1990 Vol. 5, No. 3

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SIMILARITY RULES FOR TRANSFORMATION BETWEEN HIGH-AND LOW-SPEED MODELS OF HIGH PRESSURE AXIAL COMPRESSOR BLADINGS
Zhu Nianguo, Xu Liping, Chen Maozhang
1990, 5(3): 193-198,283.
Abstract PDF(400KB) icon1964 icon421
Abstract:
The similarty rules are based on general aerodynamic similarity theory and drived for the transformation between transonic and subsonic flows,transonic and incompressible flows,and subsonic and incompressible flows respectively,as well as in the cases of full disturbance,according to the linear small perturbation theory,the transonic small perturbation theory and the full potential equation.The validity of the similarity transformations under the inviscid assumption is verified by two dimensional and quasi three dimensional numerical simulations.The reasonableness of the basic concepts and assumptions used in the present work,such as local linearization concept,are testified by the numerical expeiments as well.This work provides an essential basis for the similarity transformation between the high and low spee models of high pressure axial compressors,and offers an approach to the profile design of a similarity low speed model and to applieation of experimental data of the similarity low speed model to the high speed high pressure axial compressor bladings.
ADAPTIVE NOISE CANCELLING TECHNIQUE AND BEARING FAULT DIANOSIS
Li Zhiwei, Fu Yi
1990, 5(3): 199-203,283.
Abstract PDF(792KB) icon1416 icon473
Abstract:
In this paper,the adaptive noise cancelling technique is used to improve the signal-to-noise ratio of vibrational data induced from a bearing.Besides the LMS(Widrow) algorithm,the VS(Variable Step) algorithm with more rapid convergent speed and with only a modest increase of computational amount is discussed in comparison with the LMS algorithm.The analysis shows that,when the adaptive noise cancelling technique is adopted to monitor and diagnose a mechanical fault,the VS algorithm is quite feasible.The adaptive noise cancelling techniques at the same state and non-same state are also conducive to identify the bearing fault,but the technique at the non-same state is more applicable than that at the same state.
SIGNAL ANALYSIS TECHNIQUE FOR SURGE-DETONATION OF A TURBOJET ENGINE
Zheng Keyang, Ge Guohuang
1990, 5(3): 204-208,283-284.
Abstract PDF(376KB) icon2283 icon376
Abstract:
A diagnosis and signal analysis method with FFT signal processer is provided for on surge-detonation of a twin-spoot turbojet engine.The experimental investigation surge-detonation under a turbulent dynamic inlet distortion conditions has been completed by this method and the measuement system and test equipement are described.From the FFT analysis are drawn the conclusions:Most of surges in the turbojet compressor under turbulent dynamic inlet distortion conditions are “drift-mode” surges and appear randomly;The turbojet is far more sensitive to dynamic inlet distortion than to steady inlet distortion.It is shown that the signal analysis method with the FFT signal processor is of great practical value in the failure diagnosis,and it is a more accurate,reliable and convenient than others.under turbulent dynamic inlet distort
INVESTIGATION OF INTELLIGENT MEASUREMENT SYSTEM FOR AERO-ENGINE EXPERIMENTS
Wang Weisong
1990, 5(3): 209-212,284.
Abstract PDF(304KB) icon1407 icon401
Abstract:
This paper presents the aero-engine intelligent measurement system which has been developed.The system is original and new in its global design.It integrates the data acquisition,realtime calibration,displacement control,angle tracking,data processing and table printing.The sustem has been used successfully in combustion and turbine testings,its technical and economic effectiveness has been proved remarkable. It is of very great practical value not only in developing aero-engine and its parts,but also in its extension to other military and civil industrial departments.
SUB-MATRICES ANALYSIS APPLIED TO ENGINE FAULTS DIAGNOSIS
Tang Genglin, Chen Daguang
1990, 5(3): 213-218,284.
Abstract PDF(404KB) icon1656 icon444
Abstract:
The Gas Path Analysis method for fault diagnosis has be known as an effective method.But the crucial problem is the necessity of measuring a number of aerothermodynamic parameters more than or at least equal to the number of component fault parameters.But the number of measured parameters is always confined by cost and the structure integrity of the engine.The Sub-Matrices Analysis presented in this paper can perform component directed fault diagnosis with less measured parameters.The feature of this method is to combine the fault parameters into groups,in each of which the number of fault parameters equals the number of measured parameters to make the fault parameters solvable with respect to measured parameters.Thus,a set of solutions may be obtained.In order to select the most possible solution(s),the criteria are also given.The numerical investigation shows that this method is very effective to diagnose the malfunction occuring in one or two components which is the usual case at an early stage of engine fault.The problems which are worth careful consideration to avoid false diagnosis are also pointed out.Application of this method to JT9D commercial engine verifies its effectiveness.
