Liquid propellant rocket engines - Their status and future.

A complete set of equations describing the conservation of mass, momentum, and energy, and multicomponent transport and chemical reaction in a liquid propellant rocket (LPR) engine are given, along with appropriate boundary conditions. The formulation includes a detailed description of the two-phase (gas/liquid) flow in terms of a moving boundary problem, with droplets and vapor occupying time-dependent subdomains of the LPR chamber. Several sensor and actuator types are considered in detail, and their mathematical descriptions are incorporated into the differential equations and boundary conditions.

The design of liquid propellant rocket engines (LPREs) necessitates the efficient cooling of the combustion chamber walls. Despite the prevalence of regenerative cooling, it frequently proves inadequate, necessitating supplementary techniques such as the injection of coolant into the combustion chamber to establish a thermal barrier. A critical deficiency in the prevailing methodologies pertains to the absence of dependable methodologies for the calculation of the parameters governing coolant flow as it transitions from the cooling path to the curtain belt. The objective of this study is to derive analytical formulas for determining the parameters of the coolant flow at the exit from the curtain belt of the cooling belt of an LPR combustion chamber. The coolant passes through tangential cylindrical channels into the annular slot, thereby forming a continuous swirling flow prior to entering the chamber. The analysis is conducted in two stages. First, a one-dimensional approximation is used to determine the temperature change in the cylindrical inlet channel. Then, the flow in the annular slot is considered using the laws of conservation of mass, momentum and energy. An analytical solution for the temperature distribution is obtained by means of simplification of the dependence for the Nusselt number. The model has been verified on the data of the RD-108 engine curtain belt with kerosene as a coolant. The results indicate that the coolant temperature at the outlet is nearly equal to the temperature of the chamber wall, thereby substantiating the practical applicability of the model. In this paper, analytical expressions for the coolant flow parameters at the outlet of the curtain belt are obtained for the first time, and an approximate method for evaluating the heat transfer in the annular slot is proposed. Despite its inability to account for alterations in fluid properties and variable wall temperature, the model functions as a framework for subsequent refinements.

Numerical study of combustion instability in liquid propellant rocket engine is presented in the paper. Time-depending two-dimensional Navier-Stokes equations were employed to describe the fluid flow in the combust chamber and solved by finite difference method with MacCormack scheme. Turbulent viscosity was taken into consideration by the subgrid scale turbulence model. Propellant spray was treated with a lagrangian tracking method and droplet source terms were incorporated into the gas phase calculation by PSIC method. Droplet initial conditions were determined by a stochastic spray model . Spray burning rate is calculated by the classic D2 vaporization law with Ranz-Marshall correlation. The coupling of pressure disturb and burning rate variation was described with Prime's twoparameter model. The actual axisymmetric combustor was transferred into equivalent annular geometry as the references do. Combustion instability analysis was performed by imposing a bomb pressure pulse on the computed steady flow field to trigger unstable combustion. By observing the pressure wave propagation , we can asses the damping capability of the engine.The computational analysis show that numerical schemes and physical models used to describe unstable combustion are the most important factors to the success of numerical study. d NOMENCLATURE a sound velocity b annular combustor width constant pressure spcific heat CP C,, drag coefficient d droplet diameter D mass diffusivity e total energy f frequency h enthalpy 1 L~ heat of evaporation f i transverse length of the annular combstor mass flux or evaporation rate M number of computational droplet NU Nusselt number pr Prantdl number Ice Reynolds number s source terms time T lempeerature X,Y coordinates p density h coductivity P I velocity components

Fault detection and diagnosis algorithms for an open-cycle liquid propellant rocket engine using the Kalman filter and fault factor methods

Non-stationary heat behaviour of liquid propellant rocket engines

Blow forming technology has been used to produce glass and plastic components with easy and simple process. In order to apply this technology to metallic system, relatively high formability of material and elevated temperature are required. Since the blow forming temperature is quite high for metallic system, the main difficulty is to obtain adequate mold material which can sustain pressure difference between inside and outside of the tool. The present study demonstrates blow forming process with IN718 and steel can be applied to manufacturing of combustion chamber in liquid propellant rocket engine with complex shape. The result shows that the developed technology to process design of high temperature blow forming by the finite element method can be applied for near net shape forming of a combustion chamber of liquid rocket engine and a cone-shaped chamber extension part with cooling channels.

