1993 Vol. 8, No. 2

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EXHAUST SYSTEM MODEL TEST AND RESEARCH
Chen Lide
1993, 8(2)
Abstract PDF(262KB) icon1580 icon391
Abstract:
At Mach numbers ranging form 0 2 to 0 35,two schemes 1/5th scale models of exhaust diffusers and elbow exhausts have been researched.The first scheme was a diffuser with wide cone angle.In order to reduce the separation losses,the concentric splitter cones were set to augment the diffusibility.The splitter cones were made by subdividing the short wide angle diffuser into a series of smaller,more optimal ones.The second scheme was a diffuser with smalle r cone angle,in order to reduce the length,the diffuser ratio was distributed between the linear diffuser and the turning vane to improve the overall performence.The results of the model test show that total pressure re covery coefficient C P of the second scheme is rather high and the pressure loss coefficient ΔP/Q 1 is reduced.At the same time,the ΔP/Q 1 of different constructions of the first scheme were measured separately,and the measurements were analysed.The function of the splitter cones was demonstrated.Se veral test curves were used to prove the certain effect of the pre swirl cascade in some diffuser angle ranging.Therefrom the beneficial design ideas are obtained.
2-D THEORETICAL ANALYSIS OF CIRCUMFERENTIAL GROOVED CASING TREATMENT
Du Zhaohui, Liu Qianzhi, Liu Zhiwei
1993, 8(2): 97-100,199.
Abstract PDF(272KB) icon1558 icon472
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On the basis of analysing the 2 D simplified extended stability model of circumferential grooved casing treatment,a new rapid practical method for engineering design and analysis is developed.An exact and time-saving program without repeat process is written for calculating the cascade flowfield on the arbitrary rotary surface.In the program the boundary element method is used to determine the initial flowfield for analysing the model.The effects of structural and performance parameters of circumferential grooved casing treatment on the stability of the compressor are discussed.This method features definiteness of physical meaning and ease in practical application.The new method and technique are provided to design and analyze the circumferential grooved casing treatment equipment.
NUMERICAL SIMULATION OF UNST EADY FLOW IN A TRANSONIC CASCADE
Zhang Yangjun, Tao Deping
1993, 8(2): 101-104,199.
Abstract PDF(281KB) icon1286 icon413
Abstract:
The effects of unsteady separation,shock and viscosity on the stall flutter become now a main obstacle to aeroelastic study.The 2 D unsteady flow in an oscillating cascade is taken as composed of two parts:the steady mean flow and the small disturbance flow.The stall flutter is predicted by solving the steady Navier-Stokes equations and linearized Navier-Stokes equations separatedly,and can be simulated numerically for different kinds of flow separation,and then the unsteady separation can be treated effectively.The calculation results show that unsteady separation,shock and viscosity have significant effects on the stall flutter.Therefore,the unsteady pressure response in the separated zone cannot be taken as zero.
EFFECT OF STEADY-STATE CIRCUMFERENTIAL PRESSURE AND TEMPERATURE DISTORTIONS ON COMPRESSOR STABILITY
Lian Xiaochun, Chen Fuqun, Wang Li'an
1993, 8(2): 105-108,200.
Abstract PDF(278KB) icon2092 icon476
Abstract:
A model for predicting the effect of inlet steady state circumferential total pressure distortion on compressor stability pressented in reference [1] has been improved by considering the variation in compressor efficiency due to the distortion.The impr oved model is applicable to case of the inlet steady state circumferenti al total temperature distortion.Based on this model,a method has been dev eloped for estimating the effects of combined pressure and temperatur e distortions on the compressor stabilit y.A numerical analysis was conducted for a two spool turbojet engine with the pr essure and temperature distortions as we ll as the combined distortions.The results show that the variation in compressor efficiency due to distortion cannot be negligible for large amplitu de pressure distortion.For the combined distortions,the pressure and temperature distortion orientations play a certain role in the compressor stall.The maximum loss in the compressor stall margin occurs at superimposed (180° overlap) combined distortion.Conversely,the effect is minimum for the opposite combined distortion.
CONTROL OF SEPARATION BY DYNAMIC AIR JETS
Ren Guopin, Zhang Shiying
1993, 8(2): 109-111,200.