HOLOGRAPHIC INTERFEROMETRY MEASUREMENT OF AXISYMMETRIC FLAME TEMPERATURE PROFILES
Xie Mingxing, Wang Zhu
1990, 5(3): 219-223,284.
Abstract PDF(442KB) icon1732 icon390
Abstract:
The Double-exposure holographic interferometry is applied to measuring the temperature profiles of axisymmetric spirit-lamp flame and hydrogen flame.A new method is used for searching and thinning interferometric fringes in the computer image processing of holographic interferograms.The temperature profiles of the flames are calculated from measurements of the fringe shift by solving Abel-type integral equation with the Laplace transform.The errors resulted from refraction of the probing rays in holographic interferometry of axisymmetric phase objects are discussed.A simple method is introduced to correct the refractive errors.The effects of the object composition and of small displacement errors of the interferometric fringes on holographic measurements of axisymmetric phase objects are considered too.
EVALUATION OF PARTICLE SIZES AND POSITIONS BY THEIR COHERENT IMAGES
Fang Jinyan, Yan Chuanjun, Wang Hongji
1990, 5(3): 224-226,285.
Abstract PDF(264KB) icon1480 icon447
Abstract:
A new approach to evaluate the particle sizes and positions has been deve loped on the basis of their coherent images.According to the experimental and calculated results about the coherent images of steady particles,the paper described the features of coherent images of particles.It is concluded that the characteristics of the coherent images are different from those in incoherent cases,and the more the image deviates from the focus,the more obvious the differences are.From theoretical analysis, the conclusion has been drawn and proved that the normalized intensity distributions of coherent images of particles are the same when their recording far-field conditions are identical.The new method to determine the particle sizes and positions is just based on this conclusion.
EXPERIMENTAL TECHNIQUE FOR INVESTIGATION ON FOREIGN OBJECT DAMAGE OF AEROENGINE
Qiao Wenxiao, Xiong Changbing
1990, 5(3): 227-228,285.
Abstract PDF(228KB) icon1670 icon526
Abstract:
A method to determine the size of notch on first stage blade of compressor with a stress concentrated factor (K t) of three is briefly introduced,which is required for foreign object damage (FOD)testing of aeroengines.It is based on 8-node and 20-node finite element appoaches respectively.The 8-node super-parameter element approach is used to calculate the stress distribution of a whole blade without and with notch,and the 20-node 3D iso-parameter finite element programm is applied to computation of local stress distribution around a notch of blade.The practical testing techniques and test results are also presented.
VIBRATION DIAGNOSIS AND VIBRATION SOURCE ANALYSIS OF AIRCRAFT ENGINE
Li Xifa, Qiu Lun, Yi Juan, Meng Zhaobing
1990, 5(3): 229-230,285.
Abstract PDF(208KB) icon1632 icon404
Abstract:
This paper presents a progress in aircraft engine vibration monitoring and diagnosis in flight.Essentiality of the in-flight vibration monitoring and vibration source analysis and the features of the system used are discussed.An airborne vibration data acquisition unit,the ground analysis equipments and the method for analysing vibration signal are given.Its applications prove that it is feasible for vibration signal recording in flight and frequency spectrum analyzing.
EXPERIMENTAL RESEARCH ON DISTORTION TOLERANCE CAPACITY OF A TWIN-SPOOL TURBOJET
Wu Yuegeng, Yun Qisheng, Ge Guohuang
1990, 5(3): 231-235,286.
Abstract PDF(342KB) icon1701 icon377
Abstract:
The inlet pressure distortion experiments of a twin-spood surbojet were completed at the ground test-ring.Under a variety of conditions of the inlet circumferential and radial pressure distortion,the distorting flow field,distorting index,critical distorting angle and sensitivity have been measured separetely by the steady and dynamic measuring systems.Meanwhile,the emphasis was placed on analysing the effect of inlet circumferential and radial pressure distortions on the performance and reliability of the twin-spool turbojet,the results of the analysis provide the experimental basis for matching between the engine and its intake.
THE CALIBRATION TECHNIQUE OF FIVE-HOLE PROBE
Lin Feng, Zhang Kunyuan, Pen Chengyi
1990, 5(3): 236-238,285.
