Combustion chamber analysis code

A good liquid propellant with high specific impulse and a high speed of exhaust gases implies a high combustion pressure and temperature and small molecular weight. However there is another important factor i.e. density of the propellant leads to larger storage tanks, eventually increasing the weight of the launch vehicle. Also the storage of liquid propellants at cryogenic temperature as in case of military operations where there is no definite time to use these weapons (rockets, missiles etc) until there is war, is a tremendous challenge. Despite all these drawbacks, high efficiency of LH 2 /LOx makes these difficulties worth coping with, when reaction time and storability are not too critical. In the present work, the simulation of a combustion chamber is carried out for LH 2 /LOx fuel/oxidizer injected at cryogenic temperature in the combustion chamber and combustion takes place at high pressure (70 atm) and high temperature (5000-8000 K) and finally gases expended in a C-D nozzle resulting high velocity and high thrust. The geometry and meshing has been done on GAMBIT and commercial code Ansys FLUENT has been used for the CFD simulation. The present study has been carried out for geometries with varying number of inlets for better mixing and for varying nozzle designs for attaining the high exhaust velocity and impulse which is the main objective of designing of a launch vehicle.

IOP Conference Series: Materials Science and Engineering

The design of liquid propellant rocket engine (LPRE) is a very complicated process; this is due to two main concerns: First, the requirements to satisfy the issues of performance, stability and compatibility. Second, the complicated, interacting processes inside thrust chamber. In this paper, an attempt to illustrate the importance of different parameters affecting performance, stability and compatibility is performed, followed by extensive study of processes inside thrust chamber. The result of processes study is developing the concept of "rate limiting process" which means that the process that can be considered the most important hence the design can be done mainly by considering it alone. This is done by developing a 1D vaporizationcontrolled model with its application to two case studies to illustrate model validation and application. It was found that the 1D model is valid as long as the vaporization process is the slowest process in this case the error in computing chamber cylindrical length is ~15%. However, if the mixing process is slow, or the reaction process in gas phase is slow as in the second case study of RFNA/Tonka250 case, the error grow and may reaches 50%

The design of liquid propellant rocket engine (LPRE) is a very complicated process; this is due to two main concerns: First, the requirements to satisfy the issues of performance, stability and compatibility. Second, the complicated, interacting processes inside thrust chamber. In this paper, an attempt to illustrate the importance of different parameters affecting performance, stability and compatibility is performed, followed by extensive study of processes inside thrust chamber. The result of processes study is developing the concept of "rate limiting process" which means that the process that can be considered the most important hence the design can be done mainly by considering it alone. This is done by developing a 1D vaporization-controlled model with its application to two case studies to illustrate model validation and application. It was found that the 1D model is valid as long as the vaporization process is the slowest process in this case the error in computing chamber cylindrical length is ~15%. However, if the mixing process is slow, or the reaction process in gas phase is slow as in the second case study of RFNA/Tonka250 case, the error grow and may reaches 50% 1. Introduction The design of LPRE thrust chamber needs to consider a large number of different phenomena that occur during the combustion process As a result there are a large number of similarity parameters-e.g. Reynolds, Prandtl, Schmidt numbers in addition to first and third Damköhler numbers-that should be used in designing and scaling of rocket combustors. Hence, scaling of LPRE with complete similarity is found to be practically impossible. A classical approach to design a liquid propellant rocket engine (LPRE) combustion chamber is Characteristic length approach. This characteristic length is defined as ratio between volume of combustion chamber *starting at injection head till the nozzle critical section) and nozzle critical area. The design procedure moves by selecting the propellant composition then choose a value of characteristic length from the experimental database (table 1). After calculating the required volume, use geometric relation to find dimensions of combustion chamber for different shapes (cylindrical, spherical, ellipsoidal, etc.). However, this approach does not show how the design parameters e.g. injector design, propellant temperature, etc. affect the performance of the combustion chamber. This makes rocket engineers to turn to the concept of rate limiting phenomena, which defined as the phenomena that control the chemical reaction or heat release process and discarding all other processes for engineering purpose.

