Thrust Chamber Research Articles

Size effects and optimization during laser directed energy deposition on high thermal conductivity copper alloys

High-copper alloy/nickel-based high-temperature alloy bimetallic structures are widely utilized in critical components, such as thrust chambers of liquid rocket engines, water cooling circuits for fusion reactor deflectors, and electromagnetic railguns. Among various methods for fabricating Cu/Ni bimetallic structures, laser directed energy deposition (LDED) is considered one of the most promising techniques for depositing nickel-based high-temperature alloys onto high-copper alloy surfaces. In the LDED fabrication process, substrates with uneven thicknesses are a common. The process is characterized by its multi-parameter, high-sensitivity, and non-equilibrium solidification characteristics. Consequently, variations in substrate size can significantly affect molten pool morphology and fabrication stability. This study examines the size effect of a single track of LDED Inconel 718 (In718) on a CuCr0.8 high-copper alloy substrate. The results indicate a pronounced size effect when using a high thermal conductivity CuCr0.8 substrate, particularly regarding substrate thickness. By reducing the thermal conductivity of the CuCr0.8 substrate, the size effect is mitigated, improving the process adaptability of LDED. These findings provide valuable data for the laser processing of surfaces with uneven thickness and high thermal conductivity. They are expected to support the fabrication of variable thrust liquid rocket engines and advance the field of interplanetary exploration.

Effect of aging temperature on the redrawn cup wall of a novel Cu-0.5Cr-0.05Zr-0.05Ti (wt.%) alloy sheet: Analyses on microstructure evolution, mechanical properties and strengthening mechanism

A solution-treated Cu-0.5Cr-0.05Zr-0.05Ti (wt.%) alloy was drawn into a large-depth cup implementing a two-stage deformation process in the present study, firstly using a ∅50 mm cylindrical punch and subsequently redrawn by a ∅34 mm punch. The redrawn cup wall was subjected to aging heat treatment up to 400 °C, 500 °C, and 600 °C for 2 h followed by water quenching to understand the microstructural evolution and its correlation with mechanical properties. These aging temperatures were adopted to simulate a similar environmental condition in which the inner liner of the thrust chamber of cryogenic and semi-cryogenic engines was exposed during satellite launching. It was identified that the above aging temperatures influenced the grain characteristics, recrystallization, and annealing twin formation significantly. The redrawn cup wall possessed semi-coherent bcc spherical Cr-rich precipitates with an average size of 1.8 ± 0.1 nm, which progressively increased to 5.5 ± 0.3 nm at 600 °C, while the precipitate volume fraction increased to a maximum of 1.2% at 500 °C. The precipitation mostly influenced the strength, while the recrystallization and twin formation governed the ductility during the aging. After an initial decrement at 400 °C, the yield strength of the aged specimen was increased to a maximum of 193 MPa with a concurrent increase in ductility to 24% upon the optimal aging at 500 °C. However, a higher increment in the strength than the ductility indicated the dominance of the precipitation. Further increase in the temperature to 600 °C reduced the strength while considerably enhancing the ductility. This was attributed to the highest recrystallization and twin fraction, as well as precipitate coarsening with reduced volume fraction. Moreover, precipitation strengthening prevailed in the aged specimens, unlike dislocation strengthening in the redrawn condition. The contribution from the former was maximum at 500 °C with a 52% share of the total estimated strength, whereas it was only 31% in the redrawn specimen.

Influence of laser absorptivity of CuCr0.8 substrate surface state on the characteristics of laser directed energy deposition inconel 718 single track

An effective method for fabricating copper-nickel bimetallic liquid rocket engine thrust chambers involves utilizing laser directed energy deposition (LDED) technology. However, the state of the substrate surface significantly impacts the LDED process. This study investigates the effects of various substrate treatments on LDED single tracks, using CuCr0.8 high-copper alloy as the substrate and Inconel 718 as the deposition material. The treatments include polishing, sandblasting, laser etching, and cold spraying. Substrate surface roughness, laser absorptivity, molten pool morphology, and microstructure were characterized, and the mechanisms of laser absorptivity change and the different LDED processes were analyzed. The results indicate that the laser-etched surface exhibits the worst surface roughness (Ra 15.20 ± 0.60 μm), the highest laser absorptivity(80.70% at 1080 nm wavelength), the largest deposition width (947.33 ± 29.85 μm), and the maximum number of fine grains among the four substrates. Additionally, the cold-sprayed surface shows the largest deposition depth (237.33 ± 39.04 μm), the minimum number of fine grains and a higher laser absorptivity (66.20% at 1080 nm wavelength). In situ observations of molten pool formation and flow during LDED was conducted using an in situ high-speed high-resolution imaging system. The mechanisms underlying the alteration in laser absorptivity primarily involve the "trapped light" effect and modifications to the surface material. This research is significant as it provides foundational insights for laser processing of highly reflective materials, offering important theoretical and practical implications for engineering applications.

