Nozzle Throat Research Articles

An in-situ ablation recession sensor for carbon/carbon ablatives based on commercial ultra-miniature thermocouples

In this research, a breaking-wire like recession sensor based on commercial, ultra-fine thermocouples (250μm diameter) was designed and tested. The proposed approach uses low cost, commercially available raw materials and processing techniques. A series of thermocouples (TCs) are positioned at a well established depth from the surface. When the temperature rises above the melting point of the Seebeck junction, the TC would break working as a position marker. The proposed ablation recession sensor was tested on carbon/carbon composites (CCCs) under a severe hyperthermal environment. Two types of CCCs with different densities were considered. Two different layouts were tested: 4- and 8-levels configuration. The obtained data showed that the proposed approach provide an accurate estimation of the ablation rate. Due to the uniqueness of the proposed sensing technique, it can be applied both on TPSs of spacecraft as well as on nozzle throats of solid rocket motors.

Flow characteristics of hydrogen gas through a critical nozzle

Critical nozzles are widely used in the flow measurement and can be used for mass flow-rate measurement of hydrogen gas. The effect of real gas state equation on discharge coefficient of hydrogen gas flow through a critical nozzle was investigated. The real gas critical flow factor was introduced which considers the effect of the real gas on discharge coefficient. An analytic solution of real gas critical flow factor of hydrogen gas calculated from the modern equations of state based on Helmholtz energy, over a wider range of temperature 150–600 K and pressure up to 100 MPa was presented. An accurate empirical equation for real gas critical flow factor was determined by the nonlinear regression analysis. The equation was in good agreement with the high-pressure hydrogen gas experimental data by Morioka and CFD solutions by Nagao and Kim. Using this equation, the discharge coefficient can be directly and accurately calculated. It indicates that the discharge coefficient of hydrogen gas should be comprehensively taken into consideration with stagnation temperature, stagnation pressure and nozzle throat diameter. A lot of detailed results about the effect of real gas state equation were obtained.

Evaluation of Nozzle Damping Characteristics by a Pulsed Method

A pulsed method is numerically provided in this paper to evaluate the nozzle damping characteristics. In linear consideration of solid rocket motor, the pulsed oscillation pressure will be attenuated in an exponentially way by the nozzle damping. Based on this method, the damping constants of three models with nozzle throat radii of Rt=25mm, Rt=30mm and Rt=35mm are numerically studied. The results show that the first axial acoustic combustion instability is easily excited when a pulse is exerted on the chamber. The pulse intensity and mean chamber pressure have no effect on the nozzle damping characteristics. The factor that can greatly influence the damping constant is the nozzle throat radius. The bigger nozzle throat radius, the greater the nozzle damping effect. The calculated results are consistent well with the theoretical values.

Analytical calculation and evolutionary regression method for isentropic exponent of hydrogen gas at the throat of critical nozzle

As one of the most promising renewable energy resources, the hydrogen has been used in the fields such as aerospace, industry, and fuel cells. Critical nozzles are widely used for mass flow-rate measurement of high hydrogen gas, since the flow measurement process is not affected by its downstream flow disturbance. The flow rule of real hydrogen gas through a critical nozzle is complicated and the thermophysical property of hydrogen at the nozzle throat is vital to the accurate measurement of hydrogen flow. In this paper, based on explicit Helmholtz energy and entropy-enthalpy equations, the basic flow parameter and isentropic volume change exponent are analytically calculated. In addition, an accurate explicit equation is determined by the nonlinear regression analysis where the ways of selection, exchange and mutation derived from evolutionary algorithm are introduced to search for optimal population individual. The regression standard deviation is 0.0089%, mean residual deviation is 0.0285%, and maximum residual deviation is 0.1781%. The result shows that it not only can rapidly find the optimal solution which has the lowest number of equation items and the great overfitting suppression capability, but also has a high computation accuracy. This algorithm can also be applied to modeling flow characteristic parameters for every other flow device.

0608 キャビテーション噴流の加工能力におけるノズルスロート長さの影響

0608 キャビテーション噴流の加工能力におけるノズルスロート長さの影響

Improvement on Main/backup Controller Switching Device of the Nozzle Throat Area Control System for a Turbofan Aero Engine

Improvement on Main/backup Controller Switching Device of the Nozzle Throat Area Control System for a Turbofan Aero Engine

Discharge coefficient of small sonic nozzles

The purpose of this investigation is to understand flow characteristics in mini/micro sonic nozzles, in order to precisely measure and control miniscule flowrates. Experimental and numerical simulation methods have been used to study critical flow Venturi nozzles. The results show that the nozzle?s size and shape influence gas flow characteristics which leading the boundary layer thickness to change, and then impact on the discharge coefficient. With the diameter of sonic nozzle throat decreasing, the discharge coefficient reduces. The maximum discharge coefficient exits in the condition of the inlet surface radius being double the throat diameter. The longer the diffuser section, the smaller the discharge coefficient becomes. Diffuser angle affects the discharge coefficient slightly.

