Showing papers in "Journal of Propulsion Technology in 2014"

Loss Sources Analysis of Shroud Leakage Flow in Highly-Loaded Turbine

Abstract: The pressing demand for future advanced gas turbine requires to identify the losses in a turbine and to understand the physical mechanisms producing them In gas turbines with shrouded blades,a large portion of losses is generated by tip shroud leakage flow In order to analyze the loss sources and physical mechanisms induced by shroud leakage flow,a computational study for a highly-loaded turbine was presented and discussed Three dimensional steady multistage calculations using mixing plane approach and entropy analysis method were performed Detailed tip shroud geometry is taken into consideration which overcomes most of the limitations of simple shroud empirical correlation in modeling leakage flow over tip shroud Results show that turbine with shrouded blades show 0 9% higher in efficiency than the unshrouded one Furthermore,cavity loss,leakage loss,mixing loss and incidence loss are the four major components of shroud related losses Loss mechanisms breakdown analysis demonstrates that various loss sources account for different portions according to the tip clearance levels Therefore,full calculation of shroud leakage flow is necessary in turbine performance prediction and the shroud geometric features need to be considered seriously during turbine design process

High Strain-Rate Viscoelastic Constitutive Model for HTPB Propellant

Abstract: In order to research mechanical behaviors of HTPB propellant at high strain-rate,split Hopkinson pressure bar( SHPB) tests at different temperature(- 40 ~ 25℃) and strain rate( 700 ~ 2050s- 1) were conducted and stress-strain curves were obtained. The results show that the mechanical properties of HTPB propellant are sensitive to temperature and strain rate,and stress increases gradually with decreasing temperature and increasing strain rate. Based on Burke model,a constitutive model that incorporates nonlinear hyperelastic and viscoelastic material response for HTPB propellant was developed for high strain-rate impact loading conditions. Predictions of the stress-strain response were made using the constitutive model. The good agreement between the predicted and the experimental results indicates the validity of constitutive model for different temperature and high strain-rate. The constitutive model can provide the theoretical basis for structural integrity of solid propellant grain under ignition transient pressure loading.

Investigation on Relation Between Secondary Flow and Loss on a High Loaded Axial-Flow Compressor Cascade

Abstract: Numerical investigation was carried out to study the relation between secondary flow and loss generation in a high-loaded compressor cascade. Based on a series of qualitative analysis,a new quantitative analysis model has been derived further to attain the loss contribution of various sources,and the mechanism and effects of secondary flow on loss generation were finally conformed. The results show that loss due to secondary flow does not exceed 50% under a wide range of inlet attack angles. The indirect effect of secondary flow on low speed region is identified to dominate its contribution to loss with its direct dissipating remaining small magnitude order. A detailed research about the mechanism of the steep loss increase near corner stall also indicates that change in direction of cross flow,which is the change of indirect effect of secondary flow on low speed region,is mainly responsible as the incidence increases.

Research on Continuous Rotating Detonation Wave Propagation Process( II) : Two-Wave Collision Propagation Mode

Abstract: Experimental research on the H 2 / air continuous rotating detonation wave has been carried out on two model engines,which use the injector-injector and slot-injector impinging injection schemes respectively. Tangentially injected H 2 / O 2 hotshot jet has been used to ignite the engines successfully. Based on the PCB measurements and optical observations,the detonation wave propagation process has been analyzed. Both transient and long-time lasting two-wave collision propagation phenomena have been observed. Under the two-wave collision mode,there are two detonation waves propagating along the opposite direction in the combustor,colliding with each other periodically. Its high frequency pressure distribution presents different oscillation characteristics,which is mainly influenced by the angle between the pressure sensor location and the collision point. It is believed that to realize this rotating mode there should be an annular zone of fresh H 2 / air mixtures located at the top of the detonation wave.

