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In this paper, a novel two-stage hydrocarbon fuel heating and delivery system was used to more accurately simulate the state of fuel in the cooling channel of the regenerative cooling engine. Mildly cracked (<10%) experiments of n-dodecane were conducted by using the new heating system and pyrolysis products were sampled and analyzed. The ignition delay time and laminar flame speed of the cracked products were different from those of the surrogate gas fuel. When the temperature is 1100 K, the ignition delay time of the cracked products are ~50% higher than that of surrogate gas fuel. The laminar flame speed of the cracked products is about 30% lower than that of surrogate gas fuel at the equivalence ratio of 1.0. The supersonic combustion experiments of mildly cracked and different equivalence ratios n-dodecane were also employed at Mach number 3.0 of the isolator entrance, with total temperature ~1550 K and total pressure ~1.5 MPa. The static pressure along the axis of combustor was measured, while the velocity and temperature of the airflow at the exit of combustor were also measured by the tunable diode laser absorption spectroscopy (TDLAS). The distributions of airflow velocity, static temperature, and total temperature along the combustor at different conditions were analyzed through a one-dimensional analysis method. The deviations of velocity and temperature of the airflow are less than 4% and −6%, respectively, which indicates the calculation is in great accordance with the TDLAS detections. The experimental data will be helpful for validating the simulation of supersonic combustion, and studying the influence of mildly cracked fuel on supersonic combustion characteristics. [ABSTRACT FROM AUTHOR]

The article presents preliminary results of the development and research testing of a rocket engine chamber with a thrust of 100 N operating on fuel components: gaseous oxygen-gaseous hydrogen, designed to be used as the main engine of a small upper stage for launching a payload weighing up to 150 kg into target orbits. The main features of the engine chamber are its fabrication by selective laser melting from 12X18N10T steel powder and its regenerative cooling with gaseous hydrogen. Calculation and experiments confirmed the possibility of regenerative cooling of the chamber in the nominal mode. In addition, the current results of the development of a technique for optical recording of the process of structural material removal from the chamber during tests using different optical filters are presented. It is shown that there is a correlation between the brightness of the obtained frames and the hydrogen flow rate. It is also shown that the afterburning of the removed material particles, in contrast to high thrust engines, occurs mainly in the tail of the jet.

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In this paper, flow resistance and convective heat transfer of Al-kerosene nanofuels are studied experimentally. Al-kerosene nanofuels with mass fractions of 0.5 , 1 , and 2 g/L are prepared and applied as the flow medium for flow and heat transfer experiment via a heating facility. The experiment results indicate that the addition of aluminum nanoparticles has significant influence on both flow resistance and heat transfer performance. Compared to kerosene experiment, friction coefficient, heat transfer coefficient, and Nusselt number of Al-kerosene nanofuels all increase with different increasing rates. With a mass fraction of 1 g/L, the increase rate of friction coefficient was 11%, while the increase rate of heat transfer coefficient and Nusselt number is 19% and 12%, respectively. In order to evaluate the overall flow and heat transfer performance of Al-kerosene nanofuels, a performance evaluation criteria (PEC) is evaluated, and the present experimental results prove that the addition of aluminum nanoparticles gave a gain to the overall thermal performance of kerosene. The present study is aimed to provide useful references for regenerative cooling improvements. [ABSTRACT FROM AUTHOR]

Heat transfer enhancement based upon micro-disturbing structures becomes increasingly important for thermal protection of air-breathing propulsion systems. To explore the influence mechanism of different structures on heat transfer in regenerative cooling channels of scramjet engines, flow structures and heat transfer of supercritical-pressure n-Decane in mini-channels with dimple and micro-rib arrays are numerically investigated. Results indicate effective heat transfer enhancement inside the channels depends more on efficient transverse turbulence disturbance rather than longitudinal or perpendicular turbulence disturbance. Severe thermal stratification generates strong shear layers in channels, where transverse secondary flows bring much intenser flow mixing than longitudinal secondary flows, especially in transcritical temperature region where the best heat transfer enhancement is obtained. With same characteristic dimensions, longitudinal micro-disturbing structures like dimples have lower flow resistance, but transverse micro-disturbing structures like micro-ribs achieve stronger heat transfer enhancement. Besides, the performance of micro-ribs is much more sensitive to the dimensional change. Therefore, micro-ribs have larger working excursion for enhancing performance factor than dimples. With optimal structure parameters: hr = e = 0.3 mm, pr = 3 mm, hd = 0.225 mm, dd = 1.5 mm, overall performance factor in micro-ribbed channel is 1.14 times of that in dimpled channel. Those conclusions provide effective support for optimization design for the regenerative cooling structures.

