1990 Vol. 5, No. 2

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NUMERICAL SOLUTION OF 3-D HYBRID PROBLEM IN TURBOMACHINERY
Chen Naixing, Dong Ming, Zhang Jialin
1990, 5(2): 97-102,185.
Abstract PDF(394KB) icon1484 icon456
Abstract:
A new numerical method for solving aerodynamic hybrid problem of fully three dimensional(3-D)flow in turbomachinery is presented.This method is based on non orthogonal coordinate system and contravariant velocity components and employs the approximate factorization multigrid scheme in conservative form.It has been proved that this scheme improves the numerical computation convergence speed and the computational precision. In the present hybrid solution method,the meridional positions(z,r)at all grid points are kept unchanged in the proecss of iterations.
EXPERIMENTAL INVESTIGATION ON VORTICAL TYPE FLOW IN COMPRESSOR CASCADES
Tang Yanping, Chen Fang, Chen Maozhang
1990, 5(2): 103-112,185-186.
Abstract PDF(635KB) icon1598 icon434
Abstract:
A series of experiments to investigate the vortical flow in compressor cascades have been completed with a measuring and visualizing system.The phenomena observed include the gereration of corner vortexes and the interaction between corner vortexes and boundary layers on the blade suction surface in the compressor cascades.Based on these expermental results,the flow topology and vortex structure related to the flow pattern have been deduced at moderate incidences.However,at high incidences some important changes for the flow pattern appear and vortex breakdown and reversed flow represent the main characteristics of the vortical flow.Furthermore,the dynamic effects in the compressor cascade,such as the effects of the vortical flow on loading and losses, have been investigated.The results obtained are as follows: (1)At moderate incidences,a high loss region in the corner appears along with the formation of the corner vortex system.The principal loss mechanism comprises:(a)the three dimensional seperation of the corner vortex and secondary vortexes;(b)mixing of fluids with low and high momentum in the corner vortex system. (2)At high inidences,the reversed flow caused by vortex breakdown and strong adverse pressure gradient makes an important contribution to the loss increase. (3)At high incidences,the excessive increase of loading near the end wall appears mainly by the front part of the blade section. From the above mentioned research a physical model for the vortical flow in compressor cascades is obtained.
APPLIEATION OF LEANED BLADES TO AN AEROENGINE
Gao Jingman
1990, 5(2): 113-117,186.
Abstract PDF(324KB) icon1870 icon508
Abstract:
In order to match the low pressure compressor with the high pressure one,the outlet guide vane blades of the low preseure compressor have been designeed to be leaned circumferentially and radially.The two types of the leaned blades have been tested in an aeroengine.The results show that the leaned blades obviously improve the compressor performance in the hub region,decrease the radial pressure ratio gradient and control the flow at the end wall.
THE INFLUENCE OF THE INLET MACH NUMBER ON THE BOUNDARY LAYER DEVELOPMENT ON TURBOMACHINERY BLADE SURFACES
Zhu Nianguo, Xu Liping
1990, 5(2): 118-122,186.
Abstract PDF(331KB) icon1689 icon491
Abstract:
When Reynolds number and the blade surface pressure coefficient distribution are kept constant,the influence of the inlet Mach number on the boundary layer development on the blade surfaces of turbomachinery is discussed with numerical simulation as well as an order of magnitude analysis,based on the integral boundary layer equation.The effects of surface curvature and the free stream turbulence intensity on the development of the boundary layer are also estimated by the numerical simulation.The results show that at the same Reynolds numbers and surface pressure coefficient distributions,the variation of the non dimensional the integral parameters boundary layer with the inlet Mach number is very small,up to the critical Mach number,the changes are usually within the engineering tolerance.The variation caused by the surface curvature difference is of the same order of that due to Mach number,however the free stream turbulence level can cause remarkable changes in the development of the boundary layer.
STUDY ON ATOMIZATION MODEL AND COMPUTATION OF PREFILMING AIRBLAST ATOMIZER
Gan Xiachua, Zhao Qishou
1990, 5(2): 123-128,187.