Simulation and Calculation of Fuel Temperature in Aircraft Fuel Tank Based on MATLAB/Simulink

Abstract: The fuel temperature is one of the key parameters of the aircraft fuel tank flammability assessment. In order to study the temperature distribution of the aircraft fuel tank and the correctness of the heat transfer model,the MATLAB / Simulink software platform was used to build an aircraft fuel tank thermal simulation model. Entering boundary conditions corresponding to flight experiments,the fuel temperature inside the fuel tank at the compute nodes can be obtained through numerical simulation. The results show that under three different sailing conditions,the simulation results of fuel tank thermal model could preferably coincide with the flight test results and calculation error can be controlled within a certain range. Using this fuel tank thermal model in thermal numerical simulation,aircraft fuel tank thermal characteristics can be obtained accurately,and it can be used for optimization of the fuel tank structure and internal heat source component arrangement in the stage of aircraft design.

Numerical Simulation of Detonation Initiation in Quiescent Combustible Mixtures with Block-Structured Adaptive Mesh Refinement Method

Abstract: In order to obtain a profound understanding of the mechanism of detonation initiation in quiescent combustible mixtures using hot jet,further development for the open-code program AMROC of blockstructured adaptive mesh refinement method was adopted for fine structure numerical simulation of detonation initiation in quiescent combustible mixtures with hot jet. Simulations were conducted on a small-scale nested parallel computation system on LINUX platform. The initiation process was specified as three stages and their flow field characteristics were analyzed. Results indicate that for hot jets in certain conditions the initiation process is strongly dependent on the walls of the tube in which reflecting effects induce local combustion and detonation subsequently. With two-time reflection on the head face and the upper wall of the tube,detonation near the upper wall is amplified and spread into the whole channel. When local detonation is formed at both the upper and lower walls,two triple-wave points are generated,and collide between the upper and lower walls inside the tube. Thus a stable periodic detonation wave is formed.

Numerical Simulation of Inlet Distortion with Interceptor with Different Turbulence Models

Abstract: In order to get a more accurate turbulence model,three turbulence models have been adopted to simulate inlet distortion with interceptor. Compared with experimental data,the precision of each model to simulate inlet distortion with interceptor was investigated. The flow field distortion computed by each model is different.The area of high-pressure is large by using renormalization group k-e model and the area of transition section is large by using Spalart-Allmaras model. It is shown that the angle range of low-pressure has better agreement with experimental data when the interceptor height is 48% of the inner diameter of the distortion generator.However,when the interceptor height is 20% of the inner diameter,the angle range of low-pressure is smaller than experimental data. The value of distortion index is more accurate by using realizable k-e model.

Experimental Investigation of Dynamic Characteristics of Oblique Shock Train in Mach 5 Flow

Abstract: The structure and dynamics characteristics of oblique shock train in a simple duct model are investigated experimentally in a Mach 5 hypersonic wind tunnel. Back pressure of model is produced and varied by the close of two ramps at the end of model. Measurements made include high-speed schlieren imaging and simultaneous fast-response wall pressure along the length of model. Results show that the leading edge of shock train propagates upstream,and translates to be asymmetry with the increase of back pressure. The asymmetry state maintained throughout the last whole propagation process. Varying the rise speed of back pressure did not show any discernible effects with regard to propagated speed of the leading edge of oblique shock train. There are two kinds of motion in the shock train propagation process: steadily forward and rapidly forward. The characters of shock train propagation are decided by the fluid structure,the pressure distribution of wall,the magnitude of back pressure, and the distance to the isolator exit, etc. Different wall pressure spectral characteristics of oblique shock train in duct are explored with different back pressure. But in the same case,all the pressure spectral characteristics of transducers in the shock train region are similar. In Case 1,the dominant frequency is f1=512Hz;in Case 2,the dominant frequency is f1=578Hz,the secondary frequency is f2=260Hz that is close to the acoustic resonance frequency. Both the dominant frequencies obtained in the experiments are higher than those of the theoretical prediction of Piponniau model.

Research on Ramjet-Mode Combustion Organization by Using Vaporized Kerosene Injection in RBCC Combustor

Abstract: Considering the active cooling requirement of RBCC engine under ramjet mode,researches on combustion organization based on vaporized kerosene injection in RBCC combustor have been performed.Under the low total temperature of incoming airflow condition at ramjet mode,the reliable ignition and stable combustion of secondary kerosene fuel by using low mass flowrate fuel-rich primary rocket plume as piloting flame has been accomplished.Moreover,the peak pressure in the combustor by injecting supercritical kerosene increased about 10%in comparison of room temperature kerosene.Meanwhile,when primary rocket was closed,flameholding was only achieved by using supercritical kerosene injection.The CFD simulation has been employed to obtain the detailed information of flowfield characteristics and combustion organization in RBCC combustor under different fuel injection strategies,which provides fundamental basis for further optimizing engine performance.The results indicate that the thermal coking in the expansion flowpath has been formed due to the concentrated heat release behind the strut injectors.Reasonable heat release distribution along the combustor by changing fuel split is important to achieve high combustion efficiency.

