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Viscous effects on an ideal gas flow in a supersonic convergent–divergent nozzle are a well-studied subject in classical gas dynamics. However, the ideal assumption fails on fluids that exhibit complex behaviours such as near-critical-region and non-ideal dense vapours. Under those conditions, a realistic equation of state (EOS) plays a vital role for a precise and realistic computation. This work examines the problem for solving the quasi-one-dimensional viscous compressible flow using a realistic EOS. The governing equations are discretised and solved using the fourth-order Runge–Kutta method coupled with a state-of-the-art EOS to calculate the thermodynamic properties. The role of the Grüneisen parameter along with viscous and real gas effects and their influence on the sonic point formation are discussed. The study shows that the flow may not achieve the supersonic regime for any pressure ratio depending on the combination of that parameter and the normalised friction factor. In addition, the analysis yields the discharge coefficient and the isentropic nozzle efficiency, which may achieve maximum values as a function of the stagnation conditions. Finally, the study also evaluates the formation and intensity of normal shock waves by using the Rankine–Hugoniot relations, which now depend on the real gas and viscous effects in opposition to the inviscid solution. Moreover, the methodology used captures the sonic point and shock wave position by a space marching algorithm using the Brent method for scalar minimisation. Experimental data available in the open literature corroborate the approach.

Hybrid rocket motors have great development potential due to their outstanding thrust adjustment flexibility and long-term operation ability. However, nozzle erosion during the motor operation can cause an increase in the throat area of the nozzle, which leads to a decrease in combustion chamber pressure and nozzle efficiency. Therefore, a performance prediction model for hybrid rocket motors considering nozzle erosion has become a key technology that must be addressed when developing hybrid rocket motors. This study uses dynamic grid technology to simulate the regression of the combustion surface and nozzle erosion, which fits well with experimental values. The behavior of high-energy particles in the combustion chamber is simulated through a discrete phase model. Notably, distinctive behavior is observed in Al and Mg droplets, with Al droplets exhibiting incomplete vaporization in the combustion chamber while Mg droplets completely vaporize. A ground firing test using the Φ336 mm hybrid rocket motor lasting 200 s is conducted. The results show that the dynamic numerical simulation, accounting for nozzle erosion, substantially enhances performance prediction accuracy. The average deviation in motor thrust remains below 1.8%, and the combustion chamber pressure deviation stays under 2.6%, confirming the precision of the model. Ultimately, both simulation and experimental outcomes indicate a gradual decrease in specific impulse and characteristic velocity over the long-term operation, attributed to the gradual deviation of the oxygen-fuel ratio. This research provides valuable insights for guiding hybrid rocket motor design and optimizing design parameters to improve overall performance. This model can achieve long-duration and high-precision performance predictions for hybrid rocket motors.

High pitch-chord ratio turbine stages are a novel class of high enthalpy drop impulse turbines. Significant limitation of traditional design approach for such turbines is a low partial admission rate due to low volumetric flow rates. The main idea of high pitch-chord ratio design is in increase of the stage load coefficient with decrease of the flow angles. The latter allows to provide higher partial admission rate and, hence, higher stage efficiency compared with the traditional design approach. As a consequence, efficiency of the high pitch-chord ratio stages varies between 65% and 75%, whereas partially admitted stages efficiency tend not to exceed 60%. High pitch-chord ratio nozzles have a rectangular cross section. This feature leads to a vortex pair occurrence in a nozzle flow path. An interaction of these vortices for transand supersonic stages leads to additional losses and comparably low nozzle efficiency. The present paper aimed at first, establishing the design criteria which are essential for the nozzle efficiency, and, second, finding the optimal ratios of the proposed criteria. CFD simulation approach is used for the study. The gained deliverables are helpful for increasing of the nozzle velocity ratio.

In last two decades, the upheaval in refrigeration sector has led to replacement of conventional refrigerants with non-ozone depleting and low global warming potential refrigerants. Most of alternate refrigerants which have low global warming potential, have usually low coefficient of performance (COP). To overcome the problem of low COP, a novel Vortex Tube integrated Vapour Compression cycle (NVTCR) has been proposed. The paper emphasizes on improvement in thermodynamic performance of a simple Vapour Compression cycle (VCR) by utilizing a vortex tube in the scheme of VCR. The proposed cycle operates at three pressure levels viz. lowest pressure (in evaporator), intermediate pressure (in vortex tube) and highest pressure (in condenser). A computer program has been formulated in EES software for the analysis. The results have been computed for NVTCR and compared with VCR and Vortex Tube ycle based on Keller model (VCRK) for R134a. The results show that the COP of NVTCR (with liquid vapour heat exchanger (LVHE)) is higher than the COP for both VCR and VCRK. The COP (correponding to Te = -25[degress]C and 10[degress]C respectively) of VCR, VCRK, NVTCR and NVTCR (with LVHE) varies between 1.59 to 4.16, 1.7 to 4.24, 1.65 to 4.20 and 1.74 to 4.28 respectively. The effect of various parameters like evaporator temperature, condenser temperature, vortex tube nozzle efficieny ([[eta].sub.v]), cold mass fraction ([micro]), bifurcation of cold mass fraction stream (z) for subcooling in heat exchangers (installed after condenser and separator) and effectiveness of liquid vapour heat exchanger ([[epsilon].sub.lvhe]) on COP of NVTCR have been examined. The effect of [[epsilon].sub.lvhe] on COP of NVTCR is maximum among the parameters [[epsilon].sub.lvhe], [[eta].sub.v], [micro] and z considered for analysis. It is observed that percentage increase in COP of the NVTCR (with LVHE) with respect to COP of VCR varies by 4.46% and 10.05% corresponding to [[epsilon].sub.lvhe] equal to 0.1 and 1 respectively.