Abstract PDF(920KB) icon1662 icon465
Abstract:
An a tomization model for frefilming airblast atomizer is formulated on the basis of analysis and experiments,and put in to numerical computation.A breakup criterion of liquid film is provided to judge the film disintegration extent.Quantitative relations between the wavelength frequency spectrum of the film wave and its correspondingly formed droplet size number are consi dered.A Trace Statistical method is presented to perform the model computation.A numerical example of atomization of a three dimernsional axisymmetric atomizer has been completed.The consistence of the computational results with the experimental data demonstrates that this model is feasible.
EXPERIMENTAL STUDY OF THE EFFECT OF DENSE SPRAY ON DROP SIZE MEASUREMENT BY LIGHT SCATTERING TECHNOLOGY
Li Weiming, Zhang Yan, Jin Rushan
1990, 5(2): 129-134,187.
Abstract PDF(387KB) icon1798 icon459
Abstract:
The effect of dense spray on drop size measurement by light scattering technology was studied by using Malverm instrument with five duplicated internal mixing airblast atomizers aligned with laser beam.The correction factor data for multiple scattering were obtained.By regression analysis,an empirical equation was obtained which correlated the correction factor as a function of obscuration(OBS),Sauter Mean Diameter under dulite spray condition SMDo and drop size distribution parameter for Rosin Rammler distribution under dilute spray condition No. The experimental data showed definitely that the correction factor is not only a function of OBS,SMDo,as proposed by Lee Dodge,but also is a function of No.The correlation fits the experimental data very well,and can be used for practical purposes to correct the data from Malvern drop sizer at high obscuration conditions.
THE MECHANISM OF LIQUID-COOLED VANE FOR HIGH-TEMPERATURE TURBINE
Liu Dezhang, Yuan Baoning
1990, 5(2): 135-140,187-188.
Abstract PDF(449KB) icon1681 icon390
Abstract:
In this study the flow pattern and heat exchange coefficient the have been obtained,and relation between them have been determined in the circumstance of vapour lquid two phase boiling of water coolant through a even heated straight tube.The influences of the coolant flow parameters on the flow pattern and heat exchange have been studied by changing the coolant flow speed,or the flowrate at the entrance under certain heating conditions.As a result,the best flow pattern for the heat exchange coefficient in which the annular flow zone is the longest without “burn down”.Then the empirical formulas have been found for designing.The results of numerical calculation are in good agreement with the data of tests.
MEASURING TECHNIQUE FOR INFRARED RADIATION OF METALLIC SURFACES
Ren Linghua, Zhen Jirui, Lei Ying
1990, 5(2): 141-144,188.
Abstract PDF(285KB) icon1581 icon397
Abstract:
A testing principle,method and apparatus for measuring infrared radiative energy of metallic surfaces are presented.The infrared radiation characteristics of irons and stainless steels have been investigated experimentally.The effects of the temperature on the infrared radiative energy of metal materials are considered.The experimental results illustrate that this testing technique is applicable and the measuring apparatus is reliable.
EXPERIMENTAL STUDY OF HEAT TRANSFER IN TURBULENT SEPARATED FLOW OVER A FORWARD-FACING STEP
Duan Xinran, A.Baron, F.K.Tsou, W.Aung
1990, 5(2): 145-149,188.
Abstract PDF(1653KB) icon1531 icon405
Abstract:
An experimental study of heat transfer in turbulent separated flow over a forward facing step is reported in this paper.The present work deals with a quite small forward facing step model (step height H=4.0mm),therefore a special technique that can change the relative locations of heat flux gauge from the step is applied successfully. The investigation is focused on two recirculation zones.Measurements of heat transfer are performed in Ludwieg Wind Tunnel by the transient technique,which demonstrates a new method of experimental research in this field.Within the range of Reynolds number 11800 to 32200,the distribution of wall heat flux and the distribution of Nusselt number are obtained.Those distributions show that the heat transfer coefficient increases to the first maximum value near the step in the front recirculation zone and the second maximum occurs at the reattachment point in the downstream recirculation zone where this point is fixed roughly at X/H=2.5 for the five different Reynolds numbers.The second maximum is larger than the first for same Reynolds number.At last a correlative equation between Reynolds and Nusselt numbers is presented.