Abstract PDF(966KB) icon2272 icon424
Abstract:
A Calibration technique is developed to make the five-hole probe applicable to both definite and variable Mach numbers.A rake composed of 9 five-hole probe is calibrated.The diamelter of the probe is 2.5mm and the width of the rake is 100mm.The range of yaw angle is from -30° to +30°,so is the range of pitch angle.The minimum calibration Mach number is 0.15 and the maximum is 0.45.The standard deviations of yaw angle and pitch angle of each probe are less than 1.0°.The experimental results shows that this technique is suitable for engineering measurement.
THE GOVERNING EQVATIONS AND TURBULENCE MODELS OF COMPRESSIBLE TURBULET FLOW IN A ROTATING FRAME
Hua Yaonan
1990, 5(3): 239-244,286.
Abstract PDF(416KB) icon2629 icon406
Abstract:
The governing of equations compressible turbulent flow in the rotating frame are established by the aid of the mass weighted average method.The transport equations for the turbulent stresses,the turbulent kinetic energy and the dissipation rates are derived from the obtained equations.The from of these equations set up in a rotating frame is different from that in an ordinary inertia reference system.The rotating effects caused by Coriolis and the centrifugal forces result in the appearance of the additional iterms in these equations.They may not be ignored in solving the Navier-Stokes equations and studying the turbulence models.Five turbulent models and their mathematical expressions are considered,and their revised forms taking account of the rotating effects are discussed.
ROTATING STALL AND SURGE IN AXIAL FLOW COMPRESSOR
Tang Diyi, Guo Jie, Li Lijun, Qiao Weiyang
1990, 5(3): 245-250,286.
Abstract PDF(385KB) icon1837 icon448
Abstract:
Supported by the funds of the national natural science,we have tested and meassured the instability behavior of compressors in detail with hot-wire,hot-film probes and transient pressure 3-hole probes on single rotor,single stage and multi-stage compressors respectively.After analyzing signals with the power spectra and correlation analysis,it is found that there are two kinds of disturbances—peripheric oscillation and axial oscillation in the flow.On the basis of this fact,we propose the concept that the flow separation from airfoils results in peripheric oscillation and axial oscillation simultaneously,and provide a new model to describe the flow destabilizing mechanism. For the stability problem of multistage axial flow compressors,it is of practical value to investigate its stability limits,inclusive of the rotating stall margin.The preliminary results of this investigation show that the stability limit in the high speed regionis equivalent to a surge line,that the stability limit the middle speed region is indeed a stall line,and that the stability limit in the low speed region has two branches,one is the stall line of the whole compressor,the other is the rotating stall inception margin of the first stage.Both are of the same importance from the viewpoint of engine safety.Finally the mathematical methods for predicting the stability limits are discussed.
STUDY ON THE SECONDARY FLOW AND IT′S CONTROL IN COMPRESSOR
Tao Deping, Pen Zeyan, Wei Xinlu
1990, 5(3): 251-256,287.
Abstract PDF(259KB) icon1703 icon452
Abstract:
The present paper is a final report of the study on the secondary flow in compressors.A systematic experimental investigations on the secondary flow of both the C4 and BC-6 airfoil cascades have been completed.A end-wall corner flow model has been provided.The C4 airfoil variations of the outlet angle along the blade height have been calculated by the secondary flow theory.They approach to the experimental results.The influence of the tip clearance on the blade stall flutter is also investigated in the flutter experiment of the BF-1 compressor rotor.An experimental study on the reduction of secondary flow is carried out by using both “Heading Edge Sweep” and “End Bending” techniques.The total pressure loss coefficients of these two cascades decrease significantly.The results of the whole compressor flow field calculation of the compressor “End Bending” cascades are in good agreement with the experimental data.Some important conclusions are drawn to reveal the real flow mechanisin in the compressor and to develop new compressor flow physical models.It is of great value for engeerning application of the secondary flow control.
EXPERIMENTAL INVESTIGATION ON THE FLOW THROUGH A COMPRESSOR END-BENDING CASCADE
Wu Guohua, Ren Liyun, Pen Zeyan, Yan Ming
1990, 5(3): 257-260,287.
