Thrust Chamber Research Articles

Throttling simulation and control of liquid oxygen/liquid methane rocket engine pressurized by the electric pump

This article presented simulation and control of a liquid oxygen/liquid methane variable thrust rocket engine pressurized by the electric pump in the throttling process. To get a proper injection pressure drop, the variation of the injection area was investigated and determined. In the control scheme, the combustion pressure and mixture ratio were the targets. To simplify the control scheme, the pressure and mixture ratio were correlated to the oxygen mass flow rate and methane mass flow rate, respectively. By controlling the propellant mass flow rate, the target engine thrust was reached. In addition, the influence of mass flow rate measurement points on the control effect was analyzed. The simulation results showed that the designed controller can reduce the overshoot of the critical engine parameters. Therefore, the maximum temperature of the thrust chamber wall after the regenerative cooling was decreased. Consequently, the thrust control errors were within 3% except for the points where the thrust changed suddenly, which can meet the mission requirements.

Qualification of thermal barrier coating for semi-cryogenic engine thrust chamber through strain-controlled fatigue tests

The combustion chamber of semi-cryogenic engines experiences elevated temperatures which are beyond the capability of most structural materials. To solve this issue, an effective cooling mechanism to be implemented. In one of the developmental launch vehicles of Indian Space Research Organisation, the wall temperature in the combustion chamber is reduced by means of employing regenerative cooling along with thermal barrier coating applied on the inner wall made of special grade copper alloy. This special coating made of layers of Nickel and Nickel-Chromium also protect the thrust chamber from oxidation. The coating forms an integral part of the inner wall and are subject to a complex pattern of loads involving both thermal and pressure loads. This paper presents details about the qualification process used to evaluate integrity of the coating. Firstly, the stresses and strains on the inner wall during the engine operating conditions are arrived at based on detailed finite element analysis. Test specimens are fabricated from copper alloy and the thermal barrier coating as in the actual thrust chamber is applied on it. The specimens are tested under tensile and compressive loading that will induce the expected strains obtained from finite element analysis. The tests are done by means of strain-controlled fatigue tests at elevated temperature in tensile and compressive strain conditions. The integrity of the coating is assessed by means of Non-Destructive Test after each cycle of test and based on these, modifications were made for better performance. The modified coating was again tested for the same loading conditions and evaluated the coating integrity.

Regenerative Cooling Comparison of LOX/LCH4 and LOX/LC3H8 Rocket Engines Using the One-Dimensional Regenerative Cooling Modelling Tool ODREC

Due to the extreme temperatures inside the combustion chambers of liquid propellant rocket engines, the walls of the combustion chamber and the nozzle are cooled by either the fuel or the oxidizer in what is known as regenerative cooling. This study presents an indigenous computational tool developed for the analysis of heat transfer in regenerative cooling of such rocket engines. The developed tool incorporates a one-dimensional (1-D) combustion analysis to calculate the thermophysical properties of the combustion gas. Basic engine properties were calculated and used to generate a thrust chamber profile based on a bell-shaped nozzle. The hot gas side was analyzed using 1-D isentropic flow assumptions, along with heat transfer correlations. The coolant side was evaluated using the hydraulic analysis in the axial direction and the heat transfer analysis in the radial direction. Thermophysical properties and the phase of the coolant were determined using the given property tables and the instantaneous state of the coolant. This flexible and computationally less demanding tool was used to analyze two small-scale engines utilizing liquid hydrocarbon fuels, which are used in modern rocket propulsion. The wall cooling analyses of a liquid oxygen (LOX)/liquid methane (LCH4) engine and a liquid oxygen (LOX)/liquid propane (LC3H8) engine are presented. Fuel and oxidizer were used separately as coolants for both engines, and both of them experienced phase change. Results reveal the advantage of the high mass flow rate of the oxidizer in cooling performance. In addition, the results of this study show that the cooling of the LOX/LC3H8 engine is somewhat more challenging compared to the LOX/LCH4 engine.