2018

Model high-speed combustor on gaseous hydrocarbon fuel, prepared for experiments in T-131 wind tunnel of TsAGI, is presented. Experiments are projected to create an experimental database for validation of calculations and physical models of turbulence and combustion. Geometry of combustor and prepared measurements are described. The main subject of paper is preliminary calculations of this combustor. Numerical methods for 2D and 3D URANS calculations are described. Special attention is given to numerical techniques allowing fast calculations of 3D unsteady flow development in the combustor. Approach to parallel realization of Fractional Time Stepping (FTS) technology is described. One way of In Situ Adaptive Tabulation (ISAT) of kinetic equations solution during the calculation is presented. Possible gasdynamic structure of flow in the combustor with the flame stabilization both in subsonic and in supersonic regime is described. Asymmetrical stationary solution (for the symmetrically-expanding duct) and symmetrical solution with flame oscillations are found and analyzed. PREPARATION OF THE EXPERIMENTAL MODEL Today it is impossible to imagine the creation of perspective aircraft without supplementation of the experiments with numerical simulation. However, modern possibilities to calculate practical flows with combustion in aircraft engines are essentially limited by the huge computer cost for calculation of 3D viscid turbulent flows with finite-rate reactions [1,2] and by the low accuracy of the available models of turbulence, of chemical kinetics, and of turbulence/combustion interaction [3-5]. In 2017, the scientific laboratory "Studies and development of physical models and numerical technologies for description of different combustion regimes in aircraft engines" has been created in Propulsion department of TsAGI under the support of Russian Ministry of education and science. Goals of the laboratory are the development and validation of physically-grounded models for various combustion regimes in air-breathing engines, as well as the creation of special software for use in the cycle of aerodynamic design for new aircraft engines. The laboratory develops and improves physical and mathematical models of turbulent combustion, oriented to calculations in the framework of RANS (Reynolds-Averaged Navier-Stokes) and LES (Large Eddy Simulation). These models are implemented into computer codes, specially adjusted for concrete class of flows to get the best prediction of flow characteristics. Such adjustment is based on experimental data for flows of the considered class. To create the basis for such activities, the new "fire" aerodynamic experiments are prepared in TsAGI. Experiments will be performed on the unique high-speed wind tunnel T-131 (see detailed information about wind

Journal of Thermal …, 2001

This work presents a numerical simulation of the flow field in a liquid propellant rocket engine chamber and exit nozzle using techniques to allow the results to be taken as starting points for designing those propulsive systems. This was done using a Finite Volume method simulating the different flow regimes which usually take place in those systems. As the flow field has regions ranging from the low subsonic to the supersonic regimes, the numerical code used, initially developed for compressible flows only, was modified to work proficiently in the whole velocity range. It is well known that codes have been developed in CFD, for either compressible or incompressible flows, the joint treatment of both together being complex even today, given the small number of references available in this area. Here an existing code for compressible flow was used and primitive variables, the pressure, the Cartesian components of the velocity and the temperature instead of the conserved variables were introduced in the Euler and Navier-Stokes equations. This was done to permit the treatment at any Mach number. Unstructured meshes with adaptive refinements were employed here. The convective terms were treated with upwind first and second order methods. The numerical stability was kept with artificial dissipation and in the spatial coverage one used a five stage Runge-Kutta scheme for the Fluid Mechanics and the VODE (Value of Ordinary Differential Equations) scheme along with the Chemkin II in the chemical reacting solution. During the development of this code simulating the flow in a rocket engine, comparison tests were made with several different types of internal and external flows, at different velocities, seeking to establish the confidence level of the techniques being used. These comparisons were done with existing theoretical results and with other codes already validated and well accepted by the CFD community.

Energy Procedia, 2014

It is widely recognized that air-fuel mixing, combustion and pollutant formation inside internal combustion engines are strongly influenced by the spatial and temporal evolution of both marco-and micro-turbulent scales. Particularly, in spark ignited engines, the generation of a proper level of turbulence intensity for the correct development of the flame front is traditionally based on the onset, during the intake and compression strokes, of a tumbling macro-structure. Recently, in order to both reduce pumping losses due to throttling and develop advanced and flexible engine control strategies, fully variable valve actuation systems have been introduced, capable of simultaneously governing both valve phasing and lift. Despite the relevant advantages in terms of intake system efficiency, this technology introduces uncertainties on the capability of the intake port/valve assembly to generate, at low loads, sufficiently coherent and stable structures, able therefore to promote adequate turbulence levels towards the end of the compression, with relevant effects on the flame front development. It is a common knowledge that 3D-CFD codes are able to describe the evolution of the in-cylinder flow field and of the subsequent combustion process with good accuracy; however, they require too high computational time to analyze the engine performance for the whole operating domain. On the contrary, this task is easily accomplished by 1D codes, where, however, the combustion process is usually derived from experimental measurements of the in-cylinder pressure trace (Wiebe correlation). This approach is poorly predictive for the simulation of operating conditions relevantly different from the experimental ones. To overcome the above described issues, enhanced physical models for the description of in-cylinder turbulence evolution and combustion to be included in a 1D modeling environment are mandatory. In the present paper (part I), a 0D (i.e. homogeneous and isotropic) phenomenological (i.e. sensitive to the variation of operative parameters such as valve phasing, valve lift, intake and exhaust pressure levels, etc.) turbulence model belonging to the K-k model family is presented in detail. The model is validated against in-cylinder results provided by 3D-CFD analyses carried out 830 Vincenzo De Bellis et al. / Energy Procedia 45 ( 2014 ) 829 -838 with the Star-CD TM code for motored engine operations. In particular, a currently produced small turbocharged VVA engine is analyzed at different speeds, with valve actuations typical of full load and partial load (EIVC) operations, as well. The proposed turbulence model shows the capability, once tuned, to accurately estimate the temporal evolution of the in-cylinder turbulence according to the engine operating conditions. In the subsequent part II of the same paper, the developed turbulence model will be employed within a quasi-dimensional fractal combustion model.