Abstract PDF(229KB) icon2036 icon367
Abstract:
Results about the effects on separation of an aerofoil at high angle of attack by air jets producing streamwise vortex of alternating circulation and spanwise vortex fluctuating up and down above the surface are given.Combined effects of the above two kinds of jets are also investigated.The actions of the mentioned jets both cause the retardation of separation and earlier reattachment.The suitable design and arrangement of the jets gain reduction of 76% separation bubble length.
DYNAMIC SIMULATION OF HIGH-SPEED ROLLER BEARINGS
Li Jinbiao, Wu Linfeng
1993, 8(2): 112-116,200-201.
Abstract PDF(374KB) icon1728 icon402
Abstract:
The ge neral fatigue theory of rolling bearings is unable to predict the failure du e to skidding and skewing of rolling ele ments,which often occur in high speed roller bearings.Accordin g to the hydrodynamic lubrication and EH L theory,a model of interaction among rolling elements is est ablished.The dynamic model of roller bearings is also established according to the second Newton law.B y means of the program provided in the papar,it is possible to calculate the load distribution,oilfilm thickness and fatigue life of roller bearings,and to simulate dynamically the skidding of rollers and cage,the skewing and axial motion of rollers.Hence,a new effective tool of design and failure analysis is provided for high speed roller bearings.
EXPERIMENTAL STUDY ON TURBULENT TWO-PHASE FLOW IN A DUAL-INLET SIDE DUMP COMBUSTOR
Liao Changming, Shen Xiong, Zhou Lixing
1993, 8(2): 117-120,201.
Abstract PDF(260KB) icon1756 icon380
Abstract:
The measurements of turbulent gas particle two phase flows in two types of side dump combustors have been made.One of the combustors is of blade inlet type,the other is of central tube tangential inlet type.Experimental investigations have been carried out by using a TSI 2 dimensional LDV system.The time averaged and RMS fluctuation velocity of gas and particle phases in axial and tangential directions have been measured.It can be seen that strong recirculating flows of two phases exist in the head zone of both combustors,and swirling flows of two phases exist in the tangential inlet combustor;the time averaged velocity slip between two phases and the fluctuation velocity of two phases are sufficiently high.These flow characteristics can be favorable to ignition,flame stabilization and combustion intensification.
ACOUSTIC EXPERIMENTS OF TWO SCALED-MODEL PROPELLERS ON THE GROUND
Li Xiaodong, Sun Xiaofeng, Hu Zong'an, Zhou Sheng
1993, 8(2): 121-124,201.
Abstract PDF(264KB) icon1354 icon379
Abstract:
Propeller acoustic exper iments have been completed on the ground.The experimental results show that the aerodynamic and acoustic performances of 4 bla des ARA D propeller are better than that of the 3 blades NACA64 propeller.The sound field of propellers has directional properties,and i t is dominated by discrete tones,especially the 1～5×BPG tones.So,increase of the blade number and application of the advanced aerofoil w ill be very effective to improve aerodynamic and acoustic performances of the propeller.The best way for propeller noise reduction and aircraft fuselage noise suppression is a try to reduce the propeller discrete tones,especially the 1～5×BPF tones.
OPTIMAL DESIGN METHOD FOR CHANNEL SIZE OF THRUST CHAMBER COOLING PASSAGE
Chen Jie
1993, 8(2): 125-128,201.
Abstract PDF(640KB) icon1569 icon505
Abstract:
High combustion chamber pressure engine scheme is to be used in new generation of liquid rocket engine with propellant LOX/hydrocarbon.For this kind of engine thrust chamber cooling is one of key technical problems.An optimal design method for geometric size of regenerative cooling channel is provided.The minimum of cooling pressure loss has been taken as an optimal objective.The results of a typical calculation show t hat the cooling pressure loss may be decreased 50% by using the optimal method.
A CALCULATION OF SECONDARY FLOWS AND DEVIATION ANGLES IN MULTISTAGE AXIAL-FLOW COMPRESSORS
Li Shiming, Chen Maozhang
1993, 8(2): 129-132,202.
Abstract PDF(278KB) icon2076 icon528
Abstract:
A secondary flow model and a method for calculating blade deviation angle have been developed for multistage axial flow compressors.This model is a modification of Adkins-Smith model.The secondary flow vorticities are modeled based on the cal culated meridional flow including different kinds of spanwise mixing effects.The secondary flows are obtained on the fanlike surface downstream the blade rows with a standard separation of variables approach which yields a complex series.The secondary flow model includes effects of blade passage vortex,blade trailing edge vortex,blade en d clearances,blade end shrouding,blade boundary layer and wake centrifugation.The deviation angles are related to the calculated cross passage secondary velocities and a modified Carter ’s rule.Good agreement has been obtained between experimental data and the calculated results in the cases of design-point-type applications for which significant regions of separated flow are not present.