High-Density Liquid Hydrocarbon Fuels for Aerospace Propulsion:Synthesis and Application

Abstract: High-density liquid hydrocarbon fuels can extend the flight range and increase the payload of vehicles and have been widely utilized in aerospace propulsion. In this paper,some typical fuels,such as JP-10,T-10,RJ-7,were briefly introduced. In particular,high-density fuels developed in Tianjin University,including HD-01,fuels with density 1g/ml,alkyl-diamondoids,quadricyclane and high-density biofuels,were presented. In addition,the hypergolic ignition of quadricyclane was reported for the first time. To overcome the ignition and cooling problems under hypersonic cruise,monodispersed Pd and Pt nanoparticles that can be dispersed in fuels stably were synthesized as catalysts. The ignition temperature of JP-10 was decreased by Pd nanoparticles and the heat sink via cracking was increased by Pt nanoparticles.

Numerical Simulation of Flow Field in Nozzle of Hot Water Rocket Motor

Abstract: In order to deeply understand the working performance in the hot water rocket motor,a numerical model of the flow field in the nozzle of hot water rocket motor is established. And the numerical model is validated via a sample case. According to the study of the flow field in the nozzle of the motor,it is found out that the pressure firstly drops to the saturated pressure corresponding to the initial temperature in the convergent section and then continuously decreases along the axial direction. Besides, the phase change occurs at the nozzle throat,where the main flow would change into a two-phase flow,and vapor volume fraction at the exit is over99%. Then,the two-phase flow is transferred from a subsonic condition into supersonic flow after the throat because of the changing of sound speed. The flow process in the nozzle can be divided into three steps,which are single-phase flow step,flash evaporation step due to pressure derease and expand-accelerating step,respectively. Though it was similar to the conventional chemical rocket motor,it had a certain complexity because of the phase change.

Design and Experimental Investigation of Ejector Rocket Chamber for RBCC Propulsion System

Abstract: A GO2/kerosene ejector rocket chamber was designed and tested to meet the requirements of the RBCC propulsion system It was required to operate at chamber pressure 2MPa and 16 mixture ratio,while the mass flow rate of the rocket was changed from 95g/s to 285g/s The mass flow rate controlling,ignition,injection and thermal protection of the faceplate and chamber body have achieved reasonably good performance during the reacting flow tests and non-reacting flow tests Then the performance of the rocket was investigated in the RBCC combined tests As the results shown,the rocket has achieved efficiency 88%~98% C*,and it is influenced by the mass flow rate,mixture ratio,momentum flux ratio and pressure decrease of the rocket The higher momentum flux ratio and pressure decrease in a suitable range,the better performance of the atomizing,mixing and burning will attain Although the pressure decrease of the kerosene is only 11% of the design value,the GO2/kerosene Gas Center Swirl Coaxial(GCSC) injector can achieve efficienicy at least 88% C*by effective aerodynamic effects The purge flow should be controlled in tests and taken into account when the performance of the rocket is analyzed,as there will be efficiency decrease by 3% C*when the purge flow reaches 5% of the rocket mass flow rate When the performance of the rocket is compared in the separate tests and combined tests,the rocket isfound to attain almost efficiency 7% higher C*in combined test,because of its longer characteristic length

Adaptive Life Extending Control of Aircraft Engine in Whole Life

Abstract: Because of performance degradation of aeroengine components,a single life extending controller,which aims to extend component life,may excessively restrict the engine basic performance,thus making the engine not meet the normal needs. In order to solve this problem,based on the analysis of its basic performance and component life under engine performance degradation,appropriate life extending control strategy for different levels of performance degradation was designed. And a double-layer control system structure was used to coordinate switch of multiple controllers. The simulation results show that adaptive life extending control system can choose appropriate life extending control strategy based on current engine performance status,which can extend component life and maintain the basic performance of the engine in the whole life.