Supercritical CO2 (SCO2) is regarded as a promising supplementary coolant for traditional regenerative cooling at extremely high flight speed (Ma ≥8) due to its superior heat and mass transfer capability. Using sCO2 as the coolant, mini-ribs is introduced to the rectangular regenerative cooling channel to solve the difficulties in transferring enough heat from near walls to core mainstream at a heat flux of 5 MW/m2 in this paper. With the thermophysical properties near the supercritical point interpolated in the solver, flow structures and heat transfer performance in the cooling channels are obtained using a validated k-ω SST turbulence model. Effects of pitch ratios, channel materials and acceleration and buoyancy effects are also investigated in terms of the local bulk heat transfer coefficients, Fanning friction factors and entropy generation numbers by heat transfer and fluid flow. Overall, the existing of mini-ribs improves the heat transfer coefficient and reduces the entropy number by heat transfer at the cost of the increased fluid friction and the corresponding entropy number. Compared with the smooth channel, the overall enhanced factor can reach 1.73 with an overall thermal performance factor of 1.43 and the maximum wall temperature decreases by 27.57 % in the case of P/e = 10. When the solid channel material has higher thermal conductivity, it leads to a more uniform temperature distribution and lower entropy generation number by heat transfer with the enhanced transverse heat conduction within solid walls which also leads large heat transfer coefficients. Larger accelerations result in stronger heat transfer and larger fluid friction by strengthening the buoyancy effect which is more obvious when the state approach the critical point. This work provides a good reference for the application of supercritical CO2 in a regenerative cooling channel of scramjet engine at extremely high heat flux.

A comparative study of supercritical heat transfer in a regeneratively cooled rocket engine was conducted using three-dimensional numerical simulations for two channel geometries: rectangular and square. Various constant heat fluxes, flow velocities, and operating pressures were imposed to study their effect on heat transfer, pressure losses, and the conditions under which heat transfer deterioration HTD occurs. The results show that a rectangular channel is more efficient in terms of heat transfer than a square channel, with a more pronounced difference at high heat fluxes and low velocities, these conditions, in fact, favored the occurrence of heat transfer deterioration, particularly in the square channel. Increasing the flow velocities to reduce the wall temperature and prevent thermal deterioration was accompanied by a significant increase in pressure losses, these pressure losses are greater in the rectangular channel, despite its advantages in terms of heat transfer. Operating pressure also plays an important role in heat transfer, increasing the pressure results in a decrease in wall temperature.

In this study, the flow and heat transfer of supercritical hydrogen flowing inside a rocket engine cooling channel with arc ribs and variable heat flux is presented using numerical simulations. The arc ribs are examined in straight and thrust chamber-shaped channels. The results indicate that channel shape and rib geometry play an important role in the heat transfer. The highest performance evaluation criterion η is recorded for a channel with 90° arc ribs of h r / D H = 1/8 in both convex and concave orientations. Compared with a smooth channel, the performance of a channel with 90° convex arc ribs is 13–62% higher (depending on location) for a straight geometry, but it is only 14.7–27.7% higher in a thrust chamber-shaped geometry. The thermal stratification is moderate in the straight channel, but it is intense in the thrust chamber-shaped channel because of the onset of flow separation at the junction of the convergent-divergent section. • Supercritical H 2 heat transfer in straight and thrust chamber-shaped cooling channels with arc ribs are studied. • Under varying heat flux, 90° arc ribs outperform smooth channels by 13%–62% in straight a channel. • In contrast, the 90° arc ribs are considerably less effective in thrust chamber-shaped channels. [ABSTRACT FROM AUTHOR]

Analyzing thermal liquid-based cooling systems is crucial for improving performance in fields such as fuel cells, energy storage, and spacecraft. For this purpose, we are going to simplify the 1D archetype model, considering Kapitza resistance, which aids rapid cooling system prototyping. To simplify the understanding of engine cooling physics, the regenerative coolant is used as an efficient liquid coolant. We have found that the fluctuations in the Nusselt number are crucial for precise thermal predictions using the continuum model. Moreover, the temperature and the coolant's exit temperature are notably influenced by the channel's geometry, and there exists an inverse relationship with the Nusselt number. Furthermore, considering temperature jump and Kapitza length, which relates to surface roughness, is essential, especially at the nanoscale. These factors are crucial for the precise modeling of cooling systems, and our parametric research template, grounded in the continuum model, provides valuable insights to enhance modeling. [ABSTRACT FROM AUTHOR]

