Rocket Thrust Chambers Research Articles

Mass minimization approach for the optimal preliminary design of CMC inner liners in rocket thrust chambers

In the past decade, the world has witnessed a new space race, driven by a growing commitment to reducing the environmental impact of space missions. This has led to the widespread adoption of liquid-propellant rocket engines, which offer several advantages over their solid-propellant counterparts. One key advantage is their reusability, which not only helps to reduce the generation of space debris but also makes space exploration cheaper. To further enhance the performance of liquid rocket engines, researchers have been exploring innovative cooling techniques and advanced materials. Among these materials, Ceramic Matrix Composites (CMCs) have shown great potential in reducing the overall engine weight when used instead of high-tech metal alloys, resulting in lower fuel consumption and emissions during launches. This paper focuses on the mass minimization of inner liners made of CMCs in rocket thrust chambers. At this aim, a computationally efficient preliminary design approach, based on an analytical one-dimensional thermo-mechanical model, is proposed. A case study of mass minimization of an inner liner of rocket thrust chamber is also presented and discussed, by considering five different CMC materials.

Multi-objective optimization of the aerodynamic performance of butterfly-shaped film cooling holes in rocket thrust chamber

Abstract This study uses Multi-Island Genetic Algorithm (MIGA) and three-dimensional Computational Fluid Dynamics (CFD) software to optimize butterfly-shaped film cooling holes in the upper-stage rocket engine thrust chamber. The goal is to meet thermal protection and thrust requirements at high altitudes without re-ignition. To facilitate an all-encompassing worldwide search, the holes in the optimized design remain at set dimensions. Film continuity and stability at the nozzle outlet are greatly impacted by the hole structure. Inlet and divergence angles have little effect on thrust, according to regression research, but lip height (de) and outlet width (β) have a big impact on cold gas ejection, which affects cooling and thrust. Optimized results lead to a 20.49 K decrease in the monitoring section’s average wall temperature and a 52.8 N boost in thrust by reducing interference between supersonic airflow and extending film stability.

An analytical formula evolved from thermal-structural coupling model for life prediction and structural design of reusable thrust chambers

The thermal-structural analytic model is essential for the life prediction of reusable rocket thrust chambers. The common numerical model based on the finite element method is computationally intensive. In order to simplify the life prediction method and improve the calculation efficiency, a thermal-structural analytical formula is proposed. The one-dimensional analytical heat transfer algorithm for accurately calculating the temperature of the thrust chamber is established, and Porowski’s beam model for calculating the structural deformation is adopted. Finally, the analytic formula for life prediction is directly established by reasonable simplification. This is an algebraic solution describing the relationship between life and the cooling channel parameters, and the computational efficiency is improved by about 12,000 times. The calculated temperature and deformation of the inner wall are verified by the finite element method with an error of less than 10 %, indicating the proposed analytical formula is accurate. Further, the analytical formula is applied to design the cooling channel parameters according to the given target life. The structural design method is validated by numerical simulation and experimental data. The analytical formula and the structural design method provide new ideas for the analysis and design of reusable rocket engines, which have broad application prospects.

Thermal performance of MMH/NTO rocket thrust chamber based on pintle injector by using liquid film cooling

A liquid rocket engine needs to be protected against damage of the combustion chamber wall during operation, being exposed to high temperature and pressure gas conditions. It is in urge to design a simple and efficient thermal protection method to cool the wall of the thrust chamber. In present study, a liquid rocket engine CFD model with cooling holes in the upstream side of the thrust chamber wall was established and the propellant combustion performance of monomethylhydrazine (MMH) and nitrogen tetroxide (NTO) was simulated by a simplified multi-step chemical reaction mechanism. The liquid cooling performance of thrust chamber wall was numerically investigated by using the discrete phase model coupled with the wave breakup model and the convection/diffusion controlled model, in which the liquid film coolant is regarded as the discrete phase particles by considering the entrainment and evaporation of liquid coolant droplets in hot combustion gas. The effects of pintle head diameter Dp, oxidant slot height h0 and coolant consumption ratio on liquid cooling performance were discussed. The results demonstrated that smaller Dp and h0 can help to improve the characteristic velocity of the thrust chamber. As the coolant ratio increases from 18% to 24%, the maximum wall temperature of the thrust chamber is reduced by 11% (212 K) with only a 1.3% reduction in characteristic velocity (combustion performance). The present numerical method and results can provide important guidance for liquid film cooling design of the liquid rocket engine based on pintle injector.