Study on Two-Phase Pressure Drop of Methane during Flow Boiling in Mini Channel

For LOX/LCH4 variable thrust rocket engine, the propellant methane is traditionally selected as the coolant in regenerative cooling channel (RCC). With the decrease of engine thrust, the mass flow rate of coolant methane decreases gradually. At low engine thrust, the coolant methane is usually in a subcritical state. The heat transfer deterioration of subcritical methane occurs in RCC, which may cause thrust chamber wall ablation. The two-phase pressure drop data of methane are crucial parameters for the design and optimization of RCC. But it is rarely to find such measured frictional pressure drop data of methane in open published literature. The two-phase pressure drop of methane during flow boiling in the single mini channels with the diameters of 2.0 mm are investigated systematically. Effects of the mass flux (582.19~1755.48 kg/m2·s), inlet pressure (0.56~3.55 MPa), heat flux (53.25~318.68 kW/m2) on the frictional pressure drop of methane are discussed. The results show that the frictional pressure drop of methane during flow boiling increases with mass flux and inlet pressure at the experimental conditions, and heat flux shows weak effect on the frictional pressure drop. The comparisons of the experimental data with the predicted value by existing six correlations are analyzed. Contrary to the conventional channels, homogeneous model yields better prediction than five separated flow models. Present experimental results can provide reference for the design and optimization of RCC in LOX/LCH4 rocket engine.

Thermal-structural analysis for reusable thrust chamber using temperature-dependent CuZr alloy model adapted to cryogenic to elevated temperature

The inner liners of reusable regeneratively-cooled thrust chambers endure significant thermal and mechanical loads, resulting in doghouse deformation and even cracks during cyclic operations. To correctly reveal the failure mechanism of the thrust chamber inner liner, this study investigates the material properties for CuZr alloy and develops a temperature-dependent material model spanning from cryogenic to elevated temperatures. Firstly, we experimentally obtain the thermal and mechanical properties of the material in the temperature range from 77 K to 900 K. The experimental data reveal that the material properties vary considerably with temperature. Then, to precisely characterize the mechanical behavior of CuZr alloy under cyclic loading across a wide temperature range, a material model that combines Voce nonlinear isotropic hardening model with Chaboche nonlinear kinematic hardening model is proposed, and temperature-dependent parameters of this model are calibrated and verified. Finally, the developed material model is applied to perform the thermal-structural analysis on a reusable thrust chamber, with the aim of investigating its failure mechanism. The results demonstrate the capability of the developed model to accurately replicate the doghouse deformation of the inner liner and quantitatively capture the ratcheting effect.

A system-level multi-field coupling algorithm for regenerative cooling thrust chamber of a LOX/methane rocket engine

A systematic and hierarchical multi-physics coupling simulation method is proposed to model the combustion, combustion-gas transonic flow, coolant transcritical flow, and transient heat transfer phenomena inside the regenerative cooling thrust chamber subsystem of a LOX/methane rocket engine. By combining this algorithm with a self-developed platform, two distributed parameter modules are developed. Based on the ground test conditions of a LOX/methane engine, a verification case is established, and transient simulation studies are conducted. Comparisons with steady-state test data show that the calculated errors for coolant temperature rise and pressure drop are not higher than 4%. The simulation results reveal the locations where the transition from liquid to supercritical state occurs within the cooling jacket and provide three peril points of wall temperature on the thrust chamber. This study comprehensively reveals the occurrence of transcritical transient processes inside the cooling channels during the regenerative cooling process, laying the foundation for future entire system simulation.

NUMERICAL SIMULATIONS OF THE INFLUENCE OF VARIOUS PARAMETERS ON HEAT TRANSFER AND FLOW SEPARATION IN SUPERSONIC COOLED NOZZLES

In rocket engine nozzle design, flow separation is essential and, given the high temperatures and pressures in the thrust chamber, regenerative cooling is critical for maintaining the nozzle wall's integrity. This passage provides a summary of an in-depth numerical analysis of boundary layer separation and heat transfer within a 30°-15° cooled nozzle. The performance of the SST-V turbulence model under these conditions is assessed numerically. A variety of factors are investigated, including wall temperature, turbulent Prandtl number, and constant specific heat ratios (spanning 1.31 to 1.4 for constant fluid properties of N2O, CH4, Cl2, and air). Furthermore, variable specific heat ratios (from 1.39 to 1.66 for variable fluid properties of air, CH4, O2, and Helium) are examined, along with other parameters that affect the location of separated flow and the local wall heat transfer.

