1986 Vol. 1, No. 2

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WORKING PRINCIPLES OF INTAKE FENCES
Zhang Shiying, Yang Gaoming
1986, 1(2): 97-102,181.
Abstract PDF(745KB) icon4278 icon1255
Abstract:
Air intake with fences installed at the entrance is investigated at low speed. With fences at a high angle of attack the steady and dynamic distortions are effectively reduced and the total pressure recovery is increased at the exit.These improvements are shown to be closely related to the reduced extent of separation along leeward wall of the intake by the fences at the angle of attack. A flow model is suggested to describe the flow in the channels formed by the fences. With this model in mind, the mechanism of the fences can be easily imagined and the performance of the slanted inlet with fences can be predicted. The effects of the position, lengths and the number of the fences on the performance of the inlet are investigated. The difference between a slanted and a pilot inlet both with fences are compared and an intake with an S-duct and fences are evaluated. Combined with less radar reflection the multi-channel inlet is shown to be attractive. The above related investigation will be helpful in intake fence design.
DESIGN OF SWIRL SIMULATORS
Peng Chengyi, Lin Feng, Zhang Kunyuan
1986, 1(2): 103-106,181-182.
Abstract PDF(521KB) icon4086 icon1334
Abstract:
Response of turbo-jet engines to the swirling flow at engine face has recently been a matter of major concern, and this response can only be evaluated on the testbed with the help of swirl simulators. A method for design of such simulators is presented. Swirl blades that generate wake vortices are installed on the inner surface of the air inlet. As the arrangement of blades may vary in blade number and circumferential position, as well as in blade geometry and angle of attack, there is no want of factors by which the characteristics of the flowfield can be controlled. It is assumed that the wake vortices are well formed at the cross section of the duct corresponding to the trailing edges of the blades. The structure and positions of these vortices have to be determined, the flowfield at this cross section can then be computed. Since the velocities ofthe vortex centers are known, their respective positions at a nearby section slightly downstream can be determined. The flowfield at this new section can then be computed. Thus,the computation goes on step by step from the blade section up to the final section where the desired flow pattern is to be obtained. The design work begins with a tentative choice of blade arrangement. Comparison of computation with the desired result will guide the modification of blade arrangement. A computer, which displays a new flow pattern each time a correction is made of the blade arrangement, greatly facilitates the design work. Swirl patterns given in the literature of real aircraft inlets have been simulated by this method and test results are shown to be satisfactory.
CONVERGENCE OF PERFORMANCE CALCULATION OF TWIN SPOOL TURBOJET AND TURBOFAN
Zhu Xingjian, Wang Xueyu, Yan Qingjin, Zhang Xueliang
1986, 1(2): 107-110,182.
Abstract PDF(449KB) icon1628 icon455
Abstract:
The calculation of steady-state performance for twin spool turbojet and tur-bofan can be performed by formulating a multi-variable nonlinear equation set with parameters.
A METHOD FOR CALCULATION OF FLOW PROCESS IN AN AXISYMMETRIC STRAIGHT-WALL ANNULAR DIFFUSER
Hua Guangshi
1986, 1(2): 111-116,183.
Abstract PDF(527KB) icon1412 icon446
Abstract:
A complete method for calculation of the steady compressible flow in an axisymmetric atraight-wall annular diffuser with boundary layer correction based on streamtube method has been developed.
VIBRATIONAL CHARACTERISTICS OF THE CANTILEVER ROTOR WITH A CRACKED SHAFT
Gu Jialiu, Wang Nengqan
1986, 1(2): 117-120,183-184.
