1987 Vol. 2, No. 1

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DEVELOPMENT OF AIRCRAFT ENGINE COMBUSTORS
Zhou Xiaoqing
1987, 2(1): 1-7,87.
Abstract PDF(967KB) icon1706 icon555
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The principal requirements of combustors for advanced, high-tech aircraft engines are reviewed in this paper. Key technologies are briefly described and new development approaches are introduced.Combustors will be developed continuously in the direction of improving performances, shortening length, increasing reliability and durability, reducing exhausting gas pollution, and utilizing alternating fuels, the last one is the key to the success of wide application of industrial gas turbines.Various new combustor concepts are in study both abroad and home, includingdouble-annular combustors, vorbix, axially or radially stagged combustors, varying geometry combustors, twin recirculation combustors, horse shoe combustordome, vortex buring afterburner, barchane dune flame-holder and so on. Among the key technologies for advanced combustors, improving spray quality and combustor-dome fuel-air mixture preparation, and developing advanced coolingschemes and heat-resisting materials are in the first priorities.Rapid progress of computer and laser technologies strongly influences a varietyof science and technology fields, combustor is no exception. The main aspects of computer modeling and laser diagnostics of combustion processes are briefed, and a new empirical/multidimensional analytical design methodology is surveyed, which is believed to become a new approach to studying, designing and developing advanced combustors.
FUNCTIONS OF TURBINE BLADE COOLING EFFECTIVENESS EXPERIMENTATION
Qiu Xuguang
1987, 2(1): 8-12,87-88.
Abstract PDF(583KB) icon1330 icon466
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The functions of available equipment for cooling effectiveness of turbine blade are surveyed, and the corresponding experimental methodology is considered, which is divided into two classes, by comparison and by simulation. Stress is put on the simulating experimentation.
EXPERIMENTAL INVESTIGATION OF LOCAL HEAT TRANSFER CHARACTERISTICS BY IMPINGEMENT COOLING
Ren Linghua, Zheng Jirui
1987, 2(1): 13-17,88-89.
Abstract PDF(536KB) icon1734 icon429
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The simulation of the local heat transfer characteristics of the leading edge region of a turbine blade by impingement cooling of concave surface with a row of circular air jets has been investigated experimentally.The effects of Reynolds number, the dismensionless distance between the jet and the target, the compound geometrical configurations both of the impingement tube and target on the local heat transfer coefficients are discussed. The special distributions of local heat transfer coefficients are also analyzed.It is found from a large amount of experiments that the heat transfer coefficientdistribution is not a simple Guass function. The Guass function fits the experimental data satisfactorily only at the stagnation region,while the local heat transfer coefficients in t.he jet region at the wall can be expressed in exponent functions. A set of empirical formulas which describe the distribution of the local heat transfer coefficients by impingement cooling have been obtained from the experimentaldata by means of least square approach.
APPLICATION OF SHOCK TUNNEL TO HEAT TRANSFER EXPERIMENTS ON TURBINE BLADE
Li Jingmei, Zhao Runming, Wang Jinan
1987, 2(1): 18-22,89.
Abstract PDF(521KB) icon1815 icon483
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Transient wind tunnels and their supplementary instrumentation have been used to examine the flow and heat transfer around hypersonic vehicles. The purpose of this paper is to describe their application to turbine cascade testing and refinements of its heat transfer measurement techniques.
APPLICATION OF LOW BTU GASES TO AEROENGINE MODIFIED FOR INDUSTRIAL USE
Jiao Shujian
1987, 2(1): 23-28,89.
Abstract PDF(655KB) icon1280 icon402
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At present, we are confronted with the realistic conditions to modify aeroengines for burning low BTU gases.The combustion characteristics of these gases are presented in this paper. Some difficulties encountered in burning of these gases in gas turbine combustors are considered and their solutions are proposed through several typical examples of combustor configuration.
PREDICTION OF STALL FLUTTER OF STEAM TURBINE BLADES
Yang Xiaodong, Tang Zhiming, Zhou Sheng
1987, 2(1): 29-32,90.