Spectroscopic Observation of He Arcjet Plasma Expanding Through a Converging and Diverging Slit Nozzle

An arcjet plasma generator with a converging and diverging slit nozzle was constructed. This plasma source allowed us to directly observe the arc plasma in the discharge section, which provided useful information about a transition from ionizing thermal plasma to recombining phase. Spatial distributions of the electron temperature and density in the rectangular shaped anode nozzle were evaluated by visible emission spectroscopy. The temperature and density for a discharge current of 20 A were determined to be ~0.18 eV and ~3.7×1013 cm-3, respectively, at the nozzle throat. The values on the jet axis were compared with those calculated by the gas dynamic theory on one dimensional slit nozzle.

Combustion of Frozen Nanoaluminum and Water Mixtures

Steady-state strand burner and laboratory-scale static fire motor experiments were used to determine the relative performance and viability of an environmentally friendly solid propellant composed of only nanoaluminum and frozen water. The nominal size of the nanoaluminum particles was 80 nm. The particles were homogeneously mixed with water to form pastes or colloids and then frozen. The measured parameters include burning rates, slag accumulation, thrust, and pressure. A system scaling study was performed to examine the effect of the size of the small-scale motors. The equivalence ratio was fixed at 0.71 for the strand burner and the laboratory-scale motor experiments. The effect of pressure on the linear burning rate was also examined. For an equivalence ratio of 0.71, the mixture exhibited a linear burning rate of at a pressure of 10.7 MPa and a pressure exponent of 0.79. Three motors of internal diameters in the range of 1.91–7.62 cm were studied. Grain configuration, nozzle throat diameter, and igniter strength were varied. The propellants were successfully ignited and combusted in each laboratory-scale motor, generating thrust levels above 992 N in the 7.62-cm-diam motor with a center-perforated grain configuration (7.62 cm length) and an expansion ratio of 10. For the 7.62 cm motor, combustion efficiency was 69%, whereas the specific impulse efficiency was 64%. Increased combustion efficiency and improved ease of ignition were observed at higher chamber pressures (greater than ).

The effects of two gas flow streams with initial temperature and pressure differences in cold spraying nozzle

In order to inject powder into the nozzle in cold spraying, the pressure of the powder carrier gas at the nozzle inlet must be equal to or higher than the pressure of the propulsion gas, and the temperature of the powder carrier gas needs to be lower than the temperature of the propulsion gas to prevent from buildup or clogging by particles in the nozzle. The gas flow behavior resulting from the lower temperature and the higher pressure of powder carrier gas has an important effect on the particle acceleration in the nozzle. Several issues related to the gas flow will be examined through numerical simulation, including the initial pressure differential, the nozzle throat diameter, and the prechamber length between the injection tube outlet and the nozzle inlet. For different initial pressure differentials, nozzle throat diameters and prechamber lengths, the mixing conditions of the two gas flow streams in the nozzle are quite different and the negative effect of the inadequate mixing of the two gas flow streams in the nozzle on the particle acceleration is obvious. It is found that using lower differential pressure, larger nozzle throat diameter and longer prechamber for a certain diameter of powder injection tube can enhance the particle acceleration in the nozzle.

A Validated Numerical-Experimental Design Methodology for a Movable Supersonic Ejector Compressor for Waste-Heat Recovery

The aim of this paper is to develop the technical knowledge, especially the optimum geometries, for the design and manufacturing of a supersonic gas–gas ejector for a waste-heat driven vehicle cooling system. Although several studies have been performed to investigate the effects of geometrical configurations of gas–gas ejectors, a progressive design methodology of an ejector compressor for application to a vehicle cooling system has not yet been described. First, an analytical model for calculation of the ejector optimum geometry for a wide range of operating conditions is developed, using R134a as the working fluid with a rated cooling capacity of 2.5 kW. The maximum values of entrainment ratio (ω) have been estimated by correlation of the main parameters in a nondimensional form. The optimum values of nozzle throat diameter (dnt) and mixing chamber diameter (dmc) thus obtained are used as a starting point for the computational fluid dynamics (CFD) optimization covering a wide range of geometrical configurations. To assess the effect of various dimensional quantities, an optimization technique has been proposed for calculation of the most efficient geometry of the target ejector for manufacturing. Using a vehicle cooling system as a test case, the final optimized dimensions are reported and discussed. An experimental validation confirms the CFD results and the ejector performance with a normalized deviation of 5% between observed and simulated results, demonstrating that the methodology is a valid ejector design tool for a wide range of applications.

