Cryogenic Rocket Research Articles

Demonstration of charge-hold-vent (CHV) and no-vent-fill (NVF) in a simulated propellent storage tank during tank-to-tank cryogen transfer in microgravity

The space exploration from a low earth orbit to a high earth orbit, then to Moon, Mars, and possibly asteroids and moons of other planets is one of the biggest challenges for scientists and engineers for the new millennium. The enabling of in-space cryogenic rocket engines and the Lower-Earth-Orbit (LEO) cryogenic fuel depots for these future manned and robotic space exploration missions begins with the technology development of advanced cryogenic thermal-fluid management systems for the propellant transfer line and storage tank system. One of the key thermal-fluid management operations is the chilldown and filling of the propellant storage tank in space. As a result, highly energy efficient, breakthrough concepts for quenching heat transfer to conserve and minimize the cryogen consumption during chilldown have become the focus of engineering research and development, especially for the deep-space mission to Mars. In this paper, we introduce such thermal transport concepts and demonstrate their feasibilities in space for cryogenic propellant storage tank chilldown and filling in a simulated space microgravity condition on board an aircraft flying a parabolic trajectory. In order to maximize the storage tank chilldown thermal efficiency for the least amount of required cryogen consumption, the breakthrough quenching heat transfer concepts developed include the combination of charge-hold-vent (CHV) and no-vent-hold (NVF). The completed flight experiments successfully demonstrated the feasibility of the concepts and discovered that spray cooling combined with hold and vent is more efficient than the pure spray cooling for storage tank chilldown in microgravity. In microgravity, the data shows that the CHV thermal efficiency can reach 39.5%. The CHV efficiency in microgravity is 6.9% lower than that in terrestrial gravity. We also found that pulsing the spray can increase CHV efficiency by 6.1% in microgravity.

Two-phase flow characteristics of cryogenic propellant in filling the head cavity of liquid rocket engine

The filling of the head cavity downstream of the valve with liquid oxygen in a cryogenic rocket engine is a crucial stage in ignition. The work presented here intends to study the two-phase flow characteristics in the head cavity through visualization experiments and simulations, which are carried out using liquid nitrogen as the substitution fluid. The geometry used comes from the scaling of the actual head cavity in the engine. The temperature of liquid nitrogen (TLN2,s) supplied for the head cavity is higher than the saturation temperature in the head cavity (Tsat,c) with room pressure as the initial state, which is proved in the pro-cooling pipeline upstream of the valve before the filling process. Flash boiling is inevitable due to the sudden drop in pressure after the valve is opened. Then, the pressure in the head cavity rises in two steps. The production of gaseous nitrogen leads to the first step, and the mist flow is visualized because liquid nitrogen is broken up and carried by gaseous nitrogen in the form of droplets. The increase in the content of liquid nitrogen in the head cavity causes the second step, and the bubbly flow occurs as the gaseous nitrogen is dispersed as small bubbles in the continuous liquid nitrogen. The transition from mist flow to bubbly flow is led by the decrease in the superheat of liquid nitrogen (ΔT = TLN2,s- Tsat,c). The transition occurs at ΔT = 4 K ∼ 5 K and the minimum achievable superheat is 1.3 K when the mass flow rate of liquid nitrogen supplied is 5.4 kg/s ∼ 6.9 kg/s. Furthermore, the gas nitrogen content in the head cavity is mainly determined in the first step of the pressure rise. Reducing the temperature of liquid nitrogen supplied can weaken the flash boiling, and increase the content of liquid nitrogen in the head cavity.