Abstract PDF(429KB) icon1768 icon428
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The failure of the steam turbine last stage blades by abrupt increment of dynamic stress in some conditions has been analyzed, and it is confirmed that the main cause of this breakdown is self-excited vibration (flutter) of blades. The actuator disk method has been transplated from aeronautics to predict the steam turbine blade flutter. The actuator disk method only deals with the unsteadyflow field in front of and behind the cascade.The governing equation of the flow field is the linearized Ruler's Equation. To determine the constants of the equation solution, the cascade loss is correlated with the experimental data.The comparisons between the predictions and experimental results are acceptable.It is demonstrated that the abrupt increment of the blade dynamic stress at small volumetric flow is caused by large negative incidence stall flutter. It is discovered from the computations that at design conditions there may be a kind of high loaded unstall flutter which still hasn't been defined.
DESIGN AND EXPERIMENTAL INVESTIGATION OF SECONDARY BURNER IN AN INERT GAS EXTINGUISHING SET MODEL DQ-1000
Shi Xueqing
1987, 2(1): 33-38,90-91.
Abstract PDF(636KB) icon1866 icon391
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A new secondary burner has been successfully designed and tested for DQ1000inert gas extinguishing set which is to put out an underground fire, especiallya fire in coal mine tunnel. Its main design features are low gas flow velocity, two-zoned fuel supply, bypass torch ignition, multirow flame holders, injection of water into combustion zone, and water cooling system in both outer casing and exhaust section etc. The test results demonstrate Its good performances,such as flexible ignition capacity, stable combustion, high combustion efficiency up to 99%, safety and reliability. Particularly, oxygen content in volume is lower than 3 % in exhaust gases. The set was certificated in 1985 and passed the test in underground fire zone. The practice has proved that extinguishinga big fire by means of inert gas is an effective method at present.
A METHOD FOR COMPRESSOR AXISYMMETRIC CHARACTERISTICS BY THE AID OF ROTATING STALL PARAMETERS
Tang Guocai, Zhang Huimin
1987, 2(1): 39-44,91.
Abstract PDF(577KB) icon1487 icon408
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A new method is provided to obtain the axisymmetric characteristics of an axial compressor by the aid of its rotating stall parameters. The pricipal differenceof the method from the method of Koff and Greitzer (K.G) is its utilization of non-uniformity along circumferential direction induced by rotating stall with the inertial correction incorporated to gain the axisymmetric characteristics in the flow range,in which the rotating stall usually occurs. The charateristics of a typicalexample follow quite reasonable trend. This method exhibits a wider applicablerange because it is not restrained by the requirements for the three time scales as specified in K. G's method and also due to the fact that the stall parameters are easier to be acquired than the surge parameters.
ENGINEERING EVALUATION OF STEADY OPERATING MARGIN OF A DUAL-ROTOR COMPRESSOR
Jiang Hefu
1987, 2(1): 45-48,91.
Abstract PDF(435KB) icon1865 icon400
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The characteristics of the steady(surge-free)operating margin of a dual-rotor compressor in a turbojet engine are discussed, and the practical technical requirements are given for engineering evaluation of the steady operating margin. For the dual-rotor compressor, there is a great influence of the ratio n2/n1(H. P. rotor speed to L.P.rotor speed)on the steady operating margin. In order to guarantee the steady operation of the compressor in the whole flight envelope, it is necessary to check the surge-free operating margin not only in a specified range of the L.P. rotor speed but also in a specified range of n2/n1. A numerical example of a turbojet engine is given to illustrate the above requirements.
MEASUREMENTS AND ANALYSIS OF SURGE,STALL AND SURGE POINT PARAMETERS IN TURBOJET COMPRESSORS
Li Wenlan, Chen Fuqun, Wang Zongyuan, Cong Mengzi
1987, 2(1): 49-54,92.
