1997 Vol. 12, No. 1

  Previous Issue | Next Issue 
Display Method:
A STUDY ON MECHANISM OF ACOUSTIC CONTROL OVER AIRFOIL FLOW IN UNIFORM INCOMING FLOW
Zhong Fangyuan, Dong Guoping, Ju Hongbin
1997, 12(1): 1-5,103.
Abstract PDF(802KB) icon1643 icon396
Abstract:
An experimental study on the acoustic control of separated flow on a 2-dimensional airfoil at low speed ( Re =3—6×105) and high angle of attack was completed at the Turbine  Aero-Acoustics Laboratory,SJTU.Based on the comparison of test results with or without the acoustic control in uniform incoming flow,the benefits of the acoustic control (improvement of lift and drag coefficients),the effects of the parameters of exciting sound (sound pressure level SPL and frequency fs ) and the variation of flow parameters in the boundary layer on the suction side of airfoil (mean flow velocity,turbulence and the spectrum of fluctuating velocity) have been investigated,and some valuable conclusions have been obtained for exploring the mechanism and exploiting the acoustic control.Two threshold values of SPL for the exciting sound have been defined.There is only one optimum frequency,corresponding to the fundamental frequency of acoustic cross-resonance of the testing tube,at which the SPL of exciting sound and the enhancement of lift coefficient will approach maximum value simultaneously.It is evident that the optimum frequency of exciting sound depends not only on the cross-resonance of the testing tube,but also on the dimension of Large Scale Orgnized Structure Vorticity (LSOSV) induced by the exciting sound.Finally,a tentative idea on the non-linear analysis of sound-turbulence interaction in the separated boundary layer has been proposed.
AN EXPERIMENTAL INVESTIGATION ON FLUTTER OF A HELICOPTER BLADE MODEL
Xia Pinqi, Xu Guiqi
1997, 12(1): 6-8,103.
Abstract PDF(234KB) icon1873 icon411
Abstract:
An experimental technique and method for flutter research of helicopter blade model are proposed.In the result of a series of experiments the relationships of blade flutter margin with rotor rotational speed,effective CG position of blade in chord and torsional stiffness of blade have been obtained.Some important conclusions have been reached.
FLOW MEASUREMENTS AND ANALYSIS OF ROTATING STALL IN AN AXIAL FLOW FAN
Ge Ning
1997, 12(1): 9-12,103-104.
Abstract PDF(976KB) icon1749 icon379
Abstract:
The flow measurements of part-span stall were accomplished in an axial flow fan test.The measurements include laser anemometer measurement of the stall cell in the mean flow with an ensemble averaging technique and instantaneous flow measurement with hot-wire and flow visualisation technique.The mean flow field measurement indicates that the stall cell appears mainly in the front of the rotor,while the flow field in the tip region of the rotor features essentially the cross blockage of the tip clearance flow.The instantaneous flow measurement shows that the kinetics of the stall cell can be described by using two fundamental concepts of dipole.
A 3D AERODYNAMIC STABILITY ACTIVE CONTROL MODEL FOR AXIAL FLOW COMPRESSOR
Cao Renjing, Zhou Sheng
1997, 12(1): 13-16,104.
Abstract PDF(304KB) icon1783 icon382
Abstract:
A new three-dimensional aerodynamic stability model for an axial flow compressor to study the effects of radial disturbance on the rotating stall is presented.The calculated results show that the stalling characteristics of the compressor is closely related with its cascade characteristics including total pressure rise versus the flow rate and inlet flow angle.Based on the 3D aerodynamic stability model,the related active control technique and the open-loop transfer function have been provided,in which the three dimensional flow effects are involved.The response analyses of the open loop systems for the 3D stability active control model with step impulse were conducted in different flow ranges.
INITIAL DISTURBANCE FORMS OF ROTATING STALL IN AXIAL FLOW COMPRESSOR
Li Jun, Xu Gang, Zhu Junqiang, Liu Zhiwei
1997, 12(1): 17-20,104.