Study on unsteady evacuation characteristics of multi-layer insulation during launch/in orbit of the cryogenic propellant tanks

The MLI + foam is used for the thermal protection of the cryogenic propellant. However, the MLI generates unsteady evacuation when the cryogenic rocket is launched and traveled in space. The insulation performance of the MLI is determined by the evacuation level of the interstitial gas. Aiming at the unsteady evacuation of the MLI, the interlayer pressure node algorithm was developed. The transient evacuation characteristics of the interlayer pressure were predicted. The effect of layer number, slit rate, and layer density on evacuation was investigated. The evacuation and insulation performance of MLI were optimized. The results show that the innermost pressure of the 60 mm thickness (th60) MLI dropped to less than 0.5 Pa when the evacuation was conducted for more than 25 h. The pressure of the innermost layer and 10th layer is greater than 1 order of magnitude higher than that of the outermost layer. The evacuation rate of th20 is faster than that of th40 and th60. The innermost layer pressure value of the th20 is 0.168 Pa, which is 10.47 times higher than that of the outermost layer. The innermost layer pressure of the th20 is 48.3 % and 64.8 % lower than that of th40 and th60, respectively. The gas conduction decreased by 61.22 % when the opening pattern was optimized. These results can provide an important reference for the thermal protection design of cryogenic propellants.

Surface tension effects on cryogenic liquid injection dynamics in supercritical environment

The injection of cryogenic fluids into environments where the prevailing conditions are supercritical in comparison to the critical point of the injected cryogenic fluid is encountered in cryogenic rocket engines, and novel engine architectures such as the recuperated split cycle engine. The physical characteristics of cryogens injected into supercritical environment are rather unclear. While surface tension is usually assumed to be absent/negligible for supercritical fluids, recent experimental research has identified the existence of surface tension and its effects on liquid hydrocarbons in supercritical environment. This research work proposes an alternative computationally simple adaptive surface tension algorithm for the simulation of a liquid injected into supercritical environment. The numerical simulations presented here correspond to single- and binary-specie cases of iquid nitrogen and liquid methane respectively, undergoing phase transition post their injection into supercritical conditions. Following a critical review of related numerical works, this paper begins with a brief explanation of the physics behind the surface tension effect in a binary-fluid interface in which a supercritical fluid is involved and we present why this effect is of relevance to supercritical cryogenic jets? Then, the rationale and specifics of the the new modelling framework based on adaptive surface tension is discussed along with its implications. The results of the numerical simulations of low-temperature vs near-critical temperature iquid nitrogen and liquid methane injection dynamics revealed the drastically different fluid- and thermo-dynamics at play in these two cases. The role of surface tension at these conditions is also explored.

Cfd Analysis in Coaxial Injector of a Cryogenic Rocket Engine And Evaluation of Atomisation Characteristics Using Simulant Fluids

A Study of atomisation and mixing of 2-phase propellants using simulant fluids injected from a coaxial injector of an operating cryogenic rocket engine is carried out and presented along with the CFD analysis of flow within it. The original injector jet (Jet 3) designed, modeled and analysed has swirl for LOX (water) alone with GH2 (air) having no swirl. Injector jet having inner casing and outer sleeve is modeled using I-DEAS software and analysed for inner core fluid domain (liquid) and annular fluid (air) domain. Results of flow analysis inside jet revealed asymmetric nature of air flow at exit due to the air fluid domain itself is not axisymmetry in the original jet. This non-axisymmetry of 2-phase fluids at the exit of injector jet will obviously affect the characteristics of atomisation and mixing of fluids. Data obtained from atomization experiments done on this jet using stimulant fluids (water in LOX side and air in GH2 side) such as particle distribution and Sauter Mean Diameter (SMD), fairly matches with the predictions made using CFX software and associated codes which enabled to evaluate the analysis software. Further, the air flow at exit has been made axisymmetric by modifying the air passage in the injector jet (Jet 3 Mod) and the effect of this modification on atomisation and mixing of 2-phase fluids is predicted using the CFX software. In combustion better efficiency with less instability is to be attained. For meeting both simultaneously articles size distribution after injection has to be neither very finer nor coarse. Comparison of predicted data on atomisation revealed that the modified Jet (Jet 3 Mod) is relatively better to give stable combustion with higher performance.

