1989 Vol. 4, No. 4

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A METHOD FOR AERODYNAMIC DESIGN CALCULATION OF AXIAL GAS TURBINE STAGES WITH COOLING AIR MIXING
Huang Zhonghu, Wang Yueqi, Yang Jinfu
1989, 4(4): 301-304,388.
Abstract PDF(307KB) icon1793 icon21
Abstract:
This paper persents a method for aerodynamic design calculation of axial gas turbine stage with the mixing of cooling air.The mass flow mixing and the energy mixing of the cooling air with the main gas flow are considered directly in the continuity and the energy equations,while the momentum loss resulting from the mixing is considered indirectly.The losses of blade rows and their distributions in the radial direction are determined by means of the loss models provided in the paper.The influences of different cooling types of blades and the non uniformity of the inlet total temperature and total pressure in the radial direction have been considered in the calculation program.The method could be applied to multi stage air cooled gas turbine.The basic calculation method adopted is the streamline curvature method.As an example,a design calculation of an axial turbine stage are presented.The calculation results show that the effect of the cooling air mixing on the turbine efficiency is significant.
AN AERODYNAMICAL DESIGN AND CALCULATION METHOD FOR GAS TURBINE WITH COOLING AIR MIXING
Shi Jing, Li Junshan, Wang Piyqia, Jin Shi
1989, 4(4): 305-309,388.
Abstract PDF(344KB) icon1776 icon15
Abstract:
In order to conduct aerodynamical design and calculation for a air-coolded turbine,the coolant and gas flow mixing correlations,which satisfied the mass.momentum and energy equations,are derived in rotational coordinates.The computer program has been developed.Calculational results indicate that gas flow parameters change considerably in the case of great coolant mass flow ratio.It is necessary to consider the influence of coolant flow in detail.
DESIGN OF AN APPROXIMATELY CONSTANT DEGREE OF REACTION TURBINE STAGE
Cui Jiya, Lin Pingji, Li Yuchun
1989, 4(4): 310-312,390.
Abstract PDF(261KB) icon1620 icon24
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This paper is a successful trial of an approximately constant reaction stage designed by through-flow streamline curvature method for the first time tension splines.The tip to root reaction difference within 0.1 is attained mainly by the nozzle vanes with positive lean angle of 15° at leading edge and a reverse to usual exit angle twist of 11.6°.To avoid undesirable tip clearance increase in hot running state which would lead to excessive leakage loss,the shroud diameter is kept essentially constant with only a slight taper from nozzle inlet to axial clearance mid-section and then turning outward to the rotor shroud constant diameter.The variations of main gasdynamic parameters,total pressure loss coefficients and stage efficiency along the blade height are shown in figures together with those for a corresponding conventional stage of same meridional geometry for reference purpose which is seen giving even a little (0.20%) higher efficiency.Nevertheless,owing to lack of accurate loss model,especially in regard to leakage loss and shock loss,the efficiency results including their comparison are subject to experimental verification,although the design of constant reaction in general is credible.
AN APPROACH FOR CALCULATING STEADY SUBSONIC AND TRANSONIC BLADE TO BLADE FLOWS
Wang Pingqia
1989, 4(4): 313-318,388.
Abstract PDF(408KB) icon1405 icon13
Abstract:
A software is designed to calculate steady subsonic and transonic blade flows.It can analyse the features of the flow field according to the boundary conditions and automatically choose the numerical method which is suitable to obtain the accurate results and save calculation time.When the coordinates on blade surface at inlet and outlet are given,user can determine Mach number or pressure distribution on blade surface,Mach number or pressure contours in all of the flow field.Two kinds of the numerical methods are programmed in this software:Streamline curvature method and Time-matching method.The former directly starts from steady Euler equations,since the multi-value problem in transonic flows was solved by the author,the streamline curvature has been successfully used to compute transonic flow,especially it gets rapid covergence when the flow is subsonic at outlet.The later method starts from unsteady Euler equations but with constant total enthapy.The steady flow is calculated as the asymptotic limit after a sufficient long time of solution of the unsteady equations.This method with appropriate scheme and shock capturing technique is accurate to compute transonic flow with shock,but in the case of subsonic at outlet,the streamline curvature method can reduce the calculation time approximately 60%.Three numerical examples are given in this paper.Both the calculation methods give comparable results and the results agree well with the measurements.
