1993 Vol. 8, No. 1

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STUDY OF OPTICAL VISUALIZATION METHODS FOR EXHAUST JET OF ROCKET ENGINE
Yan Dapeng, He Anzhi, Miao Pengcheng, Wang Hailin
1993, 8(1): 1-5,89.
Abstract PDF(351KB) icon1581 icon463
Abstract:
Two optical methods are studied which are used to visualize the flow field of exhaust free jet and jet impingement from a real solid rocket engine.One is equal thickness fringe F P interferometry which is constructed with two mirrors about 200mm in diameter.It features that the optical layout is co optical light path and seperated form,the distance between two mirrors is more than 20m and the resistance to vibration distortion is strong.Another is moire deflectometry which has an aperture over 400mm in diameter.In this method a big diameter collimated light beam is used to explore the flow field,then it is reduced and recollimated into a small diameter recollimated light beam and detected by a two gratings system to produce moire fringes.It is characterized by large aperture and simplicity of the method.The experimental results show that the equal thickness fringe F P interferometry can be used to visualize the flow field of the exhaust jet from a real solid rocket,and the moire deflectometry can be even used to visualize the extensive flow field of rocket exhaust jet impingement besides the exhaust free jet.
NUMERICAL SIMULATION OF INTERACTION BETWEEN SUPERSONIC MAIN STREAM AND TRANSVERSE JET
Zhou Weijiang, Ma Hong, Li Feng
1993, 8(1): 6-10,89-90.
Abstract PDF(339KB) icon1512 icon516
Abstract:
The laminar N S equations have been solved for a supersonic flow with a transverse sonic jet.The numerical algorithm used here is a second order windward TVD scheme.The conditions of supersonic main stream are: Ma ∞=2 5,T ∞ =130K, Re ∞=5×10 5.In sonic jet, T j=243(T j/T ∞=1 87),P j/P ∞ =43 0.The grid nodes are 121×95 with compressed in both a transverse direction near the wall and a straemwise direction near the injector.The detail waves and vortex structures are obtained,the flow field pictures and pressure distribution along plate are compared with experiments in a good agreement.The calculation also demonstrates that the TVD scheme not only has high capability of shock capturing,but also can be applied to efficient simulation of the complex vortex structure.
COMPUTATION OF FLEXIBLE-WALL AIRFOIL FLOW USING N-S EQUATIONS
Ma Xia, Xiao Linkui, Jiang Jinggui, Chang Mianchun
1993, 8(1): 15-18,90.
Abstract PDF(352KB) icon1552 icon486
Abstract:
The 2 D Reynolds averaged compressible unsteady full Navier Stokes equations are solved by using the Beam Warming scheme with Baldwin Lomax turbulent model about the flexible wall airfoil.A method is proposed which can treat the locally changing surface of airfoils.The adaptive surface in the shock region can influence the shock/boundary layer interaction.Comparison has been made between the flow fields over NACA0021 airfoil with and without the flexible wall.Since we use the time marching method to get the steady results,the main difficulty in the present case is the treatment of the flexible wall boundary which changes its shape as the solution proceeds.The deflections of the flexible wall as function of the pressure over it can be found by the mechanics of the material.We proposed a method of locally changing grids to avoid the global interpolation.The locally changed grids maintain the good quality of the original grids.The calculated results show that a strong shock splits into two weak shocks in comparison with the case without the control.
A MATHEMATIC MODEL AND A CALCULATION METHOD FOR COOLING AIR MIXING IN A TURBINE
Yang Hong, Wang Zhongqi, Feng Guotai
1993, 8(1): 19-20,90.
Abstract PDF(292KB) icon1933 icon389
Abstract:
A mathematic model is proposed for predicting the effect of injected cooling air on cooled turbine performance.The model takes the gas flow as the mainstream flow and regards the cooling air as a consecutive source with continuous distribution in the flow field.A numerical method is presented for determining the intensity of the consecutive source.According to the model,the coupled aerothermodynamic equations reflecting the effect of the cooling air mixing in general form are derived in a relative coordinate system.As an example,the model is used in the throughflow calculation for the cooled turbine and the governing equations are derived along S 2 stream surface in a non orthogonal curvilinear coordinate system.By defining a quasi stream function which satisfies the continuity equation,a line relaxation calculation method is developed for the quasi stream function.It makes full use of the existed calculation method without considering the cooling air mixing and is easy to be programed.In accordance of the mathematic model a formula of entropy variation due to the cooling air mixing is derived in the light of the second law of thermodynamics.By means of the present calculation method,an air cooled gas turbine stage is analyzed.The results show the reasonableness of the mathematic model and the feasibility of the calculation method.