A meta network pruning framework for remaining useful life prediction of rocket engine bearings with temporal distribution discrepancy

Accurate remaining useful life prediction (RUL) is important for the reliability and safety of liquid rocket engines. In this paper, a meta network pruning framework with attention augmented convolutions is proposed for RUL prediction. To address the problem of distribution discrepancy in engineering data under transient working conditions, a data-driven distribution matching strategy is designed. Besides, in view of the prediction accuracy and computation complexity of the model, an iterative meta network pruning algorithm, which automatically calculates the meta-gradients of each convolutional kernel according to the chain rule, is developed to identify, and then delete the unimportant connections in the framework. The method is verified on a high-precision cryogenic rocket engine bearing experiment platform under liquid nitrogen and received better performance than benchmark algorithms.

The Mechanism Research of Low-Frequency Pressure Oscillation in the Feeding Pipe of Cryogenic Rocket Propulsion System

In the propulsion system of cryogenic liquid rockets, low-frequency pressure oscillation is a long-standing issue occurring in its feeding pipe, and is not conducive to the normal operation of the rocket. Its mechanism and excitation process are not very clear due to the limitation of the existing numerical method and the difficulty of the real dynamic experiment. Inspired by the periodic suck-back flow phenomenon of steam condensation, the fluctuation of the two-phase interface might be the crucial factor to initiate the low-frequency pressure oscillation. To simulate this interfacial characteristic of cryogenic propellant, a novel numerical model is proposed to predict the mass transfer rate weighted by the interfacial curvature. Aiming at the oxygen jet condensation simulation, the low-frequency pressure oscillation phenomenon is obtained successfully with the excitation frequency of 10.6 Hz, consistent with the natural frequency of the engine test run. It is conducted so the low-frequency pressure oscillation is caused by the periodic condensation of the continuous oxygen vapour plume, along with an oxygen suck-back flow phenomenon. In addition, the results indicate that both the oxygen and liquid oxygen mass flux promote the rise in the frequency of pressure oscillation. These conclusions provide theoretical instructions for the design and operation of the propulsion system of a cryogenic liquid rocket.

Effects of microgravity on cryogenic spray quenching of a simulated propellant storage tank with enhancement by surface coating and flow pulsing

In-space cryogenic propulsion will play a vital role in space exploration mission to the Moon and future mission to the Mars. The enabling of in-space cryogenic rocket engines and the Lower-Earth-Orbit (LEO) cryogenic fuel depots for these future manned and robotic space exploration missions begins with the technology development of advanced cryogenic thermal-fluid management systems for the propellant transfer line and storage tank system.One of the key thermal-fluid management operations is the chilldown of the propellant storage tank. As a result, highly energy efficient, breakthrough concepts for quenching heat transfer to conserve and minimize the cryogen consumption during chilldown have become the focus of engineering research and development, especially for the deep-space mission to Mars. In this paper, we introduce such thermal transport concepts and demonstrate their feasibility in space for cryogenic propellant storage tank chilldown in a simulated space microgravity condition on board an aircraft flying a parabolic trajectory. In order to maximize the storage tank chilldown thermal efficiency for the least amount of required cryogen consumption, the breakthrough quenching heat transfer concepts developed include the combination of cryogenic spray quenching cooling, Teflon thin-film coating on the simulated tank surface, and spray flow pulsing. The boiling heat transfer physics that supports the concepts is clearly explained. The completed flight experiments successfully demonstrated the feasibility of the concepts and discovered that spray cooling is the most efficient cooling method for the tank chilldown in microgravity. In microgravity, the data shows that the chilldown thermal efficiency can reach 30% with Teflon coating alone. However, Teflon coating together with flow pulsing was found to substantially enhance the chilldown efficiency further to reach the 50% goal.

A multi-head attention network with adaptive meta-transfer learning for RUL prediction of rocket engines

Accurate prediction of remaining useful life (RUL) is necessary to ensure stable and safe operations for rocket engines. The paper proposed a multi-head attention network coupled with adaptive meta-transfer learning for RUL prediction. By combining the convolution-based branch with an attention-based branch, the multi-head attention network is proposed for accurate RUL prediction of cryogenic bearings in rocket engines under the steady stage. In addition, an adaptive model-agnostic meta-transfer learning strategy is developed to further improve the performance under small sample circumstances with adaptive hyper-parameters. To demonstrate the superiority, the proposed method is compared with typical benchmark algorithms using real monitoring data from a high-precision cryogenic rocket engine experiment platform. Results indicate that the proposed method achieves better performance compared with existing models under multiple evaluation indexes.

