Plume Flowfield Research Articles

Investigation on a coupled Navier–Stokes–Direct Simulation Monte Carlo method for the simulation of plume flowfield of a conical nozzle

SUMMARYTo assess the plume effects of space thrusters, the accurate plume flowfield is indispensable. The plume flow of thrusters involves both continuum and rarefied flow regimes. Coupled Navier–Stokes–Direct Simulation Monte Carlo (NS–DSMC) method is a major approach to the simulation of continuum‐rarefied flows. An axisymmetric coupled NS–DSMC solver, possessing adaptive‐interface and two‐way coupling features, is investigated in this paper for the simulation of the nozzle and plume flows of thrusters. The state‐based coupling scheme is adopted, and the gradient local Knudsen number is used to indicate the breakdown of continuum solver. The nitrogen flows in a conical nozzle and its plume are chosen as the reference case to test the coupled solver. The threshold value of the continuum breakdown parameter is studied based on both theoretical kinetic velocity sampling and coupled numerical tests. Succeeding comparisons between coupled and full DSMC results demonstrate their conformities, meanwhile, the former saves 58.8% computational time. The pitot pressure evaluated from the coupled simulation result is compared with the experimental data proposed in literatures, revealing that the coupled method makes precise predictions on the experimental pitot pressure. Copyright © 2014 John Wiley & Sons, Ltd.

Coupled Navier–Stokes/Direct Simulation Monte Carlo Simulation of Multicomponent Mixture Plume Flows

A coupled particle-continuum simulation method is presented to compute the multicomponent mixture plume flow from the attitude-control engines of a satellite through the use of the Navier–Stokes solver and direct simulation Monte Carlo method. In this study, the implicit hybrid flux iteration scheme for the axisymmetric compressible Navier–Stokes equations is constructed to solve the inner flowfield of the nozzle. The Navier–Stokes computing technique is implemented considering slip boundary theory for the near-continuum slip flow near the nozzle exit. Direct simulation Monte Carlo methods for the axisymmetric core plume and three-dimensional far-field plume flows of multispecies mixture gas are developed by using a radial weighted factor, by tracking the molecular motion trajectory, and with secondary Cartesian and unstructured cell processing techniques. A multiregion decomposition and hybrid Navier–Stokes/direct simulation Monte Carlo algorithm with two-way and one-way coupling is established to solve the internal and external flow of the thruster, including the near-field, far-field, backflow, and gas-surface contaminated regions. After constructing the coupled Navier–Stokes/DSMC simulation scheme, the present method is applied to solve the plume flowfield and impinging contamination effects from the satellite attitude-control engine and solar array panel; the simulation results correspond well with the experimental measurements from the low-density wind tunnel and the theoretical predictions. To study the contamination effects produced by the multicomponent mixture plume, the current method is employed to simulate a five-component mixture plume flowfield of tens of meters from two representative attitude-control engines installed in different locations of a satellite in orbit. The results of the nonreacting multicomponent flow with equivalent mole fractions can be achieved with a one-component gas simulation if the species is assigned the properties of the mixture. The deposition rate of the plume flows produced by the two teamwork engines can be superimposed on each other. This method can be used to efficiently predict the impingement contamination effects of the gas-fired mixture plume flow on the satellite and solar array panels.

Effects of rotor downwash on exhaust plume flow and helicopter infrared signature

The effects of rotor downwash and exhaust direction on plume flow field, rear-fuselage temperature distribution and helicopter infrared signature were numerically investigated. The internal flow inside IR suppressor originated from engine exhaust nozzle and the external flow around helicopter airframe originated from rotor downwash were computed in a coupled mode to determine the temperature distributions on the helicopter skin and in the exhaust plume. Based on the skin and plume temperature distributions, a forward–backward ray-tracing method was used to calculate the infrared radiation intensity from the helicopter with a narrow-band model. The results show that the exhaust plume takes on strong downwards deflection to the rear-fuselage, as well as to the rotor rotational direction under the action of rotor downwash. The rotor downwash has a complicated influence on the infrared radiation distribution of helicopter. It is benefit for reducing the infrared radiation intensity when the exhaust is ejected in oblique-turned or lateral-turned mode. While for the up-turned exhaust mode, the exhaust plume could heating the helicopter rear-fuselage and the infrared radiation intensity may be enhanced under the action of downwash.