Multi-objective multidisciplinary design optimization of liquid-propellant engines thrust chamber based on a surrogate model

The design of liquid-propellant engines (LPEs) has several challenges in setting the performance parameters. Accordingly, optimizing the design of a thrust chamber is of considerable importance that has been considered in several research projects. Previous research has focused on multidisciplinary design optimization (MDO). However, despite these efforts, the main issues remain. The present paper proposes a multi-objective multidisciplinary design optimization based on an efficient adaptive surrogate model of the thrust chamber to address these issues. The proposed method introduces a practical multidisciplinary optimization method based on an adaptive surrogate model that uses the moving least squares methodology, CCM, sensitivity analysis, and the elite multi-objective genetic algorithm (NSGAII). Due to the high importance of specific impulse and thrust-to-weight ratio, these two functions were used as target functions and in the NSGAII framework, the Pareto frontier was drawn for them. The proposed method is applied to a thrust-chamber engine test case. The results show that the target performance of the engine is improved, the specific impulse value is increased by 3.4 s, and the thrust-to-weight ratio is increased by 4%. These values represent significant advances in LPE engine design. The results obtained in this study indicate the potential of the proposed method in solving large-scale thrust-chamber design optimization problems.

Influence of the thermal conductivity of different CuCr0.8 substrate state on the formability of laser directed energy deposition Inconel718 single track

One effective method for fabricating liquid rocket engine thrust chambers involves using laser directed energy deposition (LDED) to produce Cu/Ni bimetals by depositing Ni-based superalloys onto a High Strength and High Conductivity (HSHC) copper alloy. However, it is difficult to fabricate strengthening layer to the surface of HSHC Cu alloy due to its extremely high thermal conductivity. This paper investigates the influence of states of the HSHC CuCuCr0.8 substrate e.g., as-built of LPBF, rolled and annealed, and direct age after LPBF at 480 °C for 4 h, on the formability of In718 single track by LDED. The results indicate that the thermal conductivity is significantly influenced by the state of the HSHC CuCuCr0.8 substrate. At room temperature and 300 °C, the order of the thermal conductivity from low to high is as-built of LPBF, rolled and annealed, and direct age after LPBF at 480 °C for 4 h, while at 500 °C, the as-built of LPBF substrate has the lowest thermal conductivity, while the rolled and annealed substrate has the largest thermal conductivity. Thermal conductivity has a significant effect on the formability of LDED. The LPBF-ed substrate has the best formability, with a minimum Ra value of 9.72 ± 0.12 μm, a maximum deposition depth of 193.00 ± 4.32 μm, and a minimum grain size of 5.2 ± 2.6 μm because of its smallest thermal conductivity. These findings demonstrate that using an as-built LPBF-ed substrate can improve the LDED formability of the In718 single track compared with the other two substrates. This paper presents a new approach for LDED on alloy surfaces with high thermal conductivity.

Thermal Behaviour of the Cooling Jacket Belonging to a Liquid Oxygen/Liquid Methane Rocket Engine Demonstrator in the Operation Box

The cooling jackets of liquid rocket engines are composed of narrow passages surrounding the thrust chambers and ensure the reliable operation of the engine. Critical conditions may also be encountered, since the cooling jackets of cryogenic engines, such as those using LOX/LCH4 propellants, are based on a regenerative strategy, where the fuel is used as a refrigerant. Consequently, deterioration modes near where pseudocritical conditions are reached or low heat transfer coefficients where the fuel becomes a vapour and must therefore be managed. The verification of the cooling jacket behaviour to consolidate the design solutions in all the extreme points of the operating box represents a very important phase. The present paper discusses the full characterization of the HYPROB (HYdrocarbon PROpulsion test Bench Program) first unit of the final demonstrator, (DEMO-0A), by considering the working points within the limits of the operating box and comparisons with the nominal conditions are given. In this way, a full understanding of the cooling system behaviour, affecting the working of the entire thrust chamber, is accomplished. Moreover, the design strategy and choices have been confirmed, since the verifications also include potentially even more extreme conditions with respect to the nominal ones. The investigation has been numerically performed and supported the thermo-structural analyses accomplished before the final firing campaign, completed in December 2022. Since little information is available in the literature on LOX/LCH4 engines, suggestions are given as to the organization of the numerical simulations, which support the design of such rocket engine cooling systems.

