A theory and experiments for detecting shock locations (original) (raw)

Three dimensional investigation of the shock train structure in a convergent–divergent nozzle

Three-dimensional computational fluid dynamics analyses have been employed to study the compressible and turbulent flow of the shock train in a convergent–divergent nozzle. The primary goal is to determine the behavior, location, and number of shocks. In this context, full multi-grid initialization, Reynolds stress turbulence model (RSM), and the grid adaption techniques in the Fluent software are utilized under the 3D investigation. The results showed that RSM solution matches with the experimental data suitably. The effects of applying heat generation sources and changing inlet flow total temperature have been investigated. Our simulations showed that changes in the heat generation rate and total temperature of the intake flow influence on the starting point of shock, shock strength, minimum pressure, as well as the maximum flow Mach number.

Computational Analysis of Shockwave in Convergent Divergent Nozzle

The objective of this paper is to computationally analyse shock waves in the Convergent Divergent (CD) Nozzle. The commercial CFD code Fluent is employed to analyse the compressible flow through the nozzle. The analysis includes static pressure, temperature, Mach number and density of the flow for different nozzle pressure ratios (NPR i.e., the ratio between exit pressure of the nozzle to ambient pressure). The results are compared with the analytical results of quasi-one dimensional equation. The flow characteristic before and after the shock is discussed.

Shock waves formation from a converging-diverging nozzle

Gas Dynamics

The project is about studying the whole rocket. We will work on comparing between analytical and theory results on the converging diverging (C-D) nozzle, also called de Laval nozzle for an isentropic flow, a normal shock wave in the diverging section, an oblique shock wave and an expansion at the exit. Then we will simulate the flow at the upper surface of the rocket by CFD, considering the upper area as diamond shaped. The aim of the simulation is to show the oblique shock wave created by supersonic speed and doesn’t intersect with the rest of the rocket, reducing overall the drag.

Shock structure in separated nozzle flows

Shock Waves, 2009

In the case of high overexpansion, the exhaust jet of the supersonic nozzle of rocket engines separates from nozzle wall because of the large adverse pressure gradient. Correspondingly, to match the pressure of the separated flow region, an oblique shock is generated which evolves through the supersonic jet starting approximately at the separation point. This shock reflects on the nozzle axis with a Mach reflection. Thus, a peculiar Mach reflection takes place whose features depend on the upstream flow conditions, which are usually not uniform. The expected features of Mach reflection may become much difficult to predict, depending on the nozzle shape and the position of the separation point along the divergent section of the nozzle.

IJERT-Experimental and Computational Analysis of Shock Structure in a Supersonic Contoured Nozzle

International Journal of Engineering Research and Technology (IJERT), 2020

https://www.ijert.org/experimental-and-computational-analysis-of-shock-structure-in-a-supersonic-contoured-nozzle https://www.ijert.org/research/experimental-and-computational-analysis-of-shock-structure-in-a-supersonic-contoured-nozzle-IJERTV9IS090045.pdf Flow separation in a supersonic contoured nozzle whilst operating under over expanded regime is an inevitable fluid dynamic phenomenon. The flow field comprise formation of complex shock structures and its interaction with the viscous boundary layer. Profound number of researches on various types of contoured nozzle profiles have been undertaken to better understand the said phenomenon. The present study is focused on further understanding the fundamentals associated with formation of shocks and its structural transformation under varying NPR for a defined area ratio. Flow visualization utilizing Schlieren photography has enabled to capture the details of shock and its structure along with other phenomenon viz boundary layer separation, shock boundary layer interaction, aftershocks etc. The locations of shock captured experimentally have also been ascertained with computational generated data for various NPR and the results have been found quite comparable.

Experimental and Computational Analysis of Shock Structure in a Supersonic Contoured Nozzle

International Journal of Engineering Research and, 2020

Flow separation in a supersonic contoured nozzle whilst operating under over expanded regime is an inevitable fluid dynamic phenomenon. The flow field comprise formation of complex shock structures and its interaction with the viscous boundary layer. Profound number of researches on various types of contoured nozzle profiles have been undertaken to better understand the said phenomenon. The present study is focused on further understanding the fundamentals associated with formation of shocks and its structural transformation under varying NPR for a defined area ratio. Flow visualization utilizing Schlieren photography has enabled to capture the details of shock and its structure along with other phenomenon viz boundary layer separation, shock boundary layer interaction, aftershocks etc. The locations of shock captured experimentally have also been ascertained with computational generated data for various NPR and the results have been found quite comparable. Keywords—Nozzle, superson...