Abstract PDF(292KB) icon1956 icon496
Abstract:
The process of rotating stall developing in a single-stage axial-flow compressor was investigated experimentally.By means of eight dynamic pressure transducers equally located circumferentially in the inlet section of the compressor,the dynamic pressure signals of the rotating stall developing process were measured.The data have been analyzed with the SFC (Spatial Fourier Coefficient) method and the variation of the circumferential phase of initial disturbance signals with time have been obtained.So,it was found that there are two initial disturbance forms of rotating stall,rotating waves and small stall cells.In order to study the factors influencing the initial disturbance forms,the compressor was tested under 55 conditions combining four types of compressor configuration,two solidities of rotor,three rotation speeds,different IGV outlet angles .According to the experimental results,the initial disturbance form of single-stage compressor without IGV was rotating waves.For the isolated rotor without IGV,the forms were rotating waves at high rotor speeds,and small stall cells at low rotor speed.For the compressor with IGV,negative prewhirl angles of IGV enhanced the rotating waves,and increased the pretelling time of stall.Decreasing the solidity of rotor almost did not affect the form of stall initial disturbance.
INFLUENCE OF TIP CLEARANCE ON AERODYNAMIC-AEROACOUSTIC PERFORMANCE OF SKEWED-SWEPT ROTATING BLADES
Cai Na, Li Di, Zhong Fangyuan
1997, 12(1): 21-25,105.
Abstract PDF(924KB) icon1680 icon480
Abstract:
Noise characteristics of tip clearance in a fan and engineering aerodynamic-aeroacoustic performance are obtained from skewed-swept rotating blades.Two methods of tip clearance modification have been carried out:(1)Experimental results are applied to decreasing the radial tip clearances of the rotating blades,thereby the aerodynamic efficiency increases about 3-4 percent,the average noise is reduced by 2-3 dB(A),and the steady working margin is widened over 20 percent;(2)Aerodynamic-aeroacoustic characteristics are calculated with a revised formula.
NUMERICAL SIMULATION FOR STIFFNESS OF FLEXIBLE WAVY RING SUPPORT
Liu Fangjie, Liu Rongqiang, Dai Xingjian
1997, 12(1): 26-28,105.
Abstract PDF(253KB) icon1842 icon453
Abstract:
The numerical simulation for the stiffness of a flexible wavy ring support is presented.The relationship that connects the supprot stiffness of the wavy ring to that of the ring's separated arc segment is found out.A stiffness analysis model of the segment is set up and the displacement field of the segment under a normal load is solved by using the finite element program "APOLANS".The stiffness analysis results are compared with the test data obtained from stiffness measurement in a factory in good agreement.This study provides an analytical method for the design of flexible wavy ring support.
AN EXPERIMENTAL INVESTIGATION ON VIBRATION OF A ROTOR WITH THERMAL BOW
Hu Bigang, Ren Pingzhen, Feng Guoquan
1997, 12(1): 29-32,105.
Abstract PDF(636KB) icon1700 icon422
Abstract:
In order to investigate the vibrational response of a turbine rotor with thermal bow,a test rig simulating the rotor was set up,on which the temperature field and the thermal deformation of the rotor were measured.The critical speed of the rotor and response to the thermal bow were investigated.The experimental data are in good agreement with results caculated by means of the transfer-matrix method.
A COMPARATIVE TEST APPROACH FOR THERMAL FATIGUE LIFE OF HIGH-HEAT BLADE UNDER THERMAL SHOCK
Peng Zhiyong, Lu Wenlin
1997, 12(1): 33-36,106.
Abstract PDF(543KB) icon2138 icon463
Abstract:
A new comparative test approach is provided for the thermal fatigue life experiments of turbine blade under thermal shock.The tested blades and reference blades were mounted in a common vane cascade.Moreover,the position of these two kind of blade will be swapped each other regularly during the experiment.A group of comparative tests has been completed by this way for turbine nozzle blades made from analogue materials .The test results show that the approach is reliable,and the test cost is cut down greatly.