Heat Transfer Analysis and Structural Optimization for Spiral Channel Regenerative Cooling Thrust Chamber

There is currently a lack of efficient heat transfer analysis methodologies for spiral channel regenerative cooling that has been increasingly applied in liquid rocket engines. To figure out the heat transfer characteristics of the spiral channel regenerative cooling thrust chamber, a simple 1D method based on the traditional semi-empirical formula after correcting the flow velocity is proposed. The accuracy of this approach is verified by the 3D numerical simulation. The verified method is further used to analyze the distribution of inner wall temperature in the test case and optimize the channel’s parameters. The research shows that the maximum inner wall temperature cooled by the spiral channel is 8.5% lower than that of the straight channel under the same channel size and boundary condition, indicating that the application of the spiral channel significantly improves the cooling effect. In addition, the 1D model combined with the second-order response surface model is employed to optimize the channel width, channel height, pitch, and inner wall thickness aiming for the best cooling effect. The calculated maximum temperature of the inner wall after the structure optimization is 586.6 K, which is 29.8% lower than the initial structure before optimization.

Upgraded design methodology for airframe/engine integrated full-waverider vehicle considering thrust chamber design

An upgraded airframe/engine integrated design approach for a full-waverider vehicle considering thrust chamber design is proposed. Compared with the existing design approach of the full-waverider vehicle, the isolator, the combustor, and the nozzle are designed complementally. First, the airframe and inlet are generated using the full-waverider theory, the isolator with the centroid offset variable cross-section is designed using the geometric morphing method, and the maximum thrust nozzle is obtained by the method of characteristics. Coupled with a circular cross-section combustor and air rudders, a complete full-waverider vehicle including the thrust chamber is generated. Second, the correctness and the availability of the design method is verified through numerical simulations under the design condition. Finally, the aerodynamic performance of the full-waverider vehicle is analyzed under a wide-speed range in the moment equilibrium state. The optimal lift-to-drag ratio of the entire vehicle is 4.03 appearing at 4° angle of attack. Owing to the rudder deflection, the optimal lift to drag ratio reduces from 4.03 to 2.44, and the corresponding angle of attack changes as well. Moreover, the inlet unstart phenomenon occurs with the decrease of the Mach number and with the increase of the angle of attack.

A Coupled Heat Transfer Calculation Strategy for Composite Cooling Liquid Rocket Engine

To better understand the characteristics of coupled heat transfer in liquid rocket engines, a calculation scheme is proposed in this paper. This scheme can simulate the coupled heat transfer processes, including combustion and flow in the thrust chamber, radiation heat transfer, heat conduction in the wall, heat transfer of coolant flow in the cooling channel, and gas film cooling in the thrust chamber wall. The numerical method used in each physical area, the data transfer method between each computing module, the strategy of data transfer on the coupling interface, the calculation process, and the convergence criterion are all introduced in detail. The calculation scheme was verified by analyzing a water-cooled nozzle. Then, the coupled heat transfer calculation was carried out for a liquid rocket engine using a propellant composed of unsymmetrical dimethylhydrazine and dinitrogen tetroxide. Two working conditions were analyzed: whether the gas film cooling was performed or not. The results showed that the algorithm successfully indicated the protective effect of the gas film on the wall surface, and the calculation results were reasonable. It played a guiding role for the coupled heat transfer of the liquid rocket engine using a composite cooling method.

Orifice Flow Dynamics in a Rocket Injector as an Excitation Source of Injector-Driven Combustion Instabilities

To investigate a hypothesis of the orifice flow-induced instability in rocket engine thrust chambers, a single liquid oxygen (LOX) injector with an optically accessible orifice module was used for experiments, with water as a simulant for LOX. The unsteady pressure downstream of the orifice was measured using high-speed piezoelectric sensors under cavitating and non-cavitating intra-injector flow conditions. The cavitating orifice flows were directly visualized via backlight imaging with a high-speed camera through the optically accessible orifice module. Cavitation initiated at the cavitation number of 2.05, and the downstream bubble cloud formation started below 1.91. The unsteady pressure spectrum arising from cavitation comprises multiple peaks over a broad frequency range, which can cause low- and high-frequency instabilities. The dominant frequencies from cavitation decrease with increasing pressure drop, while the frequencies during non-cavitating flow increase. The non-cavitating orifice flow excites the second longitudinal acoustic mode of the injector tube. The acoustic mode excited by the non-cavitating flow becomes stronger when the pressure peak in the range of whistling phenomenon is close to the first longitudinal acoustic mode. In conclusion, the excitation mechanisms of the orifice-induced instability for the cavitating and non-cavitating flows were well identified, despite the limitations of water as a simulant for LOX.