Sensor Fault Detection for EHA System Based on Adaptive Kernel Principal Component Analysis

Abstract: For the problem that it is hard to choose the kernel function of kernel principal component analysis(KPCA),a novel adaptive KPCA method for sensor fault detection is presented. The kernel function is modified adaptively according to the training date,so the kernel function can adapt to the given training date. Besides,the common method for data standardization processing is modified,an‘equalization'processing method is presented,which can eliminate the influence caused by amplitudes and dimensions of different variables, and the entire information of the training data can be reflected at the same time. Finally,an application of the proposed method is given in the sensor fault detection for airborne Electro-Hydrostatic Actuator(EHA) system,and the simulation results show that the proposed method is more advanced than the general KPCA,and it has much nicer performance in fault detection.

Non-Uniform Entrance and Non-Equilibrium Chemical Effects on Scramjet Nozzle Performance

Abstract: Using frozen and non-equilibrium chemical 3D models,kerosene-fueled scramjet nozzle flow was numerically simulated with a uniform or non-uniform entrance and RNG k-e turbulence model,in order to obtain the effects on the nozzle performance. The numerical results show that the non-uniformity of nozzle flow is composed of asymmetric expansion and non-uniform distribution in the entrance which has effects on the nozzle thrust,lift and pitching moment. Compared to the uniform entrance,the nozzle thrust increases about 1. 2% with a non-uniform entrance according to the simulated results. The non-equilibrium chemical effect to the nozzle performance is visible. The numerical value of nozzle thrust with non-equilibrium chemical model is 3% ~ 4% greater than that with frozen model.

Numerical Simulation of Open Structure of Rear Blunt Body in Advanced Vortex Combustor

Abstract: The numerical simulation has been carried out to investigate the effects of the rear blunt body with different open angle and slot size on the advanced vortex combustor(AVC).The results show that AVC has better flow characteristics when there is a slot in the rear blunt body. By selecting the appropriate open angle and slot size to optimize the open structure of the rear blunt body in advanced vortex combustor,stable combustion with lower pressure drop can be achieved. This new structure combustor can improve the cavity temperature effectively and make temperature distribution more uniform. For the rear blunt body of AVC structure,when half open angle θ is 50° and slot size is 2mm,AVC can achieve the best performance.

Research on Continuous Rotating Detonation Wave Propagation Process (I) : One Direction Mode

Abstract: Through changing the mass flow rates of air and hydrogen,H 2 / Air continuous rotating detonation wave has been realized in a wide range of test conditions. Based on the PCB and observation results, the detonation wave propagation process of one-direction rotating mode has been analyzed. For this rotating mode,the number of detonation waves is mainly influenced by the total mass flow rate of the propellant,and all the detonation waves propagate along the same direction at the same time. There are three kinds of propagation mode: one-wave,two-wave and hybrid one / two wave,which have their corresponding test condition domains. The detonation wave propagation frequency of one-wave and two-wave mode are in the ranges of 5. 05 ~ 5. 8 kHz and 8. 6 ~ 9. 9 kHz respectively,with the mean rotating velocity of 1510 ~ 1735m / s and 1280 ~ 1480m / s,respectively. For the hybrid one / two wave mode,the number of detonation waves changes during the test process. When the test condition locates in the middle of the condition domains of onewave and two-wave modes,the detonation wave propagates stably. Otherwise,the propagation direction may change during the test process.

Preparation,Characterization,and Effects on Thermal Decomposition of ADN of Nano Al/GAP Composite Particles

Abstract: In order to prevent aluminum nanoparticles further oxidation and inactivation in the air,aluminum nanopowders(n-Al)were pretreated with silane coupling agents,then coated with glycidyl azide polymer(GAP)under nitrogen atmosphere Nano-Al/GAP composite particles(n-Al/GAP)were obtained The surface morphologies and structures were characterized by scanning electron microscope(SEM),transmission election microscope(TEM),energy spectrum instrument(EDS),X-ray diffraction(XRD) instrument,Fourier transform infrared(FTIR)and X-ray photoelectron spectroscopy(XPS) The thermal decomposition processes of composite systems,such as ADN,n-Al/ADN and(n-Al/GAP)/ADN,were investigated using differential scanning calorimetry(DSC) The results show that the coupling agent plays a role as a bridge between GAP and aluminum nanopowders,and the core-shell composite particles are observed n-Al and n-Al/GAP have little influence on the liquefaction temperature of ADN,but the decomposition temperature increases significantly,and the effect of n-Al/GAP is more significant