MULTISCALE EIGENSYSTEM ASSIGNMENT FOR VIBRATION CONTROL OF FLEXIBLE STRUCTURES
He Erming, Chen Xinhai
1997, 12(1): 37-39,106.
Abstract PDF(258KB) icon1590 icon392
Abstract:
A method of multiscale eigensystem assignment is presented for vibration suppression of flexible structures.The method can alter the original modes by appropriate feedback forces,so that the sensitive parts of the flexible structure reach their steady states at faster rates than other parts,while vibration in sensitive parts is reduced to allowable level.The new control scheme can quickly remove the vibration energy from the sensitive parts of the flexible structure to meet the requirements of high performance.By selecting a model transformation matrix T in Eq(1),the vibration energy of the controlled system is redistributed over the sensitive and insensitive parts.The resulting active damping conrol gain is K2=(T-1 )TG2T-1,where the elements of G2 are solved from the ideal characteristic equations (5),in which ( n2-3n+2)/2 elements of G2 are determined artificially.Finally,a simple example demonstrates the effectiveness of this new control scheme.
STRUCTURAL OPTIMIZATION OF AN INTEGRATED COMPRESSOR ROTOR
Lei Xianhua, Jiang Hefu, Wang Lushang
1997, 12(1): 40-42,106.
Abstract PDF(337KB) icon1749 icon472
Abstract:
An optimization model in the integrated compressor rotor of gas turbine engine is built up.The minimum weight is selected as objective function.The stress is analysed by Axisymmetric Cyclic Thermal Elastic-Plastic-Creep Finite Element Method.The structural size,local static strength, local low-cycle-fatigue life,and disc burst speed are used as constraints.At the same time, the configuration method, automatic meshing,mesh checking,sensitivity analysis, and post-processing interface are studied.Finally,an Integrated Compressor Rotor Optimal Design Program (ICRODP) is worked out.A rotor has been analysed by this program with success. The weight is decreased significantly to meet the requirements of the strength and the low-cycle-fatigue life for the rotor.
STABLE UNBALANCED RESPONSE OF MULTIROTOR WITH SFD AND TAPERED GEAR ENGAGING
Ren Pingzhen, Yanghaiyan
1997, 12(1): 43-45,107.
Abstract PDF(253KB) icon1645 icon442
Abstract:
The paper deals with a method to calculate the stable-unbalanced response of a multirotor system,supported on Square Film Damper (SFD) and engaged with the tapered gears,to bending and torsioning vibrations.The influence of SFD and engaging gear of an abnormal structure engine rotor on the stable-unbalanced response of bending and torsion coupling has been studied.A transfer matrix is deduced which is able to determine the gear's unstable response to engaging forces (in 3 directions) and torque.The transfer matrix couples bend and torsion between not only tapered gears,but also abnormal structure and engine shaft.A formula is set up for calculating the potential energy,kinetic energy and energy distribution in the multirotor system.A general computer program is worked out to calculate the engery and the stable-unbalanced response of the multirotor with the tapered gear and SFD.
A SIMPLE ALGORITHM FOR VISCOUS COMPRESSIBLE FLOWS
Zhang Moujin, Wong C.M.and, Pomfret M.J.
1997, 12(1): 46-50,107.
Abstract PDF(792KB) icon1779 icon380
Abstract:
Based on the SIMPLE algorithm,a three-dimensional pressure-based finite volume calculation procedure for compressible flows in arbitrary-shaped domains is presented.The pressure is evaluated by the compressible form of the pressure-correction equation of SIMPLE algorithm.The present method is used to compute a number of cases ranging from subsonic flows to supersonic flows.The computed results are in good agreement with the available published experimental and computed data.
DOWNSTREAM FLOW AND HEAT TRANSFER OF A FILM COOLING FLARED HOLE INCLINED AT 60°
Xu Hongzhou, Liu Songling, Xu Duchun
1997, 12(1): 51-55,107.