NONUNIFORM EROSIVE BURNING IN STAR-CORE GRAIN IN A SOLID-PROPELLANT ROCKET ENGINE
Xu Shaomin, Li Fengliang
1990, 5(2): 150-154,189.
Abstract PDF(342KB) icon1749 icon484
Abstract:
The experimental investigation of erosive burning in one lobe of the star core grain in a solid propellant rocket engine was carried out.The experimental results show that the erosive burning rate in the star recess is higher than that at the star point.The nonuniformity of erosive burning is dependent on the average flow velocity parallel to the burning surface,the chamber velocity parallel to the burning surface,the chamber pressure and the geometry of the cross section.From comparison between the calculational results and experimental data it is concluded that the nonuniformity of the erosive burning comes mainly from the nonuniformity of the velocity distribution.
EFFECT OF CATALYSTS ON THERMAL DECOMPOSITION OF HMX/AP/HTPB PROPELLANT
Zhu Hui, Zhang Reng
1990, 5(2): 155-158,189.
Abstract PDF(282KB) icon1772 icon458
Abstract:
The effects of some catalysts on the thermal decomposition of HMX/AP/HTPB propellant are studied with the differential scanning calorimetry(DSC).The experiments demonstrate that the mixture of HMX/AP/=1/1(weight)has specific decomposition characteristics.The catalysts(such as PbO2, Na2S2O8, Fe3O4, CO2O3 and Pb Cr2O4)which are able to increase the burning rate of the propellant,primarily rais6e chemical reaction rate in gaseous phase.
PARTIAL SIMILARITY AND A REAL-TIME MODEL OF TWIN-SPOOL GAS TURBINE
Wang Xueyu, Zhu Xingjian
1990, 5(2): 159-162,189.
Abstract PDF(305KB) icon1694 icon461
Abstract:
Partial similarity of twin spool gas turbine means that the high pressure rotor can be considered as a single spool gas turbine operating under a specifical circumstance,the high and low pressure turbines and exhaust nozzle are considered a single components of a combination.In formulation of a real time gas turbine model,this method reduces dimensions of the dynamic parameters of the model.Therefore,calculation speed and accuracy of the model are good enough for real time modeling.Thus the model is capable of microcomputer digital control of gas turbine and of real time estimation of mal functioned parameters in the gas turbine monitoring system.
AN EXPERIMENTAL INVESTIGATION OF ROTATING STALL IN A CENTRIFUGAL COMPRESSOR WITH VANELESS DIFFUSER
Chu Wuli
1990, 5(2): 163-165,190.
Abstract PDF(225KB) icon1776 icon470
Abstract:
This paper deals with experimental investigation of unsteady flow in a radial vaneless diffuser.The circumferentially uniform discharge at the exit of the diffuser changes into a periodic circumferentially fluctuating flow state.The experimental measurements show that the properties of rotating stall in the diffuser are very obvious.The results of the investigation indicate that two type of rotating velocity patterns occur successively as the flow rate decreases.Their rotational speeds are different and the pattern with low rotational speed firstly appears.
VALUATION OF FLOW-INDUCED FORCE IN LABYRINTH SEAL
Fan Feida
1990, 5(2): 168-168,190.
Abstract PDF(262KB) icon1601 icon400
Abstract:
The continuity equation and momentum equation of the flow through a sectorial element in labyrinth seal is linearized according to the presumption of the small disturbance and solved with the aid of matrix method.Then the distribution of the pressure on the cylindrical surface of rotor is obtained.Finally,the flow induced force on rotor is evaluated by means of integration.Numerical example indicates that as the number of seal chamber increases,the flow induced force increases while the leakage flow rate decreases;when the seal clearance increases,the flow induced force decreases while the leakage flow rate increases;but both the flow induced force and leakage flow rate decrease when the exit pressure increases.