A comparative study of supercritical heat transfer in a regeneratively cooled rocket engine was conducted using three-dimensional numerical simulations for two channel geometries: rectangular and square. Various constant heat fluxes, flow velocities, and operating pressures were imposed to study their effect on heat transfer, pressure losses, and the conditions under which heat transfer deterioration HTD occurs. The results show that a rectangular channel is more efficient in terms of heat transfer than a square channel, with a more pronounced difference at high heat fluxes and low velocities, these conditions, in fact, favored the occurrence of heat transfer deterioration, particularly in the square channel. Increasing the flow velocities to reduce the wall temperature and prevent thermal deterioration was accompanied by a significant increase in pressure losses, these pressure losses are greater in the rectangular channel, despite its advantages in terms of heat transfer. Operating pressure also plays an important role in heat transfer, increasing the pressure results in a decrease in wall temperature. [ABSTRACT FROM AUTHOR]

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. [ABSTRACT FROM AUTHOR]

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.

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. • A finite volume model system for system-level transient transcritical fluid flow. • Fundamental equation of state. • Transonic flow, transcritical flow, heat transfer, and combustion multi-physics coupled. • Modeling regenerative cooling thrust chamber of a LOX/methane rocket engine. [ABSTRACT FROM AUTHOR]

Regenerative cooling is of great significance to secure the thermal structure and promote the flight Mach number range of the SCRamjet. Interconnection structure (ICS) plays a key role in improving the coolant flow distribution and heat sink utilization. In this work, the flow and heat transfer behaviors of the hydrocarbon fuel in parallel regenerative cooling channels with ICS are numerically investigated. The ICS improves the flow distribution and alleviates the local heat transfer deterioration. The influences of ICS configuration mainly consist of two aspects: (a). inter-channel pressure communication; (b). transverse mass transfer. The maximum wall temperature falls by 117.48 K/6.96% with the ICS introduced. Different sizes and positions of ICS are also studied. ICS with too small size cannot provide enough space for pressure communication and transverse mass transfer. While ICS with too large size leads to local heat transfer deterioration. The optimal Ф value to achieve the lowest heated wall temperature is Ф = 5 in this work. Regarding the position of ICS, it affects the local heat transfer deterioration through flow distribution and thermal load distribution. Ps = 50% (ICS locates at the middle of the heated section) presents the optimal cooling effect in this work. At last, the ICS configurations are universal to different heat flux distributions. The maximum wall temperature (Case qf2) decreases by 137.72 K/8.15% compared with Case C1 (without ICS). [ABSTRACT FROM AUTHOR]

Regenerative cooling by endothermic hydrocarbon fuel (EHF) is one of the most promising techniques for thermal management of supersonic or hypersonic aircraft. How to maintain the fuel working in proper states is an important issue to maximize the cooling potential of EHT. This work proposes a novel working state map, including risking zone (RZ), thermal cracking zone (TCZ), supercritical zone (SupZ) and subcritical zone (SubZ), to differentiate possible working states of an EHF during regenerative cooling. Using n-decane flowing in a circular tube as an example, the boundaries of four zones are determined by numerical simulation covering different heat fluxes (0.2–4.0 MW·m−2) and mass flow rates (0.5–10.5 g·s−1) under two operating pressures (3.45 and 5.00 MPa). Empirical correlations for three boundary lines are obtained and the maximum cooling capacity is identified, as well as the identification of the pressure effect. The revelation of such new perspective of regenerative cooling is of great implication to the design and optimization of cooling system for future thermal management.

A one-dimensional model of the cooling process occurring in a hybrid-rocket nozzle working with cryogenic oxygen flowing through helical channels is addressed in this work. Although the regenerative cooling system in rocket engines involves further complexity, it is here investigated as an option for the suppression of graphite-nozzle erosion in hybrid rockets. The model is based on the solution of the conservation equations of mass, momentum, and energy of the coolant. The Modified Benedict-Webb-Rubin equation of state was used for its accuracy in both the fluid phases of oxygen. The conservation equations for the coolant flow are coupled with a one-dimensional heat transfer model for the evaluation of the nozzle wall temperature at both the cold and hot sides, which has allowed assessing the developed model prediction capabilities by means of the data collected from three engine hot firings. Thus, parametric analyses were carried out to show the effect of the number of helical channels and evaluate the performance improvement obtained in supercritical conditions. • Study of a hybrid rocket nozzle cooling system based on cryogenic oxygen flowing through helical cooling channels. • 1D modelling of the nozzle cooling with cryogenic oxygen flowing through helical channels. • Validation on firing tests by transient simulations. • Effect of the number of channels in a helical configuration. [ABSTRACT FROM AUTHOR]