A simplified one-dimensional transient heat transfer model for rocket thrust chamber

The heat transfer model is essential to rocket engine thrust chamber, for it determines the accuracy of chamber wall temperature and heat flux distribution. In order to predict the transient temperature field during startup and shutdown, we introduce a simplified one-dimensional model of transient heat transfer for thrust chamber. The heat transfer process is divided into two parts: the convective heat transfer from hot gas to the inner wall and the heat conductance through the chamber wall. The Bartz formula modified with constant coefficient C is used to characterize the convective heat transfer and the one-dimensional transient heat transfer model of slab is the calculative foundation of the heat conductance. Finally, we propose a strategy to iteratively solve the model. To verify the model, it is implemented in a series of hot run experiments for a LOX/Methane heat sink thrust chamber. The average error of wall temperature evaluated at different locations of the thrust chamber inner wall can be reduced up to a 1.76%. The application of our novel method provides an accurate and efficient tool to calculate chamber wall temperature and heat flux distribution in the thrust chamber design. In addition, through inverse analysis of the proposed model, we can also estimate the near-wall gas temperature, which is difficult to obtain in a hot run test, by acquiring wall temperature data.

Investigations on an All-Oxide Ceramic Composites Based on Al2O3 Fibres and Alumina–Zirconia Matrix for Application in Liquid Rocket Engines

High performance ceramics, particularly Ceramic Matrix Composite (CMC) materials found their way into liquid rocket engines. Yet, so far, mainly carbide or nonoxide CMCs have been of interest. This paper explores the potential and challenges of oxide–oxide ceramic matrix composites (OCMCs) for application in rocket thrust chambers. Therefore, strength, leakage and hot gas tests are conducted with material samples. A particular focus lies on the application of coatings to seal the permeability inherent to the material. Furthermore, prototypes in the form of flame tubes, ceramic chambers with nozzles and ceramic chambers with graphite inlays are developed and investigated experimentally in test firings. The results show that a recrystallised glass of a Y-Al-Si-O compound can successfully create an impermeable coating of the OCMC without affecting its damag-tolerant behaviour. However, the prototype developments show that it is still very challenging to manufacture even slightly complex structures without critical failures. Nevertheless, OCMC structures of relatively simple geometries showed promising results in hot firings and could be used as a lightweight housing, while the inner contour of the chamber and nozzle are realised, e.g., by a graphite inlay of appropriate quality.

Numerical analysis of self-excited tangential combustion instability for an MMH/NTO rocket combustor

This study presented the numerical simulation of the tangential combustion instability in a hypergolic liquid bipropellant rocket thrust chamber, which applied fuel liquid film cooling method and unlike impinging injectors. The liquid spray was modeled using Lagrangian approach, while the gas was regarded as Euler phase. Stress-blended eddy simulation and finite rate/eddy–dissipation model were adopted to simulate the turbulent combustion process. Consistent with the experiment results, this work successfully simulated the transformation of tangential combustion instability from standing mode to spinning mode. The mean pressure, amplitude and frequency of limit cycle oscillation were in good agreement with the experiment. There was a detailed analysis about the flow field, Rayleigh index, and driving mechanism of the combustion instability. It was found that the oscillation began with hot spots of heat release rate due to the interaction between the spray of impinging injectors and cooling fuel jet. More than that, cooling fuel jet also contributed to drive the oscillation. In the standing mode, injectors in the inner and outer rings drive the oscillation together, while the spinning mode is mainly driven by injectors in the outer ring. The pressure wave is subsonic and its Mach number is close to 1. It was shown that the pressure wave contained a complex structure divided into three parts. This led to the in-phase of the pressure along the axial direction and the double-peak characteristic of the downstream pressure signal. Besides, a positive feedback closed-loop system associated with periodic oxidizer/fuel ratio was believed to sustain the combustion instability. The oscillation can be maintained when pressure, heat release and oxidizer/fuel ratio are coupled together. The analysis results indicate that rotating detonation is an implication to tangential combustion instability.

Forecasting thermoacoustic instabilities in liquid propellant rocket engines using multimodal Bayesian deep learning

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.

Effect of the Addition of Graphene Nanoplatelets on the Thermal Conductivity of Rocket Kerosene: A Molecular Dynamics Study.

