1D heterogeneous dual-phase model of a cryogenic liquid Hydrogen (LH2) propellant tank pressurization system was developed. Simulations were carried out using SINDA/FLUINT version 6.3, a finite difference, lumped parameter tool, to determine the quantity of gaseous Helium (GHe) for pressurization of the fuel tank in the cryogenic upper stage of a rocket. As there was no feasibility of establishing a lab-level twin of the total system, the model was validated on a sub-system level with data from other flight measurements. The design strategy was to identify the worst-case scenario so that the proposed active pressurization system would cater to most adversities in flight. The non-dimensional GHe mass and time required for pressurization were determined to be in the range of m* = (0.058 to 0.145) and Δτ = 0.18 × 10–6 to 0.38 × 10–6 respectively. Post-flight simulations indicated that m* = 0.1 of GHe was required to pressurize the tank by ΔP = 0.174 in Δτ = 0.32 × 10–6

A high-pressure LOx-Methane ( LOx - CH_4 ) rocket engine is characterized by transcritical injection and supercritical combustion phenomena which pose various difficulties during numerical modeling. Near the critical point region, thermophysical properties undergo highly non-linear behavior, which can impact the flow and flame evolution. In this study, a comprehensive thermodynamic modeling of LOx - CH_4 engine is conducted to capture such a non-linear phenomenon. A real gas-based thermodynamic and high-pressure transport framework is invoked to represent transcritical injection and supercritical combustion. Multiple real gas models are tested to evaluate thermophysical properties over the entire range of operations. Our study captured non-linear behavior near critical point operation of LOx - CH_4 engine. We identified a high-pressure thermodynamic and transport modeling framework that can be utilized for computational fluid dynamics (CFD) simulations.

The exhaust plume, emanating from the spacecraft thruster and composed of hot gases, expansively extends into the vacuum of space, frequently interacting with the spacecraft's structure that are often inclined relative to the thruster. Therefore, the heat flux and pressure resulting from plume impingement on inclined spacecraft surfaces are required to be determined. In this study, we perform steady-state CFD modeling to analyze the plume expansion into vacuum and the effect of plate inclination on the impingement heat flux and pressure. The CFD model incorporates a temperature jump boundary with a partial slip condition at the plate. The investigation is conducted based on the experimental conditions reported in literature. The CFD model with partial slip predicts the peak heat flux within 1

Fractal-based methods are applied to three full-scale liquid engine combustors. The structure function is derived in each case, and it is found to have a single or dual slope, depending on whether the combustion oscillations are stable or unstable, respectively. The fractal spectrum, which is derived from the structure function, is characterized by four spectral characteristics, viz., the apex location, the mean width, the mean height, and the differential height. The evolution of these characteristics is tracked with respect to time and is shown to be correlated to the onset of instability. The variation of the apex location over a limited range of scales is studied as a novel metric to anticipate instability in lieu of the Hurst exponent. Another novel metric that is based on the variation of slope of the structure function across two different scale ranges is also formulated. It is found that the metrics can be used to provide universal thresholds for forewarning of instability in earth-storable and cryogenic liquid rocket engines. The metrics could be used in the design of shutdown and control systems for liquid engines operating near combustion stability boundaries.

This study investigates the performance of a pulse detonation engine (PDE) utilizing methane-air mixtures, where detonation is initiated in a single-tube configuration using Shchelkin spirals. The experiments focused on five 400 mm long spirals with blockage ratios ranging from 0.08 to 0.17. The methane-air equivalence ratio varied in the experiments between 0.5 and 3.1. The measured detonation shock pressures and velocities closely matched the Chapman-Jouguet values, with both peaking at equivalence ratios between 0.9 and 1.2. The thrust of the PDE was estimated based on test data. The study shows that detonation shock pressure and velocity are only marginally affected by the blockage ratio, while its impact on the frequency of the detonation pressure wave is negligible. These findings highlight critical insights into how the design of the Shchelkin spiral governs the deflagration-to-detonation transition (DDT) and shapes the overall detonation dynamics in PDEs.