A Numerical Study of Suspension Injection in Plasma-Spraying Process

Suspension feedstock in plasma spraying opened a new chapter in coating process with enhanced characteristics. The suspension carrying sub-micron up to few micron-sized particles is radially injected into an atmospheric plasma plume. Understanding the trajectory, velocity, and temperature of these small particles upon impacting on the substrate is a key factor to produce repeatable and controllable coatings. A three dimensional two-way coupled Eulerian-Lagrangian scheme is utilized to simulate the flow field of the plasma plume as well as the interactions between the evaporative suspension droplets with the gas phase. To model the breakup of droplets, Kelvin-Helmholtz Rayleigh-Taylor breakup model is used. After the breakup and evaporation of suspension is complete, the solid suspended particles are tracked through the domain to determine the characteristics of the coating particles. The numerical results are validated against experiments using high-speed imaging.

무인항공기 노즐 형상 변화에 따른 IR 신호 영향성 연구

The effects of various nozzle configurations on infrared signature are investigated for the purpose of analysing the infrared signature level of aircraft propulsion system. A virtual subsonic aircraft is selected and then a circular convergent nozzle, which meets the mission requirements, is designed. Convergent nozzles of different configurations are designed with different geometric profiles. Using a compressible Navier-Stokes-Fourier CFD code, an analysis of thermal flow field and nozzle surface temperature distribution is conducted. From the information of plume flow field and nozzle surface temperature distribution, IR signature of plume and nozzle surface is calculated through the narrow-band model and the RadThermIR code. Finally, qualitative information for IR signature reduction is obtained through the analysis of the effects of various nozzle configurations on IR signature.

항공기 비행환경에 따른 플룸 IR 신호 영향성 연구

The plume infrared signature effects at various flight conditions of aircraft were investigated for the purpose of reducing infrared signature level. The nozzle of a virtual subsonic unmanned combat aerial vehicle was designed through a performance analysis. Nozzle and associated plume flowfields were first analyzed using a density-based CFD code and plume IR signature was then calculated on the basis of the narrow-band model. Finally, qualitative information for the plume infrared signature characteristics was obtained through the analysis of the IR signature effects at various flight conditions.

항공기 IR 신호 분석을 위한 다양한 비행 조건에서의 노즐 열유동장 해석

Aerothermodynamic flowfields of aircraft engine nozzles are computationally investigated at various flight conditions for infrared signature analysis. A mission profile of subsonic unmanned combat aerial vehicle is considered for the case study and associated engine and nozzles are selected through a performance analysis. Computational results of nozzle and plume flowfields using a density-based CFD code are analyzed in terms of thrust, maximum temperature, length and optical thickness of plume. It is shown that maximum temperature, length, and optical thickness of nozzle plume increase for lower altitude and higher Mach number.

CFD assessment of intake fraction in the indoor environment

In this paper we use a validated CFD model to calculate the inhalation exposure, expressed as an intake fraction (iF), of a seated person in an office with different contaminant sources, a floor diffuser, and a ceiling vent. These sources include the floor, the walls, a desk, and the human body. First, experimental data is used to determine the correct turbulent Schmidt number for the computational model to predict the transport of the species in an indoor environment. It was found at a turbulent Schmidt number of ∼0.9 produced the best fit when compared to experimental data. Then, the iF was calculated for two representations of the computer simulated person (CSP): a CSP with detailed surface geometry, and a simplified CSP with multi-block geometry. It was found that the simplified multi-block geometry is not adequate for predicting iF because it radically changes the flowfield of the thermal plume in the breathing zone (BZ). Next, the effect of personal ventilation systems on iF was investigated. The results show that such systems can reduce the iF by an order of magnitude compared with conventional mixing and displacement ventilation systems. Finally, a comparison of iF results were made for a surface body temperature of 32 °C and 28 °C. It was found that a 4 °C change in body surface temperature influenced the iF by less than 10%.

