Experimental study of a novel coolant feed system utilizing waste heat of rocket engines for pressurization

A static analysis of thermal-mechanical coupling for full liquid rocket engines based on decoupled computation of multi-physics fields

Numerical study of flow and heat-transfer enhancement in lattice-core regenerative cooling channels for liquid rocket engines

Fatigue-resistant structural optimization and vibration testing of a spherical tube fitting in liquid rocket engines

This study presents an experimental investigation of the flow and combustion dynamics of a single LOX/CH 4 liquid-centered swirl injector under high-pressure transcritical conditions, conducted on the cryogenic MASCOTTE test bench. As a preliminary step, the interaction between hydrodynamic and combustion phenomena in a stable operating regime is examined. The structure and evolution of the flame and dense phase are analyzed using synchronized high-speed imaging at 13 kHz, with OH* chemiluminescence capturing the reactive zone and backlighting visualizing the dense phase. Time-averaged and instantaneous fields are provided to highlight the flame structure and flow behavior. Spectral analysis, conducted via Fast Fourier Transform (FFT), is performed, and complemented by video filtering using a combination of averaged Inverse Discrete Fourier Transform (IDFT) and phase-averaging techniques. The backlighting recordings reveal a dominant symmetric mode aligned with the preferred instability mode frequency of the dense jet. The corresponding Strouhal number, ranging between 0.25 and 0.5, is consistent with classical hydrodynamic predictions and is identified as a symmetric Kelvin–Helmholtz instability. Similar modal behavior is observed in the OH* chemiluminescence signal, sharing the same dominant frequency and mode shape. These coherent structures are confined to the reaction zone and dense phase region, with no significant modes persisting downstream. In the absence of significant pressure oscillations, it is inferred that the flame dynamics are primarily driven by the dense phase hydrodynamics. These findings enhance the understanding of coupled hydrodynamic-combustion instabilities in transcritical swirl injectors and offer insights relevant to the design and control of next-generation rocket propulsion systems. Novelty and significance statement The novelty of this work lies in its contribution to the very limited experimental data on transcritical LOX/CH 4 swirl flame dynamics available in the literature. Backlighting imaging and OH* chemiluminescence measurements reveal a dominant symmetric mode in the flow oscillations. Additionally, the state-of-the-art Strouhal number definition for non-reactive swirl flows provides a good estimate of the corresponding dominant frequency, which aligns with the preferred instability frequency of the dense jet. This analytical frequency estimation enhances understanding of the coupling between hydrodynamic and combustion instabilities in transcritical swirl injectors, representing a key step toward the development of safe and reliable liquid rocket engines.

This study investigates the control of methane/air inverse diffusion combustion using surface dielectric barrier discharge (SDBD) plasma technology to enhance methane fuel combustion performance in rocket engines. Under lean combustion conditions (equivalent ratio, Φ = 0.76), forward SDBD dissociates methane C–H bonds via high-energy electrons, generating CH3 radicals and forming a stable conical flame at 16 kV, while reverse SDBD suppresses turbulence to reduce flame height by 41.7%. At an optimal equivalence ratio (Φ = 1), the reverse structure achieves flame height reduction from 130 mm to 94 mm, whereas the forward structure exacerbates flame nonuniformity due to aerodynamic effects. In rich combustion (Φ 1.5), both plasma configurations inhibit methane inverse diffusion combustion, with the forward structure prone to causing flame instability. Analysis confirms that oxygen content is critical to the divergent control effects: forward SDBD excels in high-oxygen environments for combustion enhancement, while reverse SDBD is more effective for flow control in low-oxygen conditions. This research provides experimental insights and technical references for optimizing plasma-assisted combustion in rocket engines.

Regenerative cooling for hydrogen rocket engine: Fundamental design and thermal analysis

Efficient thermal analysis of film-cooling effects in reacting flows of throttleable liquid rocket engines

Koopman theory assisted transfer learning for anomaly detection of liquid rocket engines in frequency domain

Hybrid rocket engines offer a compromise between safety, controllability, and performance, making them attractive for small-scale propulsion systems. However, oxidizer selection remains a critical early-stage design decision that cannot be determined solely from ideal thermodynamic metrics. This study presents a comparative analysis of three oxidizers—nitrous oxide (N2O), gaseous oxygen (GOX), and liquid oxygen (LOX)—for a 1 kN-class hybrid rocket engine using HDPE fuel under identical operating conditions. Equilibrium combustion performance was first evaluated using NASA Chemical Equilibrium with Applications (CEA) to determine optimal oxidizer-to-fuel ratios and theoretical specific impulse. These results were subsequently refined using Rocket Propulsion Analysis (RPA) to incorporate finite combustion chamber geometry and non-ideal nozzle expansion effects. The equilibrium analysis predicts maximum specific impulses of approximately 260 s for N2O/HDPE and nearly 300 s for oxygen-based systems. However, finite-geometry modelling indicates that practical performance is reduced by approximately 5–8%, yielding delivered specific impulses of about 275 s for GOX and 272 s for LOX. The results demonstrate that although oxygen (GOX and LOX) provides higher thermodynamic performance, the practical advantage of LOX over GOX becomes marginal at the kilonewton scale. Consequently, oxidizer selection for small hybrid engines should be treated as a system-level trade-off involving performance, infrastructure complexity, and operational safety.