Purpose. This study aims to improve the accuracy of methods for determining fin efficiency. Its goal is to derive calculation relationships for finning coefficients that account for transverse temperature non-uniformity within the fin cross-section. Design / Method / Approach. The article presents the reduction of the heat conduction equation for a fin to a dimensionless form, based on dimensional analysis of the variables involved. Further development relies on analyzing the results of numerical simulations and their subsequent generalization. To this end, the gradient descent method is applied, minimizing the quadratic error function. Findings. A criterial dependence has been formulated to complement the derived heat conduction equation. Test calculations and comparisons with numerical simulations in Ansys Fluent confirm an improvement in calculation accuracy when using the proposed equation. Theoretical Implications. This paper addresses factors previously neglected in the analysis of heat transfer in fins. The results of the study thus complement existing approaches to determining finning coefficients. Practical Implications. The derived criterial relationship will enhance the accuracy of heat transfer calculations in the chambers and gas generators of liquid rocket engines. Originality / Value. The paper introduces an original criterial relationship that accounts for temperature non-uniformity across the fin cross-section. Incorporating this factor improves calculation accuracy, highlighting the practical value of the developed equation. Research Limitations / Future Research. This study focuses on rectangular fins; therefore, the proposed model is not applicable to fins with variable thickness in the ducts of liquid propellant rocket engines (LPREs). Developing a fin model without these limitations will be the objective of future research on this topic. Article Type. Applied Research. PURL: https://purl.org/cims/4.299

The inducer is an axial pump that is part of the propellant injection system of Liquid Propellant Rocket Engines (LPRE). It is located at the inlet of the turbopump assembly and is critical for designing high performance LPREs. Its geometric and operational characteristics allow it to operate at low inlet pressures, delaying the appearance of cavitation and allowing the propellant tanks to operate at lower pressures. This allows the tanks to be lighter due to a reduced wall thickness requirement. The inducer also needs to operate harmoniously with the other components of the turbopump, especially with the main impeller which is located just downstream in the system. Therefore, it is important that the flow conditions at the inducer inlet and outlet are known and integrated with the turbopump and tank design. The present work aims to develop a methodology for inducer design based on literature established methods in order to obtain geometry and evaluate the flow conditions in liquid-propelled rocket engine inducer pumps. This work will assess outlet flow and pressure conditions in a way that it is possible to match them with the main impeller inlet. Performance criteria are evaluated in terms of the outlet pressure coefficient, flow coefficient and efficiency focusing exclusively on non-cavitating conditions. Two established analytical methods were implemented, one to provide inducer geometry in terms of system operational requirements and another, from National Aeronautics and Space Administration (NASA), for performance prediction based on geometrical and operational parameters. Further analysis is complemented by simulating the generated geometry in a CFD software. The methods were validated using published experimental data and the performances of the analytical, numerical and experimental results were compared. Results showed that the 3D turbulent CFD simulations provided very good agreement of efficiency. Satisfactory results were obtained for the general trends of characteristic curves over a range of flow rates and the spanwise distribution of key performance parameters near design point. The pressure coefficient was significantly overestimated. The results of the analytical models showed good agreement with simulated CFD results, indicating appropriate calibration of loss coefficients.

For low cost, high thrust, space missions with high specific impulse and high reliability, inert weight needs to be minimized and thereby increasing the delivered payload. Turbopump feed system for a liquid propellant rocket engine (LPRE) has the highest power to weight ratio. Turbopumps are primarily equipped with an axial flow inducer to achieve the high angular velocity and low suction pressure in combination with increased system reliability. Performance of the turbopump strongly depends on the performance of the inducer. Thus, for designing a LPRE turbopump, demands optimization of the inducer geometry based on the performance of different off-design operating regimes. In this paper, steady-state CFD analysis of the inducer of a liquid oxygen (LOX) axial pump used as a booster pump for an oxygen rich staged combustion cycle rocket engine has been presented using ANSYS® CFX. Attempts have been made to obtain the performance characteristic curves for the LOX pump inducer. The formalism has been used to predict the performance of the inducer for the throttling range varying from 80% to 113% of nominal thrust and for the different rotational velocities from 4500 to 7500 rpm. The results have been analysed to determine the region of cavitation inception for different inlet pressure.

In the Republic of Korea, research on staged-combustion cycle liquid propellant rocket engines (LPRE) is proceeding to improve efficiency of rocket engines. Recently oxidizer-rich preburner using single triplex injector is developed in relation to the main injector development and combustion tests have been performed. This preburner is designed to operate in nominal conditions with the combustion pressure of 10 MPa, OF ratio of 60. For a stable ignition, LOx is fed in two steps. Triethylaluminum-Triethylborane (TEAB) is used as hypergolic fuel for ignition, supplied through a fuel injector. Despite the small amount of fuel flow rate and high pressure condition, the combustion pressure was stably maintained around 10 MPa as designed. As a result of Fast Fourier Transform (FFT) of the combustion chamber dynamic pressure, 1L mode frequencies related to the acoustic instability and hydraulic resistance exist in the combustion chamber. But their amplitudes are less than 1% of the combustion pressure and it does not affect the combustion. Therefore combustion test is stably completed. In the near future, coupled tests with uni-element triplex injector preburner and uni-element gas/liquid injector main combustion chamber will be carried out.