EXPERIMENTAL STUDY ON TURBULENT JET IN A CONFINED CROSSFLOW
Xu Minghou, Han Caiyuan, Hu Pingfan
1993, 8(1): 21-23,91.
Abstract PDF(196KB) icon1679 icon460
Abstract:
A turbulent triangular jet issuing perpendicularly into a confined gas particle two phase flow was investigated experimentally.When jet to crossflow velocity ratio R equaled to 2 5,mean and fluctuating velocity components were measured by the Three Dimensional Particle Dynamics Anemometry (3 D PDA).The PDA signals were evaluated to yield the turbulent intensity and the particle concentration.The measurement results show that the turbulent intensity in the wake of the crossflow is enhanced particularly on the edge of the recirculation zone.There are two particle concentration peaks,one of them is in the recirculation zone,and the other is on the edge of the zone.
INVESTIGATION OF STATIC PRESSURE FIELD IN AN S-SHAPED AIR INTAKE
Weng Peifen, Guo Rongwei
1993, 8(1): 24-28,91.
Abstract PDF(253KB) icon1647 icon468
Abstract:
The static pressure fields in an s shaped diffuser with rectangular cross section have been measured at different incidences.The results show that the static pressure field at the duct exit is non uniform.The higher the angle of attack is,the larger the distortion of the static pressure field.The pressure near the boundaries of the model is higher than that near the core,which is mainly caused by both the distorted total pressure field and the distorted flow velocity field.Further investigation on whirling flow field demonstrates that it is neccessary to take into serious consideration the deviation between the experimental measurements and the weighted averages by the use of the wall static pressures of the model at high incidence.
RESEARCH OF ONSET OF ROTATING STALL FOR MULTIPLE BLADE ROWS
Zhu Jiunqiang, Liu Zhiwei
1993, 8(1): 29-33,91-92.
Abstract PDF(330KB) icon1785 icon473
Abstract:
An inception criterion of rotating stall in multistage axial flow compressors is derived for the unsteady,two dimensional and incompressible flow model according to the small disturbance stability theory.By means of a three vector equation system,the correlative equations of the undetermined parameters are thus deduced when the small disturbance theory is applied to multiple blade rows.The prediction of the rotating stall onset boundary in the multistage axial flow compressors has been accomplished with the computer program designed by the authors.A theoretical prediction approach is proposed with the aid of the prediction method of stable cascade characteristics.A good agreement between the analytical and experimental results indicates that the theoretical analysis is believable.As compared with the calculation results,the influences of the undetermined parameters,such as stall cells and axial distance between two stages,on the prediction values are discussed.
ESTIMATION OF FLOW LOSS AT DESIGN POINT IN AN AXIAL-FLOW COMPRESSOR
Zhang Changsheng, Wang Xiaoli
1993, 8(1): 34-36,92.
Abstract PDF(1452KB) icon1745 icon508
Abstract:
On the basis of statistical data and analysis of axial flow compressors,the ratio of trailing edge momentum thickness to chord length θ/c and trailing edge form factor H are determined by the use of equivalent diffusion factor D eq ,and then the profile loss of the blade is obtained after appropriate corrections.Finally,the total loss distribution along the blade span of the axial flow compressor is estimated according to an improved end wall loss model and a shock loss model.An extensive comparison of the estimated loss data with the original design and test data shows that the loss model presented here is feasible for normal blade profile,double circular arc blade profile and multi circular arc blade profile of axial flow compressor,and the accuracy of the estimation is improved greatly.As a part of S 2 flow surface program,a computer subprogram has been worked out based on this approach.It gives satisfactory results of the flow field computation and is also very simple and practical.The subprogram can meet all requirements of aerodynamic design and analysis for an advanced axial flow compressor.
EXPERIMENTAL STUDY ON UNSTABLE BEHAVIOR OF AXIAL COMPRESSION SYSTEM
Li Jun, Liu Zhiwei
1993, 8(1): 37-40,92.
