different axial angles and staggered arrangement were proposed as new liquid film arrangement.Different radial angles have impact on the film cooling effectiveness in two row-jets cooled cases.The case of in-line and staggered arrangement are almost the same in the region before the second row of jets,but a staggered arrangement has a higher film cooling effectiveness from the second row of jets.

关键词： liquid film cooling numerical study rocket transcritical semi-implicit method for pressure linked equation(SIMPLE) film cooling effectiveness

Numerical study on similarity of plume infrared radiation between reduced-scale solid rocket motors

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Chinese Journal of Aeronautics 2016年第4期29卷 924-933页

作者： Zhang Xiaoying Chen Huandong Sino-French Institute of Nuclear Engineering and Technology Sun Yat-sen University School of Electricity South China University of Technology

This study seeks to determine the similarities in plume radiation between reduced and full-scale solid rocket models in ground test conditions through investigation of flow and radiation for a series of scale ratios ranging from 0.1 to 1. The radiative transfer equation(RTE) considering gas and particle radiation in a non-uniform plume has been adopted and solved by the finite volume method(FVM) to compute the three dimensional, spectral and directional radiation of a plume in the infrared waveband 2–6 μm. Conditions at wavelengths 2.7 μm and 4.3 μm are discussed in detail, and ratios of plume radiation for reduced-scale through full-scale models are examined. This work shows that, with increasing scale ratio of a computed rocket motor, area of the hightemperature core increases as a 2 power function of the scale ratio, and the radiation intensity of the plume increases with 2–2.5 power of the scale ratio. The infrared radiation of plume gases shows a strong spectral dependency, while that of Al_2O_3 particles shows spectral continuity of gray media.Spectral radiation intensity of a computed solid rocket plume's high temperature core increases significantly in peak radiation spectra of plume gases CO and CO_2. Al_2O_3 particles are the major radiation component in a rocket plume. There is good similarity between contours of plume spectral radiance from different scale models of computed rockets, and there are two peak spectra of radiation intensity at wavebands 2.7–3.0 μm and 4.2–4.6 μm. Directed radiation intensity of the entire plume volume will rise with increasing elevation angle.

关键词： Infrared radiation Plume Reduced-scale model Similarity Solid rocket motor

Thermal state calculation of chamber in small thrust liquid rocket engine for steady state pulsed mode

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Chinese Journal of Aeronautics 2019年第2期32卷 253-262页

作者： Alexey Gennadievich VOROBYEV Svatlana Sergeevna VOROBYEVA Lihui ZHANG Evgeniy Nikolaevich BELIAEV Department of rocket Engines Moscow Aviation Institute School of Astronautics Beihang University

This paper presents a method of thermal state calculation of combustion chamber in small thrust liquid rocket engine. The goal is to predict the thermal state of chamber wall by using basic parameters of engine: thrust level, propellants, chamber pressure, injection pattern, film cooling parameters, material of wall and their coating, etc. The difficulties in modeling the startup and shutdown processes of thrusters lie in the fact that there are the conjugated physical processes occurring at various parameters for non-design conditions. A mathematical model to predict the thermal state of the combustion chamber for different engine operation modes is developed. To simulate the startup and shutdown processes, a quasi-steady approach is applied by replacing the transient process with time-variant operating parameters of steady-state processes. The mathematical model is based on several principles and data commonly used for heat transfer modeling: geometry of flow part, gas dynamics of flow, thermodynamics of propellants and combustion spices, convective and radiation heat flows, conjugated heat transfer between hot gas and wall, and transient approach for calculation of thermal state of construction. Calculations of the thermal state of the combustion chamber in single-turn-on mode show good convergence with the experimental results. The results of pulsed modes indicate a large temperature gradient on the internal wall surface of the chamber between pulses and the thermal state of the wall strongly depends on the pulse duration and the interval.

关键词： Combustion chamber Film cooling Mathematical model Nonstationary thermal mode Small thrust liquid rocket engine Steady pulse mode Thermal state

Improvement in Time Efficiency in Numerical Simulation for Solid Propellant rocket Motors(SPRM)

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推进技术 2018年第1期39卷 92-99页

作者： Valeriy BUCHARSKYI ZHANG Li-hui WAN Yi-lun Propulsion System Design Dept Dnipro National University Dnipro 49000 Ukraine School of Astronautics Beihang University Beijing 100191 China

The main purpose of the present work is to study the possibilities of reducing calculation time while maintaining the validity in the numerical simulation of the combustion product flow in SPRM chamber.Three ways of decreasing the calculation time-the use of numerical methods of high accuracy order,the reduction in spatial dimension of the problem,and the use of physical features of the processes in SPRM chamber while constructing a calculation model-were considered.Presented calculation data show that the use of these approaches makes it possible to reduce the time for solving the problems of SPRM simulation significantly(up to 100times).Also conclusions about the applicability of the mentioned above approaches in SPRM design were made.