Sensing and actuation in a liquid propellant rocket engine

We present a method that combines multiple sensory modalities in a rocket thrust chamber to predict impending thermoacoustic instabilities with uncertainties. This is accomplished by training an autoregressive Bayesian neural network model that forecasts the future amplitude of the dynamic pressure time series, using multiple sensor measurements (injector pressure/ temperature measurements, static chamber pressure, high-frequency dynamic pressure measurements, high-frequency OH* chemiluminescence measurements) and future flow rate control signals as input. The method is validated using experimental data from a representative cryogenic research thrust chamber. The Bayesian nature of our algorithms allows us to work with a dataset whose size is restricted by the expense of each experimental run, without making overconfident extrapolations. We find that the networks are able to accurately forecast the evolution of the pressure amplitude and anticipate instability events on unseen experimental runs 500 milliseconds in advance. We compare the predictive accuracy of multiple models using different combinations of sensor inputs. We find that the high-frequency dynamic pressure signal is particularly informative. We also use the technique of integrated gradients to interpret the influence of different sensor inputs on the model prediction. The negative log-likelihood of data points in the test dataset indicates that prediction uncertainties are well-characterized by our model and simulating a sensor failure event results in a dramatic increase in the epistemic component of the uncertainty, as would be expected when a Bayesian method encounters unfamiliar, out-of-distribution inputs.

Deep neural network approach for fault detection and diagnosis during startup transient of liquid-propellant rocket engine

<div class="section abstract"><div class="htmlview paragraph">This paper studies the “Experimental testing of spray characteristics of swirler injector without guide vanes in Liquid Propellant Rocket Engines”. The purpose of a fuel injector is to inject and control the flow of the propellants into the combustion chamber. It consists of Tangential ports, a swirls chamber, a converging spin chamber, and a discharge orifice. Kerosene as fuel and liquid oxygen as oxidizer were employed here. To pressurize the propellants, nitrogen gas is used and the pressurized pipeline is controlled with an open/close ball valve and measure the inlet pressure using a pressure gauge. When a propellant comes through the inlet has a tangential velocity and it causes the propellants to swirl inside the swirl chamber at the exit the propellant comes with rotational momentum and forms a liquid film and then the friction between the propellant and air accelerates and disrupts the liquid film and converts into very tiny droplets. The exit mass flow rate, spray cone angle, and spray cone length with different inlet pressures and inlet diameters are measured. A graph is plotted between inlet pressure, and spray cone angle, penetration length. The theoretical mass flow rate is calculated using Bernoulli’s equation, taking the coefficient of the discharge as one and plotted a graph with the given data and comparing it with the experimental data. The accurate atomization increases the combustion efficiency, burns the propellant completely without leaving any unburnt propellant in the combustion chamber, and minimizes combustion instability.</div></div>

This paper presents the results of studies along the principal research lines of Viktor V. Pylypenko, an Academician of the National Academy of Sciences of Ukraine, obtained by his pupils over the recent years. The following lines are considered: the development of a theory of cavitation oscillations in liquid-propellant rocket engine (LPRE) pumps, the study of low-frequency dynamics in liquid-propellant rocket propulsion systems (LPRPSs), the analysis and assurance of liquid-propellant launch vehicle pogo stability, the mathematical simulation of dynamic processes in solid-propellant rocket engines, the development of pneumatic vibration protection modules, and the improvement of drilling technology by using cavitation hydrovibrator. A hydrodynamic model of cavitating LPRE pumps was developed further: it was verified using theoretical and experimental transfer matrices of cavitating pumps, and two new coefficients were introduced: the time of disturbance transfer delay due to the presence of cavities and the cavitation resistance distribution coefficient; from the results of dynamic tests of 26 pumps, the cavity elasticity was determined for pumps with extended ranges of their geometric parameters and operating conditions; a pump choking mechanism was developed, and it was shown that the pump choking characteristic is a specific nonlinearity involving a critical cavity flow; a mechanism of hard excitation of cavitation self-oscillations was proposed, and they were mathematically simulated. A theory of LPRPS low-frequency dynamics was developed further: procedures were developed for determining the effect of a nonsimultaneous engine startup and external and internal factors on LPRE startup thrust spread; a mathematical simulation was conducted to study sustainer engine startup and shutdown transients in the common feed system of the sustainer engine and the liquid-propellant thrust system of the Cyclone-4M launch vehicle’s upper stage. A study was conducted on pogo vibrations in the prototype Cyclone and Dnipro launch vehicles unstable for pogo vibrations during the first-stage engine operation.