Abstract PDF(819KB) icon1934 icon456
Abstract:
The flow field and heat transfer downstream from a film cooling flared hole has been investigated experimentally.The jet hole with a flared angle of 15° or 30° was oriented so that its streamwise injection angle with respect to the test surface is 60°.A delicate five-hole probe and thermocouple were applied to measurement of the three-dimensional flow field and the heat foil method was used to obtain surface heat transfer distribution.Experiments indicate that the longitudinal velocity contours in the downstream boundary layer of the flared exit have three basic forms.With increasing area of the jet exit,the flared hole increases the spreading of the jet and decreases the strength of the longitudinal vortex appreciably,in result the film cooling effectiveness much greater than that of a circular hole.For the high blowing ratio and big flared angle,the flared hole jet still produces a successful heat protection downstream and high cooling effectiveness.
APPLICATION OF LIQUID JET TO REINFORCEMENT OF SUPERSONIC COMBUSTION
Hu Yuli, Liu Jinhua, Ling Wenhui, Liu Ling
1997, 12(1): 56-60,108.
Abstract PDF(738KB) icon1535 icon371
Abstract:
An experimental investigation on effects of oscillating shock generated by liquid jet on reinforcement of hydrogen-fueled supersonic combustion in Mach 2 airstream was completed.According to the one-dimensional calculation,it is found that supersonic combustion can be reinforced by the oscillating shock generated by liquid injection normal to supersonic flow at some distance from fuel injectors.And it is also discovered that the nearer the liquid injector is to the fuel injectors,the more combustion efficiency is enhanced by the oscillating shock.
INFLUENCE OF COMBUSTION TIME DELAY ON RESPONSE OF LOW-THRUST POWER SYSTEM WITH HIGH COMBUSTOR PRESSURE
Shen Chibing, Lu Zhixin, Chen Qizhi
1997, 12(1): 61-65,108.
Abstract PDF(337KB) icon1627 icon494
Abstract:
The mathematical model of a low-thrust power system with high combustor pressure is set up.The numerical computation for this system and analysis of the influence of combustion time delay on this system have been carried out with the characteristic method and the fourth-order Runge-Kutta method.The computed results show that the response of the power system becomes slower,the overshoot of engine performance parameter becomes greater and the time required for reaching the steady state becomes longer,if the combustion time delay is greater.When orbit control and attitude control rocket engines share the same propellant feed system,thus attitude-control rocket engine is affected more greatly by the combustion time delay.Finally,some methods are proposed to decrease the combustion time delay.The conclusions drawn in this paper provide the basis for improving its response characteristics.
TRANSFER RULE FOR DISCRETE HOLE FILM COOLING
Li Feng, Zhang Qingfan, He Jiade, Wang Minsheng, Liu Kun
1997, 12(1): 66-70,108-109.
Abstract PDF(384KB) icon1666 icon399
Abstract:
A lot of experiments has been carried out to research discrete hole film cooling effectiveness.The effects of the geometric parameters of the discrete hole and the flow parameters on the film cooling effectiveness have been analyzed.A model of covective heat transfer for discrete hole film cooling was created by the heat balance.Six different perforated plates were investigated for this purpose.The data measured are fitted by the least square approximation into some relationships which are useful to design of afterburner heat shield.
NUMERICAL CALCULATION OF FLOWFIELD INSIDE A SMALL ENGINE ANNULAR REVERSE-FLOW COMBUSTOR
Li Gaiqi, Lai Shouchang, Yan Chuanjun
1997, 12(1): 71-74,109.