Numerical simulations of the firing of H2–O2 igniters in the presence of an air-O2 crossflow are conducted with a specific focus on air-heater configurations with faceplates. Simulations are conducted for three igniter chamber pressures (22, 41 and 55 bar) to investigate penetration lengths of igniter flames into the crossflow. Two alignments of the air-heater faceplate are assessed: one where the igniter-injector jets exhibit a direct interaction (Case 1), and one where the igniter jet is fired in-between two adjacent injector jets (Case 2). For all three igniter chamber pressures, air-heater faceplate heating is lesser for Case 1 compared to Case 2 because the igniter flame is naturally deflected away from the faceplate here. Moreover for Case 1, the 41 bar igniter chamber pressure shows the least peak heating of 68 W/cm2. For Case 2, with the 22 bar igniter chamber pressure, the igniter flame is deflected markedly towards the faceplate because of entrainment-instigated suction between two injector jets. This shows the highest peak heating on the faceplate of 467 W/cm2. When the igniter chamber pressure is increased, the extent of deflection reduces, as does the peak heating on the faceplate. Notwithstanding, hot-spots near the centre of the faceplate are still evident even at higher igniter chamber pressures.

The optimization of injector dynamics is essential for enhancing performance, reliability, and safety in rocket engines. This study examines the dynamics of an in-house-designed, coaxial swirl injector for a high-pressure LOx-CH4 rocket system, presenting a novel methodology for a swirl injector design that incorporates both flow and geometric parameters. A comprehensive parametric study, based on Bazarov's analytical framework, is conducted to assess the influence of geometric design variables on injector dynamics. The findings indicate that increasing the vortex chamber length (Lv) reduces the dynamic response and shifts the resonance peak to lower frequencies. An optimal radius ratio (Rrt) between 1.2 and 1.8 is identified to achieve a balanced dynamic response. Additionally, variations in the convergence angle ((3) have minimal impact on the injector's overall dynamics, with a 45 degrees angle found to be optimal for manufacturing purposes.This study offers crucial design recommendations for developing stable swirl injectors with an optimized dynamic response, providing valuable insights for improving injector performance and ensuring stability in high-pressure rocket systems.

An effervescent injector, as an emerging alternative to traditional pressure swirl atomizers, generates tiny droplets at lower injection pressures. An experimental investigation of the nitrogen-water atomization process in the inside-out effervescent injector is presented. The analyses are intended to figure out geometrical parameter influence on the performance of injectors through the measurement of the size of atomized droplets and spray distributions. This will eventually lead to the exploration of the application of this injector in air-breathing propulsion where ram air mixes with the fuel in the mixing chamber. This study examines the dependency of gas flow rate variations from 0.06 g/s to a maximum of 2 g/s, which will chock the water flow. The Sauter mean diameter (SMD) for various geometrical parameters and gas-to-liquid mass ratios (GLR) were measured using a Malvern particle analyzer. A high-speed camera was used to study internal flow regime variation and initialization of atomization by the bubble-breaking process. The experimental findings reveal that reliable conclusions regarding the SMD and distribution parameters cannot be drawn from a single geometrical configuration. To achieve optimal injector performance, the geometric parameters must be carefully tailored to the desired flow characteristics. Specifically, the exit orifice diameter has a significant impact on the SMD, while the distribution shows varying effects depending on the geometric combinations. Furthermore, liquid choking can be effectively controlled by adjusting the gas flow rate, opening up new possibilities for deep throttling.

