2019 Vol. 34, No. 8

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Numerical study on heat transfer mechanism of brush seal based on three-dimensional solid modeling
2019, 34(8): 1633-1643. doi: 10.13224/j.cnki.jasp.2019.08.001
Abstract PDF(3740KB) icon783 icon395
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A model for solving the heat transfer characteristics of brush seals based on three-dimensional solid modeling was established. The distribution characteristics of the flow field and temperature field of the brush seal were analyzed by verifying the accuracy of the numerical model. The influences of the pressure ratio, the rotational speed and the interference on the maximum temperature of the brush wire were studied. The heat transfer mechanism of the brush seal was revealed. The results showed that the end-discharge brush wire and the rotor face were mainly concentrated in the high temperature area of the bristle. The maximum temperature of the brush wire increased with the increase of the pressure ratio, rotational speed and interference, in which the interference had obvious influence on the maximum temperature. When the amount of interference increased from 0.1 mm to 0.7 mm, the maximum temperature increased 1.61 times; The main source of heat of the brush seal was generated by friction between the brush and the rotor surface, the heat transfer form of the brush seal included conduction heat and convection heat exchange. The friction heat flow entered the brush wire bundle and the rotating shaft through the heat conduction form. When the contact force between the brush wire and the rotor increased, the frictional heat flux increased, and the maximum temperature rose. The friction heat flowed through the heat transfer between fluid and solid transmission, and the main form of heat dissipation was the leakage of air taking away part of the heat.
Experiment on flow boiling heat transfer in 1.002 mm tube under hypergravity
2019, 34(8): 1644-1651. doi: 10.13224/j.cnki.jasp.2019.08.002
Abstract PDF(1445KB) icon758 icon381
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An experiment investigation of R134a flow boiling heat transfer in a 1.002 inner diameter mini-tube under hypergravity was carried out, with the hypergravity environment being achieved using a centrifugal acceleration machine. Results showed that a remarkable difference existed in flow boiling heat transfer between hypergravity and normal gravity. The heat transfer under hypergravity was stronger than that under normal gravity in the gravity of 3.16g (g=9.8 m/s2). With the increase of gravity, heat transfer coefficients increased at first and then decreased. The transition took place around 1.1g-1.4g. At 3.16g, some heat transfer coefficients became smaller than those at normal gravity. The effects of vapor quality on the heat transfer varied with gravity. The applicability of correlations for flow boiling heat transfer under normal gravity to hypergravity was also evaluated, and those having good prediction performances for hypergravity were identified.
Simulation analysis and optimization on test method of combustor airflow distribution
2019, 34(8): 1652-1662. doi: 10.13224/j.cnki.jasp.2019.08.003
Abstract PDF(5681KB) icon859 icon422
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In order to specify the rationality of the method of blocked holes in the current combustor flow distribution test, the numerical simulation was conducted according to the actual test procedure of a can combustor flow distribution, the test method error and the root of the generation were analyzed, the optimizational test method of associating blocked holes and accumulated blocked holes of single side was proposed, and the validation was carried out by numerical simulation. Results indicated that there were definite method errors in the flow distribution test using blocked holes method, the effects of this method on the measurement error of primary holes, dilution holes and film cooling holes on the liner were greater, while the effect on dome measurement error was smaller; the simultaneous neglect of passing flow condition of tested holes and downstream holes was the main reason; the measurement error can be reduced significantly by associating blocked holes and accumulated blocked holes of single side for the flow distribution measurement of primary holes, dilution holes and film cooling holes, which can be reduced by about 6%, 12% and 9%, showing definite improvements; There was method error before and after the test method was optimized, the static pressure difference of import and export sections of each holes on the liner changed when some holes were blocked, which was the primary cause for the existing method error, but the variation of the static pressure difference and its influence on the measurement results were smaller after the test method was optimized.
Laminar combustion characteristics of chlorella oil/RP-3 kerosene mixed fuel
2019, 34(8): 1663-1670. doi: 10.13224/j.cnki.jasp.2019.08.004
Abstract PDF(2954KB) icon755 icon426
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The laminar combustion characteristics of chlorella oil and its blends with RP-3 kerosene were examined at initial pressure of 0.1 MPa, initial temperature of 450 K and equivalence range of 0.8-1.2 conditions. The results showed that with the increase of equivalence ratio, the ignition delay period of chlorella oil decreased and the stretched flame speed increased. Compared with RP-3 kerosene, the peak value of unstretched flame speed of chlorella oil was more shifted to the region of rich mixture, and the unstretched flame speed of chlorella oil was more sensitive to equivalence ratio. It was observed that with the increase of equivalence ratio, the unstretched flame speed of chlorella oil varied significantly. With the increase of chlorella oil addition, the peak value of unstretched flame speed shifted to right, the peak unstretched flame speed of the mixed fuel with 50% chlorella oil and 50% RP-3 appeared near the equivalence ratio of 1.4. It was found that compared with chorella oil and RP-3 single fuel, the Marksein length of the mixed fuel was more larger, indicating the combustion of the mixture with chlorella oil and RP-3 is more stable.