ABSTRACT In this study, an ad hoc analytical model is developed to describe the axial profile of the time-averaged heat release rate in the diffusion flame of a Liquid Rocket Engine coaxial injector operating under transcritical conditions (resulting in a simple density stratification, without vaporization). The model is derived from the Burke-Schumann diffusion flame model with the infinitely fast combustion assumption, augmented by adding turbulent diffusivity and velocity dispersion. It is therefore built upon a series of physically motivated assumptions which are systematically validated against experimental and numerical data. The predictive capability of the model is then assessed through detailed comparisons with Large Eddy Simulation results from the literature, demonstrating its ability to capture key features of the flame structure as it gives an accurate representation of the flame profile and effectively captures the total time-averaged heat release rate. Furthermore, a preliminary extension explores the ability of the model to predict the effects of forcing amplitude and frequency on the flame response. The model exhibits some key features of the flame response to acoustic forcing. Designed to be integrated into flame transfer function models, this work represents a milestone toward fully analytical flame transfer functions.

Abstract The heat transfer efficiency of the regenerative cooling channel (RCC) plays a crucial role in rocket engines. In the present work, subcritical methane heat transfer in RCCs of varied geometries was analyzed. A calculation method for the geometric design of a phase-change-aware RCC was established, and an experimental correlation method was employed to unveil the thermal features of subcritical methane. Moreover, the thermal behaviors of the methane coolant in RCCs of varied channel aspect ratios (ARs) were analyzed, and it was found that the proposed dimension and channel number of the RCC met the requirements for thermal protection of the thrust chamber. An increase in the channel AR from 1.60 to 4.93 resulted in a decline in the maximum gas-side wall temperature and the local peak at the throat by 24.7 K and 36.8 K, respectively, but meanwhile led to a 29.8% increase in the pressure loss of the channel. Increasing the channel number could substantially reduce the wall temperature at the gas side, and a higher AR would improve the cooling performance of the RCC under an acceptable pressure loss. The research here is expected to provide insights and a practical guide for optimized RCC design in liquid oxygen/liquid methane (LOX/LCH4) rocket engines.

Abstract As a critical element within liquid rocket engine turbopumps, the end face seal plays a vital role. The stability and dynamic behavior of the rotor system are profoundly impacted by the hydrodynamic pressure effects emerging from the fluid film within the seal clearance. This study aims to investigate the liquid film's dynamic properties and their impact on the rotor system's transient behavior during both rapid acceleration and steady-state phases. Initially, a calculation model for the fluid film's equivalent stiffness and damping coefficients is developed using the small perturbation technique based on the Reynolds equation. This model explicitly accounts for axial and angular disturbances of the sealing ring, analyzing how these coefficients shift with varying parameters like film thickness and rotation speed. Findings indicate that higher rotational speeds intensify the dynamic pressure effect, leading to an overall elevation in stiffness coefficients across all directions. Subsequently, a finite element model representing a specific liquid rocket engine turbopump is constructed. By integrating time-variant seal dynamic parameters into the rotor's equations of motion, a coupled seal-rotor dynamic model is established. Numerical simulations were performed to evaluate the system's transient response during a rapid startup sequence (accelerating to 40,000 r / min within 2 seconds). The simulation outcomes highlight how seal parameters affect rotor stability and vibration magnitude, underscoring the necessity of accurately incorporating end face seal dynamics in rotor analysis. Consequently, this work offers valuable theoretical insights for the design and dynamic analysis of liquid rocket engine turbopumps.

Vibration fatigue reliability-based robust design optimization of turbine blades in rocket engines using Kriging surrogate model with sequential sampling

Abstract The integration and complexity of key equipment, such as the liquid rocket engine (LRE), render anomaly detection (AD) with multi-source data paramount for safe and effective operation. However, in real-world conditions, sensor failures frequently lead to incomplete data, which significantly degrades model performance. Consequently, existing AD methods are often ineffective when applied to incomplete multi-source data. To address this challenge, we propose a virtual signal-based depthwise separate network. The proposed network can perform AD on incomplete data without relying on imputation that requires strong distribution assumptions. Specifically, a channel-independent embedding layer augmented with virtual signals is designed to mitigate the performance degradation caused by incomplete inputs. In particular, the virtual signal is defined as a set of learnable parameters to characterize the normative patterns of sensors. Subsequently, the combined embeddings enable the proposed model to perform intra-channel feature extraction and inter-channel feature fusion, respectively. Finally, a message-passing reconstruction head is introduced, which reconstructs the observed channels utilizing the structure information learned by virtual sensors. Extensive experiments on real LRE datasets have demonstrated that the proposed method outperforms existing methods in both complete and incomplete conditions. Specifically, the proposed method achieves an F 1-score of 0.9975 on complete data and maintains a robust F 1-score of 0.8510 even under a 90% missing rate, outperforming the best baseline model by approximately 10% in severe missing scenarios.