Compressing infrared spectrum of exhaust plume by wavelets

AbstractA study on multivariate calibration for the infrared spectrum of rocket exhaust plume was presented. As samples taken in the data set, the apparent infrared radiative properties of the high‐temperature plume flowfield consisted of variable concentrations gas components and were obtained by using a flux method combined with a narrow‐band model and Mie theory. The discrete wavelet transformation as a pre‐processing tool was carried out to decompose the infrared spectrum and compress the data set. The compressed data regression model was applied to simultaneous multi‐component concentrations for determination of the exhaust plume. The compression performance with several wavelet functions at different resolution scales was studied, and the prediction reliability of the compressed regression model was investigated. Numerical experiment results show that the wavelet transform performs an effective compression preprocessing technique in multivariate calibration and enhances the ability in characteristic extraction of the exhaust plume infrared spectrum. Using the compressed data regression model, the reconstructing results are almost identical when compared to the original spectrum, and the original size of the data set has been reduced to about 5% while the computational time needed decreases significantly. © 2009 Wiley Periodicals, Inc. Heat Trans Asian Res; Published online in Wiley InterScience (www.interscience.wiley.com). DOI 10.1002/htj.20280

RANS Analyses of Turbofan Nozzles With Internal Wedge Deflectors for Noise Reduction

Computational fluid dynamics (CFD) was used to evaluate the flow field and thrust performance of a promising concept for reducing the noise at take-off of dual-stream turbofan nozzles. The concept, offset stream technology, reduces the jet noise observed on the ground by diverting (offsetting) a portion of the fan flow below the core flow, thickening and lengthening this layer between the high-velocity core flow and the ground observers. In this study a wedge placed in the internal fan stream is used as the diverter. Wind, a Reynolds averaged Navier–Stokes (RANS) code, was used to analyze the flow field of the exhaust plume and to calculate nozzle performance. Results showed that the wedge diverts all of the fan flow to the lower side of the nozzle, and the turbulent kinetic energy on the observer side of the nozzle is reduced. This reduction in turbulent kinetic energy should correspond to a reduction in noise. However, because all of the fan flow is diverted, the upper portion of the core flow is exposed to the freestream, and the turbulent kinetic energy on the upper side of the nozzle is increased, creating an unintended noise source. The blockage due to the wedge reduces the fan mass flow proportional to its blockage, and the overall thrust is consequently reduced. The CFD predictions are in very good agreement with experimental flow field data, demonstrating that RANS CFD can accurately predict the velocity and turbulent kinetic energy fields. While this initial design of a large scale wedge nozzle did not meet noise reduction or thrust goals, this study identified areas for improvement and demonstrated that RANS CFD can be used to improve the concept.

Combined Analysis of Thruster Plume Behavior in Rarefied Region by Preconditioned Navier-Stokes and DSMC Methods

(Received August 11th, 2008) Satellite attitude is usually controlled by plume exhaust from thrusters into the vacuum of space. To study the plume effects in the highly rarefied region, the Direct Simulation Monte Carlo (DSMC) method is usually used, because the plume flow field contains the entire range of flow regime from the near-continuum near the nozzle exit through the transitional state to free molecular state at the far field region from the nozzle. The purpose of this study is to investigate the behavior of a small monopropellant thruster plume in the vacuum region numerically by using the DSMC method. To obtain more accurate results, the preconditioned Navier-Stokes algorithm is introduced to calculate continuum flow fields inside the thruster to predict nozzle exit properties, which are used for inlet conditions of DSMC method. As a result, the plume characteristics in the highly rarefied flow, such as strong nonequilibrium near nozzle exit, large back flow region, etc., are investigated.

Effect of a magnetic field in simulating the plume field of an anode layer Hall thruster

In this study, we present axisymmetric simulations of xenon plasma plume flow fields from a D55 anode layer Hall thruster. A hybrid particle-fluid method is used for the simulations. The magnetic field surrounding the Hall thruster exit is included in the calculation. The plasma properties obtained from a hydrodynamic model are used as boundary conditions for the simulations. The electron properties are calculated using the Boltzmann model and a detailed fluid model, collisions of heavy particle are modeled with the direct simulation Monte Carlo method, and ion transport in the electric field uses the particle-in-cell technique. The accuracy of the simulation is assessed through comparison with various measured data. It is found that a magnetic field significantly affects the profile of the plasma in the detailed model. The plasma has a potential of 80 V at 10 mm from the thruster exit in the case of zero magnetic field, which decreases to 60 V when the magnetic field is included. Results predicted by the detailed model with the magnetic field are found to be in better agreement with experimental data.