Abstract PDF(555KB) icon2077 icon461
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This paper presents the testing and measuring method of three kinds of flow instability test(rotating stall and surge),the test results and analysis of them for a turbojet engine with "clean" or distorted inlet flow. The tests have been carriedout as follows. ( 1 )testing of the"upper"and"lower"stall-point on the stall line of the engine with and without distorted inlet flow; ( 2 )tests of surge in the engine with "clean" or distorted inlet flow at n= 1.0, 0.933, 0.9, 0.87 and with rated Anb (area of first-stage turbine nozzle) or Anb decreased by 12%; (3)tests of static-forced surge(decreasing the area of engine nozzle very very slowly)and dynamic-forced surge for an engine with"clean" or distorted inlet flow. The results from the tests comprise.
A PREDICTION METHOD FOR OPTIMUM VELOCITY RATIO OF AIR JET VORTEX GENERATOR
Zhang Fengbei, Sheng Chunhua
1987, 2(1): 55-60,92-93.
Abstract PDF(761KB) icon1677 icon423
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The selection of the velocity ratios for air jet vortex generator is considered. The concept of the optimum velocity ratio with the maximum vortex induced velocity at the wall,Ropt·v ,and the concept of the optimum velocity ratio maximizingshape factor of boundary layer ,Ropt·H, are introduced.Based on the empiricalrelationships summarized from the measured data of the side-inclined jet/ crossflow interaction flow field, the formulas forRopt·v may be deduced as follows Ropt·v = 3.22(X/D + 0.3019)-0.219 R≤1.8Ropt·v = 2.75(X/D + 0.3019)-0.111 R≥1.8The predicted values of Ropt.Hare determined according to the experimental results of the reverse flow region behind the jet hole in association with the Ropt·v empirical formulas. They agree with the measurements of the boundary layer at the flate plate.
SPACE-FLIGHT TESTS OF THE APOGEE BOOST MOTOR FOR STW EXPERIMENTAL COMMUNICATIONS SATELLITE
Gu Xuelin
1987, 2(1): 61-64,93.
Abstract PDF(436KB) icon1383 icon390
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The apogee boost motor (ABM) for the STW experimental communications satellite is a solid composite propellant rocket motor. Its case is a glass/epoxy filament wound composite structure. Its pro pliant is of HTPB type. The materialof its nozzle throat is carbon/carbon composite, and the exit cone is wound by a highly silica-filled phenolic/glass tape. Its ignition system consists of ignitor and safe and arm device. A series of motor tests have proved that the ABM has high reliability and performance accuracy. The reliability of the ABM is 0.9551 with the confidence being 0.8.This work predicts and analyses the potential accidents of the ABM during space-flight, and proposes various approaches to eliminate them. The calculationsbased on the parameters gathered before and after space flight show that the total impulse is 1258.3 kN·s and the specific impulse is 2834.1 N·s/kg. Both of them are in good agreement with those from ground-firing tests with a variation of the total impulse within the range of ±1%. It is shown that the ground-firing tests are accurate and efficient.
FLIGHT TEST RESEARCH ON STATIC CHARACTERISTICS OF AFTERBURNER FUEL CONTROL SYSTEM FOR TURBOJET ENGINES
Tao Liyi, Chen Zhanping, Fan Siqi, Wu Chihua
1987, 2(1): 65-68,93-94.
Abstract PDF(507KB) icon1938 icon477
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The operation and function of the pressure ratio controller and the air pressure controller in the afterburner fuel control unit (AFCU) are investigated on the basis of the data obtained from flight tests and laboratory experiments of the four AFCU models for turbojet engines. Our preliminary opinions are presented in respect to this moot subject which has caused controversy for years.It is found that there are three operational conditions of the afterburner fuel control system. The fuel flow rate fed to the afterburner is controled independently by the pressure ratio controller only during ground operations or in low altitude flight at high velocities, while the air pressure controller does not function. Within the major range of the flight envelope, both the pressure ratio and the air pressure controllers function and control the fuel flow rate jointly. The air pressure controller may function and control the fuel flow rate alone merely in high altitude (near ceiling) flight at low velocities.
EXPERIMENTAL INVESTIGATION ON COOLING EFFECTIVENESS OF HIGH-TEMPERATURE AIR-COOLED TURBINE BLADE
Wu Shishen, Zhang Zhi jun
1987, 2(1): 69-71,94.
