永磁同步发电机超速实验设计

介绍超速实验中保持发电机机端电压恒定方法的原理,并推导了定子d轴电流的大小与超速倍数的表达式,从而能够方便地用于工程实践。

隐极式高速永磁电机转子强度分析

针对隐极式高速永磁电机转子强度问题,推导了高速实心磁钢转子强度设计的解析表达式,提出一套有效的高速转子过盈量的工程计算方法。以一台10 kW,额定转速60000 r/min的永磁同步电机为研究对象,对转子各部分的应力分别进行了解析和有限元分析计算,两种方法的计算偏差在2%以内。研究表明,对于实心磁钢转子,磁钢径向和切向应力大小均随磁钢半径的增加而增加,且在磁钢圆心处,径向和切向应力相等,随着半径的增大,切向应力开始小于径向应力。对样机进行了空载超速测试,结果表明,基于该方法设计的高速永磁转子强度满足设计要求。

Self-recovery ability of stress-damaged salt rock experiment

After being compressed to different plastic deformation stages, the salt rock samples with lateral stress damage of 0.2, 0.3, 0.4, and 0.5 were selected. Ultrasonic technology was used to monitor the wave velocity variation law of stress-dam-aged salt rock during the self-recovery experiment under different temperatures to analyze the influence of initial stress damage and temperature during the self-recovery of salt rock. The experiment shows that the change of salt rock axial wave velocity is smaller than that of lateral wave velocity. The sample ultrasonic velocity is positively correlated with the time of self-recovery, and the damage had been recovered to a certain extent. In the first 200 hours of self-recovery stage, the salt rock lateral damage recovers fast, and then the damage remains almost unchanged. The value of lateral stable damage is positively correlated with the value of lateral initial stress damage. With the increase of temperature, the recovery of lateral damage speeds up and the value of stable damage decreases; the axial damage of salt rock almost remains unchanged during the self-recovery experiment.

Application of Rainbow Schlieren Deflectometry for Axisymmetric Supersonic Jets (Comparison of Experiments with Numerical Analysis)

This paper analyzes the correctly-expanded supersonic jet from a convergent-divergent axisymmetric nozzle by using numerical simulation of turbulent flow.And the calculated density distributions in this flow are compared with the present experimental data using rainbow schlieren deflectometry.The value of the density from the experimental data agrees well with the results calculated by this simulation.Therefore,the present method of the measurement using rainbow schlieren deflectometry is useful for the measurement of the density of the correctly-expanded supersonic jet.

Experimental and Numerical Evaluation of the Performance of Supersonic Two-Stage High-Velocity Oxy-Fuel Thermal Spray (Warm Spray) Gun

The water-cooled supersonic two-stage high-velocity oxy-fuel (HVOF) thermal spray gun was developed to make a coating of temperature-sensitive material,such as titanium,on a substrate.The gun has a combustion chamber (CC) followed by a mixing chamber (MC),in which the combustion gas is mixed with the nitrogen gas at room temperature.The mixed gas is accelerated to supersonic speed through a converging-diverging (C-D) nozzle followed by a straight passage called the barrel.This paper proposes an experimental procedure to estimate the cooling rate of CC,MC and barrel separately.Then,the mathematical model is presented to predict the pressure and temperature in the MC for the specific mass flow rates of fuel,oxygen and nitrogen by assuming chemical equilibrium with water-cooling in the CC and MC,and frozen flow with constant specific heat from stagnant condition to the throat in the CC and MC.Finally,the present mathematical model was validated by comparing the calculated and measured stagnant pressures of the CC of the two-stage HVOF gun.

Experimental study of a supersonic turbulent boundary layer using PIV

Particle image velocimetry was applied to the study of the statistical properties and the coherent structures of a fiat plate turbulent boundary layer at Mach 3. The nanoparticles with a good flow-following capability in supersonic flows were adopted as the tracer particles in the present experiments. The results show that the Van Driest transformed mean velocity profile satisfies the incompressible scalings and reveals a log-law region that extends to yld=0.4, which is further away from the wall than that in incompressible boundary layers. The Reynolds stress profiles exhibit a plateau-like region in the log-law region. The hairpin vortices in the streamwise-wall-normal plane are identified using different velocity decompositions, which are similar to the results of the flow visualization via NPLS technique. And multiple hairpin vortices are found moving at nearly the same velocity in different regions of the boundary layer. In the streamwise-spanwise plane, elongated streaky structures are observed in the log-law region, and disappear in the outer region of the boundary layer, which is contrary to the flow visualization results.