Research on process optimization and rapid prediction method of thermal vibration stress relief for 2219 aluminum alloy rings

Abstract 2219 aluminum alloy rings are important part of liquid cryogenic rocket fuel tanks. Residual stress is inevitably introduced in the forming process of the rings due to the nonlinear thermomechanical coupling conditions, which will affect its mechanical properties, fatigue properties, corrosion resistance, and dimensional stability. Thermal vibratory stress relief (TVSR) has great potential in reducing residual stress, and process optimization of TVSR is necessary to further improve its application, but it is rarely reported. In this study, process optimization of roll formed 2219 aluminum alloy rings is conducted. The influence of vibration amplitude, vibration time, vibration frequency, heating time, holding time, and cooling time on TVSR treatment are investigated. Results show that the maximum equivalent residual stress of 2219 aluminum alloy rings can be reduced by 93.6% after optimized TVSR treatment. With the increase in vibration time, heating time, holding time, and cooling time, the maximum equivalent stress decreases. However, the increase in the vibration amplitude results in an increase in the maximum equivalent stress. Further, a genetically optimized artificial neural network intelligent optimization algorithm is applied to quickly predict the TVSR effect of 2219 aluminum alloy rings.

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

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

Prediction of Wall Heat Fluxes in a Rocket Engine with Conjugate Heat Transfer Based on Large-Eddy Simulation.

Although a lot of research and development has been done to understand and master the major physics involved in cryogenic rocket engines (combustion, feeding systems, heat transfer, stability, efficiency, etc.), the injection system and wall heat transfer remain critical issues due to complex physics, leading to atomization in the subcritical regime and the interactions of hot gases with walls. In such regimes, the fuel is usually injected through a coaxial annulus and triggers the atomization of the central liquid oxidizer jet. This type of injector is often referred to as air-assisted, or coaxial shear, injector, and has been extensively studied experimentally. Including such injection in numerical simulations requires specific models as simulating the atomization process is still out of reach in practical industrial systems. The effect of the injection model on the flame stabilization process and thus on wall heat fluxes is of critical importance when it comes to the design of wall-cooling systems. Indeed, maximizing the heat flux extracted from the chamber can lead to serious gain for the cooling and feeding systems for expander-type feeding cycles where the thermal energy absorbed by the coolant is converted into kinetic energy to drive the turbo-pumps of the feeding system. The methodology proposed in this work to numerically predict the flame topology and associated heat fluxes is based on state-of-the-art methods for turbulent reactive flow field predictions for rocket engines, including liquid injection, combustion model, and wall treatment. For this purpose, high-fidelity Large Eddy Simulation Conjugate Heat Transfer, along with a reduced kinetic mechanism for the prediction of chemistry, liquid injection model LOx sprays, and the use of a specific wall modeling to correctly predict heat flux for large temperature ratio between the bulk flow and the chamber walls, is used. A smooth and a longitudinally ribbed combustor configuration from JAXA are simulated. The coupling strategy ensures a rapid convergence for a limited additional cost compared to a fluid-only simulation, and the wall heat fluxes display a healthy trend compared to the experimental measurements. An increase of heat transfer coherent with the literature is observed when walls are equipped with ribs, compared to smooth walls. The heat transfer enhancement of the ribbed configuration with respect to the smooth walls is coherent with results from the literature, with an increase of around of wall heat flux extracted for the same chamber diameter.

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

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

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

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

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

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

Simulation Study of Forced Circulation Pre-cooling for Hydrogen

Due to the low boiling point for cryogenic propellant, the cryogenic rocket engine and delivery pipelines must be fully cooled before startup. Pre-cooling methods largely influence the complexity of the pre-launch rocket test procedure and the adaptability of delayed launch. Therefore, it is needed to comprehensively consider amounts of factors. In this paper, by comparing three pre-cooling methods in detail, selection principles are given. Finally, a model of the forced circulation pre-cooling system for hydrogen is developed with AMESim. According to the system characteristics, it is found that the tank pressure, the speed of the circulating pump and the pressure in the isolated cavity influence the cooling effect. The simulation results show that the pressure difference between the outlet pressure of the circulating pump and the pressure in the isolated cavity is the key factor, and the bigger, the better.

