The 31st Research Institute of CASIC

Beijing, China

The 31st Research Institute of CASIC

Beijing, China

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Li S.,The 31st Research Institute of CASIC | Zheng R.-H.,The 31st Research Institute of CASIC | Zheng R.-H.,Beijing Institute of Machinery | Ma G.,The 31st Research Institute of CASIC | Qi Z.-M.,The 31st Research Institute of CASIC
Tuijin Jishu/Journal of Propulsion Technology | Year: 2017

For studying the sealing performance of the gas pressure system used in ramjet fuel tank after long term storage, leakage rate of varied sealing structures was theoretically calculated on the basis of conductance calculating formula, to study its sealing mechanisms and impacts of various negative factors. The loss pressure recurrence formula was deduced from the definition of the leakage rate to evalauate the sealing performance after long stroage. The results, which agree well with the experimental data, show that the pressure will lose no more than 0.15‰ from the initial value after 10 years. It demonstrates that the sealing performace of the current structures satisfy the requirements. © 2017, Editorial Department of Journal of Propulsion Technology. All right reserved.


Zhao W.,The 31st Research Institute of CASIC | Liu Z.-D.,The 31st Research Institute of CASIC | Li S.-B.,Beihang University
Tuijin Jishu/Journal of Propulsion Technology | Year: 2014

In order to reveal the unsteady flow mechanism of compressor blade boundary layer under rotor/stator aerodynamic interaction, numerical simulation was applied to investigate the unsteady boundary layer flow of the stator at the mid-span in a high loading transonic fan stage. Critical parameters, including skin friction, turbulence kinetic energy and static pressure fluctuation on the blade surface, were analyzed in detail to describe the unsteady effects of the high loaded stator boundary layer based on the model of rotor wake/potential-stator boundary layer interaction. The numerical results show that the bypass transition would occur in the stator boundary layer during the impinging of the rotor wake. And also, a pressure fluctuation would propagate with the speed of sound along the blade boundary layer after the leading edge was impinged by the wake. The distribution of the pressure and friction along the whole pressure side and the leading edge of the suction side were affected by the pressure fluctuation propagation. The pressure fluctuation effect was determined by the pressure gradient and the blade surface shape.


Liu Y.-X.,Beijing Institute of Technology | Nie L.-C.,The 31st Research Institute of CASIC | Zhang J.,Beijing Institute of Technology | Yang X.,Beijing Institute of Technology
Tuijin Jishu/Journal of Propulsion Technology | Year: 2015

In order to improve effects of gas generator flow regulation, a pressure closed loop control system was built. The working principle was analyzed, and the dynamic model of the system was established. According to the characteristics of the system model, linear active disturbance rejection controller (LADRC) was designed to control it. Because of the extended state observer, LADRC can compensate the effect caused by parameter fluctuation on the output well in a larger range. It improves the speed of response and the control accuracy. Simulation results in different working conditions show that LADRC shortens 30% adjustment time, and reduces about 50% output caused by interference. It has better dynamic characteristics than the traditional controller. © 2015, Journal of Propulsion Technology. All right reserved.


Gao X.,The 31st Research Institute of CASIC | Li D.-J.,Beijing Institute of Technology | Zhu S.-M.,Beijing Institute of Technology | Man Y.-J.,Beijing Institute of Technology
Tuijin Jishu/Journal of Propulsion Technology | Year: 2013

Different leading edge shock on hypersonic curved ramp two-dimensional inlet directly influences the inlet performance, specially, the flow coefficient. A detailed analysis and investigation was performed on the variation characteristics of leading edge shock generated by two different inlets response to the variation of inlet attack angle and inflow Mach number. The calculation results reveal that for wedge-isentropic two-dimensional inlet, the angle between the leading edge shock and compression surfaces decreases in advance, then increases with increase of the inlet attack angle. However, the angle always decreases with the increase of the free stream Mach number, whereas, for two-dimensional hypersonic curved shock compression inlet with the law of constant pressure gradient, the angle between the leading edge shock and compression surfaces decreases with increase of both the attack angle and the free stream Mach number. According to theory and calculation analysis, different oblique shock and Mach wave interaction mode causes different characters of leading edge shock generated by two different inlets response to the variation of inlet attack angle and inflow Mach number.


Nie L.-C.,The 31st Research Institute of CASIC | Liu Z.-M.,The 31st Research Institute of CASIC | Liu Y.-X.,Beijing Institute of Technology
Tuijin Jishu/Journal of Propulsion Technology | Year: 2013

In order to improve the accuracy of gas generator flow regulation for solid fuel ramjet, the pressure close loop control was designed. To reduce the pressure overshoot and to ensure the consistency of pressure changes, fuzzy integral control algorithm was designed. Simulation and experimental results show that this method has less pressure overshoot, faster response and better dynamic characteristic in different working condition.