Abstract PDF(485KB) icon1739 icon449
Abstract:
The vibrational characteristics of the high speed cantilever rotor with a cracked shaft are investigated to find an optimum vibration monitoring method for a turbine rotor. The model studied is established based on the structure of a real turbine rotor.The transfer matrix-direct integration method proposed by the first author in reference (2) is applied to the theoretical analysis. In consequence it is possibleto avoid making any assumption about the whirling orbit, and modification of some real parameters and whirling orbit of the disk is performed sequentially in the process of numerical integration. In experimental investigation some measures are taken to create a crack very similar to a real crack on the shaft.The results of Fourier analysis of the unbalance responses obtained from both calculations and experiments show that in signature of vibration of the cracked rotor exist very strong 2nd anc 3rd harmonics. The triple-frequency harmonic increases with the rotating speed and reaches maximum at the speed being equal to 1/3 of the critical speed.Since the second harmonic can be caused by the other factors such as misalignment,as a result of this research the authors suggest that operating the rotor in the vicinity of 1/3 of the critical speed and monitoring the triple-frequency component be the best method to detect a crack for this kind of turbine.
EFFECTS OF THREE CENTRES OF BLADE ON FLUTTERING
Zhang Chengsheng, Song Zhaohong
1986, 1(2): 121-124,184.
Abstract PDF(393KB) icon1676 icon491
Abstract:
The effects of three centres, i.e. aerodynamic centre, centre of gravity and centre of torsion, of blade on fluttering have been studied. A structural model of the coupling between bending and torsion is presented in the theoretical analysis.According to their dependence of the displacements of the blade vibration, the aerodynamic loads are considered as a function of the velocities of the blade vibration. The complex eigenvalue equations are resolved by iteration approach .The conclusions drawn from the study are as follows: the blade with the forward centre of torsion is less stable than the blade with the backward one; the blade with the forward centre of gravity is more stable than the blade with the backward one. In the case of the blade with the foreward torsion centre,its stability is improved when the aerodynamic centre takes place between the gravity centre and the torsion centre. As far as the blade with backward torsion centre is concerned, the blade will be unstable when the aerodynamic centre is located between the gravity centre and the torsion centre. The analysis of these three different kinds of blades demonstrates that the blade with the aerodynamic centre located at the leading edge of the blade is the best one.
PROPAGATION OF COMBINED STEADY CIRCUMFERENTIALTOTAL PRESSURE AND TEMPERATURE DISTORTIONTHROUGH AN AXIAL ROTOR
Zhang Huimin, Tang Guocai, Hu Jun
1986, 1(2): 125-130,185.
Abstract PDF(543KB) icon1751 icon364
Abstract:
Although the effects of combined steady circumferential total pressure and temperature distortion on performance have been investigated analytically and experimentally in recent years, the other urgent problem concerning the interrelation between them in the process of propagation through an axial compressor (including multistage and two-spool axial compressor)has been somewhat neglected.On the basis of previous work (4) and the approach proposed by the authors in dealing with the distortion propagation, two cases of combined steady circumferential total pressure and temperature distortion are investigated, first, how the variation of the total temperature distortion amplitude influences the attenuation of the total pressure distortion under the conditions of no phase difference; second, how it does when there is phase difference between them.Numerical experiment and analysis show that the amplitude of the total temperature distortion and the phase angle between them have great influences upon the attenuation of the total pressure distortion, within a certain range of the total temperature amplitude, the existence of the total temperature distortion is expedient to the attenuation of the total pressure distortion. The worst case is when the two are just opposite in phase and the total temperature distortion amplifies the total pressure distortion instead of attenuating it.
A METHOD FOR COMPUTATION OF VISCID/INVISCID INTERACTION ON TRANSONIC COMPRESSOR CASCADES
Jiang Haoxin, Chen Maozhang, Cui Jiya
1986, 1(2): 131-136,185-186.
Abstract PDF(565KB) icon1712 icon368
Abstract:
An inviscid/viscous interacton method has been developed to predict the blade-to-blade flow in transonic compressor cascades. The method intends to be used primarily for relatively high camber blades where the suction surface boun-dary layer may be slightly separated near the trailing edge. The inviscid flow was calculated by means of AF2 finite-difference scheme on a newly constructed H-type grid system[3], and the viscous flow by the aid of an integral method for laminar and turbulent boundary layers[6] . An integration calculation was then carried out to determine the blade deviation angle and the total pressure loss coefficient. One of the authors has been involved in the recent transonic experimental work of one set of MCA cascades in Sichuan.The numerical predictions obtained have been compared with test data on four different transonic cascades in reasonably good agreement.