Effects of Cavity on the Performance of Dual Throat Nozzle During the Thrust-Vectoring Starting Transient Process.

The dual throat nozzle (DTN) technique is capable to achieve higher thrust-vectoring efficiencies than other fluidic techniques, without compromising thrust efficiency significantly during vectoring operation. The excellent performance of the DTN is mainly due to the concaved cavity. In this paper, two DTNs of different scales have been investigated by unsteady numerical simulations to compare the parameter variations and study the effects of cavity during the vector starting process. The results remind us that during the vector starting process, dynamic loads may be generated, which is a potentially challenging problem for the aircraft trim and control.

Thrust and efficiency model for electron-driven magnetic nozzles

A performance model is presented for magnetic nozzle plasmas driven by electron thermal expansion to investigate how the thrust coefficient and beam divergence efficiency scale with the incoming plasma flow and magnetic field geometry. Using a transformation from cylindrical to magnetic coordinates, an approximate analytical solution is derived to the axisymmetric two-fluid equations for a collisionless plasma flow along an applied magnetic field. This solution yields an expression for the half-width at half-maximum of the plasma density profile in the far-downstream region, from which simple scaling relations for the thrust coefficient and beam divergence efficiency are derived. It is found that the beam divergence efficiency is most sensitive to the density profile of the flow into the nozzle throat, with the highest efficiencies occurring for plasmas concentrated along the nozzle axis. Increasing the expansion ratio of the magnetic field leads to efficiency improvements that are more pronounced for incoming plasmas that are not concentrated along the axis. This implies that the additional magnet required to increase the expansion ratio may be worth the added complexity for plasma sources that exhibit poor confinement.

Development of non-eroding rocket nozzle throat for ultra-high temperature environment

The powder processing methods including powder metallurgy (P/M) and powder injection molding (PIM) techniques for tungsten (W)–rhenium (Re) were employed to produce a W–Re rocket nozzle. The composition of W–Re was determined by 25wt.% of Re to avoid the formatting brittle sigma (σ) phase. The samples for analysis of the densification behavior on sintering were prepared by die pressing and cold isostatic pressing (CIP). The feedstock for the PIM process was produced by mixing the W–25wt.% Re powder and binder system based on a wax-polymer with an optimum solid loading through the twin-extruder mixer. The injection molded specimens were debound to extract and decompose the binders via the solvent and thermal debindings. The debound samples were sintered in a hydrogen atmosphere. After sintering, hot isostatic pressing (HIP) was carried out in an argon atmosphere to enhance the density.The dilatometry experiments were performed to analyze and predict a densification behavior during sintering. The master sintering curve (MSC) model was used to characterize the densification behavior with a minimal set of preliminary experiments. The mechanical properties were evaluated through microstructure and chemical composition measured by EDX–SEM and X-ray diffraction (XRD).Finally, the eroding test was conducted using the W–25wt.% Re rocket nozzle produced by PIM under the high temperature. After carrying out erosion tests, the erosion rate, hardness and microstructure were evaluated.

Study on the Obstruction Effect for a Large Explosion Shock Wave Tube Used the Ideal Nozzle Theory

This paper presents a calculation on the obstruction effects for the given large explosion shock wave tube using the ideal nozzle the theory. The relationship among Mach number, Mach number ratio, dynamic pressure ratio in the nozzle throat and blocking area ratio are established according to the fundamental equations of one-dimensional steady flow, which can be taken as the reference of blocking limit design.

Design and experiments of plasma jet igniter for aeroengine

A plasma jet ignition technology was studied for aeroengine combustor. The advantages of compact structure and advanced performance of air-cooled plasma jet igniter had been tested and verified in the opening test. The plasma jet igniter could produce a continuous plasma jet, stable and reliable ignition. The influence factors of plasma jet ignition aerodynamic and structure were studied in the opening test. Continuous plasma jet was closely related to inlet pressure and flow, simultaneously to the igniter nozzle geometry and throat size. Based on the stable continuous plasma jet, some methods were explored in order to reduce plasma output power, optimize the structure design, and improve the thermal protective. The plasma jet igniter applied to aeroengine combustor was identified initially. For combustion chamber with the igniter, altitude ignition performance were experimented for the inlet pressure of plasma ignition from 10kPa to 50kPa, the flow of plasma jet not more than 0.20g/s, and energy output of ignition from 800W to 1500W. The test results were compared with that of conventional aeroengine high energy ignition system. The results show that the plasma jet igniter is better than the conventional one.