关键词： Numerical simulation Solid propellant rocket motor Calculation time reduction

Overview of Physical and Chemical, Operational Properties of Nitrous Oxide Used as a Propellant for Low-thrust rocket Engines

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火炸药学报 2019年第1期42卷 17-31,I0001页

作者： A.A.Boryaev Saint Petersburg State University of Architecture and Civil Engineering

The results of a system analysis of the efficiency of nitrous oxide(N_2O) as a propellant component for small space vehicles(SSV) were presented. A criterion for mass efficiency of the SSV propulsion system(PS) is determined. The current global state-of-the-art of SSV PSs is shown. The application field of nitrous oxide in SSV PSs is calculated and mass efficiency of N_2O application is quantitatively determined. An overview of physical and chemical as well as operational properties of nitrous oxide as a promising, non-toxic component of rocket propellant is provided. Main physical and chemical constants of gaseous and liquid nitrous oxide; chemical properties of N_2O, thermal stability of N_2O, catalytic decomposition of N_2O, a mechanism of decomposition of N_2O, catalysts for decomposition of N_2O, ballast additives to N_2O, application of nitrous oxide, nitrous oxide as a rocket propellant, production of nitrous oxide, toxicity of nitrous oxide, fire hazard of N_2O, requirements to equipment when handling N_2O; storage and transportation of N_2O are considered. It is demonstrated that nitrous oxide is a chemical compound meeting the requirements to rocket propellants, including those related to the environmental friendliness of propellants. With 75 references.

关键词： rocket propellant nitrous oxide physical and chemical properties

Internal ballistic simulation of multi-burning-rate solid rocket motor based on parameterized feature CAD modelInternal ballistic simulation of multi-burning-rate solid rocket motor based on parameterized feature CAD model

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航空动力学报 2015年第12期30卷 3017-3028页

作者： CHEN Wei LIANG Guo-zhu School of Astronautics Beijing University of Aeronautics and Astronautics

Internal ballistic simulation(IBS)method of multi-burning-rate solid rocket motor(SRM)was developed based on 3-D burning regression method by parameterized feature CAD model(PFCADM)and lumped parameter,in consideration of time-dependent,erosive-burning-effect from internal ballistic numerical algorithm.By driving multi-parameter CAD model based on PFCADM,the approach is capable of conducting the geometric regression simulation of various grain combinations of complex configurations with different burning rates.Through suitably simplifying the internal ballistic numerical algorithm,the problems of coupling geometric regression simulation of sub-grains of different burning rates and high computational consumption of internal ballistic calculation were solved.One tri-burning-rate grain motor,which had been firing-tested,was used as the validation case of simulation.The results show that,with the 3-D grain regression model and sufficient accurate internal ballistic algorithm,the method realizes IBS of the case in low computationalconsumption prediction of its performance within the accuracy of 2% during 1hclock-time.The application of the method provides a practical approach to aid SRM design of multi-burning-rate grain.

关键词： solid rocket motor internal ballistic simulatio multi-burning-rate grain design unsteady flow

Development of Kabila rocket: A radioisotope heated thermionic plasma rocket engine

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Chinese Journal of Aeronautics 2015年第2期28卷 427-433页

作者： Kalomba Mboyi Ren Junxue Liu Yu School of Astronautics Beihang University

A new type of plasma rocket engine, the Kabila rocket, using a radioisotope heated thermionic heating chamber instead of a conventional combustion chamber or catalyst bed is intro- duced and it achieves specific impulses similar to the ones of conventional solid and bipropellant rockets. Curium-244 is chosen as a radioisotope heat source and a thermal reductive layer is also used to obtain precise thermionic emissions. The self-sufficiency principle is applied by simultane- ously heating up the emitting material with the radioisotope decay heat and by powering the differ- ent valves of the plasma rocket engine with the same radioisotope decay heat using a radioisotope thermoelectric generator. This rocket engine is then benchmarked against a 1 N hydrazine thruster configuration operated on one of the Pleiades-HR-1 constellation spacecraft. A maximal specific impulse and power saving of respectively 529 s and 32% are achieved with helium as propellant. Its advantages are its power saving capability, high specific impulses and simultaneous ease of storage and restart. It can however be extremely voluminous and potentially hazardous. The Kabila rocket is found to bring great benefits to the existing spacecraft and further research should optimize its geometric characteristics and investigate the physical principals of its operation.

关键词： Plasma Radioisotopes rocket Self-sufficiency principleThermionic Thermoelectric power