49th AIAA/ASME/SAE/ASEE Joint Propulsion Conference, 2013

This study details the reactive flow simulation of a single injector LOX/CH 4 rocket similar to a comparable device, the CIRA Sub Scale Bread Board (SSBB) planned within the context of the HYPROB program funded by the Italian Ministry of University and Research (MIUR). In order to identify a suitable numerical method for use in the design of the SSBB calculations have been performed on the NASA sponsored Penn State LOX/CH 4 uni-element rocket engine for which heat flux data are available. The main features of this rocket are discussed in a former article relating to the experimental setup, methodology and results. The current reactive flow model is based on a twostage injector-chamber and chamber-throat-nozzle approach. In the first stage the injector-chamber region is modelled using a pressure based solver approach. In the second stage the converged results for the chemical distribution at the outlet calculated in the first stage are used as upstream boundary conditions to model the chamber-throat and nozzle region using a density based solver. In the first stage reactive flow is modelled using the Peng-Robinson equation of state whereas the second stage assumes an ideal gas. The reason for this two-stage approach is to reduce CPU time and circumvent problems associated with numerical stability which occur when modelling the complete domain in a single stage and which is validated by direct comparison of predicted wall heat fluxes with experiment. Comparison is also made between the results of this two-stage method and calculations for the complete configuration using different chemistry models. These simulations are performed with a commercial CFD solver to solve the steady state axi-symmetric Navier-Stokes equations. The Laminar Finite Rate and the Eddy Dissipation Concept models are used for the turbulent chemical kinetic coupling and the Jones-Lindstedt model is used for the chemical reactions for liquid oxygenmethane combustion. The results presented indicate the qualitative effects of different turbulence and chemical kinetic models on flow structures as well as a quantitative comparison of wall heat fluxes for different model combinations which are compared with experimental data.

Energy Procedia, 2014

It is widely recognized that air-fuel mixing, combustion and pollutant formation inside internal combustion engines are strongly influenced by the spatial and temporal evolution of both marco-and micro-turbulent scales. Particularly, in spark ignited engines, the generation of a proper level of turbulence intensity for the correct development of the flame front is traditionally based on the onset, during the intake and compression strokes, of a tumbling macro-structure. Recently, in order to both reduce pumping losses due to throttling and develop advanced and flexible engine control strategies, fully variable valve actuation systems have been introduced, capable of simultaneously governing both valve phasing and lift. Despite the relevant advantages in terms of intake system efficiency, this technology introduces uncertainties on the capability of the intake port/valve assembly to generate, at low loads, sufficiently coherent and stable structures, able therefore to promote adequate turbulence levels towards the end of the compression, with relevant effects on the flame front development. It is a common knowledge that 3D-CFD codes are able to describe the evolution of the in-cylinder flow field and of the subsequent combustion process with good accuracy; however, they require too high computational time to analyze the engine performance for the whole operating domain. On the contrary, this task is easily accomplished by 1D codes, where, however, the combustion process is usually derived from experimental measurements of the in-cylinder pressure trace (Wiebe correlation). This approach is poorly predictive for the simulation of operating conditions relevantly different from the experimental ones. To overcome the above described issues, enhanced physical models for the description of in-cylinder turbulence evolution and combustion to be included in a 1D modeling environment are mandatory. In the present paper (part I), a 0D (i.e. homogeneous and isotropic) phenomenological (i.e. sensitive to the variation of operative parameters such as valve phasing, valve lift, intake and exhaust pressure levels, etc.) turbulence model belonging to the K-k model family is presented in detail. The model is validated against in-cylinder results provided by 3D-CFD analyses carried out 830 Vincenzo De Bellis et al. / Energy Procedia 45 ( 2014 ) 829 -838 with the Star-CD TM code for motored engine operations. In particular, a currently produced small turbocharged VVA engine is analyzed at different speeds, with valve actuations typical of full load and partial load (EIVC) operations, as well. The proposed turbulence model shows the capability, once tuned, to accurately estimate the temporal evolution of the in-cylinder turbulence according to the engine operating conditions. In the subsequent part II of the same paper, the developed turbulence model will be employed within a quasi-dimensional fractal combustion model.

The industrial and scientific communities are devoting major research efforts to identify and assess critical technologies for new advanced propulsive concepts: combustion at high pressure, as well as the replacement of hydrogen with a hydrocarbon, has been assumed as a key issue to achieve better propulsive performances and lower environmental impact and to reduce the costs related to ground operations (propellant handling infrastructure and procedures) and increase flexibility. Starting from this background, the present work describes a CFD analysis (based on the numerical solution of RANS equations for chemically reacting flows) performed to evaluate the effects of different thermo-chemical modeling assumptions on the performances of a high-pressure LOx/CH4 rocket thrust chamber.

52nd AIAA/SAE/ASEE Joint Propulsion Conference, 2016