Increasing the efficiency of liquid rocket engines (LRE) is inextricably linked to the intensification of cooling processes in heat-critical elements such as combustion chambers and nozzles. Additive Manufacturing (AM) technologies, in particular selective laser fusion (SLM), make it possible to create cooling paths with complex internal channels and heat transfer intensifiers. However, the surface microrelief formed in this case has a significant effect on hydrodynamic resistance and heat transfer, which is not always taken into account in existing design methods. An experimental assessment of the effect of surface roughness, characteristic of additive manufacturing, on the formation of a dynamic boundary layer under conditions simulating flows in the cooling channels of liquid propellant is considered. The sample plates made of AlSi10Mg aluminum alloy were made using the SLM method: with controlled roughness and flow intensifiers. The surface roughness profile was evaluated on a TR 110 profilometer. The research was carried out on a specially designed aerodynamic installation that simulates the flow in a rectangular channel. The velocity fields in the wall area were measured by the pneumometric method at a fixed air flow rate. The errors were estimated and the data approximated. It is established that the roughness of SLM surfaces leads to a significant restructuring of the velocity profile in the boundary layer compared to the hydraulically smooth case: a decrease in its thickness and a change in the shape of the velocity diagram are observed. The largest deviations were recorded for the plate with intensifiers. Power-law approximating dependences are obtained for each type of surface, where the exponent varies depending on the roughness parameters. It is shown that the use of boundary layer models for smooth surfaces in liquid propellant calculations in relation to additive parts can lead to incorrect estimation of hydraulic losses and, as a result, to errors in predicting the temperature state of the wall. The obtained dependences are necessary for the development of adjusted calculation methods that take into account the actual topography of the surfaces formed by AM methods, which will increase the reliability and efficiency of the design of liquid propellant cooling systems.

The renewed interest in hydrogen peroxide-based space propulsion systems has highlighted the persistent issue of catalyst degradation during long-term operation. Although several studies have investigated the underlying causes of this phenomenon, effective regeneration techniques capable of restoring catalytic activity have not yet been clearly demonstrated. This study investigates the mechanisms responsible for performance degradation and proposes a viable regeneration strategy for palladium-based catalysts. Experimental analyses were conducted on a batch of commercial Al2O3/Pd pellets subjected to multiple firing cycles in a 10 N-class hybrid mini-thruster. Monitoring of the propulsive performance revealed a progressive decline in catalytic activity, ultimately preventing ignition of the hybrid rocket engine. To characterize the degradation mechanisms, the pellets were examined through visual inspection, static hydrogen peroxide decomposition tests, and Temperature Programmed Reduction (TPR) analysis. The results indicated significant surface oxidation of palladium, leading to reduced decomposition efficiency. A chemical regeneration procedure based on sodium borohydride (NaBH4) treatment was subsequently developed to restore catalytic performance. The regenerated pellets were tested under the same experimental conditions that had previously led to ignition failure. Their propulsive performance was then compared with both the degraded pellets and a new batch of equivalent catalysts. The results demonstrate that the regeneration process successfully restored the catalytic activity to levels comparable with the original state, enabling stable and efficient hybrid combustion. These findings confirm the role of surface oxidation in catalyst degradation and demonstrate that targeted chemical treatment can significantly extend catalyst lifetime. The proposed regeneration strategy offers a practical method to reduce costs of ground-based experimental campaigns and support the future deployment of hydrogen peroxide-based propulsion systems in space applications by providing insights into the mechanisms that can degrade the performance of palladium catalysts.

An enhanced intelligent-driven learning framework for adaptive multi-objective optimization control of pintle solid-fuel rocket engines

Porous supercritical transpiration cooling characteristics of triply periodic minimal surface

ABSTRACT This study investigated the influence of the pressure‐cure process on the pore characteristics of nitrate ester‐plasticized polyether (NEPE) propellant using micro‐computed tomography technology. The mesostructure characteristics of the NEPE propellant were obtained under various curing pressures, including atmospheric pressure (0 MPa), 1, 1.5, and 2 MPa. The results demonstrated that the pore characteristics of the NEPE propellant exhibited significant changes under pressure cure. Specifically, the number of pores with a sphericity ranging from 0.61 to 1, indicating a more rounded morphology, increased, while the proportion of micro‐pores with a volume below 0.5 µm decreased. The micro‐pores coalesced into larger voids, and the pore distribution exhibited a distinct clustering tendency. The NEPE propellant produced by the pressure cure process exhibited an increased elongation rate, along with enhanced tensile strength and elastic modulus, as well as a significant reduction in porosity. However, when the applied pressure exceeded a certain threshold, like 2 MPa, the mechanical properties of the pressurized‐cured propellant no longer improved but instead decreased. When the pressure applied during the pressure cure process was in the range of 1–1.5 MPa, the porosity of the propellant decreased, the numerical distributions of pore sphericity and volume became more concentrated, and the mechanical properties were optimized to the greatest extent. Analysis of the pore characteristics and mechanical properties of NEPE propellant produced via pressure cure provides a guiding foundation for optimizing the solid rocket engine propellant manufacturing process in the aerospace industry.