Knocking control of piston-type aviation kerosene engine based on dual-spark plug ignition strategy
2019, 34(8): 1671-1676. doi: 10.13224/j.cnki.jasp.2019.08.005
Abstract PDF(2132KB) icon809 icon423
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Focusing on dual-spark plug ignition forms, test studies were conducted on a 650 mL piston-type single engine using kerosene as fuel. Result showed that the ignition advance angle of dual spark plug synchronous ignition delay can effectively restrain detonation, meanwhile, the amplitude of combustion heat release rate decreased gradually, and the whole combustion phase was delayed gradually. With the increase of ignition phase difference of ignition advance angle, the knock intensity decreased gradually. By matching the ignition advance angle of the major spark plug with the ignition advance angle of the auxiliary spark plug, the power performance of the engine can be further improved.
Numerical calculation of flow characteristics of hybrid vortex reduce structure combing de-swirl inlet
2019, 34(8): 1677-1687. doi: 10.13224/j.cnki.jasp.2019.08.006
Abstract PDF(5001KB) icon904 icon399
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A hybrid vortex reducer combining the de-swirl inlet hole and the vortex tubes was proposed, the effect of drag reduction and bleed air was numerically studied, and the effects of rotating Reynolds number and dimensionless inlet mass flow on internal flow field structure and pressure loss were analyzed. It was found that in the inflow structure of the hybrid vortex reducer, the static pressure decreased gently in the radial direction and was evenly distributed in the circumferential direction. With the ascension of the dimensionless inlet mass flow or the rotating Reynolds number, the total pressure drop of the inflow structure showed a monotonous increasing trend. Especially, under the conditions of higher rotating Reynolds number and lower dimensionless mass flow, outstanding drag reduction was observed and the corresponding turbulence parameter range of 0.106 4-0.324 5 was obtained. Compared with the simple disk cavity, the pressure loss of the de-swirl vortex reducer and the tube vortex reducer was reduced by 62.5% and 60.5%, respectively, and the hybrid vortex reducer can reduce the pressure loss by 80.4%, showing a better drag reduction.
Numerical study on flow field characteristics of RQL combustor under different quenching structures
2019, 34(8): 1688-1698. doi: 10.13224/j.cnki.jasp.2019.08.007
Abstract PDF(8063KB) icon746 icon394
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The Reynolds averaged Navier-Stokes (RANS) method was used to simulate the steady flow process in rich-burn/quench/lean-burn（RQL）combustors of different quenching structures. The flow field structure of different cross sections was analyzed to study the effects of momentum flux ratio and quenching hole arrangement. Results showed that the jet deflected after entering the RQL combustor and the deflection direction was related to the swirling direction. The recirculation zone length, jet depth and high kinetic energy zone increased with the increase of the momentum flux ratio. When J≥80, a high kinetic energy zone appeared downstream the quenching jet, and gradually merged with the upstream. When J≥120, the length of the recirculation zone was substantially no longer changed. In addition, the non-central opposed jet structure C3 had a larger turbulent kinetic energy zone with better mixing.
Unsteady aerodynamic characteristics of ducted coaxial propeller unmanned aerial vehicle
2019, 34(8): 1699-1707. doi: 10.13224/j.cnki.jasp.2019.08.008
Abstract PDF(6012KB) icon835 icon405
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Based on the sliding grid technique and the SST(shear stress transport)k -ω turbulence model, a CFD(computational fluid dynamics) numerical method for the flow field of ducted coaxial propeller UAV(unmanned aerial vehicle) was established, and its effectiveness was verified by wind tunnel experiment results. The dynamic flow field of the ducted coaxial propeller UAV was numerically simulated. The influences of the factors such as duct, flight velocity, propeller speed and angle of attack on the unsteady aerodynamic characteristics of the ducted coaxial propeller UAV were analyzed. The calculation results show that the established CFD numerical method is suitable for simulating the dynamic flow field of ducted coaxial propeller UAV. The existence of duct significantly weakens the tip vortex, trailing edge vortex, propeller wake contraction, blade-vortex interaction and vortex-vortex interaction, and reduces the required power of the ducted coaxial propeller UAV. The lift and drag of the ducted coaxial propeller UAV increase as the forward velocity increases. The lift of the ducted coaxial propeller UAV increases and the drag decreases as the propeller speed increases. The drag of the ducted coaxial propeller UAV increases and the lift increases first and then decreases as the angle of attack increases.