ANALYTICAL FORMULAS FOR LOGARITHMIC MEAN TEMPERATURE DIFFERENCE AND HEAT-EXCHANGE AREA
Wang Feng, Liu Zhiwei, Han Zhenxing
1990, 5(2): 169-170,190.
Abstract PDF(992KB) icon2800 icon409
Abstract:
In order to save energy and metallic material,it is very important to calculate the logarithmic mean temperature difference between the two kinds of fluids and the superficial area of the heat exchanger accurately.The traditional calculational method is too complicated,but the error of simplified method is too significant,so we deduce the formulas for logarithmic mean temperature difference between the cold and hot fluids in a heat exchanger in consideration of their thermophysical properties and the formula for sueperficial area of the exchanger.The formulas are easy to use and give accurate results.
A CALCULATION METHOD FOR INFLAMMABILITY LIMIT OF MIXTURE WITH NONINFLAMMABLE GASES
Su Dawei
1990, 5(2): 171-172,191.
Abstract PDF(213KB) icon1269 icon395
Abstract:
This paper provides a calculation method which can estimate the inflammability limit of mixture with noninflammable gases with more accuracy than the available methods.
MEASUREMENT OF TRANSIENT BURNING RATES OF COMPOSITE PROPELLANTS IN FAST DEPRESSURIZATION
Wang Junye, Sang Bingcheng
1990, 5(2): 173-175,191.
Abstract PDF(234KB) icon1273 icon389
Abstract:
A continuous laser beam is transmitted through the burning surface of a propellant strand and converted in to voltage signal by photocell.The voltage differentiation with respect to time is proportional to the burning rate of the propellant strand.The experimental results show that the burning rate at the start of depressurization remains unchanged,and then decreased with very low frequency oscillation.During the whole depressurization process the transient burning rate is not higher than the quasi steady burning rate at the corresponding pressure and oscillates around quasi steady burning rate,except for a short time after the start of the depressurization.
OPTIMIZATION DESIGN OF THE SOLID ROCKET MOTOR
Fang Guoyao, Guo Zhiyong
1990, 5(2): 176-178,191.
Abstract PDF(241KB) icon1668 icon493
Abstract:
This paper presents the optimization design of a two stage solid rocket motor for a air to air missile.There are eight variables in the optimization process.The mathematical model of the solid rocket motor is formulated.A variable simplex method has been used.Finally,the optimum results for a real solid rocket motor are obtained.
LINEAR DYNAMICS OF SUPERSONIC RAMJET
Qian Boshun, Zeng Qingfu
1990, 5(2): 179-181,191.
Abstract PDF(239KB) icon1534 icon484
Abstract:
This paper presents a mathematical model of a supersonic ramjet.The basic thermoaerodynamical equations of ramjet components are linearized in conformity with small perturbation.Then the transfer functions of the components are deduced on the S domain.With the method of subsystem in series,the dynamical mathematical model of the ramjet with the inlet normal shock position,the parameters along duct and thrust is formulated under the perturbation of upstream Mach number,temperature,pressure and downstream fuel flow rate of the ramjet.A numerieal example of digital imitation of a real ramjet is completed and gives reasonable results.
SELF-TUNING CONTROL OF VARIABLE THRUST LIQUID PROPELLANT ROCKET ENGINE
Zhang Yulin, Lu Zhenling, Chen Qizhi
1990, 5(2): 182-184,192.
Abstract PDF(242KB) icon1669 icon465
Abstract:
A self tuning controller is constructedfor a liquid propellant rocket engine with variable thrust.To verify the effectiveness of the control method,the digital simulation of controlling the variable thrust liquid propellant rocket engine with this controller is presented.The results show that the control quality of the engine system is greatly improved by the self tuning controller.

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