This study presents results of predicting the transport properties of hydrocarbon aviation fuels and their decomposed products after pyrolysis. Twenty-seven pure substances and two types of mixture, including both low and high molecular weight hydrocarbons as well as hydrogen, are considered. The specified temperature and pressure ranges, 300 to 1000 K and 0.1 to 5.0 MPa, respectively, correspond to representative operating conditions of a hydrocarbon aviation fuel that circulates as a coolant in the regenerative cooling system of a hypersonic vehicle and include the critical temperatures and pressures of most of the hydrocarbon fuels of interest. Four methods are adopted for the prediction of viscosity and thermal conductivity; the Brule-Starling method is used to predict viscosity, the Modified Propane TRAPP method for thermal conductivity, and the Methane TRAPP, Propane TRAPP, and Chung et al. methods are used for both transport properties. A comparison of the total average values concludes that the Chung et al. and Brule-Starling methods perform best in predicting the viscosity of all substances ranging from hydrogen to high molecular weight hydrocarbons in the temperature and pressure ranges specified in the current study. The quantified comparison by the total average also confirms that the Modified Propane TRAPP method best predicts the thermal conductivity of all of the 29 substances over the set temperature and pressure ranges, although the Propane TRAPP and Chung et al. methods offer a similar level of accuracy. [ABSTRACT FROM AUTHOR]

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. [ABSTRACT FROM AUTHOR]

At an extremely high Mach number, the regenerative cooling of traditional kerosene cannot meet the requirement of the heat sink caused by aerodynamic heating and internal combustion in a scramjet propulsion system. As a supplement of traditional regenerative cooling, supercritical CO2 is regarded as an effective coolant in severe heating environments due to its excellent properties of heat and mass transportation. In this paper, the heat transfer and flow structure characteristics of regenerative cooling in a rectangular channel using supercritical CO2 are analyzed numerically using a validated model. The effect of heat flux magnitude, nonuniform heat flux, acceleration and buoyancy and flow pattern are considered to reveal the regenerative cooling mechanism of supercritical CO2 in the engine condition of a scramjet. The results indicate that the heat transfer deterioration phenomenon becomes obvious in the cooling channel loaded with relatively high heat flux. Compared with the cooling channels loaded with increased heat flux distribution, the maximum temperature increased for the channel loaded with decreased heat flux distributions. When larger acceleration is applied, a relatively lower wall temperature distribution and higher heat transfer coefficients are obtained. The wall temperature distribution becomes more uniform and the high-temperature region is weakened when the coolants in adjacent channels are arranged as a reversed flow pattern. Overall, the paper provides some references for the utilization of supercritical CO2 in regenerative cooling at an extremely high Mach number in a scramjet. [ABSTRACT FROM AUTHOR]

During the cracking process of hydrocarbon fuel, the heat would be converted into internal energy and chemical energy. It is necessary to study the conversion relationship between internal energy and chemical energy in the above-mentioned energy conversion process, which is of great significance for the efficient use of fuel heat sinks in the regenerative cooling process. Therefore, the influences of pressure, mass flow rate, wall heat flux, and heating rate on the energy conversion were investigated, and an empirical prediction model was developed. The numerical results revealed that the influence of pressure on heat sink mainly focuses on the low and medium conversion. The chemical heat sink decreases with the increase of the mass flow rate and the heating rate, while the chemical heat sink increases with the increase in the wall heat flux. In addition, based on the analysis of the numerical data, an energy conversion correlation is developed by the Buckingham PI theorem, and most of the data calculated by this correlation can satisfy the deviation bandwidth of ±10%. [ABSTRACT FROM AUTHOR]

Regenerative cooling using hydrocarbon fuel requires thermal decomposition to increase heat absorption, but a high temperature is required to initiate decomposition and catalysis is expected to enhance decomposition at lower temperature regime. The effects of catalysis on thermal decomposition and coking of normal-dodecane and methyl-cyclo-hexane within a tube heater were experimentally investigated. Platinum (Pt) was spread on the inner surface of a stainless tube, and the inner surface of the tube was heated to target values using an electric heater. Decomposed components were accumulated and analyzed, and heat input was evaluated based on temperature gradient within the stainless tube. The results were compared with those without tube coating to show that the Pt-coating lowered the onset temperature of decomposition by 100 K for both fuels. Additionally, heat absorption per decomposed fuel amount was affected by the Pt-coating and the Pt-coating sizably reduced coking generation. [ABSTRACT FROM AUTHOR]