Study of Plume Characteristics of a Stationary Plasma Thruster

Electron density and temperature of the plume are measured by a double Langmuir probe in an experimental chamber. A numerical model based on both particle-in-cell scheme and direct simulation Monte Carlo hybrid method is developed to simulate the flow field of plume. The equation for plasma potential is solved by alternative direction implicit technique. The simulation is verified by comparing the computational results with the measured data. The study indicates that the electron temperature of flow field is about 2 eV and the electron density is about 2.5 × 1016 ∼ 5 × 1015 m−3 at the central line with a distance of 0.3 ∼ 1.0 m downstream of the thruster exit. The model can well predict the flow field parameters of the steady plume. The efforts of this paper are referable for further investigation.

Investigation of Soot Combustion in Underexpanded Jet Plume Flows

A new overlay technique to model and simulate the soot combustion in rocket plumes is proposed and examined. An efficient methodology for the accurate prediction of soot oxidation is discussed so that the sensitivity of soot combustion to the complex plume flow characteristics may be better understood. The accurate prediction of soot combustion in rocket plumes affects the plume radiation pattern and the location of the maximum radiation in underexpanded rocket plumes. A continuum approach for the soot overlay problem is presented for the example problem of a single-nozzle axisymmetric plume representative of the Atlas rocket. The overlay calculations are applied to a third-order-accurate converged Navier-Stokes result for a chemically reacting gas flow at the altitudes of 21, 30, and 40 km. Results are presented to demonstrate the sensitivity of soot oxidation to the initial spatial distribution of soot at the nozzle exit plane and the altitude dependence of soot combustion in the plume flowfield for various pressure ratios. It is shown that the majority of the soot combustion occurs in the plume shear layer due to the higher temperatures and penetration of oxygen into the flow. Also, there is more combustion at lower altitudes due to increased availability of oxygen for combustion, even though the shear layer at the lower altitudes is thin.

Plume Modeling and Application to Mars 2001 Odyssey Aerobraking

Covers advancements in spacecraft and tactical and strategic missile systems, including subsystem design and application, mission design and analysis, materials and structures, developments in space sciences, space processing and manufacturing, space operations, and applications of space technologies to other fields.

논문 : 유한속도 화학반응을 고려한 초음속 로켓의 플룸 유동장 해석

케로신/액체산소 추진기관을 갖는 초음속 로켓의 플룸 유동장을 9 화학종 14 반응 모델과 연계된 레이놀즈 평균 Navier-Stokes 방정식을 이용하여 해석하였다. 유한속도 화학반응이 플룸 유동장에 미치는 영향을 고찰하기 위하여 그 결과를 화학적 동결유동 해석 결과와 비교하였다. 계산은 상용 CFD 소프트웨어인 FLUENT 5를 이용하여 수행하였다. 반응 유동 해석 결과는 노즐 내부에서의 화학반응에 따른 연소가스의 온도 증가로 인해 전체적으로 동결유동에 비해 더 높은 온도장을 나타내었다. 플룸에서의 모든 화학반응은 전단류와 배럴 충격파 반사지점 후방의 고온 영역에 국한되어 일어났으며 본 해석의 경우 플룸내에서의 유한속도 화학반응이 유동에 미치는 영향은 미약한 것으로 나타났다. 그러나 본 연구에서 이루어진 유한속도 화학반응을 고려한 플룸 해석을 통하여 플룸에서의 주된 화학 반응 및 이들의 반응 메커니즘을 확인할 수 있었다. A supersonic rocket plum flowfield of kerosene/liquid-oxygen based propulsion system has been analysed using the Reynolds-averaged Navier-Stokes equations coupled with a 9-species 14-reaction finite-chemistry model. The result were compared with chemically frozen flow solution to investigate the effect of finite-rate chemistry on the plume flowfield. The computations were performed using a commercial CFD software, FLUENT 5. The finite-rate chemistry solution exhibited higher temperature caused by the reactions within the nozzle. All the chemical reactions within the plum were dominated only in the shear layer and behind the barrel shock reflection region where the temperatures are high and the effect of finite-rate chemical reactions on the flowfield was found to be insignificant. However, the present plume computation including the finite-rate chemical reaction within the plume has revealed major reactions occurring in the plum and their reaction mechanisms.