COMPUTATION OF TRANSONIC FLOW IN A PLANE CASCADE WITH AN UNFACTORED FLUX SPLITTING IMPLICIT METHOD
Huang Dongtao, Shen Mengyu, Zhang Yaoke
1989, 4(4): 319-324,389.
Abstract PDF(407KB) icon1641 icon12
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MacCormack′s flux splitting method has proven itself guite fast converging in numerical simulation of flows in converging-diverging nozzle.In this paper a development on the basis of this method is presented for Euler equations in order to obtain numerical solutions of inviscid transonic flow through a two-dimensional turbine cascade.In the present finite-area flux splitting implicit method the Gauss-Seidel line relaxation recommended by R.W.MacCormack is applied to solving the implicit descretization equations.But we abandon the predictor-corrector two-step method to cut down the computational time in every time-step,because it cannot enhance the accuracy of the steady soluton in flux-splitting methods.In our practical computation,in order to improve the accuracy of the solution,the right side of the difference equation is discretized to be of second order accuracy,the blade wall boundary conditions and the numerical results show that the present method retains the effective convergence rate (only about 30 timesteps to attain a steady solution)and yield numerical solutions conformable with the experimental results.
A RELAXATION METHOD FOR TRANSONIC POTENTIAL FLOWS THROUGH 2-D CASCADE WITH LARGE CAMBER ANGLE
Wu Guohua, Peng Zeyan
1989, 4(4): 325-328,389.
Abstract PDF(1195KB) icon1711 icon12
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A new trnsonic relaxation method is presented which can be used for computation of transonic,potential flows through two-dimensional cascade with large camber angle.A non-orthogonal mesh system composed of streamlines and straight lines parallel to y-axis is employed here.The governing equation expressed by the streamline coordinate system is solved in physical plane.Because of the streamline coordiate system,the governing equation is greatly simplified and the formulation of the finite difference scheme is also made correctly and easily.In the case of the cascade with large camber angle (especially for turbine) direct formulation of difference scheme from full-potential equation (instead of perturbation-potential equation) is suggested. The numerical experimentations show that both the convergence and the stability of the finite difference scheme proposed here are satisfactory.The compuation results with acceptable accuracy can be obtained after 60 to 90 relaxation steps.The numerical examples also show that the results are in good agreement with experimental data and analytical solution (for Hobson-airfoil).
EXPERIMENTAL INVESTIGATION ON THE PERFORMANCE OF AN ANNULAR NOZZLE CASCADE of A HIGHLY-LOADED TRANSONIC TURBINE-STAGAE
Zhou Shiying, Li Jinnian, Huang Zhonghu, Yang Xuezhen, Zhang Tiansong
1989, 4(4): 329-332,389.
Abstract PDF(295KB) icon2255 icon18
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An annular nozzle cascade of a highly-loaded transonic turbine-stage has been tested in annular cascade wind tunnel,and its two plane cascades at the hub-and mid-sections—in a plane cascade wind tunnel.Their loss performance,exit flow angle,exit flow velocity as well as the flow velocity distribution on the blade-profile surfaces obtained in both cases are analyzed comparatively.It is proven that the experimantal results of the annular nozzle cascade are comprable with those of the plane cascades,particularlly,in quite good agreement at the mid-section.But at the hub-section a obvious difference between the experimental of both results occurs because of the local flow complexity in the annular nozzle cascade.