Accurate calculations of the heat transfer and the resulting maximum wall temperature are essential for the optimal design of reliable and efficient regenerative cooling systems. However, predicting the heat transfer of supercritical methane flowing in cooling channels of a regeneratively cooled rocket combustor presents a significant challenge. High-fidelity CFD calculations provide sufficient accuracy but are computationally too expensive to be used within elaborate design optimization routines. In a previous work it has been shown that a surrogate model based on neural networks is able to predict the maximum wall temperature along straight cooling channels with convincing precision when trained with data from CFD simulations for simple cooling channel segments. In this paper, the methodology is extended to cooling channels with curvature. The predictions of the extended model are tested against CFD simulations with different boundary conditions for the representative LUMEN combustor contour with varying geometries and heat flux densities. The high accuracy of the extended model's predictions, suggests that it will be a valuable tool for designing and analyzing regenerative cooling systems with greater efficiency and effectiveness. [ABSTRACT FROM AUTHOR]

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. [ABSTRACT FROM AUTHOR]

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. [ABSTRACT FROM AUTHOR]

This study developed a combustor with silicon carbide (SiC) microchannels cooling (SiC MC) at the walls by the motivation of the excellent physical properties of SiC and its promising application in aerospace areas. Its thermal and combustion performance was experimentally explored by the comparison of a combustor with stainless steel microchannel cooling (STS MC). Results showed that the combustor with SiC MC induced much more complete combustion, an 18–20% increase in combustion efficiency, and a 5.3–5.6% decrease in CO emissions than the STS MC counterpart. Besides, the effects of coolants on the performance of combustor with SiC MC were also evaluated. [ABSTRACT FROM AUTHOR]

Thermal protection is one of the crucial issues for the advanced propulsion systems of Reusable Launch Vehicles. New service requirements for materials, such as high strength, low density, low thermal expansion, high thermal stability, etc., are raised for the thermal structure with the increasing demand of flight Mach numbers and thrust‐to‐weight ratio. Carbon fiber‐reinforced ceramic composites, which generally meet the aforementioned requirements, show great potential for various applications and they have been widely applied in the thermal protection for hypersonic vehicles. This paper gives a comprehensive and systematic review of current research status for carbon fiber‐reinforced ceramic composites applied in the thermal structure of advanced propulsion systems. Three aspects are presented and discussed: the ceramic composites fabrication and the property characterization, the thermal performance of composite thermal structure used in practical engines, and the numerical methods for predicting mechanical and thermal properties of composites as well as the physicochemical phenomenon in the cooling channels. Firstly, the main manufacturing processes for the carbon‐reinforced ceramic composites are presented and the corresponding advantages and disadvantages are analyzed. The high‐temperature oxidation and ablation behaviors of composites are demonstrated and the improvement of oxidation and ablation resistance by introducing the ultra‐high‐temperature ceramics into C/C composites is discussed in detail. Then, several typical applications of carbon fiber‐reinforced ceramic composites (mainly C/SiC), including the work of RCI, JAXA and NASA, have been reviewed and analyzed. After that, the current research status of macroscale equivalent and multiscale numerical methods for predicting the mechanical and thermal properties of composites as well as the complex physicochemical phenomenon occurring in hydrocarbon fuels are sorted out. Finally, several potential prospects are pointed out for the future research on the thermal protection of advanced propulsion systems based on the carbon fiber‐reinforced ceramic composites. [ABSTRACT FROM AUTHOR]

The feasibility of regenerative cooling technology in scramjet engines has been verified, while the heat transfer behavior involved in the process needs further study. This paper expounds on the necessity of coupled heat-transfer analysis and summarizes its research progress. The results show that the effect of pyrolysis on heat transfer in the cooling channel depends on the heat flux and coking rate, and the coupling relationship between combustion and heat transfer is closely related to the fuel flow rate. Therefore, we confirm that regulating the cooling channel layout according to the real heat-flux distribution, suppressing coking, and accurately controlling the fuel flow rate can contribute to accomplishing the optimal collaborative design of cooling performance and combustion performance. Finally, a conjugate thermal analysis model can be used to evaluate the performance of various thermal protection systems. [ABSTRACT FROM AUTHOR]