Numerical Study on Flow Structure and Loss of Large Expansion Ratio Transonic Turbine

Abstract: The loss features of a typical large expansion ratio transonic turbine and the effects of two trailing edge cooling methods,including trailing edge ejection and pressure side ejection,on losses are investigated by numerical method. It can be found that most of the loss is profile loss which is about 65% of total loss and shock wave loss is the main source of profile loss. For the trailing edge ejection,the pressure at base region arises because of the coolant ejection which leads to decrease of the flow acceleration caused by the expansion wave. Thus the Mach number and shock wave loss are decreased. For the pressure side ejection,the trailing edge shock system is changed and the original shock wave is split into two or more than two weak shock waves which result in the decrease of shock wave loss. Both of the two trailing edge cooling methods are beneficial to reduce the shock wave loss of transonic turbine with large expansion ratio,but the pressure side ejection is more effective.

Axial Compressor Blade Automatic Optimization Design Based on a Quasi Three-Dimensional Approach

Abstract: For improving the effectiveness of axial compressor blade aerodynamic design,an axial compressor blade aerodynamic optimization design process based on the quasi three-dimensional approach is developed,which combines blade quasi three-dimensional approach with profile automatic optimization technique.Radial distribution of loss data from blade three dimensional calculation is fed back to S2 flow surface calculation,for replacing the original loss model based on the test data.With through-flow redesign,more precise calculation on S2 flow surfaces is developed.Automatic optimization design method based on parallel genetic algorithm and turbine CFD is applied to blade profile design on S1 flow surface.Replacing artificial modulation of blade aerodynamic profiles with process of numerical optimization,the optimization method reduces the dependence of artificial experience.A radial stacking rotor blade of axial compressor first stages is designed using the design process above.With total pressure ratio objective of 2.07 and mass flow-rate objective of 6.3kg/s,rotor designed shows isentropic efficiency about 90%,and reaches stall margin of15.5%.The proximity of objectives from throughflow design is satisfactory,too.Sweeping-bending design is conducted with 3D numerical optimization,which shows that the efficiency increase by 1%from the stall point to the working point.

Experimental Study on Impinging Velocimetry of Gel Simulants Based on PIV

Abstract: In order to study the atomization of gelled propellant,the effects with different pressure drop and impinging angle on atomization of impinging injector with gelled propellant were studied by a Time-Resolved Particle Image Velocimetry(TR-PIV),in which the injection impinging angles were 45°,60°,75°,90°and 120°,the pressures drop ranged from 0.4 MPa to 0.8MPa.The experimental results indicate that the distribution of atomization velocity is symmetry by the impinging axis and has one peak.The further from the strike spot,the smaller the atomization velocity and more uniform.The effective impinging speed of gel propellant is improved by increasing impinging angle and pressure drop.That is because the liquid kinetic energy converted to the energy of liquid broken is enhanced,and the atomization quality is improved.Finally,the gel propellant prepared in laboratory can be fully atomized when the effective impinging velocity is above 27.9m/s.

Numerical Investigations on Aerodynamic and Mechanical Performance of Long Blade Based on Two-Way Fluid-Structure Coupling Approach

Abstract: Detailed numerical investigation of a typical long blade based on the two-way fluid-structure coupling approach was conducted in order to study the effects of fluid-structure method on aerodynamic and mechanical performances Aerodynamic performance of long blade was analyzed using mesh deformation and Reynolds-Averaged Navier-Stokes( RANS) solutions The mechanical performance of long blade with the damper shroud and snubber was conducted using finite element analysis approach with consideration of the surface aerodynamic pressure and nonlinear contact between adjacent damping tip-shroud and snubber Nodes displacement and aerodynamic pressure of blade surface were interpolated in the fluid-solid coupling surface for maintaining the energy balance Stable long blade aerodynamic and mechanical performances were obtained by several iterations The numerical results show that the difference of static pressure distribution with and without coupling is mainly located at top half of long blade Absolute outlet flow angle decreases in the range of 40% ~ 95% span with consideration of fluid-structure coupling The maximum displacement of long blade appears at the 85% span near leading edge without coupling effect The maximum dis-placement position moves from leading edge to trailing edge and the quantitative value changes from 3 471mm to 3 082mm with consideration of fluid-structure coupling The maximum Von-Mises stress of long blade appears on the first tooth upper surface near leading edge of blade pressure side Von-Mises stress distributing trend and the location of the maximum Von-Mises stress are the same with and without coupling effect The maximum Von-Mises stress increases by 0 4% with consideration of fluid-structure coupling effect