Obtaining in-situ reacted Fe–W–Ni–Cr–Cu high-entropy alloy within the interlayer of dissimilar Cf/SiC-GH3536 joint by designing non-high-entropy Cu–Ti–W composite filler

In aerospace applications, the development of high-entropy alloys (HEAs) filler is an effective strategy to obtain reliable Cf/SiC nozzles and GH3536 thrust chambers components. Considering the high costs and extended preparation times with HEA fillers, utilizing non-HEA fillers to generate HEAs within brazing seam offers significant advantages. This study successfully joined Cf/SiC with GH3536 using a non-HEA Cu–Ti–W composite filler that facilitated the in-situ formation of Fe–W–Ni–Cr–Cu HEA during the brazing process. Furthermore, the formation and accompanying martensite phase transformation were revealed. The microstructure characteristic of W reinforcements/HEA/Cu(s,s) was ultimately formed in the brazing seam. The atomic structure of the HEA, confirmed to be hexagonal close-packed, was elucidated using spherical aberration-corrected transmission electron microscopy. Additionally, a martensite phase transformation was observed in the HEA, involving two adjacent layers of (0 0 0 1) atomic planes that sheared and move a3 distance alon g [10–10] in opposite directions, resulting in the formation of M-HEA. The inclusion of HEA and M-HEA in the Cf/SiC-GH3536 joint, brazed with Cu–Ti–W composite filler, increased its shear strength to 84.8 MPa, which was 2.15 times higher than that of joints without the W reinforcements. This study offers new insights into the design of composite fillers and the application of HEAs.

Thermal recycling analysis in regenerative cooling channels based on liquid rocket engine cycles

For rocket cycles involving electric pumps, gas generators, expanders, and staged combustion, a thermal recycling method is proposed to provide a more realistic simulation of regenerative cooling systems in liquid rocket engines. This method simulates a real situation where the outlet of the cooling channel is thermally recycled through the turbine or preburner to the inlet of the combustion chamber. Supercritical fluid properties are determined using the extended Redlich–Kwong (RK)–Peng–Robinson (PR) real-fluid equation of state. The axisymmetric reacting flows of a thrust chamber with regenerative cooling channels are simulated using the flamelet-based lookup table. To account for the multi-injector effect in the two-dimensional axisymmetric simulation, a non-uniform velocity distribution is implemented, utilizing exponential distributions of each injector’s mass flow rate. For the hot gas temperature, coolant temperature, and cooling mass flowrate, the thermal recycling method is comparatively analyzed with the thermal decoupling method. The regeneration effect of the heated fuel is explored by evaluating the inflow energy, reaction energy, wall heat transfer, and the exit kinetic energy conversion. The thermal recycling method is developed for regeneratively cooled rocket engines with the expander, gas generator, staged combustion, and electric-pump cycles. Through this numerical procedure, the thermal recycling method is successfully applied to a liquid oxygen/methane engine and NASA’s Crew Exploration Vehicle nozzle with two types of cooling. Based on the results of the four types of rocket cycles, it is confirmed that the specific impulse increases by 1.5–2% due to the regenerative heat effect.

Study on the influence of hydrogen injector flow rate in hydrogen-oxygen engine thrust chamber

Study on the influence of hydrogen injector flow rate in hydrogen-oxygen engine thrust chamber

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.