Efficient Thrust Distribution with Adaptive Pressure Control for Multinozzle Solid Propulsion System

An integrated approach for chamber pressure control and thrust distribution for multiple nozzles of a solid propulsion system is presented. In this study, a solid propulsion system with variable nozzle throat area controlled by a pintle actuator is considered. To keep the hardware configuration simple and lightweight, multiple nozzles share a combustion chamber. Consequently, changing the throat area of one nozzle results in an unintended change of chamber pressure and thrust from other nozzles. Therefore, pressure stabilization and thrust distribution for multiple nozzles should be conducted simultaneously to achieve the desired performance. For the inner control loop, an adaptive sliding mode controller with feedback linearization is used for pressure stabilization. With the pressure control loop working properly, a thrust distribution algorithm into 10 nozzles for translation and attitude maneuver is developed as an outer control loop.

Experimental investigation of the adjustable ejector in a multi-evaporator refrigeration system

In this paper, the experimental investigation was carried out for the performance of adjustable ejector used in a multi-evaporator refrigeration system. The adjustable ejector with a spindle to adjust the area of nozzle throat was applied to deal with the considerable variation in primary cooling load for air-conditioning in such system. The adaptability of the adjustable ejector for the system was first evaluated by the tests and the results show the adjustable ejector can efficiently deal the problem of variable primary cooling load in the system. The tests for the performance of pressure recovery were subsequently carried out. The experimental results indicate that the Pressure Recovery Ratio (PRR) and Relative Pressure Recovery Ratio (RPRR) can reach 20% and 8%, respectively, in the system with design operating condition, which means the adjustable ejector applied in multi-evaporator refrigeration system may be a promising approach for a practical multi-evaporator refrigeration system with significant energy saving. It also can be found that PRR and RPRR are reduced with decreasing of the primary cooling load.

NUMERICAL ANALYSIS OF IDEALLY-EXPANDED SUPERSONIC JETS WITH NONEQUILIBRIUM HOMOGENEOUS CONDENSATION

Steam or moist air is used as working gas in a wide range of engineering applications of supersonic jets. In these cases, nonequilibrium homogeneous condensation may occur at the downstream of nozzle throat. The surrounding gas will be heated by the release of latent heat of condensation, and may results a change in the flowfield. The present report will describe numerical investigations predicting the effect of nonequilibrium condensation on the flow characteristics of ideally-expanded supersonic free jets. A TVD numerical method is applied to solve RANS and droplet growth equations. The predicted results are compared with the experimental data.

로켓 모터의 작동시간이 노즐 열전달 계수에 미치는 영향에 관한 연구

미사일 탄두의 정확한 제어를 위하여, 로켓 모터의 구조 및 유체/열역학적 안정성이 절대적으로 필요하다. 특히, 유도탄의 초기 선회형 로켓 모터는 작동 시간이 짧음에도 불구하고, 고온 고압에서 작동되기 때문에 노즐 목 부근에서는 삭마(削磨, Ablation)가 일어나 유동자체의 불안정이 발생하고, 열 및 기계적 응력 때문에 시스템 자체가 파국에 이르는 경우가 종종 발생한다. 이와 관련, 본 연구에서는 열응력(Thermal stress) 및 삭마는 유동 연소 가스로부터 노즐 재료로의 전열량과 재료 내부의 온도 차이에 기인된다는 판단에 따라 로켓 모터의 작동시간이 노즐 벽면 열전달 계수에 미치는 영향을 수치적으로 해석하는 것을 연구의 목적으로 하였다. 그 결과, 노즐 벽면 열전달 계수는 노즐 목 바로 직전에서 가장 크게 되고, 로켓 모터의 작동 시간이 길어질수록 열전달 계수는 감소하는 것으로 나타났다. 또, 노즐 목의 곡률반경이 작을수록, 최대 열전달 계수가 크게 되는 것으로 나타났다. To guarantee the exact control of missile warhead, it is inevitable to ensure the stabilities in the view points of structural and fluid/thermo dynamics of the rocket motor. Specially, despite of shortness in operating time of the rocket motor which is initial turning type of missile, it occurs frequently some problems of ablation at the neighborhood of the nozzle throat, with the result that the system itself gets to failure. In these connections, in the present study, the effect of the operating time of a rocket motor on the coefficient of convective heat transfer at the nozzle wall is investigated by numerical analysis. As a result, it is turned out that the heat transfer coefficient is largest at the just ahead of nozzle throat and decreases with the increase of operating time of the rocket motor. Furthermore, we found that the radius of curvature of throat becomes smaller, the maximum coefficient of convective heat transfer becomes larger.