Thrust Chamber Research Articles

Array of carbon black-based microthrusters for CubeSat applications

Access to space for small private companies requires to improve the ability to bring low-cost satellites into orbit. CubeSats offer a unique opportunity to meet these needs thanks to their reduced production times, the low manufacturing costs and ease of use. In order to be able to communicate with each other, exchange information and interact, it is necessary to place CubeSats in formation: in this context, miniature propulsion technologies, including chemical and electric propulsion, play a critical role in achieving mission requirements and maintaining satellites position. In this article, the feasibility of solid propellant micro rockets, fully integrated in an opposing array of printed thrust chambers is examined: each rocket can be fired together with the others or separately to modulate thrust. Theoretical and experimental results show that the microthruster, made of nylon and carbon fiber, have good mechanical and thermal resistance and simultaneously good performance is achieved. In particular, a microthruster with a diameter of 4 mm and a length of 6 mm, with 55 g of black powder propellant, achieves a thrust of about 3.5 N for about 7 ms.

Numerical Investigation on the Service Life of a Liquid Rocket Engine Thrust Chamber

Rocket engine thrust chambers withstand very high temperatures and thermal gradients during service that induce multiple damaging phenomena such as plasticity, low-cycle-fatigue (LCF) and creep. Numerical models can be used during the design of these mechanical components in order to simulate the main mechanical damaging processes, accounting for complex material behavior as due to non-linear hardening phenomena and viscoplasticity. This work represents an improvement upon previous research by the authors, with particular reference to the addition of the Wang–Brown fatigue criterion, to consider the effects of multiaxiality and non-proportionality of loads, and the Voce model to account for non-linear isotropic hardening. A precipitation hardened copper alloy has been considered as the material of the thrust chamber internal structure. The most critical areas resulted to be on the internal surface of the chamber and in particular in correspondence of the throat region, consistently with experimental tests available in the literature conducted on similar geometries. Results demonstrated that low cycle fatigue and thermal ratcheting (plastic instability) are the dominant damaging phenomena for the considered test case.

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.

The Influence of Thrust Chamber Structure Parameters on Regenerative Cooling Effect with Hydrogen Peroxide as Coolant in Liquid Rocket Engines

Liquid rocket engines with hydrogen peroxide and kerosene have the advantages of high density specific impulse, high reliability, and no ignition system. At present, the cooling problem of hydrogen peroxide engines, especially with regenerative cooling, has been little explored. In this study, a realizable k-epsilon turbulence model, discrete phase model, eddy dissipation concept model, and 10-step 10-component reaction mechanism of kerosene with oxygen are used. The increased rib height of the regenerative cooling channel causes the inner wall temperature of the engine increases, the average temperature of the coolant outlet decreases slightly, and the coolant pressure decreases. The overall wall temperature decreases as the rib width of the regenerative cooling channel increases. However, in the nozzle throat area, the wall temperature increases, the average coolant outlet temperature decreases, and the coolant pressure drop increases. A decrease in the inner wall thickness of the regenerative cooling channel results in a significant decrease in the wall temperature and a small increase in the average coolant outlet temperature. These findings contribute to the further development of the engine with hydrogen peroxide and can guide the design of its regenerative cooling process.

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.

Dynamic Characteristics of LOX/Kerosene Variable Thrust Liquid Rocket Engine Test System Based on General Modular Simulation Method

For the research demand of reusable LOX/kerosene variable thrust liquid rocket engine, a test system with electric displacement pumps is designed and a multidisciplinary modular dynamic simulation method based on AMESim platform is used to analyze the system. The method comprehensively considers the characteristics of complex components in the engine and realizes the fast module assembly and variable step size solution. Considering the combustion model of thrust chamber, the positive displacement pump model with complex leakage channels, and the cooling jacket heat transfer model, the component dynamic equations are deduced and the final model simulation results reveal that the system has a smooth ignition, stage turning, and shutdown process. The thrust can reach 6900 N in high working condition and the variable thrust ratio is 5 : 1. The dynamic characteristics of the system show that the performance error of main pump components is less than 5%, the maximum average temperature rise of the thrust chamber coolant is about 28°C, and the time of stage adjustment is within 300 ms, which mean the overall design of the system is reasonable. Although the accumulation of LOX before kerosene injection can adversely affect the temperature of the thrust chamber, large pressure pulses do not occur due to the ignitor’s duty flame. Moreover, the pintle injectors based on PID control can effectively stabilize the pressure drop at lower conditions. The system and the simulation method provide important support for the actual engine test and the general LRE dynamic characteristics analysis.