Abstract PDF(237KB) icon1368 icon453
Abstract:
Unstable behavior of an axial flow single stage compressor has been researched successfully by use of a simple dynamic measuring system.The character and the influence factors of the unstable behavior have been analyzed from systematic view.The experimental results demonstrate that the unstable types of the compressor could convert each other in certain conditions.The compressor rotor speed,the construction and size of the exhaust parts are the main factors influencing the convertion.The surge occurs in the compressor on condition of high rotor speed or large exhaust volume,otherwise the rotating stall appears.The decrease of the exhaust volume leads to increment of the surge frequency.There is an obvious difference of the unstable behavior between the single stage compressor and the isolated rotor compared.In the condition of high rotor speed ( >0 85),a stall phenomona of nonrecoverable stall appears.It's character is also considered.
EXPERIMENTAL INVESTIGATION ON WHIRL FLUTTER OF PROPELLER ROTOR SYSTEM
Li Zhiwei, Xu Ming, Ren Xingmin
1993, 8(1): 41-44,92-93.
Abstract PDF(285KB) icon1370 icon434
Abstract:
A whirl flutter of a propeller rotor system has been investigated experimentally in the wind tunnel.The design of whirl flutter experimental equipment and a dynamic stiffness coefficient measurement method for the supporting system are also studied.The experiments show that if the support stiffness of the propeller rotor system is reduced as a result of failure of some elements in the engine frame or mount,the threshold flight velocity of onset whirl flutter will be decreased.The results of its wind tunnel experiments exhibit the same trends as in its theoretical calculation.
STUDY OF A SUBSONIC COMBINED-LEANING TURBINE GUIDE VANE WITH TIP ENDWALL CONTOURING
Jiang Hongde, Lu Yongliang, Zhou Fanzhen, Wang Songjun
1993, 8(1): 45-48,93.
Abstract PDF(350KB) icon1907 icon373
Abstract:
A three dimensional,multigrid potential flow computation method is developed to analyse subsonic flow in steam turbine nozzle guide vane.In order to reduce the large secondary losses in the first annular nozzle guide vane with very small aspect ratio of 0 26 in a 300MW steam turbine,both tip endwall contouring and combined lean blades are used in the new design.Experimental investigation shows that the mass weighted loss coefficient of the new nozzle guide vane is 42 percent less than traditional radial one.The agreement between the measured and calculated results of the pressure distribution on the blade surface in the annular nozzle guide vane is very good.The present study is a success in developing fully 3D steam turbine nozzle guide vane with tip endwall contouring and combined leaning blade.
SUPERSONIC UNSTALLED FLUTTER IN ARBITRARY MISTUNED CASCADES
Li Qinwu, Hen Qun, Fan Feida
1993, 8(1): 49-52,93.
Abstract PDF(291KB) icon1851 icon381
Abstract:
By the use of Hamilton's principle,a set of aeroelastic equations is formulated to describe the motion of an arbitrary mistuned rotating cascade with high aspect ratio,flexible pretwisted,nonuniformed blades.A two dimensional aerodynamic theory is used to calculate unsteady force and moment coefficients for vibrating aerofoil in supersonic unstalled cascade.This set of equations is of great value to a fan blade with the inertial,structural and aerodynamic coupling in bending bending torsion.Results obtained from the aeroelastic stability analysis of a real fan stage show that amplitude and different order of mistuning and other kinds of structural properties have an apparent effect on boundaries of supersonic unstalled flutter.
RESEARCH ON 3-D DISCHARGE FLOW IN A CENTRIFUGAL IMPELLER
Xi Guang, Wang Shangjin
1993, 8(1): 53-55,93.
Abstract PDF(292KB) icon1510 icon401
Abstract:
A method for calculating 3 D discharge flow in a centrifugal impeller is developed by means of solving the incompressible steady time averaged N S equations,the continuity equation and the k ε turbulence model equation.This method is accomplished with calculation grids generated by an algebraic transformation,and thus it can be easily applied to the design analysis of centrifugal impellers.The discharge flow of a three dimensional shrouded centrifugal compressor impeller was measured for assessing the validity of the present calculation procedure.The comparison between the experimental data and the calculation results shows that the present method can predict the mean velocities.
NUMERICAL STUDY ON BLADE-TO-BLADE PASSAGE FLOW FIELD IN A CENTRIFUGAL CASCADE
Wu Xiaosong, Wu Guochuan
1993, 8(1): 56-58,94.