Simulations With Compressible Multiphase Formulation on Heat Transfer Studies in a Rocket Thrust Chamber With Experimental Validation

Abstract In the combustor for rocket engines, liquid film cooling is a widely adopted technique for restricting the structural components to the acceptable bounds for the successful operation of the propulsion systems. Multiple cooling techniques for thrust chamber walls are favored along with the coolant film in propulsion systems operating at higher chamber pressures by incorporating an ablative cooling mechanism for the nozzle. The adoption of combination will result in challenges in simulation studies due to the presence of multiphase phenomenon in the system. The current article presents the effects of these two cooling methods on a thrust chamber wall studied through compressible multiphase formulations. A majority of the previous studies have used incompressible flow equations to model coolant film behavior and associated heat transfer. The present study utilizes an approach utilizing compressible multiphase flow to simulate the behavior of the compressible hot gas flow and the incompressible coolant liquid film accounting for phase change effects, vapor diffusion into hot gas, radiation effects on coolant surface, liquid entrainment, and blowing effects due to vapor formation at interface. Film-cooling effects on ablative nozzle accounting for pyrolysis phenomenon due to material decomposition governed by Arrhenius expression are also emphasized. The outcome of the numerical model showed good agreement with available test data in literature, and results from the tests done in-house. The model described in the present study was able to support the actual evidence of liquid film cooling being able to preserve the walls of the thrust chamber from severe internal thermal environment and prevent ablation on surface of nozzle.

Laser ultrasonic testing of defects in milling groove brazed joints of thrust chamber

This paper presented a new approach for noncontact inspection of defects in milling groove brazed joints of thrust chamber with laser ultrasonic testing method and synthetic aperture focusing technique (SAFT). Firstly, laser ultrasonic testing methods for milling groove brazed joints of thrust chamber was studied. Subsequently, numerical models were constructed to analyze the influence of defects on laser-excited signals. The analysis revealed that the brazed defects caused internal waves to reflect on the rib surface, manifesting as defect echo signals preceding the outer wall echo. Through scanning setting, the obtained SAFT images illustrate the presence of the defect directly and clearly. Furthermore, an experimental system was established to detect and image artificial defects with different degrees of weld leakage. The experimental results are consistent with simulation results, validating the capability and effectiveness of the testing and imaging method. In general, the proposed laser ultrasound method offers inherent advantages of non-contact detection with high resolution and precision, and it is easy to achieve fast and automated scanning of large and complex structures like thrust chambers, demonstrating its potential for enhancing the safety and reliability of liquid rocket engines.

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.

A novel coupled algorithm for Studying the performance characteristics of water ramjet in the flight of supercavitating vehicle

When a supercavitating vehicle performs a flight, the water ramjet of the vehicle exploits the water in the surrounding environment as an oxidiser and is inevitably affected by the vehicle motions. In this study, the characteristics of the performance change in a water ramjet were investigated when a supercavitating vehicle performs a flight to gain more insights into the design of underwater power systems. Given the principle of independent cross-section expansion of cavities, a longitudinal ventilated supercavity model was built to predict the development of a ventilated cavity and its interaction with the body. A thermal calculation model was developed based on the minimum free energy method to calculate the performance parameters of water ramjets. A novel coupling algorithm was proposed based on the aforementioned model to integrate the motion of a supercavitating vehicle with the operation of a water ramjet. The integration described above was achieved by linking the water intake and body motion and calculating the performance change characteristics of the water ramjet during the flight of the supercavitating vehicle. The operating characteristics of the water ramjet with changes in the designed water-fuel ratios were investigated to provide a basis for exploring the unperturbed flight of the vehicle, depth regulation, and underwater environmental changes. The thrust of the water ramjet is less affected by external factors, and its performance feedback tends to exacerbate external disturbances at a designed water-fuel ratio lower than the optimal water-fuel ratio. A water ramjet with a designed water-fuel ratio higher than the optimal water-fuel ratio has a rapid response time, adaptive thrust, and stable combustion chamber pressure. External disturbances slightly affected the thrust and combustion chamber pressure of the water ramjet when the designed water-fuel ratio was equal to the optimal water-fuel ratio. The results of this study provide technical support for optimising the design of water ramjet.

Application of semiempirical correlations for multidimensional heat transfer in rocket engine cooling channels

Heat transfer in cooling channels of liquid rocket engine thrust chambers is an intrinsically multidimensional problem due to the geometry of the channels, often rectangular in section, and the asymmetric distribution of heat input from the hot combustion gas. In this study, the possibility of addressing this problem with semiempirical correlations for the coolant heat transfer and multidimensional numerical simulations for the heat conduction in the wall is investigated. Using experimental data from a test campaign of a cooling channel representative of liquid rocket engine conditions, semiempirical correlations are derived. Both cases of constant and variable heat transfer coefficient along the perimeter of the channel section are considered. The resulting correlations are then used to evaluate their reliability in reproducing the experimental wall temperature data from the same test campaign. For comparison, a correlation taken from the open literature and obtained with experimental data of circular heated tubes is also considered. The results demonstrate that correlations based on a constant heat transfer coefficient fail to reproduce with sufficient accuracy the test cases characterized by higher wall temperatures, while correlations based on a variable heat transfer coefficient overestimate the wall temperature in the initial sections of the channel.