Rocket kerosene plays an important role in the regenerative cooling process of rocket thrust chambers. Its thermal conductivity determines the cooling efficiency and the tendency to coke in rocket kerosene engines. In this paper, graphene nanoplatelets (GNPs) were introduced into rocket kerosene to improve its thermal conductivity. Molecular dynamics simulation was used to investigate the thermal conductivity of the composite system and its underlying thermal conductivity mechanism. Firstly, by studying the effect of the mass fraction of GNPs, it was found that, when the graphene mass fraction increases from 1.14% to 6.49%, the thermal conductivity of the composite system increases from 4.26% to 17.83%, which can be explained by the percolation theory. Secondly, the influence of the size of GNPs on the thermal conductivity of the composite system was studied. Basically, the thermal conductivity was found to increase by increasing the aspect ratio of GNPs, indicating that GNPs with a higher aspect ratio are more conducive to improving the thermal conductivity of rocket kerosene. By carefully analyzing the effect of the size of GNPs on thermal conductivity, it was concluded that the thermal conduction enhancement by adding GNPs is determined by the combined effect of the percolation theory and the Brownian motion. The results of the temperature effect study showed that the ratio of thermal conductivity to rocket kerosene increased from 1.16 to 1.26 and from 1.07 to 1.11 for the composite systems, with graphene sizes of 41.18 Å × 64.00 Å and 24.14 Å × 17.22 Å in the temperature range of 293 K to 343 K, respectively. It is further proved that the Brownian motion of GNPs has a non-negligible effect on the thermal conductivity of the composite system. This work provides microscopic insights into the thermal conduction mechanism of GNPs in nanofluids and will offer practical guidance for improving the thermal conductivity of rocket kerosene.

Impact of shear-coaxial injector hydrodynamics on high-frequency combustion instabilities in a representative cryogenic rocket engine

The excitation mechanism of a thermoacoustic instability in a 42-element research rocket thrust chamber with representative operating conditions with respect to European cryogenic rocket engines is investigated in detail. From previous research it was known that the chamber 1T mode can be excited by persistent heat release rate oscillations which are modulated by the resonant modes of the liquid oxygen injectors. The excitation source of the longitudinal injector eigenmodes is investigated in this study. Fibre-optical probes measuring the OH* dynamics from the recess volume of two injectors showed additional frequency content which could neither be explained by the chamber acoustics, nor the acoustics of the injection system. Instead, the temporal evolution of these frequencies correlate with the oxidizer flow velocity. In this work we show that the additional flame modulation originates from a hydrodynamic effect in the injection system. Even though the exact process cannot be precisely identified, an effect designated orifice whistling at the injector inlet orifice seems to be a likely candidate. Combining the new results with previous publications about this combustor, it is now possible to explain past and present observations in terms of the hydrodynamic and thermoacoustic conditions which are necessary for the combustion instability to appear. The conditions, which lead to an injection-driven excitation of the 1T mode are matching frequencies of the 2L mode of the injectors and the chamber 1T mode as well as a Strouhal number between 0.2 and 0.4 based on the length and flow velocity of the injector inlet orifice.

Analysis of a transpiration cooled LOX/CH[formula omitted] rocket thrust chamber

Transpiration cooling has been theoretically identified as a more efficient thermal protection system, compared to regenerative and film cooling techniques. The availability of porous light-weight high-temperature materials, such as ceramic matrix composites (CMCs), is making transpiration cooling an attractive solution for use in future re-usable rocket thrust chambers for cheaper access to space. In this study, a numerical framework, utilizing the industry standard computational fluid dynamics (CFD) code ANSYS CFX, for the simulation of transpiration cooled rocket thrust chambers is presented. The framework is shown to be suitable for predicting the performance, both in-terms of thermal protection and thrust produced, of thrust chambers that utilise transpiration cooling systems. A liquid oxygen (LOX) and methane (CH4) oxidiser and fuel combination is used, although the method is not limited to these substances. Foreign gas transpiration is modelled by a multi-component model. Different porous liner geometries and reservoir conditions are studied to demonstrate their effect on the performance of the thrust chamber. The numerical framework was developed to design a sub-scale transpiration cooled rocket thrust chamber ground test.