Numerical study of combustion and cooling performance of a gaseous oxygen and gaseous methane rocket combustor with the needle-bolt injector

The needle-bolt injector is a convenient method for adjusting the propellant combustion and has attracted a great deal of attention in recent years. However, the effects of injector parameters on the flow and temperature fields in the combustion chamber are rarely investigated. In this study, a new computational fluid dynamic rocket combustor model based on needle-bolt injector is established to study the combustion performance of the gaseous oxygen and gaseous methane (GOx/GCH4) rocket. Three key parameters of needle-bolt injector including the radial fuel injection hole diameter Df, the needle valve diameter Dn and the length of needle into the combustion chamber L are studied. In addition, a special coolant hole for liquid coolant cooling is proposed to reduce the high temperature gathered in thrust chamber. Compared with the chamber wall temperature of 1807 K without cooling, the wall temperature can be reduced to 1467 K after introducing 15% of water coolant. However, with a further increase in coolant ratio from 25% to 30%, the cooling effect is not obvious. The present numerical study can provide important guidance for studying combustion characteristics and cooling performance of rocket combustor with the needle-bolt injector.

Life Analysis of Reusable Liquid Rocket Engine Thrust Chamber

The thrust chamber’s inner wall suffers high temperature and pressure differences from the coolant channel, which limits the life of the rocket engine. Life prediction of the thrust chamber really plays an important role in reusable launch vehicle propulsion systems. The Porowski beam model is widely used in the life prediction of reusable liquid rocket engine thrust chambers, which calculates the life caused by fatigue, creep, and thinning after each firing cycle. In order to analyze the life of the thrust chamber, a LOX/Kerosene rocket engine is investigated in this paper. The life analysis consists of pressure and temperature differences and structural parameters. Two kinds of inner wall materials were chosen for comparison in this research: OFHC copper and Narloy-Z alloy. The results are presented to offer a reference for the design and manufacture of reusable rocket engine thrust chambers in the future.

Chemical bond formed between electrodeposits and substrate at room temperature with a novel in situ electrochemical treatment

The interfacial adhesion strength between electrodeposits and substrate is crucial in many applications, such as the thrust chamber of hydrogen-oxygen rocket engines. To achieve excellent adhesion, herein we propose a novel in situ electrochemical treatment that promotes formation of chemical bonds at the electrodeposit/substrate interface. This treatment integrates the electrochemical dissolution, substrate transport between baths and electrodeposition in a continuous homologous-acid system, thereby achieving and maintaining the activated interface. As a result of the Ni electrodeposition on the Cu substrate, Cu-Ni bonds forms between the deposits and the substrate at the nanoscale interface roughness (<200 nm), realizing the ultrahigh adhesive strength of 386 MPa. The detected diffraction peak intensities of Cu–Ni bonds, initial electrodeposition potential, and deposit–substrate adhesive strength show that the smooth interfacial micromorphology and the continuous homologous-acid environment are the keys to the formation of the Cu–Ni bonds. Qualitative models are built to further analyze and understand chemical bond formation during electrodeposition.