Abstract PDF(295KB) icon1590 icon393
Abstract:
By the comparison of secondary flow vector diagrams and 2-D iso-loss line diagrams between a end-bending cascade and a ordinary cascade,this paper reveal the mechanism that the end-bending cascade improves the flow through end wall region and reduces the loss of the end wall flow.At the same time,the end-bending cascade reduces the mass-averaged total pressure coefficient.Both directions of blade force on the suction and pressure surfaces are in favor of improving the end wall flow.The former pushs the high energy flow outside the end wall region near the pressure surface into the end wall region,the latter transfers the low energy flow within the end wall region near the suction surface to the main stream.The improvement of the end wall flow contributes to control the degeneration of the end wall flow downstream.
OPTIMIZATION OF STRUCTURAL PARAMETERS OF VENTILATINGN FANS
Jiang Zhenliang
1990, 5(3): 261-264,288.
Abstract PDF(290KB) icon1474 icon394
Abstract:
Centrifugal and axil ventilating fans are optimized by using different objective functions,different variables of design,different constraint forced conditions,different methods of optimization.The results show that they have common features.The effficiency of the optimized design is 5-10% higher than the traditional design.If the static pressure total-effectiveness is chosen as the objective function,the results will be more satisfactory.this procedure of optimization may be modisfied easily to applicable to a blower and a compressor.
EXPERIMENTAL INVESTIGATION OF RIBLETS DRAG REDUCTION FILM
Jiang Zhi, Tang Zhiming, Gao Ge, Ning Huang
1990, 5(3): 265-267,287.
Abstract PDF(259KB) icon1696 icon419
Abstract:
Based upon the new explanation of the structure of turbulent boundary layer and the fact that specially designed non-smooth surface can reduce fluid drag,the authors have designed and fabricated the Riblets Drag Reduction Film.This Film was covered on racing boats,the boats and recieved excellend awards in competition.The authors also tested the flat plates with and without the Biblets Drag Reduction Film in wind tunnel and compared their fluid drags.The results showed that the Drag of the covered plate has been reduced by 7%.
WJ5A MODIFICATION AND EXPERIMENTAL RESULTS
Li Jinghai
1990, 5(3): 268-270,288.
Abstract PDF(247KB) icon1212 icon378
Abstract:
WJ5A compressor has been modified.The performance data obtained in testing prior to and after the modification of the compressor are compared.Test results indicate that the performance of the modified compressor is greatly improved.In the case of the relative corrected speed ηnp1 =1.0,the maximum isentropic efficiency of the modified compressor ηkmax is increased by 3 6%.When ηnp1 =0 70～0 90,η*kmaxis increased by over 15%.The pressure ratio and the mass flow are changed only a little at all speeds.When ηnp1 >1 0,there is no evident variation in the compressor performance.
A POWER CORRECTION FORMULA FOR PISTON AEROENGINES
Lu Ziying
1990, 5(3): 271-272,288.
Abstract PDF(1281KB) icon2162 icon384
Abstract:
A new formula to correct the power for piston aeroengines is presented.It has the advantage of considering the influence of humidity on power in the simplified calculation.Theoretical analysis and test data show that this new formula is more accurate than that used before.
SIMULATION CACULATION AND ANALYSIS OF STEADY ERRORS IN A NONLINEAR SPEED CONTROL SYSTEM
Hu Shiming
1990, 5(3): 273-274,288.
Abstract PDF(205KB) icon1632 icon366
Abstract:
A method of the digital simulation is used to analyse the steady characteristics of the nonlinear speed control system with dead zone and saturation in detail.The correlations of the system steady errors with the dead zone,ramp segment,saturation and disturbance are considered separately.It is pointed out that the steady errors are essential in the nonlinear control system substantially.According to the characters of the system in the whole control process.There are three feedbacks:First is proportion/velocity/proportion feadback,second is proportion and velocity feedback.
EXPERIMENTAL INVESTIGATION OF OSCILLATING CASCADE AERODYNAMICS
Daniel H.Buffum, Sanford Fleeter
1990, 5(3): 275-282.
Abstract PDF(516KB) icon1940 icon452
Abstract:
A series of experiments are performed in the NASA Lewis Transonic Oscillating Cascade Facility to provide fundamental data quantifying the high subsonic and transonic steady and oscillating aerodyna mics of a biconvex airfoil cascade at realistic reduced frequency values for all interblade phase angles.This is accomplished by developing and utilizing an unsteady aerodynamic influence coefficient technique in which only one cascaded airfoil is oscillated at a time.The vector summation of the resulting airfoil surface unsteady pressures measured on a dynamically instrumented airfoil enables the unsteady aerodynamics of an equivalent cascade with all airfoils oscillating at any specified interblade phase angle to be determined.

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