Shang X.-S.,Navel Academy of Armament | Wu Z.-F.,The 31st Research Institute of CASIC | He Y.-F.,CPLA 8.1 Parachute Brigade
Tuijin Jishu/Journal of Propulsion Technology | Year: 2013

Solid rocket motor underwater test parameters change dramatically and signals often overlap severely. In order to solve the problem arisen due to this in data processing, the denoising method based on soft threshold wavelet transform was proposed and the analysis and processing of data for testing the thrust of solid rocket motor under water was performed. Comparison with conventional method shows that this method can divid dynamic data into the approximate component to reflect 'low frequency' characteristics and the detail components to reflect the 'high frequency' characteristics. This method is convenient for analysis of the motor performance test data. Variation characteristics of the data can be well retained in the simultaneous removal of test data interference. It can meet the requirements of the engineering application for data processing in solid rocket motor underwater test.


Yang S.,Wuhan Naval University of Engineering | Sun K.-Q.,The 31st Research Institute of CASIC
Tuijin Jishu/Journal of Propulsion Technology | Year: 2012

A space-time conservation element and solution element method(CE/SE) was used to simulate the formation and propagation process of two-dimensional detonation ignited by hotspot, and analyze the detonation mechanism and the factors to hotspot ignition. For the hotspot ignition, the initial pressure was 0.1 MPa and temperature was different. Implicit trapezoidal technique was established to handle stiff source term. Results show that the uniform temperature and linear temperature distribution of hotspot have different detonation mechanism, which is effected by the varied parameters in ignition zone. The large region size, the linear temperature distribution or the shape of hemisphere for the hotspot could induce detonation easily. The results offer reference to study detonation ignition.


Chen Y.-C.,Northwestern Polytechnical University | Liu X.-Y.,The 31st Research Institute of CASIC | Huang X.,Northwestern Polytechnical University | Ma X.-S.,The 31st Research Institute of CASIC | Huang X.-L.,The 31st Research Institute of CASIC
Tuijin Jishu/Journal of Propulsion Technology | Year: 2012

For overcoming the disadvantages of the model based on differential equations for scramjet performance computation, a model based on lumped parameter method was presented. Intergral equations of conservation of mass, momentum, and energy of 0D control valume coupled with chemical kinetics model, combined with critical mass flow method, were used to compute the 1D flow field and to balance the mass flow of isolator and scramjet combustor in various operation modes. The thermal throat of scramjet was captured, operation states of isolator and operation modes of combustor were found, consequently the charactristcs of scramjet were computed at various states in flight envelope. Computational results show that this model can not only overcome the disadvantages of the model based on differential equations, but also has advantage of high computation precision, rapid computation speed, and good convergency. The new model provides a good reference for performance simulation.


Sun R.-P.,Harbin Engineering University | Zhu W.-B.,Harbin Engineering University | Xu L.-Z.,The 31st Research Institute of CASIC | Guo H.-Y.,The 31st Research Institute of CASIC | Chen H.,Harbin Engineering University
Tuijin Jishu/Journal of Propulsion Technology | Year: 2012

NASA-MarkII transonic blade was selected as an example in this paper to investigate the application of the Transition k-kl-ω model in gas thermal coupling calculation of an innercooled blade, and to analysis the difference between the whole coupled method and heat transfer coefficient criteria coupled method in cooling channels. Compared with other turbulence models, the transition model is capable of predicting the boundary layer of laminar flow and the transition state more accurately. In the Transition k-kl-ω model, laminar kinetic energy was used to describe the development of disturbance in the early stage so as to avoid the empirical formula which includes flow turbulence intensity, and the "split system" was introduced into this model to describe the interaction between laminar and turbulent fluctuation. Besides, Tollmien-Schlichting wave was added in the bypass transition and natural transition source term, as a result, the computational temperature results for behind the strong shock wave agreeed better with the experimental value than that obtained with intermittent transition factor model. Using the coupled heat transfer coefficient criteria in the cooling channel heat transfer calculation, which removes the error caused by cooling channel boundary conditions, leads to better agreement with the experimental results and easier engineering applications.


Liu G.-D.,The 31st Research Institute of CASIC | Yu S.-Z.,The 31st Research Institute of CASIC | Liu F.-J.,The 31st Research Institute of CASIC
Tuijin Jishu/Journal of Propulsion Technology | Year: 2014

Determining the thermal throat in the scramjet combustor is one of the key steps to predict the overall performance of the scramjet engine, and thus accurately locating the thermal throat is of great importance in the scramjet performance evaluation. A new approach to predict the thermal throat is presented. The prediction of the thermal throat starts with the combustor exit, and then proceeds upstream by assuming Mach number as unity at each cross-section. The equations of mass flow conservation, energy conservation and gasdynamic function in terms of pressure are solved simultaneously to obtain all one-dimentional gasdynamic parameters at the cross-section. Besides, the critical gradient of combustion efficiency is computed using an analytic method, and subsequently the judgement condition for the thermal throat is given. Then the thermal throat position can be located. The new approach was applied to the scramjet model in the flight conditions from Ma=3.5 to 6. Comparing with the results of the traditional method, it shows that the new method can quickly calculate the thermal throat with good feasibility and the error between the traditional methods is less than 6%.

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