Forming limit and failure characterization of as-received and bi-axial prestrained Cu-Cr-Zr-Ti alloy sheets in the context of multistage forming

The characterization of forming limit utilizing Cu-Cr-Zr-Ti alloy sheets during the multistage forming process is imperative for successfully fabricating thrust chamber liners used in satellite launch vehicles. Therefore, the present work studies the effect of prestrain on the forming limit and failure behaviour of this alloy under different deformation paths. A stretch-forming setup was utilized to induce 8% equi-biaxial prestrain (EBP), and subsequently, these prestrained specimens were deformed till the onset of failure using the same setup. After prestraining, a shift in the failure strains was observed in the principal strain space, whereas it was found to be absent in the polar effective plastic strain, principal stress and effective plastic strain-stress triaxiality loci. The cup height along different deformation paths was marginally decreased in the prestrained specimens because of the reduction in failure strains, and the fracture occurred in the prestrained region. Further, failure analysis revealed intergranular cracks in the specimens deformed along plane strain and uniaxial deformation paths due to the presence of micron-size Cr-rich precipitates along grain boundaries. However, intergranular and transgranular cracks were present in the prestrained specimens, which were deformed equi-biaxially similar to the as-received specimens. The presence of nano-size Cr-rich precipitates inside the grains acted as nucleation sites for voids during equi-biaxial stretching, leading to transgranular cracks. It was also observed that failure in the prestrained and as-received specimens occurred at a similar effective plastic strain value due to the combined effect of void density and void size. Based on experimental observations, it was inferred that the forming limit of Cu-Cr-Zr-Ti sheets under a two-stage forming process was primarily affected by the presence of micro and nano-size Cr-rich precipitates.

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.

Effects of Different Structural Film Cooling on Cooling Performance in a GO2/GH2 Subscale Thrust Chamber

To investigate the wall cooling of the thrust chamber in an engine, two film-cooling structures, namely, a circular hole structure and a slot structure, were designed. Numerical simulations were performed to study the coupled flow and regenerative cooling heat transfer in thrust chambers with different structures. The influences of parameters such as the film mass flow rate and film hole size on wall cooling were analyzed. Experiments were conducted in a thrust chamber to validate the accuracy of the numerical calculation method. The results indicate that the slot-structured film adheres better to the wall than the circular-hole-structured film, and the film closely adhering to the wall provides better insulation against hot gas, resulting in a reduction of approximately 6% in wall temperature. When the film hole size changes, the change in circumferential wall temperature in the upstream region of the slot-structured film is more pronounced. This paper aims to provide a reference for the design of the cooling structure at the head of the thrust chamber in engineering and suggests directions for optimization and improvement.

Deterioration effect of creep on fatigue properties in QCr0.8 alloy: Damage mechanism and fatigue-creep life evaluation

The QCr0.8 alloy used in the thrust chamber liner wall of the reusable liquid rocket engine is subjected to high-temperature fatigue and high-temperature creep loading. This work investigates the effect of fatigue-creep interaction (FCI) on the fatigue life of QCr0.8 alloy. Experimental results reveal accelerated cyclic softening and decelerated stress relaxation in the alloy under FCI conditions, attributed to creep deformation and stress amplitude growth. The damage evolution mechanism of FCI is revealed by the results of the energy dispersive spectroscopy (EDS) and scanning electron microscopy (SEM). The decrease in fatigue-creep life of the QCr0.8 alloy is attributed to the fracture of the outer surface oxide layer, as well as the coupling development of creep voids and secondary cracks. Finally, a new fatigue-creep life prediction model has been developed, taking into account the increase in plastic strain energy resulting from stress relaxation and creep deformation behaviors. The study results provide substantial theoretical justification for the utilization of QCr0.8 alloy in the thrust chamber wall of reusable liquid rocket engines.