Error in rotating sound source localization caused by installation deviation of microphone array
2019, 34(8): 1708-1716. doi: 10.13224/j.cnki.jasp.2019.08.009
Abstract PDF(3999KB) icon609 icon391
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The localization error caused by three types of common array installation deviation, namely angular deviation, x-axis offset deviation and z-axis offset deviation was studied in detail in the rotating sound source localization. Research showed that when angular deviation was within the range from -10° to 10°, the position error varied approximately linearly with the angular deviation, the maximum positioning error was 0.089 m, the strength error was randomly distributed, the minimum strength error was 0.97 dB and the maximum strength error was 4.69 dB; when x-axis offset deviation was within the range from -0.1 m to 0.1 m, the position error varied approximately linearly with x-axis offset deviation, the maximum positioning error was 0.098 m, the strength error was randomly distributed, the minimum strength error was 0.91 dB and the maximum strength errors was 4.94 dB; when z-axis offset deviation was within the range from -0.1 m to 0.1 m, the maximum positioning error was 0.01 m, strength error caused by the positive z-axis offset deviation was smaller than the negative z-axis offset deviation generally, the minimum strength error was 0.81 dB and the maximum strength error was 4.51 dB. Research results can provide guidance for controlling the localization error of microphone arrays in practical applications.
Performance change of aero turbofan engine after swallowing water
2019, 34(8): 1717-1723. doi: 10.13224/j.cnki.jasp.2019.08.010
Abstract PDF(2504KB) icon869 icon468
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In order to understand the turbofan engine performance change after swallowing water，the test of one turbofan engine performance change after swallowing water was designed and carried out on the conventional testing facility. When the engine was at idling，throttling，max speed，and afterburning，the mass of swallowing water was 2%, 3.5% and 5% of the inlet air mass flow. The test results showed that: at the idling，the lowering degree of the inner and outer culvert outlet temperature of the fan did not change with the change of the mass of water swallowing，after swallowing water the engine thrust decreased about 20% at most；at the afterburning，the decreasing degree of the outer culvert outlet temperature of the fan grew with the increase of the mass of water swallowing, after water was swallowed, due to the low-pressure speed increase，engine thrust increased about 15% at most. No stalling, surge and flameout of the combustion chamber of the engine were caused during the swallowing process. Engine performance did not deteriorate after water was swallowed.
Effect of discharge resistor on performance of plasma synthetic jet actuator
2019, 34(8): 1724-1730. doi: 10.13224/j.cnki.jasp.2019.08.011
Abstract PDF(2061KB) icon654 icon367
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A plasma synthetic jet actuator (PSJA) with normal orifice was designed to study the effect of discharge resistor. The discharge characteristics and transient flow-field characteristics of actuator with 100,200,300 Ω were researched by electric parameter measurement system and high-speed schlieren technology. Results showed that the discharge process can be divided into two stages due to the existence of current limiting resistance, rapid discharge stage and slow discharge stage.This mode can continuously provide energy to the actuator while satisfying the higher initial energy injection, and effectively improve the stability of the actuator.At the same time, the smaller discharge resistance can obtain greater jet velocity, enhancing the jet flow control capability.The smaller discharge resistance led to the increase of discharge current, unstable velocity and worsening of working stability.In practical application, it was necessary to consider and determine the final resistance value comprehensively to ensure the rational distribution of energy in the two discharge stages.