Wall Heat Loads in a Cryogenic Rocket Thrust Chamber During Thermoacoustic Instabilities

A subscale, research rocket thrust chamber operating with cryogenic oxygen and hydrogen exhibits self-excited transverse-mode instabilities with amplitudes of more than 80% of the steady combustion chamber pressure (peak-to-peak) for some operating conditions. During unstable combustion, an increase in the integral heat flux into the water-cooled combustion chamber walls of 20–40% with respect to stable conditions was experienced. A model was derived to predict changes in the axial heat flux profile considering only the dependence of flame length on the amplitude of transverse acoustic oscillations. The model predicts an increase in heat flux in the upstream part of the chamber by up to a factor of 7. This drastic increase is in agreement with past observations of rocket engine failures due to instabilities, in which the structural damage is commonly observed on the faceplate and the walls adjacent to the injection plane. The model also predicts a peak increase in integral heat flux of up to about 25%. While falling short of the peak experimental value of 40%, it nevertheless suggests that flame length is the dominant influence on the distribution of thermal loads in this study.

Early detection of thermoacoustic instabilities in a cryogenic rocket thrust chamber using combustion noise features and machine learning

Early detection of thermoacoustic instabilities in a cryogenic rocket thrust chamber using combustion noise features and machine learning

Early detection of thermoacoustic instabilities in a cryogenic rocket thrust chamber using combustion noise features and machine learning.

We present a data-driven method for the early detection of thermoacoustic instabilities. Recurrence quantification analysis is used to calculate characteristic combustion features from short-length time series of dynamic pressure sensor data. Features like recurrence rate are used to train support vector machines to detect the onset of instability a few hundred milliseconds in advance. The performance of the proposed method is investigated on experimental data from a representative LOX/H 2 research thrust chamber. In most cases, the method is able to timely predict two types of thermoacoustic instabilities on test data not used for training. The results are compared with state-of-the-art early warning indicators.

Flame response to transverse velocity excitation leading to frequency doubling and modal coupling

Transverse thermoacoustic modes may occur in gas turbines, aero-engines, and rocket engines. Various scenarios of flame excitation can be observed, depending on the type of transverse mode and the location of the flame relative to the mode shape. If an acoustically compact, symmetric flame is exposed to transverse velocity fluctuations of uniform strength and direction, the resulting modulation of the overall heat release rate is invariant to the direction of the velocity perturbation. It follows that the dominant flame response occurs at twice the forcing frequency, even for an infinitesimally small oscillation amplitude. The present study proposes a modeling framework for this inherently non-linear phenomenon, which relies on a second-order kernel of the Volterra series. As a possible realization of the Volterra series, an ad-hoc model is proposed and validated with CFD for mono-frequency forcing. Furthermore, a mechanism of modal interaction is established by which frequency doubling in the flame response causes an unstable transverse mode to drive a higher order stable mode, such that at near-resonance conditions the higher order mode exhibits elevated amplitude. This mechanism can explain the observations made by Urbano et al. (2016) in a small-scale rocket thrust chamber, where a radial mode appears at exactly twice the frequency of the dominant transverse mode. A simple representative setup of a cylindrical combustion chamber is used to explain this mechanism analytically.

Multi-Phase Impact on the Heat Load Characteristics of a Multi-Element Methane-Oxygen Rocket Thrust Chamber

Aiming at the capability to numerically model relevant operating conditions in rocket thrust chambers, this research work introduces a real gas multi-phase data extension to investigate the impact of water condensation within a widely studied sub-scale research test case. Following a dense gas approach and embedded into a state-of-the-art timescale based frozen non-adiabatic flamelet combustion model able to capture the characteristic hydrocarbon chemical reaction and recombination phenomena, the pre-computed manifold incorporates the underlying physical processes to describe the properties of the fluid mixture. The modeling framework is initially evaluated against simplified non-reacting experimental studies of water condensation along a cooled wall structure in the presence of non-condensable gases to confirm its ability to predict the phase change impact on the wall heat transfer. The good agreement with experimental heat load data is further confirmed in the frame of a conjugate heat transfer simulation of a multi-element methane-oxygen rocket thrust chamber, for which the design of the regenerative cooling system leads to localized condensation effects as the wall temperature remains below saturation conditions. The associated release of the formation enthalpy coupled with a change of the fluid properties captured by the newly developed model formulation leads to higher wall heat loads compared to the numerical predictions of an ideal gas single-phase model.