CALCULATION OF TWO-PHASE FUEL DISTRIBUTION IN AFTERBURNER FLAMEHOLDER PLANE
Gu Shanjian, Wan Jian, Yang Maolin, Huang Yong
1993, 8(2): 133-137,202.
Abstract PDF(319KB) icon1921 icon406
Abstract:
The effects of the air velocity,the air temperature,the injector pressure drop and the flameholder position on the f uel distribution in the flameholder plan e were investigated.It was found that the penetration and the diffusion of the fuel vapor were differed greatly from the fuel droplets.The geometry of the flamehold er has strong influence on the air flow field,and therefore,on the two phase fuel distribution.There a re two different trajectories in a hot a ir stream.A semi empirical method has been developed to predict two phase fuel distribution in the flameholder plane,based on the trojectory diffusion model considering above factors.
SIMPLIFIED MATHEMATICAL MODE L AND DIGITAL SIMULATION OF AEROENGINE
Zhao Ziyuan, Yao Hua, Li Weiye
1993, 8(2): 138-142,202-203.
Abstract PDF(733KB) icon1916 icon393
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Two real time、large deviation mathematical models of an aeroengine are investigated to develop an airborne simulator,Full Authority Digital Electronic Control (FADEC) system.This study includes three parts:(a)to derive two mathematical models based on aerothermodynamics;(b)to estimate the engine performance by the method of digital simulation;(c)to dis cuss theoretical range of the calculated interval of the dynamic cycle and the model influenced by the fuel control regulation law.Two models are non linear、non iteration and full range.Simulation results show that the accuracy of models is quite good,and most relative errors are less than 3 percent,the biggest is about 5 percent.The calculated interval was reduced to 16ms(on the PC/AT with an Intel 80287 floating point coprocessor and C langu age programming).
AN EXPERIMENTAL INVESTIGATION OF ENDWALL FLOW CONTROL IN A COMPRESSOR PLANE CASCADE WIND TUNNEL
Li Xihong, Wu Guohua, Peng Zeyan
1993, 8(2): 143-147,203.
Abstract PDF(306KB) icon2090 icon439
Abstract:
A comaprative experimental investigation on three kinds of end treatment cascade in relat ion to an original cascade is presented. Three end treatment techniques are end bend,end forward sweep,and end sweep bend.The last one is a new kind of endwall flow control techniuqe.Experimental results s how that:(1)The end forward sweep cascade can be used to delay flow separation and improve the distrib ution of flow parameters at the cascade exit;(2)For a shrouded stator,the end bend cascade with increasing turning angle at the end of the blade would increase the loss around the endwall region and induce flow separation in the mid span region;(3)For the case when it is required to increase the turning angle in the endwall region,the end sweep bend cascade has better performance than the end bend blade.
EXPERIMENTAL INVESTIGATION ON EFFECT OF SOLID PARTICLES ON BLADE PRESSURE DISTRIBUTION IN COMPRESSOR CASCADE FLOW
Yi Jinghai, Ma Caifen, Gao Jingshi, Xu Zhong
1993, 8(2): 148-150,203.
Abstract PDF(222KB) icon1721 icon496
Abstract:
The blade pressure distribution of gas particle two phase flow in a compressor cascade has been investigated experimentally.Detailed measurements were made for three mass concentrations 1 1%,2 0% and 4 5%,and for two inlet velocities 10 and 22m/s,and for two inlet angles of attack 7° and 27 5°.The results for two inlet angles of attack 7° and 27 5° indicate that the solid phase has certain effect on the blade pressure distribution of the gas phase .With increasing inlet velocity and mass concentration,the effects of solid particles on the blade pressure distribution are enhanced and wake deviation is obvious.Also,the variation of the effects depends on the inlet angle of attack.
DYNAMIC ANALYSIS OF A GEAR DRIVE SYSTEM IN AEROENGINE
Yuan Xiangdong, Zhou Chuanrong
1993, 8(2): 151-154,204.