Unstructured grid space marching based design optimization of hypersonic vehicle back sketch lines
2019, 34(8): 1731-1740. doi: 10.13224/j.cnki.jasp.2019.08.012
Abstract PDF(5183KB) icon804 icon395
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The design optimization of hypersonic vehicle back sketch lines was investigated. A parametric geometry model was presented based on free form deform (FFD) method and Hicks-Henne function, which was capable of keeping the back surface smooth and continuous during deforming. In order to save time cost, unstructured grid space marching method was adopted to solve the Euler equations. The accuracy and efficiency of this method were analyzed by DOE methods through hypersonic cone cases. There was a good conformity between the un-structured grid space marching and time marching results of the hypersonic vehicle flow. Indicated by the comparisons, the unstructured grid space marching method had the capacity to increase the computational efficiency over 4.5 times. A response surface model was built through optimize latin hypercube sampling. The Pareto effects of design variables were analyzed. The multi-objective optimization algorithm NSGA-Ⅱ (non-dominated sorting genetic algorithm-Ⅱ) was adopted to optimize the drag coefficient and lift-to-drag ratio under the constraint of the vehicle volume change within 20%. The optimization history showed a negative correlation between the drag coefficient and lift-to-drag ratio. The drag coefficient of the optimal configuration decreased by 7.1% and the lift-to-drag ratio increased by 16.9%.
Aerodynamic optimization design of general parameters for cycloidal propeller in hover based on surrogate model
ZENG Jianan, ZHU Qinghua, WANG Kun, ZHU Zhenhua, SHEN Suiyuan
2019, 34(8): 1741-1750. doi: 10.13224/j.cnki.jasp.2019.08.013
Abstract PDF(3523KB) icon593 icon363
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A surrogate-model-based aerodynamic optimization design method for cycloidal propeller in hover was proposed, in order to improve its aerodynamic efficiency, and analyze the basic criteria for its aerodynamic optimization design. The reliability and applicability of overset mesh method were verified. An optimization method based on Kriging surrogate model was proposed to optimize the geometric parameters for cycloidal propeller in hover with the use of genetic algorithm. The optimization results showed that the thrust coefficient was increased by 3.56%, the torque coefficient reduced by 12.05%, and the figure of merit (FM) increased by 19.93%. The optimization results verified the feasibility of this design idea. Although the optimization was only carried out at a single rotation speed, the aerodynamic efficiency was also significantly improved over a wide range of rotation speeds. The optimal configuration characteristics for micro and small-sized cycloidal propeller were: solidity of 0.2-0.22, maximum pitch angle of 25°-35°, pitch axis locating at 35%-45% of the blade chord length.
Prediction of hybrid airfoil leading edge pressure distribution based on deep learning
2019, 34(8): 1751-1758. doi: 10.13224/j.cnki.jasp.2019.08.014
Abstract PDF(2225KB) icon901 icon475
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A prediction model on the leading edge pressure distribution of the hybrid airfoil based on deep learning was proposed. A convolutional neural network model (CNN) was established on the basis of the geometric feature extraction of the hybrid airfoil and the parameterization of the pressure distribution curve. A group of hybrid airfoils with different trailing edges were analyzed by a verified CFD method. The CFD results were used as the training set of the CNN. Results show that the goodness of fit of the calculation results of the two methods exceeds 0.98. The proposed prediction method based on deep learning takes 1.7 s and CFD method takes more than 50 s, computation time is greatly reduced. The proposed method can improve the computational efficiency with satisfying the calculation accuracy and it can be applied to the design processes of other airfoils.
Measurement data self-calibration fusion method
2019, 34(8): 1759-1763. doi: 10.13224/j.cnki.jasp.2019.08.015
Abstract PDF(698KB) icon750 icon375
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A measurement data self-calibration fusion method was proposed. Self-recognition self-calibration formulas and computational procedures of the systematic error (unknown input) in measurement data were given, enabling to automatically identify, estimate, compensate and correct the systematic error, so as to reduce its influence. Formulas and computational procedures of multi-sensor data fusion were established to further reduce the influence of random errors. The linear system and the non-linear system were discussed in detail respectively, and a large number of examplifications and simulations were carried out. It can be seen in two examples that the relative error of the proposed method is at least 45% less than that of the traditional method, and the calculation is simple, and well-suited for engineering applications.