AN EXPERIMENTAL STUDY OF TIP CLEARANCE EFFECTS ON THE PERFORMANCE OF AN AXIAL TRANSONIC TURBINE
Huang Zhonghu, Wang Yueqi, Han Jianyuan, Yang Jingfu, Ren Rugen
1989, 4(4): 333-336,390.
Abstract PDF(305KB) icon1706 icon16
Abstract:
In order to investigate the effects of the rotor blade tip clearance on the performance of a highly-loaded axial turbine,a single-stage axial transonic scaled-down turbine with low aspect has been tested with four rotor tip clearances.The reduced revolutions of the turbine vary from 80% to 110% and the turbine total pressure ratios vary from 2.0 to 4.0. The experimental results are presented and analysed in the paper.The tip clearance effects on the turbine efficency are predicted with five empirical and semiempirical and the results predicted fomulas are compared with the present experimental results.The experimental results show that the effect of tip clearance on the turbine efficiency is significant,but that on the turbine inlet mass flow is small.When tip clearance increases from 1.054% to 2.823%,the turbine efficiency decreases by 5.0% at the design point,but the mass flow increases only by 0.54%.
EFFECT OF VANE AND BLADE NUMBERS ON PERFORMANCE OF TRANSONIC TURBINE STAGE
Han Jianyuan, Shi Jing, Cheng Xinhua, Wang Yueqi, Zhou Shiying
1989, 4(4): 337-340,390.
Abstract PDF(276KB) icon1617 icon16
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The parameters,such as throat area,trailing edge block ratio and profile turning angle at trailing edge,has been investigated with change of vane and blade number,namely 29,26 vanes and 47,52,42 blades.The results of the effect of the vane and blade numbers on the mass flow,reaction and efficiency of transonic turbine stage are analyzed.
DESIGN AND CALCULATION OF COMPOSITE AIR-COOLED BLADES IN A HIGHLY-LOADED TRANSONIC TURBINE
Deng Huayu, Liu Cunlu, Luo Mingjun, Liu Yufang, Peng Qingming, Sheng Rongchang, Gao Yuhua, Hu Zhiqing, Wang Huage, Li Xingheng, Zou Xueyan
1989, 4(4): 341-343,391.
Abstract PDF(237KB) icon1792 icon16
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A set of calculation methods is presented for predicting coolant mass flow distribution,temperature rise、heat transfer coefficients (internal and external) and temperature field of transonic turbine blades with impingement-convection-film cooling.Various air-cooled schemes has been calculated,in result four internal air-cooling structures in turbine blades have been chosen accordingly and designed elaborately.The results of calculation according to actual parameters are compared with the measured wall temperature distributions over turbine blades.Maximum and minimum deviations of the local wall temperature over the turbine vanes and blades are 9.8% and 2%,6.4% and 1.3% respectively.The methods are verified to be credible and accurate.
CALCULATION OF COOLANT FLOW AND HEAT TRANSFER INSIDE COMPOSITE AIR-COOLED TURBINE
Deng Huayu, Liu Yufang, Luo Mingjun, Peng Qingming
1989, 4(4): 344-346,391.
Abstract PDF(1220KB) icon1765 icon17
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A calculation method is presented for iterative solution of the coolant mass flow distribution,temperature rise and internal heat transfer coefficient inside composite air-cooled turbine blade.An numerical example is given.The effects of rotation on coolant flow and heat transfer are discussed.The method has been successfully applied to the design of air-cooled turbine blades for HPPSARP project,and it can also be used in model modification and calculation check.
NUMERICAL STUDY OF CONDUCTIVE TEMPERATURE DISTRIBUTION AND RECURRENCE RELATION IN TRANSIENT COMPUTATION
Zhu Gujun, Ji Shouli, Yang Fan
1989, 4(4): 347-350,391.