The motion of liquid hydrogen in a cryogenic stage tank under microgravity conditions is numerically simulated in the present paper. The fluid motion is studied for various tank maneuvers such as tank rotation, simultaneous tank pitching and rotation, and disturbance imparted during satellite separation. The effectiveness of thrusters to settle the propellant and the effect of their firing durations on propellant availability near the drain port are determined. The propellant motion is studied using time-dependent computational fluid dynamics (CFD) simulations carried out in ANSYS Fluent. The vapor-liquid interface is captured using the volume of fluid (VOF) method. The surface tension is modeled using the continuum surface force model. The present analysis shows that capillary rise of the liquid during tank rotation takes place at slow speeds, whereas the bulk liquid motion is seen when the tank undergoes pitching and rotation simultaneously. Increasing the firing duration of the propellant settling thrusters from 5 s to 10 s is found to increase the propellant volume fraction at the drain port location from 0.89 to 0.99. Analysis shows that during spacecraft separation, the liquid-vapor interface is disturbed due to the sudden deceleration felt during separation. The liquid hydrogen blobs separate from the bulk liquid and migrate towards the tank fore-end, unlike the cases of tank rotation and simultaneous pitching and rotation. The present work clearly shows that significant liquid movement happens during spacecraft separation and tank pitching, and thus, a sufficient impulse needs to be provided by firing thrusters for settling the propellant before subsequent engine operations.

Hybrid rocket motors (HRM) are of current interest to space programmes in view of their throttleability, restartability, environmental friendliness and safety. But HRMs like the ones using liquid oxygen-hydroxyl terminated polybutadiene hybrid are prone to combustion instability. In order to develop a stable hybrid rocket motor, detailed characterisation of impinging doublet injector was carried out. The injector head is comprised of 60 identical doublets arranged in three axisymmetric rows that ensure uniform spray of liquid oxygen in the chamber. Jet impingement and fine atomization were demonstrated in the cold flow tests of the injector head. This injector performance was assessed during cold flow tests, both with and without backpressure. The measurements of droplet sizes were corroborated by a linear stability model. The hybrid motor with this injector head was successfully hot tested and stable combustion was demonstrated for a duration of 15 s. The stiffness of the injector was evaluated and found to be adequate to suppress feed coupled oscillation. The details of the configuration, design aspects and measured cold flow spray characteristics of the injector as well as hot test performance of the hybrid motor are presented. The injector configuration plays a crucial role in achieving stable combustion, making it a promising candidate for enhancing the performance of high-thrust hybrid motors.

In liquid rocket engines, cavitating venturi is used as flow control device which meters the constant flow rate at fixed inlet pressure under varying back pressure conditions. Cavitating venturi is used mainly in two configurations, fixed area cavitating venturi and variable area cavitating venturi. The use of fixed area cavitating venturi is more prominent in propellant feed lines of fixed thrust engines as well as that of test facility. In case of throttling engines, where a proper flow control is needed at different thrust level, variable area cavitating venturi is preferred. The present work investigates the cold flow characterization of a variable area cavitating venturi designed for 1000N MMH-N2O4 based Throttleable Pintle engine. The flow controlling area for the cavitating venturi was varied using a needle by mechanical means. Experiments were carried out using water as simulant fluid. Effect of back pressure on flow characteristics was studied at different stroke lengths of the moving needle at fixed inlet pressure. In addition, effect of varying inlet pressure on flow characteristics was also studied at a fixed stroke length of the moving needle.

This paper investigates combustion dynamics in a complex multi-injector element combustor using a flamelet approach in a large eddy simulation (LES) framework. The capability of computationally less expensive chemistry tabulation method to capture the interaction between unsteady heat release and acoustics is investigated. A non-adiabatic steady flamelet-based tabulated chemistry closure is invoked to simulate hydrogen–oxygen reactions in mixture fraction space. The model incorporates flow-induced non-equilibrium flame effects through scalar dissipation rate and the turbulence-chemistry interaction using a probability density function (PDF). A multi-element combustor dynamic study captures the first tangential mode close to 4000 Hz and corresponding high-frequency harmonics appropriately. Spectral analysis of the pressure variation displays similar frequency features in chamber and injector sections, suggesting the possibility of injector-chamber coupling. The coupling of the transverse pressure waves in the combustion chamber with the longitudinal pressure oscillations in the oxidizer post was probed as the reason for the pressure dynamics observed in the combustor.