Reliability based multidisciplinary design optimization of single turbine stage based on parameterized mesh deformation
2019, 34(8): 1764-1772. doi: 10.13224/j.cnki.jasp.2019.08.016
Abstract PDF(2930KB) icon1016 icon445
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On basis of free-form deformation method, a parameterized mesh deformation to morph coordinately fluid analysis mesh and structure analysis mesh was proposed, and reliability-based multidisciplinary design optimization of single turbine stage with parameterized mesh deformation was introduced. According to the extended feature of blade in spanwise direction, the mesh deformation of 3D analysis mesh was translated to the mesh deformation of 2D cross section analysis mesh at different blade height. Control plane for 2D cross section was designed on basis of topological structure of the 2D fluid analysis mesh. The coordinates of blade profile control points were calculated by aerodynamic design parameters. The movements of other control points in 2D control plane were obtained according to the movement of blade profile control points in order to improve the quality of the deformed mesh. The mesh deformation of 3D fluid analysis mesh could be carried out by the control volume composed of 2D control planes at different blade heights. Structure analysis mesh was morphed by using the same control volume for 3D fluid analysis mesh deformation in order to obtain the coordinate mesh deformation on interface mesh. Reliability-based multidisciplinary design optimization with proposed parameterized mesh deformation method was used to optimize single turbine stage. Results showed that the efficiency was improved by 4.97% and the blade life was improved by 40.86% under satisfying the reliability constraint, proving the effectiveness of the developed method.
Symmetrical ablation mechanism of exhaust valve at a certain aero piston engine
2019, 34(8): 1773-1780. doi: 10.13224/j.cnki.jasp.2019.08.017
Abstract PDF(1972KB) icon803 icon419
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By measuring actual operating temperature of exhaust valve body near the valve sealing surface, the corresponding relationship between valve operating temperature and exhaust gas temperature (EGT) at this engine was obtained, and the mechanism of symmetrical distribution of valve ablation point was analyzed in combination with the engine structural design and actual operating environment. Results showed that the valve deflection led to uneven circumferential symmetry distribution of valve sealing surface temperature and the high temperature of valve sealing surface due to excessive leaning in operation, bringing about exhaust valve symmetrical ablation. Based on this, solution of limiting the leaning range in operation with EGT as reference was put forward, and the 70 000 flight hours practical test showed that the fault has been completely solved.
Numerical simulation on cavitation flow field characteristics of squeeze film damper based on two-phase flow model
2019, 34(8): 1781-1787. doi: 10.13224/j.cnki.jasp.2019.08.018
Abstract PDF(2435KB) icon745 icon405
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A computational fluid dynamics model was established for squeeze film damper with a central feeding groove, based on the Mixture model and the Zwart-Gerber-Belamri vapor cavitation model. The unsteady gas-liquid two-phase flow field in squeeze film damper was simulated by employing the dynamic mesh technique. Numerical simulations showed that with the precession of damper inner ring, two negative-pressure strips were generated in the low-pressure zone of damper symmetrically about the central feeding groove, and there was a high vapor volume fraction in the strips. It demonstrated that pressure distribution in the low-pressure zone and vapor volume fraction were sensitive to position of inlet hole, and variation frequencies of the oil film force and flow field cavitation ratio were closely related to the number of inlet holes. The simulation of the influence of centered precession radius and frequency of inner ring on cavitation flow field shows that the increasing precession radius and frequency both aggravate the cavitation phenomenon, and the phase lag phenomenon of the cavitation ratio becomes more significant.
Load reducing mechanism under fan blade out event
2019, 34(8): 1788-1794. doi: 10.13224/j.cnki.jasp.2019.08.019
Abstract PDF(2268KB) icon705 icon386
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By using a simplified rotor finite element model，comparative analyses with and without load reduce device (LRD) design were carried out,including the critical speed,dynamic load response and the axis trajectory of the rotor. The mechanism of structural load reduction under fan blade out was studied. Analysis results indicated that due to the fuse design of the 1# bearing,the low-pressure rotor support changed from three bearings to two bearings.As a result, the critical speed of fan mode was reduced, the radius of axis trajectory was decreased,and fan rotor ran at supercritical speed,which reduced load response at high speed.The method adopted in this mechanism study can be used to determine the load threshold range of fused structure and to evaluate its load reduction effect.
AP clustering improved immune algorithm for aeroengine fault diagnosis
2019, 34(8): 1795-1804. doi: 10.13224/j.cnki.jasp.2019.08.020
Abstract PDF(3226KB) icon722 icon386
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In the process of immune algorithm training, the affinity propagation(AP) clustering and entropy weight method were introduced, the training samples were clustered and weighted, and the weights were introduced into the calculation of the sample selection threshold in the immune algorithm to solve the problem of a fixed selection threshold in the training process, which led to over fitting of the detector in a partial area, and under-fitting of the partial area. Result showed that, when the improved immune algorithm was used for the optimization of typical nonlinear functions, the iterative performance was better than the traditional immune algorithm. In most cases it was better than the particle swarm optimization algorithm and the quantum genetic algorithm, in the case of a certain type of engine fault diagnosis. The improved algorithm had a diagnostic accuracy of 98.06%, which was higher than 92.60% of the traditional immune algorithm.