Interfacial bonding and microstructural evolution in Inconel-copper bimetallic structures fabricated by directed energy deposition-arc

Directed Energy Deposition-Arc (DED-Arc) was utilized to fabricate Inconel-copper bimetallic structures, with the aim of improving manufacturing modes and reducing costs for rocket engine thrust chambers. In this paper, bimetallic Inconel superalloy (GH4169) and copper alloy (C18150) structures were deposited using different deposition strategies. The formation mechanisms of grain-size gradients and heterogeneous interfacial microstructural evolution were comprehensive investigated and further discussed. Both strategies produced high-quality bimetallic samples with tensile strength exceeding 260.2 MPa. Fractures occurred within the C18150 metal rather than at the interface, indicating successful interface strengthening of the bimetallic samples. Gradual gradients in grain size distribution and nano-hardness were observed across the interface. The crystal structures of C18150 were unaffected by GH4169, and no new phases were generated in the interfacial region. The crystallographic orientation relationship between Ni and Cu was determined to be [011]Cu//[011]Ni and (111)Cu//(111)Ni, and interfaces were strengthened by grain boundary and dislocation strengthening through. This work demonstrates an approach for fabricating large, high-performance bimetallic structures with tailored grain-size gradients and heterogeneous microstructures, suitable for rocket engine thrust chamber applications.

Optimal Liquid Engine Architecture by Performance-Cooling Tradeoff Analysis

We perform physics-based theoretical analyses on the tradeoff between the overall performance of specific impulse and the thermal protection ability of a coolant liquid film to identify the optimal configuration of a liquid rocket engine. A bipropellant thruster is set as the target, which uses a hypergolic propellant mixture of nitrogen tetroxide as the oxidant and monomethyl hydrazine as the fuel. By considering the practical nonuniform distribution of the local mixture ratio produced in the thrust chamber, the maximum specific impulse is achieved when the fuel and oxidizer spray widths become identical, allowing for the specification of the diameter of the impinging-type injector. The heat balance between the convective heat transfer from the combustion gas and the latent heat of the three-dimensionally wavy liquid film provides the optimal diameter of the combustion chamber. The longest liquid film is found to be achieved when half of the initial film is entrained by the fast gas, correspondingly mitigating the heat transfer area due to the liquid film waviness. We successfully demonstrate the optimal architecture of the liquid engine based on the tradeoff, in which an improvement of specific impulse by 1 s shortens the film length by 2 mm.

Experimental and Simulative Evaluation of a Reinforcement Learning Based Cold Gas Thrust Chamber Pressure Controller

At DLR neural networks, as potential future controller for rocket engines, are studied. A neural network-based chamber pressure controller for a simplified cold gas thruster is presented and analyzed in simulation and experiment. The goal of the controller is twofold: It can track a trajectory with different changes of setpoints and it allows to set and control a wide variety of steady state chamber pressures. The neural network gets feeding line pressure measurement data as input and calculates valve positions as output values. The training phase of the controller is done with a reinforcement learning algorithm in an EcosimPro/ESPSS simulation, that is validated with data from the corresponding experimental set up. To increase the robustness and to allow a transfer from the simulation directly to the test facility domain randomization is applied. The controller is evaluated in simulations and experiment. It was found that – in the range of physically possible operation points – the controller achieves a constantly high reward which corresponds to a low error and a good control performance. In the simulation the controller was able to adjust all required set points with a steady state error of less than 0.1bar while retaining a small overshoot and an optimal settling time. It is found that the controller is also able to regulate all desired set points in the real experiment. A reference trajectory with different steps, linear and sinus changes in target pressure is tested in simulation and experiment. The controller was in both cases able to successfully follow the given trajectory.

Influence of helix angle on heat transfer characteristics of regenerative cooling in spiral channel

A rapid solution method for spiral line and a one-dimensional (1-D) heat transfer calculation model for spiral regenerative cooling channels (RCCs) are proposed to clarify the influence of the helix angle on the heat transfer characteristics of spiral RCCs in thrust chambers. The method based on the projection principle is characterized by a simple model and the ability to generate variable helix angles spiral line. The calculation model is characterized by high accuracy, with coolant pressure drop errors of less than 5 % coolant and coolant temperature errors of less than 20 % compared to experimental data. Utilizing the calculation model, research on the heat transfer characteristics of RCCs with varying helix angles is conducted. Results indicate that the coolant temperature rise and pressure drop increase as the helix angle increases. The peak values of the gas-side wall temperature occur at the throat and the contraction section. The throat peak value decreases with an increase in the helix angle, while the contraction section peak value increases. When the helix angle is 30°, the two peak values are close as well as the highest gas-side wall temperature is minimized. The highest wall temperature is only 1111K. The research results can provide guidance for the design and processing of spiral RCCs.