A Technology for Improving Regenerative Cooling in Advanced Cryogenic Rocket Engines for Space Transportation

Regenerative cooling of thrust chamber is the unique solution for the thermal management of high heat flux generated inside the combustion chamber of Cryogenic rocket engine. Heat is transferred from combustion hot gas to coolant through the channels provided on inner copper shell, thereby cools the inner wall of the nozzle. A novel technique of providing copper foam inside the channels will act as an infinite fin and also act as barrier for coolant stratification. This will improve the heat transfer to the coolant and reduce the nozzle wall temperature. Heat transfer improvement with copper foam inserts to the coolant channel is demonstrated through experiments with simulated fluids. Experiments are conducted with simulated hot gas chamber and coolant channels using water as the coolant. Copper foam with high porosity is selected to fill the channels. Hot tests are carried out with copper foam filled coolant channels and measured the coolant temperature rise and pressure drop across the channels. Tests are repeated with similar hot gas condition, but without inserting copper foam inside the channels. A substantial enhancement in heat transfer to the coolant is observed with copper foam inserts experiments, which will reduce the wall temperature. This gives a good handle on the life cycle improvement of multi-start cryogenic engines for future space transportation systems. This paper details the specification of copper foam, hardware design, experiments and measurements, and the application of the augmentation of heat transfer coefficient in operating cryogenic engines.

Design and realization of thrust measurement system for high altitude testing of cryogenic rocket engine

Design and realization of thrust measurement system for high altitude testing of cryogenic rocket engine

Design of Regenerative Cooled Cryogenic Rocket Engine

Design of Regenerative Cooled Cryogenic Rocket Engine

Experimental investigation of pressure drop of feeding pipes of liquid oxygen in a five-way distribution cavity under different working states

The uniform distribution of liquid oxygen is essential for design of feeding system in multi-engines of cryogenic rockets. Aiming at a cryogenic launch vehicle with four engines in parallel, a five-way spherical cavity structure was used to evenly distribute liquid oxygen in the multi-branch pipes. These cryogenic rocket engines need to be controlled by opening or closing states for multi-branch pipes to adapt different flight conditions. To study effect of working states on pressure drop in the four branch pipes with the five-way spherical cavity, a large-scale test system was constructed to carry out discharging flow of liquid oxygen. The feeding pipes of full-scale model were adopted, and forty-nine sensors were arranged in the pipes to measure pressure and temperature of liquid oxygen. The constant discharging flow was maintained by a feedback system of tank pressure. The opening or closing states of branch pipes were controlled by cut-off valves at the end of each branch pipe. The liquid oxygen tests were carried out under different actual flight conditions. The results shown that the opening or closing states for four branch pipes had a dramatic effect on the pressure drop in the multi-branch pipes. When the branch pipes all drained at the same time, there was a large pressure drop in two symmetrical branch pipes which enhanced linearly with the increase of Reynolds number. The pressure drop of two branch pipes on the symmetrical sides could exceed 0.1 MPa which was far greater than the normal flow resistance, when the Reynolds number was 5.0 × 106. It was found that there was a regular swirling flow formed in the five-way spherical cavity by using CFD method. When two branch pipes were turned off, whether symmetrical sides or adjacent sides, the large pressure drop still existed, and was positively correlated with the varying Reynolds number. The maximum pressure drop in the branch pipes was close to 0.09 MPa, when the Reynolds number was 3.3 × 106. When three branch pipes were closed, the large pressure drop disappeared, and the flow resistance related positively with the Reynolds number, which accorded with the normal flow resistance. Moreover, according to data of six hot test runs and two verification tests, it was found that the inlet pressure in the main pipe had greatly influence on the pressure drop, which raised sharply as the inlet pressure increased. These results could help to provide a design reference for feeding system of multi-engines in cryogenic rockets.