Dynamic Numerical Simulation of Hybrid Rocket Motor with HTPB-Based Fuel with 58% Aluminum Additives

The addition of aluminum (Al) to the fuel is an effective way to increase the regression rate of hybrid rocket motors (HRMs). Due to its high regression rate, the impact of the regression of combustion surface on the performance of HRMs cannot be ignored. Therefore, it is significant to establish a dynamic numerical simulation model to predict the performance of HRMs. In this study, the dynamic simulation model was established based on dynamic mesh technology and was verified by a firing test. The results show that the simulation results agree well with the experimental results, and the errors of the average thrust and combustion chamber pressure are 3.4% and 1.4%, respectively. The dynamic simulation shows that with the regression of the combustion surface, the vortex of the pre-combustion chamber is divided into two vortices. The vortex near the front of the grain will increase the regression rate downstream. The results show that the addition of Al can obviously improve the regression rate of HRMs. The fuel containing 58% Al can improve the regression rate by 88.8% compared with the fuel with pure hydroxyl-terminated polybutadiene (HTPB). Moreover, due to the higher combustion temperature and the scouring of metal particles, the ablation rate of the nozzle with carbon ceramic materials reaches 0.16 mm/s. This investigation provides a valuable reference for HRMs design and simulation.

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.

The dynamic model of the high-thrust orbital maneuvering engine with storable propellant

The storable propellant rocket engine has the advantages of long-term storage and rapid use. A new design of a large-thrust orbital maneuvering engine is proposed for the first time and the dynamic model for working process is established by the multi-disciplinary dynamics simulation software AMEsim. The accurate model parameters of hydraulic and combustion components are obtained through actual unit tests. The dynamic characteristics of the engine are calculated and evaluated in detail, and the results show that the system has good robustness to disturbance parameters. The fluctuation of the propellant tank pressure has little influence on operating parameters. The peak pressure of the water hammer, especially at the tank outlet pipe, is about 17 times higher than the stable state. The starting sequence influences working conditions of the mixing ratio regulator and the thrust chamber for less than 0.1s, and has no influence on the stable working state of the engine. The research provides an effective method for designing a high-thrust orbital maneuvering engine system, which has broad prospects in low-Earth orbit transportation and military missions.

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.

Rotating Detonation Combustion for Advanced Liquid Propellant Space Engines

Rotating (also termed continuous spin) detonation technology is gaining interest in the global research and development community due to the potential for increased performance. Potential performance benefits, thrust chamber design, and thrust chamber cooling loads are analyzed for propellant applications using liquid oxygen or high-concentration hydrogen peroxide oxidizers with kerosene, hydrogen, and methane fuels. Performance results based on a lumped parameter treatment show that theoretical specific impulse gains of 3–14% are achievable with the highest benefit coming from hydrogen-fueled systems. Assessment of thrust chamber designs for notional space missions shows that both thrust chamber length and diameter benefits are achievable given the tiny annular chamber volume associated with the rotating detonation combustion. While the passing detonation front drastically increases local heat fluxes, global energy balances can be achieved if operating pressures are limited to be comparable to existing or prior space engines.

Integrated design of solar thermal propulsion system for microsatellite with liquid ammonia as propellant

This paper presents a design of solar thermal propulsion (STP) system for microsatellite with liquid ammonia as propellant. The system was equipped with two concentrators, which were respectively placed in the tank and thrust chamber for propellant supply and heating. A platelet heat exchanger was adopted to heat the propellant in the chamber, and the fluid–solid coupling effect between the wall and the gas was considered. Meanwhile, the effects of satellite mass, initial orbit, nozzle size and target temperature on the performance of STP system were analyzed. The results show that for microsatellites with a total mass of 100 kg, the STP system can fully heat the propellant to more than 2050 K, generate an intermittent thrust of about 26 N, and enable the satellite to obtain a velocity increment of more than 1470 m/s within 19 days, consuming only 42 kg of propellant, which can directly meet the transfer mission from the geostationary transfer orbit (GTO) to the geostationary orbit (GEO). The maximum velocity increment could reach more than 1950 m/s when the propellant was completely consumed; Changing the mass and initial orbit of the satellite will not affect the thrust and specific impulse. Satellites with smaller mass will spend less time and propellant during orbit transfer. The lower is the perigee height of the initial orbit, the greater is the propellant consumption, while the shorter is the time of orbit transfer; The reduction of nozzle throat size and target temperature will lead to the increase of specific impulse and the decrease of orbital transfer time, but the reduction of thrust.

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.