Thermal Analysis on Regenerative Cooled Liquid Propellant Rocket Thrust Chamber Using ANSYS

In this research work, thermal analysis has done on a rocket engine’s thrust chamber with four different shapes like, circular, rectangular with greater breath (b>d), rectangular with greater depth (d>b) and rectangular with greater depth with edges filleted coolant channel models to establish perfect suitable shape for the coolant channels in order to maintain the entire thrust chamber within the melting point of its materials. Finite Element Method (FEM) has adopted to establish the thermal behaviour of the coolant channel’s materials such as Nickel, Copper and Zirconium under the specific boundary conditions. Temperature distribution with respect to the nodal distances at thrust chamber’s inlet, throat and outlet portions with four different shapes of coolant channel models under Nickel, Copper and Zirconium materials were attained and compared with each other. Finite element results reveals that the successful configuration for generatively cooled rocket thrust chamber must have the coolant passage shape of rectangular cross section with greater depth (b>d).

Reacting Flow Simulations of a Dual Throat-Dual Fuel Thruster

Numerical simulation of the supersonic turbulent reacting flow field in a dual throat- dual fuel rocket thrust chamber is presented. Future single stage to orbit and high lift space transportation missions aspire a reliable, efficient, and cost-effective propulsion systems. The dual throat-dual fuel concept is a simple altitude compensating propulsion alternative with reusable possibilities. Turbulent reacting supersonic flow field emanating from independent thrust chambers needs to be resolved for a better understanding of the flow structures and design modifications for the performance improvement. The operation of a dual throat nozzle brings about a unique shock train in reacting supersonic flow. Two-dimensional axis-symmetric compressible reacting flow field has been solved using HLLC (Harten, Lax, van Leer, with Contact wave) scheme based finite volume Riemann solver with multi-step finite rate chemistry model for hydrocarbon/hydrogen-oxygen combustion. The computational procedure has been validated with experimental data for species distribution of a coaxial supersonic combustor. Chemical species distribution in the supersonic free shear layer is analyzed in detail to explore the nature of active reaction zones in the flow field.

Wall Heat Flux Evaluation in Regeneratively Cooled Rocket Thrust Chambers

In the present Paper, different wall heat flux evaluation methods for rocket engines are analyzed and compared. The test case for the comparison is a supercritical load point of an upper-stage engine. The evaluation of the heat flux using existing gradient methods shows unacceptable deviations, whereas significant improvements are found with an inverse heat transfer method. Using a Nusselt number correlation for the modeling of the cooling channel heat transfer coefficient, good agreement with the experimental profiles is achieved, while the number of installed thermocouples is held at a minimum. However, the choice of the Nusselt number correlation directly influences the obtained results and strongly depends on the fuel and geometry of the system. Using a simultaneous optimization of the heat flux and heat transfer coefficient, on the other hand, seems to eliminate the need for additional modeling and leads to great agreement with the experiment. A new optimization method proposed in this Paper based on a preprocessed Jacobi matrix significantly speeds up the estimation of the free parameters. With the obtained results, the placement of the thermocouples before the design of the hardware can be optimized, leading to reduction of manufacturing costs.

Measuring the velocity field of a shear-coaxial, cryogenic flame in a high-pressure rocket thrust chamber

High-speed imaging was used to visualize one of the transcritical flames in a multi-injector, sub-scale rocket thrust chamber at pressures up to 80 bar. Image correlation velocimetry (ICV) was applied to the imaging to obtain quantitative information on the flow field from the shear-coaxial injectors. ICV was used to track surface irregularities on the liquid oxygen (LOX) jet in imaging filtered to blue wavelengths. By choosing the interrogation area carefully, only the LOX jet was effectively tracked, excluding the coaxial H2 flow, and the time-averaged velocity field was reconstructed. Due to the transient nature of the tracked features, which frequently change shape or disappear, the averaged ICV result underestimates the absolute values of velocity. Therefore, the averaged values were scaled by the mean of the instantaneous velocity maxima. A second, reference measurement of LOX jet propagation speed was calculated using dynamic mode decomposition (DMD). The results were consistent following the aforementioned correction of the ICV values. Comparing the ICV fields for two different operating conditions showed a marked difference in the axial velocity distribution and lateral expansion of the LOX jet, demonstrating the potential of the method in studying injection in rocket combustion chambers.