Numerical investigation on the thermal protection mechanism for blunt body with forward-facing cavity

Numerical experiments are carried out using the standard hypersonic ballistic-type model(HB-2) to investigate the effect of forward-facing cavity on the aerodynamic heating. A general concept is proposed which utilizes the flow disturbances generated passively in the nosed subsonic region to weaken the detached shock wave. Several aspects are mainly studied, including shock shape and standoff distance, surface heat flux and pressure, flowfield feature and cooling mechanism. The numerical results indicate that shock strength and standoff distance increase with an increase in the L/D ratio of the cavity. Interestingly, a bulge structure of the detached shock associated with a deep cavity is observed for the first time. Local surface heat flux and pressure around the concave nose are much lower respectively than those at the stagnation point of the baseline model. In addition, both surface heat and pressure reductions are proportional to the L/D ratio. A negative heating phenomenon may occur in the vicinity of a sharp lip or on the base wall of a deep cavity. If the L/D ratio exceeds 0.7, the detached shock appears as a self-sustained oscillation which can be referred to as the cooling mechanism.

Experimental study on the performance of a rotational supercavitating evaporator for desalination

Efforts have been made on experimental research of a supercavitation device for desalination, which is named rotational supercavitating evaporator （RSCE）. The RSCE is characterised by the simple construction and responsive capacity control, and only requires rough filtration of the source water for scaling- and fouling-free operations. The present study has been conducted for the water flow at temperature of around 22-30℃ and atmospheric pressure as the first step for investigation of the performance characteristics of RSCE. The multiply factor extremal experiments conducted with the Box-Wilson＇s method have revealed the salinity of the condensate, the temperature of steam inside the supercavity, and dependence of the shape of supercavity on the rate of steam extraction and rotation speed of impeller. The shape of impeller and the expected supercavitating effects it generates have been confirmed by experimental results at the rotation speed of 5430 rpm （round per minute）. The design of the steam extraction openings has been approved during the evacuation of steam. The experimental dependencies have been obtained in form of statistically valid regression equations, which can be used for engineering design of RSCE.

Comparative study on aerodynamic heating under perfect and nonequilibrium hypersonic flows

In this study, comparative heat flux measurements for a sharp cone model were conducted by utilizing a high enthalpy shock tunnel JF-10 and a large-scale shock tunnel JF-12, responsible for providing nonequilibrium and perfect gas flows, respectively. Experiments were performed at the Key Laboratory of High Temperature Gas Dynamics(LHD), Institute of Mechanics, Chinese Academy of Sciences. Corresponding numerical simulations were also conducted in effort to better understand the phenomena accompanying in these experiments. By assessing the consistency and accuracy of all the data gathered during this study, a detailed comparison of sharp cone heat transfer under a totally different kind of freestream conditions was build and analyzed. One specific parameter, defined as the product of the Stanton number and the square root of the Reynold number, was found to be more characteristic for the aerodynamic heating phenomena encountered in hypersonic flight. Adequate use of said parameter practically eliminates the variability caused by the deferent flow conditions, regardless of whether the flow is in dissociation or the boundary condition is catalytic. Essentially, the parameter identified in this study reduces the amount of ground experimental data necessary and eases data extrapolation to flight.

Experimental Study on Measurement and Calculation of Heat Flux in Supersonic Combustor of Scramjet

An experimental measurement and calculation method which consist of thermal response model, convergence criteria and control algorithms, is proposed in this paper for the determination of heat flux in a scramjet combus- tot. Numerical simulations are done to evaluate the effectiveness of the proposed method, and experiments are made in the direct-connect hydrocarbon fueled scramjet combustor of Mach-6 flight for different equivalence ra- tios. The distribution of heat flux along the axial and circumferential directions can be obtained using the pro- posed method. The distribution of heat flux is uneven which is caused by the aerodynamic heating, combustion heat release and changes of section area, and the peak heat flux can be more than 2MW/m^2 during the experi- ments. Heat flux increases with the increase in equivalence ratio for the same Mach number. And axial distribu- tion of heat flux is uniform for different equivalence ratios. In addition, the combustion heat release area of the combustion chamber can therefore be concluded which is useful for guiding the structural design of the thermal protection system.