A review of vaporization models as design criterion for bipropellant thrust chambers

In the beginning of liquid propellant rocket engine development, the thrust chamber sizes were obtained, mainly, empirically. With the technological advancements over the years, several approaches have been developed in order to optimize its sizes and predict more accurately the performance. Besides the clear contribution in predicting efficiencies, the use of accurate vaporization models to optimize combustion chambers decreases losses and the number of required tests. To increase efficiencies, the chamber must be optimized. In case the chamber is too small, incomplete combustion is achieved and combustion instability may occur. In case the chamber is too large, losses due to weight and heat transfer increase and the vehicle becomes larger (leading to more drag losses). Additionally, the number of tests is reduced since models were experimentally validated and less experimental iterations are required in order to obtain the optimized design. Although there are many models, all of them reach similar conclusions, such as an increase in chamber pressure, a decrease in injected droplet size and velocity, and others, lead to a decrease in the required chamber size. Nowadays, with the advancements in computing budget, more complex and accurate models have be developed. Some of these models account for chemical reactions, turbulence effects, droplet collisions and interactions, two- and three-dimensional modeling, and others. Also, the use of CFD codes provides relevant contributions to the analytical and numerical models, especially in validating them, and, additionally, decreases the amount of required experimental tests. The main propulsive parameter that rules this phenomenon is the characteristic length, which accounts the required chamber size for the propellants to be injected, atomized, vaporized, mixed and combusted. Most of the available models neglect the atomization, mixing and combustion of the propellant, since those phenomena occur much faster compared with the vaporization. This work provides a review of those vaporization models, focusing on the main used models worldwide. Such review is of great importance in order to supply enough information and comparison between models, making possible for the researcher/engineer to choose the model that better fit its necessities, requirements and limitations.

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.

Thermostructural Numerical Analysis of the Thrust Chamber of a Liquid Propellant Rocket Engine.

The numerical simulation of rocket engine thrust chambers is very challenging as several damaging phenomena, such as plasticity, low-cycle-fatigue (LCF) and creep occur during its service life. The possibility of simulating the thermostructural behavior of the engine, by means of non-linear finite element analyses, allows the engineers to guarantee the structural safety of the structure. This document reports the numerical simulations developed with the aim of predicting the thermostructural behaviour and the service life of the thrust chamber of a liquid-propellant rocket engine. The work represents a step ahead of previous researches by the authors, with particular reference to the addition of the Smith-Watson-Topper (SWT) fatigue criterion, and to the implementation of a sub-modelling technique, for a more accurate assessment of the most critical section of the component. It was found that the equivalent plastic strains in the most critical nodes obtained through the sub-modelling technique were about 20% lower than those calculated without sub-modelling. Consistently with experimental tests from literature conducted on similar geometries, the most critical areas resulted to be on the internal surface of the chamber. The analyses demonstrated that the LCF damaging contribution was significant, with a life prediction for the thrust chamber of about 3400 cycles.

Research on thermomechanical fatigue life prediction of regenerative cooling thrust chamber

The ratcheting effect of the liner wall structure in the thrust chamber of liquid rocket engine under cyclic load is studied, and its service life is predicted. Firstly, the 3-D heat transfer analysis of the wall of the thrust chamber is carried out while explaining the mechanism of its "dog room" failure. Based on the results of the heat transfer analysis, the stress-strain evolution of the structure under cyclic load is analyzed by nonlinear plane strain finite element method. According to the finite element analysis results, the ratcheting effect of the liner wall structure is described and the ratcheting damage is calculated. The low cycle fatigue damage is calculated by using modified Morrow model, and the creep damage is also calculated using Norton model. Finally, the total damage of thrust chamber liner wall is calculated by using Miner linear cumulative damage law and the thermomechanical fatigue life of the structure is estimated. The result shows that the middle point of the lower surface of the liner wall will fail first, and the calculated service life is 33 times, which is good agreement with its actual service life. In the total damage of the weak point, ratcheting damage, creep damage and low cycle fatigue damage account for 52%, 32% and 16% respectively, indicating that the ratcheting effect is the main reason for the failure of the thrust chamber structure. The research results of this paper provide important engineering reference for the structural optimization design and rapid life prediction of regenerative cooling thrust chamber of liquid rocket engine and the future reusable aerospace power.