In this paper, a novel two-stage hydrocarbon fuel heating and delivery system was used to more accurately simulate the state of fuel in the cooling channel of the regenerative cooling engine. Different cracking degrees (9–23%) experiments of JP-10 were conducted by using the new heating system, and cracked products were sampled and analyzed. The ignition delay time of high cracking degree product is shorter than that of low cracking degree product. The laminar flame speed of JP-10 and low cracking degree products are similar, the laminar flame speed of high cracking degree products is 32% faster than that of JP-10 when the equivalence ratio is 1.0, and the laminar flame speed of surrogate gas fuel is two times faster than that of JP-10. The supersonic combustion experiments of JP-10 with different cracking degrees and equivalence ratios were also employed at Mach number 3.0 of the isolator entrance, with total temperature ∼1550 K, total pressure ∼1.5 MPa, as well as at Mach number 2.5 of the isolator entrance, with total temperature ∼1250 K, total pressure ∼0.78 MPa, respectively. The static pressure along the axis of combustor was measured, while the velocity and temperature of the airflow at the exit of combustor were also measured by the tunable diode laser absorption spectroscopy (TDLAS). The distributions of airflow velocity, total temperature and static temperature along the combustor with different conditions were analyzed through a one-dimensional analysis method. The deviations of velocity and temperature of the airflow are less than −4% and −5% for Mach number 3.0, and less than 9% and 4% for Mach number 2.5, which indicates the calculations are in accordance with the TDLAS detections to some extent at Mach 3.0. Specially, supersonic combustion of JP-10 with different cracking degree at the same fuel temperature and equivalence ratio were obtained, which benefits studying the influence of different cracking degrees fuel on supersonic combustion characteristics. Furthermore, the experimental data will be helpful for validating the simulation of supersonic combustion. • The supersonic combustion properties of JP-10 with different cracking degrees, and as well as the influence mechanism of the small amount of pyrolysis components on the supersonic combustion of the entire fuel were studied rarely. • The ignition delay time and laminar flame speed of the different cracking degrees products and surrogate gas fuel were calculated and compared, and it was found that the results of these fuels were significantly different. • Different cracking degrees products at the same fuel temperature and flowrate were obtained by using two kinds of two-stage hydrocarbon fuel heating and delivery system. • The deviations of velocity and temperature of the airflow at exit of the combustor between the one-dimensional calculation results and the TDLAS results are less than -4% and -5% for Mach number 3.0, and less than 9% and 4% for Mach number 2.5, respectively. [ABSTRACT FROM AUTHOR]

This paper introduces the concept of a thermal management system (TMS) with integrated on-board power generation capabilities for a Mach 8 hypersonic aircraft powered by liquid hydrogen (LH2). This work, developed within the EU-funded STRATOFLY Project, aims to demonstrate an opportunity for facing the challenges of hypersonic flight for civil applications, mainly dealing with thermal and environmental control, as well as propellant distribution and on-board power generation, adopting a highly integrated plant characterized by a multi-functional architecture. The TMS concept described in this paper makes benefit of the connection between the propellant storage and distribution subsystems of the aircraft to exploit hydrogen vapors and liquid flow as the means to drive a thermodynamic cycle able, on one hand, to ensure engine feed and thermal control of the cabin environment, while providing, on the other hand, the necessary power for other on-board systems and utilities, especially during the operation of high-speed propulsion plants, which cannot host traditional generators. The system layout, inspired by concepts studied within precursor EU-funded projects, is detailed and modified in order to suggest an operable solution that can be installed on-board the reference aircraft, with focus on those interfaces impacting its performance requirements and integration features as part of the overall systems architecture of the plane. Analysis and modeling of the system is performed, and the main results in terms of performance along the reference mission profile are discussed. [ABSTRACT FROM AUTHOR]

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

As a promising active cooling method for aero engines, regenerative cooling with endothermic hydrocarbon fuel plays an important role in thermal protection systems. However, simply assuming the thermal boundary conditions for the numerical simulation of regenerative cooling as either Dirichlet or Neumann boundary conditions is often unsuitable, as the combustion process exhibits complex three-dimensional (3D) characteristics. We propose a novel multi-step iterative calculation method in this work to simulate the coupled flow and heat transfer between supersonic combustion in the combustor and regenerative cooling outside. The method captures the main features of the supersonic combustion including different shockwave structures, and generates a realistic 3D distribution of heat flux as the boundary condition for the regenerative cooling study. With improved simulation, the result shows that regenerative cooling significantly reduces the peak temperature of the combustor by 58.5 % with a mass flow rate of 1 g/s. A local high-temperature region is observed on the coupled wall inside the combustor near the injector, where the flame is primarily concentrated. Increasing the mass flow rate by 100 %, from 1 g/s to 2 g/s, leads to a reduction in the maximum wall temperature by up to 15.7 %. • A decoupling method for bi-directional combustion-cooling coupling is proposed. • Coupled heat transfer performance for combustion and regenerative cooling is analyzed. • Key features of supersonic combustion process are captured and validated. • Realistic 3D heat flux distributions for regenerative cooling simulation are obtained. • The impact of mass flow rate on cooling performance and thermal cracking is evaluated. [ABSTRACT FROM AUTHOR]