Abstract PDF(309KB) icon1338 icon13
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The numerical analysis of steady and unsteady conductive temperature distributions of turbine disk and blade is presented as part of HPPSARP project.It consists of two-dimensional (plane and axisymmetric) and three-dimensional problems.This paper gives a summary illustration of its feactures and emphasis on the stability of time recurrence relations in transient computation.A variable weighted coefficient to summarize the different time schemes is proposed in the developed computing program.It is favourable to improve the accuracy and restrain the oscillation in calculation.Some numerical examples are given to verify the characters of different schemes.
EXPERIMENTAL INVESTIGATION ON COPOSITE AIR-COOLED BLADES OF HIGHLY-LOADED TRANSONIC TURBINE
Liu Cunlu, Luo Mingjun, Deng Huayu, Liu Yufang, Wang Huage, Sheng Rongchang
1989, 4(4): 351-356,392.
Abstract PDF(1168KB) icon1621 icon17
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Four sets of composite air-cooled blades of highly-loaded and high temperature transonic turbine with various schemes of film-cooling holes have been designed and made,and have successfully passed 290 hour test run.Experimental results show that the cooling characteristics of the turbine vanes and blades are superior as their cooling mass flow ratios are 2-8% and 1-6% respectively.The wall temperature of turbine vanes and blades were corrected by measuring the heat losses of the test section.The results are converted to actual engine operation conditions,according to the similarity theory.It is obtained that the cooling effectiveness of the turbine vanes and blades is 335-420K respectively,and the wall temperature distributions are uniform.The effects of the parameters of the air-cooled turbine blades on the cooling effectiveness are also considered systematically,based on a large quantity of data obtained from the experiments of four kinds of composite air-cooled transonic turbine vanes and blades.
EXPERIMENTAL INVESTIGATION ON COMPOSITE COOLING OF A TURBINE BLADE
Zheng Jirui, Ji Honghu, Kong Zukai, Wang Baoguan
1989, 4(4): 357-362,392.
Abstract PDF(394KB) icon1795 icon12
Abstract:
In view of the whole process of composite cooling on a turbine blade,this paper is composed of three parts.First the impingement-convection cooling on th internal surface of a blade leading edge is investigated,in result an empirical expression of internal-local heat transfer coefficients at the stagnation point is formulated and regularity of their variatons along the leading edge region is found.Secondly,the variations of the film-cooling effectiveness along the external surface of the leading edge are investigated according to various blowing ratios when the film jets ejected from the film holes on the blade.Thirdly,as the film jets impose some disturbances on the main stream,heat transfer is enhanced,as a result a set of empirical expressions for the external-local heat transfer coefficients are formulated.
RESEARCH ON FILM-COOLING OF TURBINE BLADE
Ge Shaoyan, Liu Dengyun, Xu Jingzhong, Yao Yongqing, Deng Suqing, Jia Jianguo
1989, 4(4): 363-367,392.
Abstract PDF(1304KB) icon1622 icon26
Abstract:
The effects of discrete film-cooling holes,wall curvature and the pressure gradient on the turbine blade film cooling are systematically studied by means of heat-mass analogy,numerical methods and flow visualization techniques.The flow and heat transfer performances of the cooling film over the flat plate,curved wall and blade surfaces are investigated.It is proven that the characteristics of discrete-hole film cooling are quite different from those of continuous-slot film cooling in the vicinity of the injection holes because the coolant jets from holes tend to separate from the cooled surfaces. An optimal blowing ratio is expected for ideal distribution of film cooling effectiveness.Wall curvature exerted an additional force on the cooling film.When momentum blowing ratio I(=ρ2U2/ρ∞U2 ∞)>1,this force tends to seperate the film from the convex surface and to keep it close up to the concave surface,but when I<1,it acts contrary.For a given film cooling scheme,the effectiveness of turbine blade deponds on the blowing ratio,wall curvature and the pressure gradient.Some of the correlation formulas,exprimental data and computational results are presented to evaluate the film cooling effectivenesses and heat transfer coefficients at the leading edge,mid-chord and tailing edge of the turbine blade.