Abstract PDF(391KB) icon1709 icon480
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This paper deals with the application of the static test results at low temperatureand low pressure to the actual engine operating conditions. In order to extrapolate the test-bed results of the cooling effectiveness of a turbine blade to actual conditions of engine operation, two methods are presented. ( 1 ) to derive the experimental relationship of cooling effectiveness by heat transfer analysis, ( 2 )to substitute the cooling effectiveness tests at low temperature and low pressurefor those at high temperature and high pressure according to the similarityprinciple. Five kinds of blades in the same shape but with different internalstructure have been tested for cooling effectiveness at various conditions. The results show that the two methods are satisfactory.
EXPERIMENTAL INVESTIGATION ON COOLING EFFECTIVENESS OF A TURBINE BLADE UNDER VARIOUS DESIGN CONDITIONS
Xuan Zhijiang, Zhang Sanduo
1987, 2(1): 72-74,95.
Abstract PDF(339KB) icon1347 icon408
Abstract:
Blade cooling characteristics under various design conditions were obtained from the turbine blade cooling effectiveness experiments of a given aeroengine and its two modifications. The results obtained provide useful data for air-cooled blade design. Structure and testing approach of the blades are also described.Experimental results are also analysed. The analysis demonstrates that the testing technique is aviable, the experimental data are amenable and satisfactory for air-cooled blade design.
COOLING FLOW TECHNIQUE FOR DESIGNING AFTERBURNER HEAT SHIELDS
Lian Kangbao
1987, 2(1): 75-77,95.
Abstract PDF(444KB) icon1731 icon407
Abstract:
By employing the air flow cooling technique, an afterburner heat shield was designed semi-empirically. It exhibited the desired advantages of long service life, low wall temperature and good performance. The cooling air flow accounted for 10% of the total airflow. The distribution of the cooling air was designed in referenceto successful designs. The practice of the aeroengine with the designed afterburner verifies that the presented technique is reliable.
HEAT TRANSFER THROUGH MULTI-TURNS OF COOLING PASSAGES IN A TURBINE BLADE
Zhang Yumin, Gu Weizao, Shen Jiarui
1987, 2(1): 78-80,95-96.
Abstract PDF(391KB) icon1262 icon396
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With the aim of cooling high-temperature turbine blade, the heat transfer characteristics and flow drag of turbulent air flow through 180 degree multiturnsof rectangular passages were investigated experimentally. The experimental data of the heat transfer coefficients both through a 180 degree turn and in a straight passage were correlated. The test results may be useful to cooling designinggas turbine blades.
A CALCULATION METHOD OF DETERMINING BLADE SECTION DRAG IN A ROTOR BASED ON WAKE MEASUREMENTS
Wang Yuzong
1987, 2(1): 81-83,96.
Abstract PDF(323KB) icon1488 icon412
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A calculation method is proposed to determine the blade section drag instead of sophisticated experimental measurement of it. Since it is possible to obtain the detailed distributions of velocity and pressure on the boundary line from the wake measurements behind a compressor rotor. By using these distributions, the blade section drag is determined by means of the momentum equation of integralrepresentation and appropriate mathematical manipulation.
1987, 2(1): 84-84.
Abstract PDF(126KB) icon1471 icon381
Abstract:
THERMAL EFFECTS ON TRANSIENT PROCESS OF A TURBOJET ENGINE
Chen Daguang, Pan Yongquan
1987, 2(1): 85-86,96.
Abstract PDF(339KB) icon1804 icon413
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The mathematical models commonly used for dynamic process of turbojet or turbofan engine consider only the rotor inertia and gas path volume dynamics. Usually the gas path volume of an aeroengine is not great, so the effect of volumedynamics on normal transient process is negligibly small. The thermal effectsinclude transient heat transfer and the component efficiency changes caused by clearance variation with temperature in the components of compressor and turbine.A single-shaft turbojet engine has been tested for cold and hot transient processes. The comparison between the cold and hot transient test results shows that the thermal effects have significant influence on the acceleration time, the changing rate of aerothermodynamic parameters with engine speed, etc. The experimentalinvestigation demonstrates that the thermal effects must be taken into account so as to establish an accurate mathematical model for the transient process of turbine engines.

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