OPTIMUM COMPUTATION FOR EXIT CONE CONTOUR OF NOZZLE WITH TWO-PHASE FLOW
Fang Dingyou
1986, 1(2): 137-140,186.
Abstract PDF(447KB) icon1432 icon384
Abstract:
To simplify the computational procedure and minimize the expensive computer time, the optimum calculation for exit cone contour of nozzle with two-phase flow is carried out by the procedure of two-dimensional two-phase (2D-2-phase) nozzle flow. This procedure consists of computations of both transonic and supersonic two-phase flow. The former provides an exact data at the initial line for the latter.The computation of supersonic two-phase flow is completed by charactristics method. The direct marching method is applied to the interior points and axis points, and the inverse marching method to the wall points. There are seventeen finite difference grids in the flowfield, but the typical grids are only five, the others can be divided into several typical grids.The influences of inertia, compression and rarefaction must be considered in the particle drag and heat transfer formulae for computing the supersonic two-phase flow.The 2D-2-phase specific impulse losses have been computed by the present procedure, and the results illustrate that the optimum initial expansion angles are about 28.5% 29.5°,and 27.0° for parabolic, circular and cubic contours respectively. The 2D-2-phase specific impulse losses of these three nozzle configurations are almost the same when the exit angles are equal and the initial angles are close for the above curves.
DYNAMIC SIMULATION RESEARCH ON DIGITAL SPEED CONTROL SYSTEM OF AEROENGINE
Tian Min, Wu Chihua
1986, 1(2): 141-144,187.
Abstract PDF(477KB) icon1286 icon457
Abstract:
This paper describes the real time dynamic simulation research on the digital speed control system of an aeroengine under small perturbation conditions. The control calculation approach has been programmed for a microcomputer. The proportional derivative control is applied to the large error region and the proportional control to the small error region. Furthermore, in the small error region, the large proportional constant control is adopted if the error increment tends to increase, and the small proportional costant control is employed when it tends to decrease. For the sake of normal operation in the digital speed control system, the influence of the high frequency disturbance, the overspeed and spillover in the microcomputer have been eliminated. Finally, the flow chart of the program and the real time simulation tests are demonstrated. The test results and some useful conclusions are also presented.
EXPERIMENTAL INVESTIGATION ON MODIFICATION OF THE WZ-6 AEROENGINE TO USE NATURAL GAS AS FUEL
Hu Yian, Zeng Miaolan
1986, 1(2): 145-149,187-188.
Abstract PDF(554KB) icon1851 icon464
Abstract:
This paper describes the experimental investigation on modification of the WZ-6 aeroengine into an industrial turbomachine (WZ-6G) which has been run for 1500 hours continuously as a powerplant in oil field. The fuel supply system of the engine has been reconstructed so that the natural gas is introduced through the three tubes radially inserted in the axial diffuser into a stationary jet ring, and therefrom injected into the main combustion zone. The ejection of thenatural gas from three rows of holes on the jet ring is peripherally uniform.Testing shows that the reasonable injection speed of the natural gas is 50 to 80m/s, nonuniformity coefficient of the gas temperature field at the outlet of the combustion chamber is about 0.22, Under the above conditions, the distributionof the temperature field in the combustion chamber meets the essential requirements and provides smooth combustion with the transparent light blue flame. The temperature of the flame tube wall is permissible.The main combustion chamber is actuated and ignited by the torch ignitors with natural gas as fuel.
RANDOM VERTEX METHOD AND SIMULATION OF VORTEX STRUCTURE BEHIND A TRIANGULAR PRISM
Xu Chenggang, Yan Chuanjun, Wang Hongji, Zhu Huiling, Tang Ming
1986, 1(2): 151-153,188.