Simulation and neural network analysis of spreading the oblique fuel jet impinging onto concave walls

A series of numerical simulations of oblique-jet impinging onto concave walls are carried out to gain insight into the formation process of the liquid fuel film in a liquid film cooling thrust chamber. Moreover, the influence of jet and wall surface parameters on spreading the fuel film is investigated. The volume-of-fluid method is employed to capture the spreading process of liquid film formation after an oblique jet impacts the curved surface. Furthermore, the neural network model is developed to analyze the effects of jet diameter (0.2 mm ≤ d ≤ 0.6 mm), jet velocity (15 m/s ≤ v ≤ 20 m/s), impingement angle (15°≤β ≤ 30°), equilibrium contact angle (45°≤θe≤75°), curvature radius (22.5 mm ≤ R ≤ 67.5 mm), and surface roughness (3.2 μm ≤ Ra ≤ 16μm) on the fuel film's spreading length and maximum width based on Radial Basis Function. The results show that the liquid fuel film gradually stabilizes from upstream to downstream during spreading and the circumferential spreading of the liquid film is at its greatest when all the forces in the circumferential direction are in equilibrium. The jet diameter has the most significant effect on film spreading among jet parameters. Doubling the jet diameter at least doubles the film width and length. The effect of jet angle and velocity on film spreading is also magnified at larger jet diameters. Compared to other surface parameters, a change in the wall contact angle has the most marked effect on the spreading of the film. Lastly, the wall radius of curvature between 40 mm and 50 mm is more favorable for obtaining larger film lengths.

Metallurgical aspect of a joint made of Inconel 718 and CuCrZr brazed with Ag-Cu eutectic filler

The realization of bimetallic structures is one of the major challenges of manufacturing industry, to this aim different technologies are employed, among these vacuum brazing presents some advantages: this technology, in fact, is relatively easy to use and the relative low temperature and low process time employed, avoid the formation of brittle intermetallics and thermally affected zones. An application of vacuum brazing technology is the realization of regenerative thrust chamber for rocket engines where cooling channels of CuCrZr are joined to Inconel 718 combustion chamber. In this work the critical metallurgical aspects occurring during brazing of Inconel 718 and CuCrZr using eutectic Ag-Cu filler are described in detail. These aspects include the formation of two distinct metallurgical zones inside the filler and an interdiffusion zone with Kirkendall porosity at the interface Inconel 718-Brazing Alloy. The analysis of fracture surface after tensile tests indicates that the interdiffusion zone is the weak point of the joint.

Transient study on filling characteristics of LOX dome

The 3D flow filling process of LOX (Liquid Oxygen) dome is studied in a liquid oxygen/methane rocket engine. The volume, flow field pressure and velocity distributions of liquid oxygen at different time are determined. The flow property of propellant and the injection characteristic time of liquid oxygen are obtained. The results show that when the LOX dome has not been filled, the LOX flow state is disorder, and flows around the annular wall in the middle dome. The static pressure near the equalizing orifice near the liquid oxygen inlet is high and the velocity is slow. When the LOX filling is completed, the pressure at the outlet of each equalizing orifice is the same. The liquid oxygen flow is in a stable state, and the flow rate decreases. The overall velocity distribution is more uniform. In addition, in the filling process of LOX dome, the time of arrival in each area is different. There are different injection delays in each area on the injection panel, and the time of arrival of liquid oxygen in the opposite area of the inlet section of LOX dome is the latest. This result leads to the inhomogeneity of injection characteristic time and flow rate in thrust chamber.