In this study, a mathematical analysis by considering the effect of an actual index of expansion clearly shows a persistent of the existence of subsonic flow after the throat to a down stream in the region of divergent part to produce a supersonic velocity at the exit of a convergent-divergent nozzle. The length of the divergent part where subsonic velocity found is dependent upon the magnitude of the nozzle efficiency and the actual index of expansion. The change in velocity from subsonic, sonic and supersonic occurs only in the divergent part while the corresponding frictionless behavior has the classical features (subsonic in the convergent, sonic at the throat, and supersonic in the divergent). This is mainly due to thermodynamic processes which result a change in enthalpy due to friction and a gain in entropy. The reference conditions are newly derived for an actual frictional flow condition. This design aspect differs in a physical manner corresponding to that from an isentropic flow. An actual nozzle shape for convergent-divergent nozzles is also investigated under a non-isentropic flow condition.

A rocket engine for space propulsion usually has a nozzle of a large exit area ratio. The nozzle efficiency is greatly affected by the nozzle contour. This paper analysed the effect of the constant capacity ratio in Rao’s method through the design process of an apogee engine. The calculation results show that increasing the heat capacity ratio can produce an expansion contour of smaller expansion angle and exit area ratio. A simple modification of Rao’s method based on thermally perfect gas assumption was made and verified to be more effective. The expansion contour designed by this method has much thinner expansion section and higher performance. For the space engine, a new extension contour type for the end section of the nozzle is proposed. The extension curve bent outward with increasing expansion angle increases the vacuum specific impulse obviously.

In this article, the J85-GE-21 turbojet engine for an altitude of 1000–8000 m, with the speed of 200 m/s and at 10, 20, and 40 °C, was provided, and then, based on the objective functions, the above system was optimized using particle swarm optimization method. For the purpose of optimization, the Mach number, compressor efficiency, turbine efficiency, nozzle efficiency, and compressor pressure ratio were assumed to be in the range of 0.6–1.4, 0.8–0.95, 0.8–0.95, 0.8–0.95, and 7–10, respectively. The highest exergy efficiency of 73.1% for different components of the engine at sea level and speed of 200 m/s belonged to the diffuser. Second and third to it were nozzle and combustion chamber with 68.6 and 51.5%, respectively. The lowest exergy efficiency of 4% belonged to the compressor, and the second to it was the afterburner with 11.6%. Also, the values of entropy production and efficiency of the second law of thermodynamics were 1176.99 and 479 K/W, respectively, prior to optimization, which were respectively changed to 1129 and 51.4 K/W postoptimization. Obviously, the entropy production is reduced, while the efficiency of the second law of thermodynamics is increased.

The iron oxide scales exfoliated from the inner wall of a boiler tube and a main steam pipe is known to cause solid-particle erosion on the control-stage nozzle. A combined experimental and numerical investigation was conducted to explore the optimization method of end-wall contouring for reducing the nozzle's erosion damage most effectively. The results indicate that increasing the end-wall contraction ratio and (or) decreasing the distance between the starting point of end-wall contouring and the trailing edge can significantly reduce the erosion-induced weight-loss of the nozzle, and can slightly improve the nozzle efficiency, irrespective of the variation in the particles size distribution and the aerodynamic parameters of a steam turbine. A main reason of erosion reduction is that the movement of loading towards the rear of the nozzle cascade caused by these contoured end walls has reduced the incident velocity of particles. In this study, the weight-loss of the nozzle was reduced by 40—50 per cent, and the nozzle efficiency was improved by 0.4—0.5 per cent by improving the end-wall contouring of the nozzle according to the methods mentioned above.

The paper presents data on the impact of new environmental requirements relating to the quality of diesel fuel on the anti-wear properties of fuel. Anti-wear additive is proposed as a material for increasing the tribotechnical characteristics of diesel fuel. This additive consists of diesel fuel with micelles contained in it, formed on the basis of molecules of solid plasticity lubrication of iron oxide (Fe3O4) - magnetite, and with surrounding molecules of oleic acid (C18H34O2). The additive has low shear resistance and increased lubricity of diesel fuel when this additive is introduced into it.

This paper presents the study which was intended to identify the optimum geometries for different operating conditions of a Coanda bases Fluidic Thrust vectoring using ACHEON nozzle (Aerial Coanda High Efficiency Orienting Nozzle). This computational study was done utilising the software ANSYS. The study is aimed at optimising the Coanda surface by identifying the most appropriate geometry of the Coanda surface for different operating conditions. The radius of curvature of the coanda surface has been modified for different studies and also different levels of truncation of the coanda surface has also been done. In this study a variation in the radius is from 52.73 mm to 62.67mm.