Abstract PDF(315KB) icon2149 icon481
Abstract:
Calculations of flow field with fuel injection and heat release inside an annular reverse-flow combustor of a small aeroengine have been performed with a 3-D elliptic turbulent reacting thermo-aerodynamic software.In the program the turbulence is simulated by the k-ε two-equations model,the combustion after vaporization is determined by a two-step chemical reaction model,the reaction rate is calculated with Arrhenius and eddy breakup model.The SIMPLE algorithm is used to solve the finite-difference equations.The complex boundary conditions are treated particularly.Predictions of gas temperature,velocity,and fuel/air ratio show that a strong vortex emerges in the primary zone of liner;the velocity circumferential distributions have a similarity;the combustor performance is satisfactory due to favourable matching between velocity and fuel/air ratio distributions.The small circumferential velocities contribute to a reasonalble velocity distribution at turbine stator.
NUMERICAL SIMULATION OF HOT FLOW IN LIQUID ROCKET ENGINE COMBUSTOR WITH AND WITHOUT BAFFLE
Zhao Wentao, Nie Wansheng, Zhuang Fengchen
1997, 12(1): 75-78,109.
Abstract PDF(294KB) icon1634 icon371
Abstract:
The hot flow fields in liquid rocket engine combustion chamber with and without baffles are simulated numerically.The gas governing equations are described by Navier-Stokes equations in the Euler coordinates.The liquid governing equations are represented in the Lagrangian coordinates.The droplet moving tracks in the hot solutions of the two cases are investigated.The flowfield parameters of the combustion chamber are compared between the two hot solutions.It is proved that the numerical treatment of the baffle is valid.
FUZZY HYPER-BODY NEURAL NETWORK AND ITS APPLICATION TO ROCKET ENGINE FAULT ISOLATION
Huang Minchao, Wu Jianjun, Chen Qizhi
1997, 12(1): 79-82,109-110.
Abstract PDF(962KB) icon1756 icon451
Abstract:
A neural network twice classifier is provided which utilizes fuzzy sets as fault patterns of a liquid propellant rocket engine.Each fuzzy set is an aggregate of fuzzy hyperbodies.A fuzzy hyper-body is an n-dimensional hyper-sphere defined by a radius and a center with a corresponding membership function at first learning of fuzzy neural network and a part of the hyper-sphere defined by an included angle,a center and a direction vector at second learning.The training isolation accuracy of the twice learning algorithm is higher than that of the once learning algorithm,and the twice learning of fuzzy neural network also enhances sensitivity to rocket fault.The twice learning algorithm can learn nonlinear fault pattern boundaries in two passes through the input data and provides the ability to incorporate new fault messages in succession and refine existing fault classes without retraining.The emulation of its application indicates that the fuzzy hyper-body neural network can be successfully employed in the fault detection and isolation of the turbo-pump feed liquid rocket engine.
TECHNICAL NOTES CREEP RELIABILITY CALCULATION OF TURBINE DISK WITH IODM
Li Quantong, Chai Weidong, Lu Wenlin
1997, 12(1): 83-85,110.
Abstract PDF(241KB) icon1746 icon378
Abstract:
Integration Order Descending Method (IODM) is a new approximate method for solving the mathematical problems in the reliability.There are some difficultes to solve the engineering problems with the IODM.This paper shows how to use IODM for this purpose and applies it to calculating the creep reliability of an aeroengine turbine disc.This paper also presents the loads that have been considered in calculaion of the turbine disk creep with ADZNA programm.
ANALYSIS OF POTENTIAL ENERGY OF COMPLEX ROTOR-CASE SYSTEM BY USING WHOLE TRANSFER MATRIX METHOD
Ren Guangming, Zhu Zigen, Li Qihan
1997, 12(1): 86-88,110.
Abstract PDF(312KB) icon1879 icon495
Abstract:
An improved transfer matrix method-Whole Transfer Matrix method was applied to calculation of the critical speeds and vibration modes of a dual-rotor system with engine casing.Its potential energy distributions corresponding to each rotor critical speed and vibration mode are calculated and analyzed by this way.Whole Transfer Matrix method takes rotors and cases of the system as a whole body in transferring,the size of the freqency determinant is determined only by the number of rotors and cases,but free from the supports.This method is of great advantage for developing general computer program.The investigation indicated that the vibration level of the rotor system can be found from potential energy distribution.Thus the effective measures can be taken to depress the vibration.