A Simplified Thermal Analysis Model for Regeneratively Cooled Rocket Engine Thrust Chambers and Its Calibration with Experimental Data

An essential part of the design of a liquid rocket engine is the thermal analysis of the thrust chamber, which is a component whose operative life is limited by the maximum allowable wall temperature and heat flux. A simplified steady-state thermal analysis model for regeneratively cooled rocket engine thrust chambers is presented. The model is based on semi-empirical correlations for the hot-gas and coolant convective heat transfer and on an original multi-zone approach for the wall conduction. The hot-gas heat transfer is calibrated with experimental data taken from an additively manufactured water-cooled nozzle that is connected to a combustion chamber either fed with decomposed hydrogen peroxide or decomposed hydrogen peroxide and automotive diesel. The thrust chamber (i.e., combustion chamber and nozzle) is designed to produce about 450 N of thrust when operating with a chamber pressure of 11 bar. For this application, the calibrated model predicts the total wall heat transfer rate very accurately and the temperature distribution within the wall structure with an uncertainty of a few tens of kelvins. This level of accuracy can be considered more than adequate for the design, and generally for engineering-type thermal analysis, of similar thrust chambers.

An investigation on the cyclic deformation and service life of a reusable liquid rocket engine thrust chamber wall

PurposeRegeneratively cooled thrust chamber is a key component of reusable liquid rocket engines. Subjected to cyclic thermal-mechanical loadings, its failure can seriously affect the service life of engines. QCr0.8 copper alloy is widely used in thrust chamber walls due to its excellent thermal conductivity, and its mechanical and fatigue properties are essential for the evaluation of thrust chamber life. This paper contributes to the understanding of the damage mechanism and material selection of regeneratively cooled thrust chambers for reusable liquid rocket engines.Design/methodology/approachIn this paper, tensile and low-cycle fatigue (LCF) tests were conducted for QCr0.8 alloy, and a Chaboche combined hardening model was established to describe the elastic-plastic behavior of QCr0.8 at different temperatures and strain levels. In addition, an LCF life prediction model was established based on the Manson–Coffin formula. The reliability and accuracy of models were then verified by simulations in ABAQUS. Finally, the service life was evaluated for a regenerative cooling thrust chamber, under the condition of cyclic startup and shutdown.FindingsIn this paper, a Chaboche combined hardening model was established to describe the elastoplastic behavior of QCr0.8 alloy at different temperatures and strain levels through LCF experiments. The parameters of the fitted Chaboche model were simulated in ABAQUS, and the simulation results were compared with the experimental results. The results show that the model has high reliability and accuracy in characterizing the viscoplastic behavior of QCr0.8 alloy.Originality/value(1)The parameters of a Chaboche combined hardening constitutive model and LCF life equation were optimized by tensile and strain-controlled fatigue tests of QCr0.8 copper alloy. (2) Based on the Manson–Coffin formula, the reliability and accuracy of constitutive model were then verified by simulations in ABAQUS. (3)Thermal-mechanical analysis was carried out for regeneratively cooled thrust chamber wall of a reusable liquid rocket engine, and the service life considering LCF, creep and ratcheting damage was analyzed.

QCr0.8 Cu alloy /S06 stainless steel bimetal structure via In718 multi-interlayer fabricated by laser powder hybrid additive manufacturing

Fabricating high strength high conductivity (HSHC) Cu alloy inner parts by laser powder bed fusion (LPBF) and high-strength steel out layers on Cu alloy surfaces by laser-direct energy deposition (LDED), e.g., laser powder hybrid additive manufacturing (LPH-AM), has shown promise in the development of rocket engine thrust chambers with large thrust forces. In this study, we successfully fabricated a QCr0.8 HSHC Cu alloy/S06 stainless steel bimetal structure via In718 multi-interlayer using LPH-AM. We investigated the relative density, defects, microstructure, and interfacial characteristics of the bimetal structure in detail. The results indicate that the sample achieved up to 99.9% relative density with no naked defects and good metallurgical bonding at the QCr0.8/In718 and In718/S06 interfaces. The LPH-AM-ed QCr0.8/In718 interface exhibited a combination of columnar and equiaxed dendrites. The vertically combined sample exhibited an average ultimate tensile strength (UTS) and break elongation (EL) of 300.3 ± 10.6 MPa and 15.0 ± 1.4%, respectively, with fracture occurring at the QCr0.8 side. This study provides a new method for fabricating complex-shaped Cu/steel bimetal components, such as rocket engine thrust chambers.