Abstract PDF(249KB) icon1508 icon384
Abstract:
Numerical simulations of the blade to blade flow field in a centrifugal impeller have been carried on by using 2 D mean steady incompressible Navier Stokes equations and k ε turbulent model in arbitrary boundary fitted coordinate system.It is shown that the underrelaxation factors of line by line iteration are lower than those in non rotational flow calculations and a more reasonable initial flow field is required due to the larger ratio of centrifugal forces in the source terms of momentum equations.The calculation results of the flow fields are in a good agreement with the LDA data with the exception of the near wall region of the blade surfaces.Some of the discrepancies in the near wall region are caused mainly by the assumption of 2 D flow and ignoring the rotational effects in the turbulence model.Additionally,the calculation precision is also affected,since inlet and outlet boundaries of the computational domain cannot be located so far away from the blades.
EXPERIMENTAL STUDY ON PROPELLER FENCE FOR EFFICIENCY ENHANCEMENT AND NOISE REDUCTION
Mao Xichang, Li Hongmin, Wang Jiebing
1993, 8(1): 59-62,94.
Abstract PDF(795KB) icon1562 icon420
Abstract:
Three blades scale model propellers with and without propeller fence along the chord direction at 88% tip blade radius are compared experimentally in the wind tunnel in respeet of aerodynamic and acoustic characteristics.Their rotational speeds vary between 1500rpm to 3000rpm.Experimental results indicate that under conditions of identical advance ratio J the efficiency of the propeller with propeller fence is enhansed 1 9%～4 9% and its lower noise level is reduced 13～20dB.It is shown that the propeller fence is one of effective measures to increase efficiency and to decrease noise of propellers.
ON MODELLING CONDITIONS FOR COOLING EFFECTIVENESS EXPERIMENTS
Qiu Xuguang
1993, 8(1): 63-66,94.
Abstract PDF(330KB) icon1293 icon395
Abstract:
According to strict theory of similarity,the modelling conditions for model experiments on cooling blades or vanes of a gas turbine are obtained from a necessary set of differential equations and single valued conditions.It is shown that when the pressure in modelling is lower than that in the actual turbine,the temperature in modelling must be lowered simultaneously in order to maintain the criterion on pl/μc PT for cooling air and for gas invariant.Therefore,it is necesary to increase the pressure and to decrease appropriately the temparature in modelling tests for the blades and vanes of high pressure turbines with high pressure ratio.
SURFACE TEMPERATURE MEASUREMENT OF TURBINE DISKS
Wu Hongdao, Qu Yuwu, Li Xungguang, Du Shengqin
1993, 8(1): 67-68,94-95.
Abstract PDF(285KB) icon1924 icon402
Abstract:
A new method of temperature measurement with a single wire thermocouple-slip ring system is first introduced to measure the surface temperature of the turbine disks in the turbojet engines and the accuracy of this method is considered in this paper.In this case the limited channels of the lip ring are fully utilized and the measured surface temperatures of the disk are actual and believable enough.The problems in its application such as the installation of the thermocouples,the cooling of the slip ring and the balance of the turbine rotor with this system are also discussed briefly.
NUMERICAL STUDY OF TURBULENT COMBUSTION MODEL
Zhao Jianxing
1993, 8(1): 69-71,95.
Abstract PDF(248KB) icon2227 icon380
Abstract:
Although the k ε g turbulent combustion model becomes widespread in engineering application,but the coefficients of the model C R,C g1 ,C g2 and expressions of the square of concentration fluctuations g used by researchers are different.In fact,the effects of these coefficients and g expressions on reaction rate and temperature profiles are significant.This paper aims at providing optimal values of these coefficients and rational g expression.Numerical calculations were carried out for recirculating reacting flow behind a bluff body burner.It is shown that C R is given by C R=(ΔT /1000) 2g 1/2 /g 1/2 max ln (1+2g 1/2 max /g 1/2 ), C g1  is defined as C g1 =2/σ f, C g2  is 2 in this prediction. The prediction is in favourable agreement with the experimental measurements.
A CALCULATION RESEARCH ON BURNING RATE OF ALUMINIZED COMPOSITE PROPELLANT
Liu Dehui, Peng Peigen
1993, 8(1): 72-74,96.
