Numerical analysis of combustion instability in liquid propellant rocket engine

The nonlinear high frequency combustion instability in liquid rocket engine is simulated in this paper. The gas governing Navier-Stokes equations are described under the Euler coordinate. The liquid phase governing equations are described under the Lagrangian coordinate. The two phase influences of gas and liquid are through the interaction of source terms of equations. The spray process in rocket engine is described with evaporation theory. The numerical method is MacCormack scheme which is two order accuracy in space and time direction. The two phase turbulent combustion steady process and high combustion instability in liquid rocket engine are simulated with the method of computational combustion dynamics. When there are spontaneous high frequency combustion instability in liquid rocket engine, the pressure in combustion chamber is oscillating near the steady state at the beginning, and as the simulated time passes, the pressure is exponentially increased until the high amplitude pressure oscillation is excited in combustion chamber because of nonlinear effects. In the three dimensional phase space of chamber pressure, mixed gas density and mixed gas velocity, the physics parameters are going round and round near the attractor, the six spiral main moving trajectories are created. There are lots of secondary moving trajectories between the main trajectories. The chaotic state is appeared far away from the attractor point. The same rule of moving trajectories is occurred in the three dimensional phase spaces of chamber pressure, mixed gas density, spray velocity and mixed gas density, mixed gas temperature, mixed gas velocity.

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.

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

Performance, weight, design, flexibility, engine cycles and reliability of liquid propellant rocket engines used in NASA and USAF vehicles

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

: This research was directed towards developing new knowledge of combustion processes in liquid-propellant rocket engines. Attention was given to the combustion of different fuels, beginning with hydrogen-oxygen but ultimately including multicomponent hydrocarbon mixtures representative of RP fuels, and to nonlinear processes of importance in combustion instability. Effects of detailed chemical kinetics on performance and stability were addressed by introducing systematically reduced chemistry that enables theoretical analysis to be completed thoroughly and accurately. The interactions between chemistry and turbulence were addressed, including the effects of the heat release on the turbulence. The orientation of the work was fundamental and aimed at predicting not only performance and acoustic response but also intrinsic instability. The results were intended to help to improve understanding of combustion mechanisms and combustion instabilities in liquid-propellant rocket motors. The most recent specific results concern induction times in hydrogen-oxygen mixtures and heal-release effects on turbulent missing.

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.

: Combustion instability in liquid rocket engines can have severe consequences including degraded performance, accelerated component wear, and potentially catastrophic failure. High-frequency instabilities, which are generally the most harmful in liquid rocket engines, can be driven by interactions between disturbances associated with transverse acoustic resonances and the combustion process. The combustion response to acoustic perturbation is a critical component of the instability mechanism, and is in general not well understood. The current paper describes an experimental facility at the Air Force Research Laboratory (AFRL) at Edwards Air Force Base that is intended to investigate the coupling between transverse acoustic resonances and single/multiple liquid rocket engine injector flames. Critical aspects of the facility will be described, including the capability to operate at supercritical pressures that are relevant to high-performance liquid rocket engines, accurately-controlled and cryogenically-conditioned propellants, and optical access to facilitate the use of advanced diagnostics. The transverse acoustic resonance is induced through the use of carefully-controlled piezo-sirens, allowing monochromatic excitation across a range of amplitudes at a number of discrete frequencies. The location of the flame within the acoustic resonance mode shape can also be varied through relative phase control of the two acoustic sources. The operating space of the facility, for oxygen and hydrogen operation, will be described. Preliminary non-reacting and reacting data will also be presented to demonstrate the quality of operation of this facility. It is anticipated that future results generated using this facility will provide both fundamental insight into the acoustic-flame interactions as well as provide a database useful for validating combustion instability models.

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.

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.

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.

Coupling of harmonic flow oscillations to combustion instability in premixed segments of triple flames