Numerical Analysis of Flowfield in a Duct Embedded Rockets

Abstract: In order to investigate the influence on the duct flow fields,ejecting performance by changing the embedded rocket geometric distribution,operational conditions and geometric parameters,the cold flow fields in a duct embedded rockets were simulated numerically by using the Favre-Averaged Navier-Stokes equations with the two-equation k-ω SST turbulence model. The effects of the positions and combustion chamber pressure of embedded rockets on the duct flow field and the secondary flow injection performance were studied comparatively. The results of the calculation of different parameters show that:(1) The secondary flow mass flux grows larger as the combustion chamber pressure of the primary rocket increases as long as the rocket expanding jet current does not attach the duct side wall,after that it is difficult to increase the secondary flow entrainments.(2) As the inflow Mach number increased,ram compression role of the secondary flow becomes important,and it can be used to modulate duct flow field.(3) The embedded rocket operational conditions,which is consistent with the inflow conditions at low vehicle flight speed,could not meet the full power demand of the rocket propelling vehicle at takeoff,and duct embedded rocket concept such as strut jet concept will be more suitable for flight phase excluding takeoff.

Ablation Characteristics of C/C Composites In Reentry Simulation Conditions

Abstract: In order to study the ablation characteristics of axial carbon rod C/C composites,the ablation experiments were conducted by reentry simulation condition ablation system based on thermal plasma. The nitrogen(N2),oxygen(O2) and air were taken as the operation gases,and the ablation characteristics and ablation rates of the specimens in different conditions were compared. The results show that the micro-morphologies and the ablation rates are different in three operation gases conditions. The linear and mass ablation rates are listed as follows: oxidation of O2(0.0423mm/s and 0.0451g/s) nitridation of N2(0.0314mm/s and 0.0338g/s) composite reactions of air(0.0215mm/s and 0.0208g/s). There are some tiny cracks in thermal influencing areas,and the ablation mechanisms of the C/C composites in three operation gases conditions are differents. They are some combinations of sublimation of carbon,carbon oxide and carbon nitrogen reaction.

Effects of Step Height on Low Emission Stirred Swirl Combustor

Abstract: A single annular lean-burn internally-staged combustor consists of pilot stage and main stage.The pilot stage is located in the center-line of the combustion chamber while the main stage surrounds coaxially the pilot stage. There exists a physical isolation called step height between the two stages. The emission tests using a single module rectangular combustor were operated under the idle condition to study the effects of step height on NOX. The commercial software Fluent was used to simulate the cold flow field and combustion performance. The results indicate that step height has a great influence on the emission of NOX. When the step height increases by 38%,the surface and volume of primary center recirculation zone will increase by 6.7% and 16.4%.However,the axial length has a decrease of the small amount and the emission of EINOX will increase by 35.1%.

Propulsive Performance of Pulse Detonation Engine with Fluidically Augmented Nozzle

Abstract: Both generalized 1-D flow model and constant volume cycle model are employed to investigate the performance of pulse detonation engine with a fluidic nozzle. Thrust coefficient of a fluidic nozzle in steady flow and average thrust coefficient of a pulse detonation engine with a fluidic nozzle during a pulse detonation cycle have been calculated. In the steady flow,fluidic nozzle shows a better performance than its baseline nozzle when nozzle pressure ratio is lower than a critical value. During a pulse detonation cycle,for a certain initial uniform combustion pressure,the fluidic nozzle with a continual second injection can augment the average thrust coefficient when the expansion ratio is higher than a critical value. Interrupted injection can improve the max average thrust coefficient for this initial uniform combustion pressure by 2.4% when initial uniform combustion pressure is 5MPa.