Abstract PDF(240KB) icon1413 icon439
Abstract:
A research has been completed by appling PEM model to combustion simulation of aluminized propellants.The effects of aluminium particles on the surface energy of the propellant are taken into consideration so simply as the heat sink PEM model of J.P.Renie et al,but a new method to determine the pseudopropellant's formulations is presented in this paper,thereby an extended and modified PEM model is developed.According to this model,the burning rates and pressure exponents of aluminized propellants are calculated.The results show that our model makes the calculated results of burning rates and pressure exponents more approach to their experimental data as compared with the heat sink PEM model of J.P.Renie et al.
FUEL CONTROL FAILURE RESEARCH FOR JT8D ENGINE
Xu Chunsheng, Yu Zhili, Ma Qianchao
1993, 8(1): 75-78,95.
Abstract PDF(1078KB) icon1530 icon380
Abstract:
A failure research of the fuel control in JT8D turbofan engine has been completed,which results in effective reduction of its unscheduled removal rate and sets up the foundation for domestic fuel control repairs.The theoretical analysis,simulation,adjustment tests,establishing and practising the diagram of failure insulation are applied to the failure research.Its economic and social benefits are obvious.
DESIGN PRINCIPLE OF CHANGE-SPEED INVOLUTE CONICAL GEAR TRAIN
Shen Qian
1993, 8(1): 75-78,95.
Abstract PDF(298KB) icon1384 icon376
Abstract:
A theory of hobbing involute conical gears on a hobber with a horizontal feed motion is described in the light of the gear engagement principle.The relations among some important parameters of the imaginary rack shaper cutter and the generated gear are analyzed,and the design method of a change speed train mainly composed of intersected involute conical gears is considered in detail.It has many special advantages such as many speed degrees only with a few gears,adjustable gear clearance by changing position of the gear along with their axes respectively.Because in designing this kind of train it is very important how to minimize the volumn and size and how to maximize the changable speed degrees and the difference between the highest degree and the lowest degree.A practical method is provided to control some key parameters to reach these goals.The present theory is proved correct and the method is shown practical by tests of a sample change speed train.
OPTIMAL TOOTH PROFILE MODIFICATION OF HELICAL GEARS FOR IMPROVING RESISTANCE TO SCUFFING
Fang Zongde, Shen Yunwen
1993, 8(1): 79-82,95.
Abstract PDF(342KB) icon1532 icon444
Abstract:
An approach is presented to increase scuffing load capacity of helical gears.Firstly,the flash temperature distribution on tooth is evaluated by using 3 D finite element method,mathematical programming technique and Block's equation.The calculated results show that the highest flash temperature occurs at the mesh begining and finishing corners on the tooth flank where scuffing could happen.Further,a tooth profile modification by a parabola is presented.Through the agency of optimization programming with an object function of minimizing the highest flash temperature,the optimal profile modification can be obtained.According to the numerical example,the scuffing limit load could be raised 2 5 times as unmodified helical gears.Finally,the influence of the supporting structure on the distribution of flash temperature is discussed.An optimal profile modification method is developed,which takes the deformation of supports into consideration.
PRIMARY RATED LIFE ESTIMATION FOR TURBINE DISC
Jia Jun, Long Shenggang, Tang Pingxuan
1993, 8(1): 83-86,96.
Abstract PDF(282KB) icon1804 icon461
Abstract:
The estimation of the primary rated life is studied for a aeroengine turbine disc.A method for this purpose is considered and put into practice.The rated life of turbine disc is predicted by local stress strain approach Larson Miller curve and Masson Linear rule.And then it is proved by tests.In the life test investigation of the turbine disc,the equivalent accelerated test loading spectrum provided according to equivalent rule (without changing dangerous position and damage characteristics) not only accelerates test,but also satisfies life test specifications.Besides,the turbine disc is enclosed in a case hotted up by middle frequency electro induction.This method improves hotting characteristics of a turbine disc tester.
RALIABILITY CALCULATION IN CONSIDERATION OF PHYSICAL PARAMETERS OF A MATERIAL AS RANDOM PARAMETERS
Lu Wenlin, Fu Shuang
1993, 8(1): 87-88,96.
Abstract PDF(201KB) icon1429 icon353
Abstract:
A method of reliability calculation for directionally solidified turbine blade has been provided,which mainly considers physical parameters as random by means of stochastic FEM and Hasofer Lind method.The method ensures convergency of its calculation,so it is feasible.The calculation results indicate that physical parameters as random ones have less influence on the failure probability than the loading and mechanical parameters.

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