Supercritical CO 2 closed-Brayton-cycle demonstrates considerable application potential on hypersonic vehicles, effectively addressing power supply issues during prolonged flights. However, the CO 2 flow rate is strictly limited by the cycle's cooling source (aviation kerosene), rendering traditional regenerative cooling inadequate for sufficiently cooling the scramjet engine. To address this challenge, this study proposes an integrated cooling scheme that combines regenerative cooling, ceramic thermal-protective layer, and fuel film cooling. A two-dimensional coupled numerical model was developed, simultaneously considering the convective heat transfer of supercritical CO 2 and the shear mixing of supersonic fuel film. The results indicate that the high-enthalpy gas imposes an extreme aerodynamic heat load on the regenerative cooling channel wall, resulting in a notable deterioration in heat transfer for supercritical CO 2. Implementing kerosene film cooling within the combustor can effectively mitigate the abnormal heat transfer behavior of supercritical CO 2 and significantly reduce viscous dissipation heat within the gas boundary layer. By optimizing the fuel film injection layout, this combined cooling scheme reduced wall heat flux by 43 % and temperature by over 400 K compared to traditional regenerative cooling, thus providing effective thermal protection for the scramjet engine. • Coupled numerical simulation between S-CO 2 cooling channel and scramjet is achieved. • Heat transfer deterioration occurred in the S-CO 2 cooling channel under aero-heating. • Even with a ceramic insulation layer, S-CO 2 cooling channel can't cool the scramjet. • Supersonic fuel film mitigates the heat transfer deterioration in the S-CO 2 channel. • Double fuel film spraying combined with S-CO 2 channel is an effective cooling scheme. [ABSTRACT FROM AUTHOR]

In the design of the nuclear rocket engine nozzles, the heat in the throat area of nozzles is so intense that the temperature may exceed the endurance of existing materials. Regenerative cooling technology is a widely employed method for effectively cooling the nozzle. The influences of a strategy involving the arrangement of ribs in the regenerative cooling channel are investigated. The flow and heat transfer characteristics in the regenerative cooling channels with three different types of ribs are compared and analyzed. In accordance with the computational results, two composite structures are proposed. Moreover, the flow and heat transfer characteristics in the bending channel with these two composite structures are calculated and analyzed. The results show that the straight and streamlined ribs can reduce the maximum temperature only when the height is lower than 1 mm. The side rib can reduce the temperature near the rib. For the composite rib, when the distance between the side rib and bottom rib is ranging from 3 cm to 5 cm, the appearance of the high temperature zone, which is located near the bottom rib can be prevented. Moreover, in the bending channel, the straight composite rib can considerably reduce the maximum temperature in the throat area by 96 K. However, the streamlined composite rib can only reduce the wall temperature near the rib. • The streamlined rib and the side rib are proposed and examined. • Two new composite rib structures are designed and analyzed. • The optimal distance of composite rib is ranging from 3 cm to 5 cm. • The straight composite rib can enhance heat transfer in the bending channel. [ABSTRACT FROM AUTHOR]

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.

This study presents results of predicting the transport properties of hydrocarbon aviation fuels and their decomposed products after pyrolysis. Twenty-seven pure substances and two types of mixture, including both low and high molecular weight hydrocarbons as well as hydrogen, are considered. The specified temperature and pressure ranges, 300 to 1000 K and 0.1 to 5.0 MPa, respectively, correspond to representative operating conditions of a hydrocarbon aviation fuel that circulates as a coolant in the regenerative cooling system of a hypersonic vehicle and include the critical temperatures and pressures of most of the hydrocarbon fuels of interest. Four methods are adopted for the prediction of viscosity and thermal conductivity; the Brule-Starling method is used to predict viscosity, the Modified Propane TRAPP method for thermal conductivity, and the Methane TRAPP, Propane TRAPP, and Chung et al. methods are used for both transport properties. A comparison of the total average values concludes that the Chung et al. and Brule-Starling methods perform best in predicting the viscosity of all substances ranging from hydrogen to high molecular weight hydrocarbons in the temperature and pressure ranges specified in the current study. The quantified comparison by the total average also confirms that the Modified Propane TRAPP method best predicts the thermal conductivity of all of the 29 substances over the set temperature and pressure ranges, although the Propane TRAPP and Chung et al. methods offer a similar level of accuracy.