Abstract PDF(368KB) icon1800 icon449
Abstract:
In this paper, the flow passing a triangular prism has been calculated by random vertex method. In order to eliminate the oscillation of the tangential velocity at the prism surface,a coefficient Kw is introduced, which is used to consider the interaction of newly generated vortex blobs at the surface. The idea of so called "gravity blob" is proposed to save the CPU time in running the program for simulation of vortex structure behind the triangular prism. The essential features of the large-scale vortex structure in the recirculation region and Karman vortex street downstream are clearly shown in the figures of vortex blob distribution obtained from calculations.
A FLIGHT-TEST STUDY ON THE TOTAL PRESSURE RECOVERY AND EXIT FLOW FIELD IN AN INLET
Miao Liang
1986, 1(2): 154-156,188-189.
Abstract PDF(386KB) icon1583 icon463
Abstract:
A flight-test study on the total pressure recovery and the exit flow field in an inlet of a supersonic jet aircraft is presented. The total pressure recoverycharacteristics and the exit total pressure distortion patterns in the transonic flight are given. The variation of the total pressure recovery coefficient andthe distortion factor with Mach number during inlet-engine matching is illustra-ted. The flighttest study provides valuable data for research on inlet-engine compatibility .
EFFECT OF WAKE-TYPE INLET VELOCITY PROFILES ON PERFORMANCE OF SUBSONIC DIFFUSER
Chen Xiao
1986, 1(2): 157-158,189.
Abstract PDF(262KB) icon1381 icon429
Abstract:
The paper discusses the experimental investigation on subsonic dif fuser. When in the front of the subsonic two-dimensional diffuser an elliptic bar is placed perpendicular to the flow stream and parallel to the diverging wall of the diffuser,the wake-type inlet velocity profile is created and the core-stream turbulence level at the entrance of the diffuser is increased, which results in flatter velocity profile implying an increase in momentum transport across the boundary layer.The value of the boundary layer shape factor is therefore decreased. Consequently,the flow separation is delayed and the separation range is reduced. As a result,the static and the total pressure recovery coefficients of the diffuser rise by 30% and by 3% respectively when the maximum Mach number of the diffuser inlet reaches 0.7.
A METHOD FOR MEASURING TWO-DIMENSIONAL UNSTEADY FLOW VELOCITY WITH X-HOT-WIRE PROBES
Tang Guocai
1986, 1(2): 159-160,189.
Abstract PDF(265KB) icon1506 icon437
Abstract:
A method is presented for measuring two-dimensional unsteady flow velocities with X-hot-wire probes.Based on the well-known calibration law incorporated with the yaw angle correction function f(θ) = cosmθ, the magnitude and the direction of the velocity are determined. The applicable range for this method and the errors caused by variation of parameter m with flow angles and velocities are also discussed.Since no approximation is introduced in the calculation,the accuracy of measurement is enhanced. This method can be employed where the magnitudes and the directions of the velocities vary to a great extent. Moreover, the formula provided is simple to use.
A COMPUTER-CONTROLLED VIBRATORY FATIGUE TEST RIG WITH PROGRAMMED LOADING FOR BLADING
Lu Qixin, Zhuang Zhongliang
1986, 1(2): 161-163,190.
Abstract PDF(345KB) icon1691 icon492
Abstract:
In order to improve the accuracy in experimental prediction of service life of axial-flow compressor blade,it is necessary to simulate random field loads in the laboratory. Based on recorded flight spectrum loading, a microcomputer system for data processing and counting was successfully utilized to derive a mathematical model for a steppedload spectrum and software programming. For simulation by programmed step loading, a computer-controlled closed cycle servo-electrical test righas been newly devised. This paper describes briefly its design and technical features. On this rig, a fatigue life distribution of certain compressor blading was automatically tested under its simulated flight block load spectrum. This advanced technique for inspection of the fatigue life and quality in blade production may be employed by blade manufacturing factory to enhance the flight reliability.