Abstract PDF(273KB) icon1352 icon402
Abstract:
A new technique of structural dynamic analysis especially for analysis of gear drive system has been developed,which combines the principle of substructure analysis with the idea of structure modification.Meanwhile the gear stiffness model about the de grees of the freedom of the structure node in which the gear relation exists has also been established.It has the advantages of both the method of substructure analysis an d vibration reanalysis of modified structure.The natural frequencies and modes calculated by the program written by author are in quite good agreement with the results calculated by finite element method.
EXPERIMENTAL INVESTIGATION ON PATTERNED BLADES OF COMPRESSOR
Miao Runtian, Wang Liangui, Qian Luhong, Tang Zhiming, Gao Ge
1993, 8(2): 155-157,204.
Abstract PDF(229KB) icon1568 icon381
Abstract:
The presented experimental investigation on patt erned blades of compressor opens up a new way for improving performance of compressor.The experiments were carried out with a HP compres sor of a twin spool turbojet engine in a multistage compressor rig.The HP compressor consists of three stages(named the 4th,5th and 6th stage).Two kinds of pattern presented in the paper,as two design,are all engraved on the surfaces of the existent 4th and 5th stage stator blades.The pattern Ⅰ is 45° cross grooves and the pattern Ⅱ is streamwards grooves.The experimental data of the efficiency η and surge margin SM are listed below as compared with the reference compressor for uniform inlet flow: [WT6BZ]Pattern ⅠPattern Ⅱ[WX(!3KG2,4KG2,5KG2,4KG2,5]△η(%)△SM(%)△η(%)△SM(%)0 9+2 5+2 2+3 5+1 31 0+2 9-0 1+3 1+2 81 05+1 1-2 1+3 7+3 31 1+1 2-3 0+1 9+3 6 The performance with circumferential distorted inlet flow is kept the same with the reference compressor.The above experimental results show that the efficiencies and surge margines of the compressor with appro priately patterned stator blades can increase significantly.
INFLUENCE OF RADIATION HEAT LOSS OF ENCLOSURE ON WALL TEMPERATURE OF JET NOZZLE
Zhang Qingfan, Wu Yanming, Zou Jianjun
1993, 8(2): 158-160,205.
Abstract PDF(221KB) icon1681 icon407
Abstract:
Using the appa rent emissivity,the net heat loss of the radiation from the enclosure surface due to passing through t he open end of the enclosure was calculated. The influence of the net heat loss on surface temperature was also d iscussed.Results showed that the net heat loss on the downstream side of the jet nozzle could reach 20～30% of total heat flux.Therefore it can not be neglected in calculating the wall temperature.
NUMERICAL ANALYSIS OF AERODYNAMIC LOSSES IN FILM-COOLED VANE CASCADE
Jiang Tao, Fan Weihong, Liu Son gling
1993, 8(2): 161-164,205.
Abstract PDF(275KB) icon1928 icon491
Abstract:
Intergal Boun dary Layer Parameter" method based on a 2 dimensional film cooling flow and heat transfer prediction program STANCOOL is provided to pred ict the aerodynamic losses in the film cooled vane cascade.The comparison of the numerical analysis with the corresponding experimental data demonstrates that the present method is more suitable to the case when blowing rates are low or the film cooling is located at thick local boundary layer.Finally this method is compared with another loss predicting method,named "TOTLOS",and it is concluded that both methods have their own advantages and extent of application.In practice it is reasonable to use the two methods in combination.
EXPERIMENTAL STUDY ON HEAT TRANSFER OF SEPARATED IMPINGEMENT JETS IN SHORT DISTANCE
Kang Ying, Qiu Xuguang, Zhou Wei
1993, 8(2): 165-168,205-206.
Abstract PDF(309KB) icon1404 icon405
Abstract:
In actual cooling construction of turbine blade the relative distance between circular jets c n/d n is less then 3 and the relative distance of nozzle exit to target surface is also less than 3.But the available empirical formulas were achieved on the basis of very large range of these geometric parameters.Obviously,these formulas may be inaccurate in the practical narrow range of th ese paremeters,in particular when the temperature of the turbine is very high,and on the surface of the blade exist many rows of film co oling holes.So an experimental research on the heat transfer of impingement jets has been carried out in the actual range of the geometric parameters and new empirical formulas have been established,which can be applied to the turbine blade o f the advanced aeroengine.
A NEW WAY TO SOLVE STEADY CREEP PROBLEM WITH FINITE ELEMENT METHOD
Mu Xiaying, Wang Jiwei
1993, 8(2): 169-172,206.