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At extremely high Mach number (Ma ≥8), kerosene is faced with issues of cracking with a limited heat sink for regenerative cooling. Supercritical CO 2 can be used as additional cooling method for regenerative cooling because of its excellent heat and mass transfer capability and it can easily convert heat into electricity for the engine electric system. In this study, pin-fins are applied to a regenerative cooling channel using sCO 2 to further enhance heat transfer at extremely high heat flux. Heat transfer and fluid flow are analyzed by the k-ω SST model considering effects of pitch ratio, solid materials and accelerations. From this study, compared with a smooth cooling channel, the pin-fin channel (Case 3) obtains a heat transfer enhancement of 3.08, a friction factor of 4.66, thermal performance enhancement of 1.84, and the maximum temperature of the heated surface is decreased by 36 % at Re = 45,000. The maximum velocity is found at the near-wall regions determined by the combined effects of temperature difference and accelerations. When the channel material is Cu with the high thermal conductivity, the maximum temperature is decreased by 37 % compared with a steel channel and the temperature distribution also becomes more uniform. • Supercritical CO2 is introduced to regenerative cooling. • Pin-fin structures are applied further enhance heat transfer. • Heat transfer enhancement by 3.08 and maximum temperature decreased by 36 %. [ABSTRACT FROM AUTHOR]

A harsh and complex thermal environment in the combustor threatens safe working of scramjets. In this study, heat transfer and flow structures of supercritical n-decane under 3 MPa in a regenerative cooling channel loaded with non-uniform heat flux distributions are investigated, including uniform, sinusoidal, increased, and decreased heat flux distributions. A verified k–ω SST turbulence model was employed, and a corresponding mesh independence study was performed. From this work, the fluid temperature at the outlet of the heated channel is only determined by the averaged heat flux, and all the regenerative cooling channels achieve the same temperature although loaded with different heat flux distributions. Compared with the fluid temperature, the wall temperature distribution is more sensitive to the variations of heat flux distribution. The regenerative cooling channels loaded with the sinusoidal heat flux distributions exist in several high-temperature regions, and the channel loaded with linear distributions changes the trend of temperature distribution. A larger temperature gradient is found in the regenerative cooling channel wall with a lower thermal conductivity. This work provides a good insight into the characteristics of the flow and temperature field of regenerative cooling channels loaded with non-uniform heat flux considering the effect of conjugate heat transfer.

Under different gravity conditions, the buoyancy can influence different effects of regeneration cooling effect of aviation kerosene.The RNG k-ε turbulence three-dimensional model is established, which is verified by comparing the wall temperature obtained by other experiments and simulations under the same conditions.Aiming at the variation of temperature distribution, secondary flow velocity, convective heat transfer coefficient and turbulent kinetic energy induced by the buoyancy of the aviation kerosene RP-3 in the horizontal pipe under different gravity conditions, the effect of jet fuel buoyancy on the heat transfer flow is studied.The research results show that the increase of gravity can significantly increase the influence of buoyancy on convection, and the turbulent kinetic energy can increase the nonlinearly, and the convective heat transfer coefficient is significantly improved.When the gravitational conditions change, the secondary flow induced by buoyancy in the direction of gravity undergoes a complicated evolution.The convective heat transfer coefficient is firstly decreased, then increased, and finally decreased.When the temperature of aviation kerosene exceeds the supercritical temperature, the convective heat transfer coefficient is changed abruptly, which can cause the significant differences in the wall temperature.

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.

At an extremely high Mach number, the regenerative cooling of traditional kerosene cannot meet the requirement of the heat sink caused by aerodynamic heating and internal combustion in a scramjet propulsion system. As a supplement of traditional regenerative cooling, supercritical CO2 is regarded as an effective coolant in severe heating environments due to its excellent properties of heat and mass transportation. In this paper, the heat transfer and flow structure characteristics of regenerative cooling in a rectangular channel using supercritical CO2 are analyzed numerically using a validated model. The effect of heat flux magnitude, nonuniform heat flux, acceleration and buoyancy and flow pattern are considered to reveal the regenerative cooling mechanism of supercritical CO2 in the engine condition of a scramjet. The results indicate that the heat transfer deterioration phenomenon becomes obvious in the cooling channel loaded with relatively high heat flux. Compared with the cooling channels loaded with increased heat flux distribution, the maximum temperature increased for the channel loaded with decreased heat flux distributions. When larger acceleration is applied, a relatively lower wall temperature distribution and higher heat transfer coefficients are obtained. The wall temperature distribution becomes more uniform and the high-temperature region is weakened when the coolants in adjacent channels are arranged as a reversed flow pattern. Overall, the paper provides some references for the utilization of supercritical CO2 in regenerative cooling at an extremely high Mach number in a scramjet.