Experimental Investigation on Starting Process of Supersonic Single-stage Air Ejector

This paper deals with experimental study of flow field of starting process in two-dimensional, single-stage supersonic ejector on different air total pressure. Schlieren pictures of flow field were taken, static pressure distribu-tions on side wall were measured. The obtained results show that, on critical pressure, the starting main shock waves in ejector oscillated back and forth between the second throat and the middle section of the mixing chamber, it causes the pressure in the second half of the mixing chamber acutely fluctuated .When the working pressure of the active flow is higher than the critical starting pressure, ejector starts normally and the inner flow-field of the mixing chamber keeps stable and the shock waves in the second throat have a certain degree of oscillation . After ejector starts, the operating pressure of the active flow may be lower than the starting pressure .

Development of the detonation-driven expansion tube for orbital speed experiments

The hypersonic flow at orbital speeds is a fundamental issue for the ground tests of aerospace crafts.The detonation-driven high-enthalpy expansion tube(JF16 expansion tube)was developed to investigate re-entry physics.A forward detonation cavity(FDC)driver was applied in the JF16 expansion tube to create stable driving flows.The sound speed ratio of the detonated to test gas was examined to minimize the magnitude of test flow perturbations.The acceleration section length,incident shock decay and diaphragms thickness were investigated in detail to obtain optimal operation parameters.Flow visualization was also carried out with schlieren system to demonstrate the test flow stability and the effective test duration.Experimental data showed that the test flow with a velocity of 8.3 km/s and a total enthalpy up to 40 MJ/kg can be generated successfully and the test duration lasts for more than 50μs.

Experimental study of unsteady supersonic flow interacting with porous cavity and jets or rods

In this study,an experiment was performed to clarify the flow field,in which the jets were normally injected into a main supersonic flow surrounded by a porous cavity,and this report figures out interaction between starting shock wave and porous cavity.In the experiment,a porous cavity is attached to a main duct and jets and rods are inserted to the main duct on the porous cavity.To reveal this flow field,the thermal tuft probe was adopted to experimentally investigate the flow in the cavity.In the experiments,the effect of the porous cavity with jets or rods on the flow field is studied by means of visualization of schlieren method with a high speed camera and measurement of cavity flow with thermal tuft probe.As a results,frequency analysis of output of the thermal tuft probe revealed that some clear dominant frequencies were confirmed when the starting shock wave existed around the porous cavity in all cases of jets and rods arrangements.Moreover,visualization of schlieren method with a high speed camera clarified that a starting shock wave had the same dominant frequencies as that of the flow fluctuation in the cavity only around the cavity.

Investigation of the Compressible Flow through the Tip-Section Turbine Blade Cascade with Supersonic Inlet

The contribution deals with the experimental and numerical investigation of compressible flow through the tip-section turbine blade cascade with the blade 54″ long. Experimental investigations by means of optical(interferometry and schlieren method) and pneumatic measurements provide more information about the behaviour and nature of basic phenomena occurring in the profile cascade flow field. The numerical simulation was carried out by means of the EARSM turbulence model according to Hellsten [5] completed by the bypass transition model with the algebraic equation for the intermittency coefficient proposed by Straka and P?íhoda [6] and implemented into the in-house numerical code. The investigation was focused particularly on the effect of shock waves on the shear layer development including the laminar/turbulent transition. Interactions of shock waves with shear layers on both sides of the blade result usually in the transition in attached and/ or separated flow and so to the considerable impact to the flow structure and energy losses in the blade cascade.

Experimental investigation of hypersonic flow induced separation over double wedges

Flow separation occurs over the compression comers generated by deflected control surfaces on hypersonic re-entry vehicles and in the inlet of scram jet engines. Configurations like a double wedge and double cone model are useful for studying the separated flow features. Flow fields around concave comers are relatively complicated and produce several classical viscous flow features depending on the combination of the first and second wedge or cone half apex angles. Particularly characteristic phenomena are mainly shock/boundary layer, shock/shock interaction, unsteady shear layers and non-linear shock oscillations. Although most of these basic gas dynamics characteristics are well known, it is not clear what happens at high enthalpy conditions. This paper reports a result of flow fields over a double wedge at a stagnation enthalpy of 4.8 MJ/kg. The experiment was carried out in a free piston shock tunnel at a nominal Mach number of 6.99. Schlieren and double exposure holographic interferometry were applied to visualize the flow field over the double wedge.