Abstract PDF(268KB) icon1354 icon462
Abstract:
In this paper,a new way to solve steady creep with finite element method is presented for temperature dependent material properties of symmetric body. According to different material properties,initial strain rate method and general tangential stiffeness method are provided.A computing program is compiled.Some examples are analysed numerically to check its convergency and applicability.
ON ENGINE PARAMETER ESTIMATION WITH FLIGHT TEST DATA
Fan Siqi, Rong Xiangjun, Li Shuren, Xin Xiaowen
1993, 8(2): 177-178,206-207.
Abstract PDF(273KB) icon2027 icon453
Abstract:
This paper introduces the parameter estimation research for the turbine inlet temperature and compressor exit pressure by using the engine variables measured under the different flight conditions.The parameter estimation method is multiple linear regression analysis.The independent variables in the regression equations have been selected from six measured engine variables based on the comprehen sive optimal principle about variable selection.The regression models which have been established only include three measured engine variables.The curves of the parameter estimation results shown in this paper demonstrate that the estimation accuracies of these mode ls are satisfactory.This paper also presents the data treatment meth ods of engine variables recorded on line with special equipment.
AN AERODYNAMIC DESIGN PROGRAM FOR CONTRA-ROTATING TURBINE CASCADES
Cai Ruixian, He Yongmei, Wei Xinglu
1993, 8(2): 179-183,207.
Abstract PDF(228KB) icon1977 icon430
Abstract:
An extremely simple,rapid and universal turbine cascade design program is develope d for contra rotating turbine cascades.With its simplest operating condition,only four data are necessary to be input for obtaining the geometry of a good enough cascade.They are just the input data for designing a cascade-the inlet Mach number and flow angle,the outlet Mach number and the specific heat ratio of working substance.Therefor e,it means that the design can be accomplished without any additiona l human decision.The program is based on the Mean-Stream-Line Method .It is given in detail how to select the independent design functions of this method in programming.This program enables to be operated with a very small computer,even a pocket computer such as the Sharp PC 1500.
R & D OF LIGHT WEIGHT GAS TU RBINE IN CHINA
Wang Zuhu, Li Huaiti
1993, 8(2): 184-186,207.
Abstract PDF(318KB) icon1456 icon465
Abstract:
This paper introduces the engineering research o f aeroderivative gas turbines in recent ten years in aspects of such as generator modified;power turbine research;alternative fuel combustibility.The advanced and applied researches aim at meeting the demands for industrial and marine propulsion,such as prolonged life,improved performance,enhanced reliability and satisfactory variability.
COUSTICAL PROPERTIES OF SOUND ABSORBING STRUCTURES AT HIGH TEMPERATURE
Liu Wei, Tang Daomin
1993, 8(2): 187-190,207.
Abstract PDF(237KB) icon1468 icon354
Abstract:
The acoustical properties of sound absorbing layers used in gas turbine exhaust silencing component at hi gh temperature with temperature gradient were investigated by analysing the effect of temperature on t he acoustical properties of porous materials.Some experimental results are given and a good agreement is found between theoretical prediction and measurements of acoustical properties at various temperatures up to 500℃.
PERFORMANCE IMPROVEMENT OF GAS TURBINE WITH STEAM INJECTION
Qian Yansheng, Li Huaiti
1993, 8(2): 191-194,208.
Abstract PDF(267KB) icon1828 icon394
Abstract:
This paper describes how to choose and modify the WZ 5 gas turbine for steam injec tion testing and how to evaluate testing parameters and programmes.Testing results indicate that when the steam is injected 3 6t/h into the WZ 5 gas turbine,its power output and thermal efficiency increase 40% and 25% respectively.Moreover,the steam injection does not dete riorate the exit temperature profile in the turbine,which ensures a long term reliability of operation for the engine.
EXPERIMENTAL RESEARCH ON A SEMIWATER-GAS-FIRED GAS-TURBINE
Zhou Kangcheng
1993, 8(2): 195-198,208.
Abstract PDF(255KB) icon1376 icon379
Abstract:
A turboshaft aeroengine WZ5 was modified into a gas turbine working with semiwater gas.The features of its modification and its fuel gas supply system are described,testing program and results of its performance tests are discussed.It is shown that its starting reliability,burning stab ility,acceleration response,distributional quality of the exit temperature field and the operating performeance are reasonable,some of them are even better than the original WZ5 engine working with aviation kerosene.

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