VIBRATION SPECTRUM ANALYSIS OF A TURBOPROP ENGINE IN STARTING PROCESS
Gao Jinghai, He Lianrui
1986, 1(2): 164-166,190.
Abstract PDF(339KB) icon1626 icon457
Abstract:
This paper describes how the FFT (Fast Fourier Transform)technique is applied to the spectrum analysis of the vibration signals recorded in starting process of a turboprop engine and what features of its spectrum have been found out. Based on analysing the vibration data recorded from 174 starting processes, the basic rules of vibration in starting of the engine are found and essential difference between normal and troubled engines is clarified. Vibration divergence in hot and cold startings and the influences of thermal deformation of the engine rotors on the vibration are also investigated.This work has accumulated data and experiences for fault diagnosis and monitoring in starting of the engine.
CONSTRUCTIONAL IMPROVEMENTS IN A TURBOPROP ENGINE
Yan Yanxiao
1986, 1(2): 167-169,191.
Abstract PDF(312KB) icon1563 icon367
Abstract:
The principal constructional improvements in the reducer, compressor , com-bustor, turbine and bearings of a turboprop engine are summarized. Practice shows that the constructional improvements in the reducer and the turbine are of the most importance in prolonging the service life of this turboprop engine.
AN EXPERIMENTAL STUDY ON DISTRIBUTION OF COLD AND HOT AIRFLOWS IN COMBUSTOR
Mo Lixiao
1986, 1(2): 170-172,180-191.
Abstract PDF(386KB) icon1696 icon466
Abstract:
The flow distribution in an annular combustor at different Mach numbers and heating ratios was investigated experimentally.An empirical relationship for coordination between the cold and the hot flow distributions has been established with the aid of the least square method based on the experimental results. The data calculated are in good agreement with the test ones.Therefore, the aim of the study is realized that the empirical formula enables us to predict the combustor flow distribution at different Mach numbers and heating ratios according to the flow distribution at a given low Mach number with cold air flow.
AN APPROACH TO SPECIFY THE INITIAL CONDITIONS IN COMBUSTION FLOW BY SIMULATING A FICTITIOUS CATALYSTIC BOUNDARY
Sheng Yiqian, Zhou Qian, Zeng Qiufan
1986, 1(2): 173-175,191-192.
Abstract PDF(363KB) icon1257 icon444
Abstract:
In numerical calculation of steady combustion flow field,the governing dif-ferential equations can be discretized with finite difference schemes and solved with iterative method. But their convergence solution can be attained only with good estimates of initial conditions.Instead of empirical estimate of the initial conditions,an approach to specify the initial conditions is presented,based on the mathematical simulation of the forced ignition. In the result, the initial conditions can be determined computationally by setting a catalytic boundary in the combustion zone.
RADIATION CORRECTION OF MEASURING TEMPERATURE IN COMBUSTOR
Lu Xinli, P.Wild
1986, 1(2): 176-177,192.
Abstract PDF(256KB) icon1774 icon432
Abstract:
In tests of combustor heat loss arises from radiation in measuring te mpera-ture with a thermocouple. Therefore,it is necessary to make a correction for it. An approach to neglect some minor factors and to simplify some coefficients in the energy balance equation is recommended in this paper. A simplified calculation method is thus deduced for this correction. The temperatures corrected by this method quite agree with the temperatures calculated from gas analysis data.
A FORMULA FOR AVERAGE CONVECTIVE COEFFICIENTIN LAMINAR BOUNDARY LAYER OVER A PLATE WITHUNHEATED LEADING EDGE
Wang Baoguan
1986, 1(2): 178-180,192.
Abstract PDF(371KB) icon1835 icon486
Abstract:
A dimensionless expression for average convective coefficient in laminar boundary layer over a flat plate with unheated leading edge is derived. By approximate integration,it is transformed into the following formula.

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