EXPERIMENTAL INVESTIGATION ON THERMAL-BARRIER PERFORMANCE OF CERAMIC COATINGS
Deng Huayu, Luo Mingjun, Liu Yufang, Liu Cunlu
1989, 4(4): 368-370,394.
Abstract PDF(241KB) icon1693 icon16
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Thermal-barrier performance of composite air-cooled vane with ceramic coatings in a highly-loaded transonic turbine has been experimentally investigated.Experimental results show that the average net thermal-barrier effectiveness on the middle section of the coated vane is 167K,but the net thermal-barrier effectivenesses at some points are measured as 145-206K.The effects of gas temperature on thermal-barrier performance of ceramic coatings are also investigated.It is found that the performance is even better at a higher temperature or higher heat flux condition.
EFFECTS OF FILM COOLING ON SURFACE PRESSURE DISTRIBUTION ON TURBINE BLADES
Luo Mingjun, Deng Huayu, Liu Bo
1989, 4(4): 371-372,393.
Abstract PDF(209KB) icon1345 icon15
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Effects of film cooling on the surface pressure distribution on composite air-cooled blades has been investigated experimentlly at medium temperatures and pressures.The effects of the ratio of cooling mass flow to gas mass flow KG,the ratio of gas temperature to coolant temperature K,and the location of film cooling holes on the surface pressure distribution are discussed.It is found from the experimental results that the KG and the location of the film cooling holes have considerable effects on the surface pressure distribution.
HEAT TRANSFER IN A WEDGED PIN-FINNED PASSAGE WITH PRESET PERFORATED PLATE
Gu Weizao, Zhang Yuming, Liu Changchun
1989, 4(4): 373-376,393.
Abstract PDF(316KB) icon1373 icon17
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An experimental research on enhanced cooling of gas turbine blades are presented,based on scaled up model heat transfer tests.The air from the radial inlet channel passes through the perforated plate and is forced into the pin-finned wedged passage with a dip angle of 3°52′ and initial width 18 6mm.At a distance 9mm from the plate is located a guide pillar row with the oval cross section and then 8 pin-fin rows.
FLOW PATTERNS AND HEAT TRANSFER OF IMPINGEMENT FLOW
Qiu Xuguang, Kang Ying
1989, 4(4): 377-380,393.
Abstract PDF(299KB) icon1728 icon14
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The various flow patterns and corresponding various characteristics of local heat transfer of impingement flow are considered for the short distance between nozzle exit and target surface.In the case of the two-dimensional slot jet,the flow patterns and characteristics of local heat transfer are remarkably influenced by these factors such as the degree of turbulance development in jet,the reaction of target surface on the jets and the compressibility of jet at high velocity.The types of multi-peak distribution of local heat transfer on the target surface depend on variation of these factors.For the row of circular jets,besides the above-mentioed factors,the effect of the three-dimensional flow of jets at short distance between nozle exits and target surface,leads to makes the local heat transfer from both sides of the stagnant line drop sharper than that of the two-dimensional slot jet.
EXPERIMENTAL SIMULATION OF IMPINGEMENT COOLING IN MIDCHORD REGION OF TURBINE BLADE
Li Liguo, Jiang Jun, Chang Haiping, Zhang Donglia
1989, 4(4): 381-386,394.
Abstract PDF(386KB) icon1667 icon18
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The simulating experiments have been completed to research the characteristics of impingement cooling in midchord region of turbine blade in the condiltons of given geometric para-meter of jet array,initial crossflow and pressure ratios of film cooling exhaust.Comparative experiments have been made between curvilinear and flate plate impinged surfaces,and impinged surfaces with and without chordwise fin.Moreover,the thermal patterns from liquid crystal indication show that the jet hole arrangement and distance between cooled surface and jet plate affect the heat transfer distributions for jet array impingement.Finally,the effects of the cooling air axially injected into guide tube on radial jets flow have also been determined by measurements.

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