Modeling of Flow and Radiation in the Atlas Plume

Modeling results and data are presented for a chemically reacting flow from a thrusting Atlas II rocket at low altitudes. High spatial resolution imagery and spectra have been obtained for a kerosene/liquid oxygen multinozzle plume at altitudes in the continuum flow regime. A numerical solution for a three-dimensional plume flow from the Atlas rocket engine is obtained using two different Navier-Stokes computational fluid dynamics codes. The influence of the flow near the rocket body on the plume structure is considered at an altitude of 40 km. The plume flowfields at 15- and 40-km altitudes are used for radiation calculations in the infrared spectral region. Calculated spectral radiant intensities and pixelated images are compared with the data extracted from the recent in-flight measurements. Comparison of the modeling with the data shows that numerical modeling is able to predict the plume structure existing at both altitudes

TEMPERATURE MEASUREMENT OF SOLID ROCKET MOTOR EXHAUST PLUME BY ABSORPTION-EMISSION SPECTROSCOPY1*

A method, which measures the temperature of solid rocket motor exhaust plume, was developed by employing an improved sodium line reversal process. The formula for calculating the temperature was improved and simplified. The temporal temperature-time distributions of the exhaust plume of double base propellant rocket motors were given by the established method. The maximum time resolution and accuracy for the temperature measurement are 40 μs and 8%, respectively. This kind of temperature measurement does not disturb the combustion flow field of exhaust plume and is simpler and more sensitive. The maximum temperatures of SQ-2 propellant exhaust plumes are 2234 K and 2202 K for 240 mm and 500 mm apart from the motor nozzle, respectively. More combustion phenomena of the exhaust plume can be measured by using an existing instrument.

Numerical Analysis of the Rarefied Plume from an Arcjet Thruster: Effect of the Nozzle Exit Shape

The plume flow field of an arcjet thruster with hydrazine decomposed gas as the propellant has been numerically analyzed to reveal a backflow field, using Direct Simulation Monte Carlo (DSMC). Hydrazine decomposed gas is approximated by a mixture of gas composed of N2, H2, and H. Four different nozzle lip shapes have been prepared to investigate their effects on the backflow, and species existing in the backflow have also been examined. The present results of mass flux at various angles measured from the plume axis have been compared with experiments, and a qualitatively fairly good agreement has been obtained.

Aerosol dynamics in near‐field aircraft plumes

A numerical model including gas phase HOx, NOx, and SOx chemistry; H2SO4‐soot adsorption; binary H2SO4‐H2O nucleation; aerosol coagulation; and vapor condensation is used to investigate aerosol formation and growth in near‐field aircraft plumes. The plume flow field is treated using the JANNAF standard plume flow field code, SPF‐II. Model results are presented for a Mach 2.4 high‐speed civil transport at 18 km altitude and 85°N latitude and a subsonic Boeing 707 at 12.2 km, 47°N. The results, based on hydroxyl radical driven oxidation kinetics, indicate that 1–2% of the emitted SO2 is converted to H2SO4 in the near‐field exhaust (1–2 s) and that for typical exhaust SO2 emission indices (≈1 g kg‐fuel) the plume is supersaturated with respect to both the pure liquid acid and H2SO4/H2O solutions. Classical nucleation theory predicts high levels of small (0.3–0.6 nm radius) H2SO4/H2O embryos. Coagulation and gas‐to‐particle conversion are followed to provide estimates for the number density of activated soot particles capable of serving as condensation nuclei for contrail formation. Results are presented illustrating the dependence of water condensation on the number density and size distribution of activated exhaust soot nuclei.