PREDICTION OF AEROENGINE FLIGHT LOAD SPECTRUM
Song Yingdong
1997, 12(1): 89-91,111.
Abstract PDF(536KB) icon1793 icon378
Abstract:
A method of aeroengine flight load spectrum prediction is developed,which uses principles of aircraft flight mechanics to turn aircraft design mission profiles into engine thrust profiles.Flight altitude and flight velocity are acquired by solving the aircraft motion equations with the aircraft motion constraints.In the result of the engine performance calculations,the other engine operating parameters can be obtained.In the flight process from take-off to landing,taking the final value of the previous operation as the initial value of the following one,the engine flight load spectrum can be calculated for a specific flight mission.As an example,the flight load spectrum of an engine used on a trainer has been calculated.
EFFECTS OF POROSITY AND OSMOTIC OIL FLOW ON FLAME STABILITY OF CERAMIC EVAPORATION CAVITY
Sun Huixian, Du Shengtong, Ma Xuiesong
1997, 12(1): 92-94,111.
Abstract PDF(236KB) icon1627 icon467
Abstract:
A series of experiments of flame stability in ceramic evaporaton cavity with different porosity at different osmotic oil flow rate have been completed in a model afterburner.Experimental results show that,in the case of the main oil supply only or with osmotic oil supply,there is an optimal porosity respectively,where extinction limit of lean poor mixed gas is the widest,with the osmotic oil suuly only,flame extinction limit of ceramic holder becomes wider as the porosity increases.Concerning the effect of the osmotic oil flow rate on the flame stability,the extinction limit becomes extended usually as the osmotic oil flow rate increases.
A FINITE ELEMENT ANALYSIS OF TRANSIENT HEAT CONDUCTION WITH WEIGHTED VARIABLE FACTOR
Yang Huitao, Han Zengxin
1997, 12(1): 95-97,111.
Abstract PDF(236KB) icon2067 icon445
Abstract:
The stability and oscillation of numerical finite element solution of transient conduction temperature field with two-point time deference scheme have been investigated.The weighted factor of the time diference scheme σ and the time step △t are correlated to the temperature oscillation respectively.Thus the functions of variable σ and △t are obtained.Its application demonstrates the simplicity in programming,high aecuracy and econmy of running time.
STUDY ON DIGITAL CONTROL OF ENGINE REHEAT TRANSIENT PROCESS
Zhang Hui, Tao Tao, Zhang Jiazhen
1997, 12(1): 98-100,112.
Abstract PDF(224KB) icon1590 icon471
Abstract:
The emphasis of this study is laid on designing a reheat-starting transient control subsystem of an aeroengine,after the main fuel control system of the engine has been designed.That is to say,the study pursued to the reheat-starting transient control method,which can coordinate the changes of the reheat fuel supply and the nozzle area,so the reheat-starting transient process does not affect the regimes of the core engine as possible,so that the engine can work safely and steadily.The optimal control method is employed with success.
A DATA ACQUISITION AND AUTO-TEST SYSTEM FOR A TURBOJET BENCH
Fan Ding, Li Yuanye, Yang Shengfa, Chen Huaimin, Li Runzhi
1997, 12(1): 101-102,112.
Abstract PDF(242KB) icon1627 icon443
Abstract:
This paper deals with the general structure,technical standards and main functions of the data acquisition and auto-programming test system for two turbojets on test bed.The closed-loop control methods for auto-testing the turbojets are considered in detail.The system has following advantages:(1)It not only has powerful functions,but also is fairly economical becaues of the double-decker arrangement of computer.(2)The auto-throttle system has fine dynamic performance and high position finding accuracy of the output throttle angle of the engine.(3)The adaptive controller is well-designed and all the control indexes of the closed-loop test for the qualified engine are improved in any climate condition.

Copyright © 《Journal of Aerospace Power》 All Rights Reserved.

Supported by: Beijing Renhe Information Technology Co., Ltd.