Effects of Two-Stage Pulse Channel Configurations on Local Heat Transfer Characteristics in Combustion Chamber of Dual Pulse Motor

Abstract: In order to study the effects of two-stage pulse channel configurations on reattachment location and relative convective heat transfer coefficient in combustion chamber of dual pulse motor,the Reynolds-average Navier-Stokes equations were solved with three-order three-step Runge-Kutta iterative algorithm by cell center upwind finite volume method and AUSM-PW scheme were implemented for spatial discretization. Also,modified SST(shear-stress-transport) turbulence model that improved the capability of predicting separation was used to simulate the turbulence flow. The results show that the maximum absolute error of subsonic flow over backward facing step and solid rocket motor combustion chamber flow are within 8.9% and 5.8%,respectively.During the second pulse phase,when the inner diameter of the second grain larger than the one of the pulse channel,as the diameter of pulse channel increases by 9.1%,the reattachment location and relative convective heat transfer coefficient decrease by an average of 28.2% and 9.6%,respectively. When the other conditions are constant,as the channel angle increases,the reattachment location and relative convective heat transfer coefficient reduce by about 3.4% and 3.1% in average,respectively while the value varies slightly with increasing the pulse channel width.

Impact of Mach Number of External Flow on Performance of Counterflow Thrust Vectoring Nozzle

Abstract: In order to find the effects of different Ma of the external flow on the performance and structure of internal flow field of counterflow thrust vectoring nozzle,referring to the experiment,numerical simulation was applied to study the flow field,vectoring angle and mass flux ratio of secondary flow under different Ma(0~ 1.5).The results show that,as the Ma increases from 0 to 0.8,the flow filed in the nozzle changes so as to change the difference of pressure across the primary flow,resulting in the vectoring angle decrease,and the mass flux ratio of secondary flow will increase quickly.As the Ma increases from 0.8 to1.5,both the flow filed and the vectoring angle have no change,and the mass flux ratio of secondary flow will increase slowly.

Experimental Research for Effects of Bearing Supply Gas Pressure on Aerostatic Bearing-rotor System Dynamic Characteristics

Abstract: In order to study the effects of bearing gas supply pressure on the dynamic characteristics of aerostatic bearing-rotor system,the corresponding experiments were carried out on the turbo-expander refrigerator test bed with the coaxial centrifugal compressor and a radial turbine rotor structure supported by aerostatic bearings The effects of bearing supply pressure on the critical speed and low frequency characteristics were analyzed by vibration test and analysis methods including bifurcation diagrams,axis center tracks,and frequency spectrum characteristics The results show that gas film half-speed whirl disappeared at 26000r/min when the bearing supply pressure was at 080 MPa,and when the bearing supply pressure was at 075 MPa,vibration amplitude of low frequency at 50000r/min only accounted for 5 percent of the corresponding operating frequency vibration amplitude,which was far smaller than the relative value(the ratio of low-frequency vibration amplitude and operating frequency vibration amplitude was 3489 percent) at the same rotational speed for 080 MPa bearing supply pressure Consequently,gas film half-speed whirl is eliminated by the increase of bearing supply pressure,which brings the increase of gas film direct stiffness The vibration amplitude of gas film whip is suppressed by the reduction of bearing supply pressure,which brings the decrease of gas film damping ratio

Effects of Different Aluminum Particle Sizes in Composite Propellant on Distributed Combustion Response and Particle Damping

Abstract: According to the aluminized composite propellants with different initial particle sizes,the characteristics of the particle damping of condensed phase combustion products and the distribution combustion response were researched. The result shows that,particle damping is dependent on the particle size distribution of the condensed phase combustion products and the oscillation frequency,which is reasonably insensitive to the initial aluminum particle sizes. On the aspect of estimating the particle damping,the particle damping computed using the mono-sized particles were different from the value in the experiment,with the relative error more than10 percent. Thus,further improvement on the prediction method is needed. On the aspect of distributed combustion response,the distributed combustion response is sensitive to the initial aluminum particle sizes in the same oscillation frequency,that is,the bigger the sizes of the initial aluminum particles are,the higher the combustion driven in the combustion,which is not good in solid rocket motor from a combustion instability point of view.