Simulation of the crack deflection mechanism for ceramic matrix composites with multilayer interphase
2019, 34(8): 1805-1812. doi: 10.13224/j.cnki.jasp.2019.08.021
Abstract PDF(2090KB) icon925 icon418
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The crack deflection mechanism in ceramic matrix composites (CMCs) with multilayer interphase was simulated by finite element method. The sub-layers of interphase were created according to their real thickness within a cylinder unit-cell model to describe the microstructure of multilayer interphase of CMCs. The finite element method was applied to analyze the stress by defining different material properties for the interphase sub-layers. The energy release rates Gd and Gp for two cases, i.e. penetration and deflection, were calculated by the virtual crack closure technique (VCCT). After all, a simulation method for crack deflection in the multilayer interphase was developed. It can be seen that the anisotropic interface is more conducive to the occurrence of crack deflection than the isotropic interface phase as the value of Gd/Gp is higher. There is more chance for crack deflection in the multilayer interface phase than the monolayer interphase with same thickness according to the distribution of Gd/Gp in the interphase. Hence, the crack deflection is easier in the (PyC/SiC/PyC/SiC/PyC) interphase compared with the (PyC/SiC/PyC) interphase.
Experimental and numerical study on axially non-uniform clearances with a squealer cavity tip
2019, 34(8): 1813-1821. doi: 10.13224/j.cnki.jasp.2019.08.022
Abstract PDF(5550KB) icon661 icon405
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Experimental method and numerical method were both used to study the effect and mechanism of the axially non-uniform clearances on turbine leakage flows with a squealer cavity tip. Two clearance heights and three clearance types: shrinking clearance, uniform clearance and expanding clearance, were investigated. The following findings were made. Firstly, there was no relation between the reduction of the leakage mass flowrate and the reduction of the total pressure loss. The leakage mass flow rate increased by 26.7% under the expanding clearance case, while the corresponding total pressure loss achieved a 2.44% reduction under small clearance and a 3.53% reduction under large clearance. Moreover, the reduction of the total pressure loss was obtained by the decrease of the scale of the passage vortex in pitch-wise and in span-wise directions and the decrease of the leakage vortex in pitch- wise direction. The arrangement of the expanding clearance strengthened the span-wise flows of the leakage flows in the squealer cavity, led to a closer reattachment line with the blade on the pressure side and a further back leakage vortex separation line, and finally restricted the leakage flows.
Mechanism of affect on stall margin of circumferential grooves with variable pre-swirl angle
2019, 34(8): 1822-1835. doi: 10.13224/j.cnki.jasp.2019.08.023
Abstract PDF(6163KB) icon634 icon381
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Numerical simulations were performed for Rotor 67 to analyze the mechanism of the stall margin(SM) improvement of circumferential grooves(CG) casing treatment, and the matching between CG and pre-swirl angle. Results showed that stall in Rotor67 was mainly attributable to the low velocity region arisen by the inference of tip leakage flow and shock wave. The mechanism of SM improvement was that the flow from the pressure side of adjacent blade caused a vortex counter-rotating to the leakage vortex(LV), which inhibited the development of LV. The most important groove(s) changed while the distribution of static pressure changed because of the varying mass flow or pre-swirl angle. A design method was presented depending on the research results.
Effects of gas nozzle and resonators on acoustic characteristics of combustion chamber
2019, 34(8): 1836-1846. doi: 10.13224/j.cnki.jasp.2019.08.024
Abstract PDF(4171KB) icon864 icon393
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In order to obtain the influence of gas-liquid nozzle and acoustic resonators on the acoustic characteristics of combustion chamber and explain the mechanism of tangential frequency differentiation in acoustic experiments, a model combustion chamber with cylindrical tubes was investigated by acoustic finite element method (FEM). The analysis results were verified by the experiment of a single injector chamber. On this basis, the influence rules of gas nozzle, orifice and resonators on acoustic frequency were established in terms of the frequency and sound pressure distribution. The phenomenon of frequency differentiation was explained by the sound pressure distribution results. Results showed that the energy of tangential mode transferred into the gas nozzle and changed the original mode of combustion chamber, when the frequency of first order longitudinal mode of gas nozzle or resonator was equal to that of the tangential mode of combustor, and the phenomenon of frequency differentiation in tangential modes emerged in acoustic experiments. The coupling relationship between gas nozzle length and diameter of orifice was discovered, enabling to make combinatorial parameter optimization during rocket engine design.

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