Numerical investigation of transient nozzle flow

Shock Waves, 2003

Abstract. The starting process of two-dimensional and axisymmetric nozzle flows has been investigated numerically. Special attention has been paid to the early phase of the starting process and to the appearance of a strong secondary shock wave. For both cases, shock intensities and velocities are obtained and discussed. The flow evolution in the axisymmetric case is proved to be more complex and the transient starting process is slower than in the plane case. Finally, the effects of changing the nozzle angle and the incident shock wave Mach number on the transient flow are addressed. It is shown that a faster start-up can be induced either by decreasing the nozzle angle or increasing the Mach number of the incident shock wave.

Study on flow characteristics and performance of baffled shock two dimensional vector nozzle

2023

The development of aerospace technology in the future depends on supersonic and hypersonic vehicles. As the national defense and military strength become more and more crucial for national growth, communication, security and stability, the high-efficiency thrust vectoring nozzle, which can significantly enhance the maneuverability and stability of supersonic and hypersonic vehicles, has always been the focus of research for military powers. This thesis paper presents a numerical study on the flow characteristics and performance of baffled shock two dimensional vector nozzle. The baffled shock vector nozzle is a type of fluidic thrust vectoring nozzle that uses secondary injection to deflect the primary flow and generate a vector angle. Fluidic thrust vectoring technology is regarded as a key technology for the development of very low detectable vehicles because of its advantages such as fast response, light weight and good stealth performance. The main objectives of this study are to investigate the effects of various parameters such as slot interval distance, slot width, injection angle, nozzle pressure ratio, secondary flow pressure ratio and outflow Mach number on the deflection angle, thrust coefficient, thrust efficiency and secondary flow ratio of the nozzle. The numerical simulations are carried out using k−epsilon turbulence model, which is validated by comparing with experimental data. The results show that the nozzle performance can be improved by optimizing the slot interval distance and width, increasing the injection angle, adjusting the nozzle pressure ratio and secondary flow pressure ratio, and controlling the outflow Mach number. The results also reveal the complex flow phenomena inside the nozzle, such as shock wave interactions, flow separation and reattachment, and boundary layer effects. The study provides a comprehensive understanding of the flow characteristics and performance of baffled shock two dimensional vector nozzle and offers some guidance for its design and optimization.

Shock detection from computational fluid dynamics results

14th Computational Fluid Dynamics Conference, 1999

In complex ow regimes, it may be di cult for an analyst to nd the location of shock discontinuities within a Computational Fluid Dynamics (CFD) solution. They do not correspond to locations where the mach number is unity, and the high gradients associated with the discontinuity can be di cult to detect because of numerical smoothing performed in order to obtain the solution. An algorithm is introduced that uses the ow physics to locate shocks in transient and steady state solutions. The test was validated with simple one and two dimensional models, then extended to more realistic three dimensional ows. A set of ltering algorithms was developed to remove any false shock indications. Results indicate that both the stationary and transient shock nding algorithms accurately locate shocks, but need ltering to compensate for a lack of sharpness in CFD discontinuities.

Large eddy simulation of shock train in a convergent–divergent nozzle

This paper discusses the suitability of the Large Eddy Simulation (LES) turbulence modeling for the accurate simulation of the shock train phenomena in a convergent-divergent nozzle. To this aim, we selected an experimentally tested geometry and performed LES simulation for the same geometry. The structure and pressure recovery inside the shock train in the nozzle captured by LES model are compared with the experimental data, analytical expressions and numerical solutions obtained using various alternative turbulence models, includingk–"RNG, k–!SST, and Reynolds stress model (RSM). Comparing with the experimental data, we observed that the LES solution not only predicts the \locations of the ¯rst shock" precisely, but also its results are quite accurate before and after the shock train. After validating the LES solution, we investigate the effects of the inlet total pressure on the shock train starting point and length. The effects of changes in the back pressure, nozzle inlet angle (NIA) and wall temperature on the behavior of the shock train are investigated by details.