Impingement of satellite thruster plume on the adjacent surfaces of the satellite generates disturbance torques and heat loads that are undesirable. It is therefore important to configure the spacecraft such that the impingement effects of the thruster plume are minimized. In the present work, Computational Fluid Dynamics (CFD) study of plume expansion in to low pressure environment is carried out, and the impact pressure and heat flux on the flat plate are determined. The simulations are carried out for the reported experimental conditions of plume impingent on a flat plate for various standoff distances of the plate from the nozzle. The Knudsen number distribution shows that the flow remains in continuum in the vicinity of the nozzle and at the impingement region on the plate. The results of the simulation are compared with the measured test data for model validation. Finally, the effect of incorporating temperature jump condition on the impingement heat flux estimation is presented. The CFD model predicts the peak pressure and heat flux within 1.6 and 11.9

This paper presents the computational methodology developed to simulate monomethyl hydrazine/nitrogen tetroxide (MMH/NTO) combustion. A three-dimensional rocket scale combustor domain with multi-element triplet injectors is utilized to study hypergolic flow and flame features. A Eulerian–Lagrangian framework is invoked for continuous phase treatment of combustion gas and discrete phase treatment for both MMH and NTO droplets. A discrete particle-based method (DPM) with finite rate chemistry is employed to study droplet injection, evaporation, and combustion. A description of flow and flame characteristics in three-dimensional RANS framework is presented in this paper. The model captures impinging jets from multiple triplet injectors, and MMH film cooling injection appropriately. It presents physical trends on the core combustion process, as well as the global evolution of temperature, pressure, and droplet spray in the combustor. The focus of the study is to develop a hypergolic combustion model which can be used to predict combustion performance under off-nominal operating conditions. The aim is to extend the model to study the combustion instability aspects of MMH/NTO-based combustors.

In the realm of reusable rocket technology, the methane–oxygen propellant combination has gained prevalence in recent years. However, understanding of design parameters affecting the combustion performance of a methane–oxygen combustor is currently limited. The present study proposes a novel analytical and computational approach to design and optimize a combustor. The design of a seven-shear coaxial injector-based combustion chamber is optimized using an in-house developed code line with the aim of minimizing chamber length while preserving chamber performance. Numerical simulations are then carried out to ensure complete combustion within the chamber. Methane–oxygen reactions are modeled using laminar finite-rate and eddy dissipation concept models. An optimized design for the combustion chamber shows maximum characteristic velocity and optimal chamber length near the oxygen-to-fuel mass ratio of 2. The computational study shows that the eddy dissipation concept model can accurately capture the effect of turbulent mixing on gaseous methane–oxygen reactions. The eddy dissipation concept model yielded notable differences in flame characteristics compared to the laminar finite-rate model. The study indicates a need to optimize injector configuration. An optimized injector configuration is proposed by varying its geometric parameters, which exhibits substantial improvement in combustion performance compared to the initial configuration, utilizing practical recess, divergence, and an increased velocity ratio.

Numerical analysis of the hot oxygen rich exhaust gas flow field during the ground test of engine subsystem is carried out in the present paper. The effect of water injection on the hot gas plume and its impingement velocity on the test bay is studied. The effect of the wind conditions such as velocity and directions on the oxygen concentration in the test facility is also studied. It is found that a single water jet injected at 500kg/s with 55m/s velocity is able to reduce the impingement velocity of the hot gas from 250m/s to 75m/s. Analysis shows that the wind up to 12m/s does not have significant effect on the oxygen concentration in the test facility. The wind blowing opposite to the hot gas flow results in oxygen accumulation at ground level of the test facility.