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Kerosene vs Methane: A Propellant Tradeoff for Reusable Liquid Booster Stages
Abstract
Kerosene and methane are two promising candidate propellants for a future reusable booster stage. This study assesses the merits of both propellants and compares their respective performance when used in a booster stage. First of all, the principal properties of both propellants are identified. An analysis of a comparable full-flow staged combustion cycle engine for each propellant follows. The final assessment is made based on the results of a performance analysis of a launch vehicle making use of these motors in reusable fly-back boosters. The use of kerosene as propellant leads to a lower booster dry mass, making it the preferred choice if no operational benefits of methane can be identified.
... The reaction velocity is defined as follows [14,15]: ...
... Method' was utilized, where the chemical affinity is supposed to be neglected at 'j' reactions [13,15]. ...
Combustion-chamber is a critical component of the propulsion engine, which is widelyused in the space industry and aeronautics. The goal of this article is to perform a numericalanalysis on the combustion process using a liquid-type propellant. The steps that must be followeduntil total combustion is achieved are emphasized. It concerns the fuel feeding phase, its injectionand the combustion operation. The amount of combustion products and the energy generated areevaluated. It has been shown that the liquid propellant may present an efficient alternative fuelthan the kerosene. In addition, the temperature of combustion does not exceed a certain limit toavoid structural problems in the chamber. The parametric survey allowed determining the range ofthe most influence factors, including the pressure, mixture richness, velocity and flow rates ofinjection for the fuel and oxidizer. The number and type of injectors revealed a considerableinfluence on the velocity and flow rates of injection. To maximize thrust force and systempropulsion, a careful selection of chamber material and ignition methods is required. A thorough inspection on the issues of walls cooling showed the necessary survey of maximum temperaturesthat may be reached during the combustion. Finally, an investigation of the thermal exchangethrough the walls will be very interesting.
... Other recent studies of different reacting flows have been introduced by Schneider et al. [15], Chikitkin et al. [16], Taghavi et al. [17], Zidane et al. [18] and Zhukov [19]. However, Paraffin and Kerosene chemical data comparisons were made with Thomas et al. [20], Wang [21] & Burkhardt et al. [22], respectively. ...
... However, the distinctions that might be found for ≠ 1 are negligible. Also, comparison with literature results shows good qualitative agreement (see [20][21][22]) in the context of propellants nature and nozzle performance. ...
This paper presents a simulation-based solution for calculating rocket engine performance with liquid-type propellants of Paraffin and Kerosene for oxidizer to fuel ratio that is given by a linear formula. The engine was divided into two main stages: combustion chamber and a nozzle. In the first phase, conditions were found in the combustion chamber, based on the assumption of equilibrium according to Barrere. Next, the flow in the nozzle was calculated based on the fluid in the combustion chamber. Three main theories were examined in order to find the flow conditions in the nozzle: Equilibrium, Frozen and Mixed flows (Bray conditions). While the latter assumes the existence of the "Sudden Freezing Point" found by Bray, so that from this point to the end of the nozzle, the flow is assumed to be frozen. The use of the proposed simulation might contribute for multiple calculations performance (e.g., fuels with multiple intermediate reactions). Comparison between both types of fuels/propellants for the three described types of flow is presented alongside CEA software results, whereas good agreement between solutions was found. Also, the greater the ratio between hydrogen and carbon atoms, the better the engine performance for a particular oxidizer. Finally, it was found that an equilibrium flow model throughout the nozzle has a better nozzle performance compared to the other types of flows.
... In addition, to increase the level of confidence in the numerical predictions, tools and models need to be validated with experimental data over the wide range of operating conditions that occur in rocket thrust chambers. Since methane shows good performance and cooling properties, as well as low-toxicity and is space storable, it is an attractive option for future space transport systems [2]. Where an approach based on equilibrium chemistry has been successfully applied for the simulation of hydrogen fueled rocket combustors [2] [3], a model based on the Flamelet assumption is more suitable for methane engines, due to the comparatively larger chemical time scales, which result in a deviation from equilibrium. ...
... Since methane shows good performance and cooling properties, as well as low-toxicity and is space storable, it is an attractive option for future space transport systems [2]. Where an approach based on equilibrium chemistry has been successfully applied for the simulation of hydrogen fueled rocket combustors [2] [3], a model based on the Flamelet assumption is more suitable for methane engines, due to the comparatively larger chemical time scales, which result in a deviation from equilibrium. However most commonly used Flamelet models do not include the effect of heatloss on the flame structure in terms of gas composition at low enthalpy levels. ...
In the present work a single-element rocket thrust chamber operated with gaseous methane (GCH4) and gaseous oxygen (GOX) is investigated numerically by employing the tabulated chemistry models of chemical equilibrium and Flamelet. Due to the low chemical reaction rates present in hydrocarbon combustion, non-equilibrium effects are needed for the correct description of the flame. Since the Flamelet model includes non-equilibrium effects in form of scalar dissipation, it is considered to be superior to the equilibrium chemistry model (ECM) in case of CH4/O2 chemistry. For this reason a comparison of the two models was undertaken and their differences were identified. Apart from the standard “frozen” Flamelet model approach, which cannot predict recombination effects close to the wall, a non-adiabatic model developed by the authors was implemented for the simulation of the test case. Significant differences between the frozen and non-adiabatic methods are observed, especially in the vicinity of the cold chamber walls. Although physically more motivating, the non-adiabatic model appears to over-predict the heat released due to recombination reactions in the boundary layer, thereby leading to a high heat flux. Moreover, the effects of multi-pressure tabulation and of turbulence-chemistry interaction (TCI) are investigated. It is found that multi-pressure tabulation is not needed for low pressure operating points (20 bar), whereas the absence of TCI over-estimates the temperatures in the chamber and TCI should therefore be included.
... The use of methane instead of kerosene solves the problems of soot formation and coking in cooling channels. In addition, methane has cheaper costs in production and storage, and better cooling properties compatible with liquid oxygen due to similar thermodynamic properties [1]. Methane is also a green propellant with low pollution to the environment and is safe to handle and store. ...
... Methane is also a green propellant with low pollution to the environment and is safe to handle and store. The rocket fuel tank size can be reduced due to the high density of methane as compared to hydrogen and a less complicated cooling system can be designed, thus providing more payload mass in return [1]. Therefore, methane is an excellent choice for upper stage and main stage engines. ...
A skeletal methane kinetic mechanism is developed for conditions relating to the combustion of undiluted methane-oxygen mixtures at high pressures. The new skeletal mechanism is based on the detailed mechanism of oxidation of alkanes by Zhukov (2009). The skeletal model has been created by eliminating unimportant species and reactions from the detailed mechanism. The reduction technique is based on the reaction path and sensitivity analyses. They allow one to determine the reactions and species that play important roles in combustion in rocket combustion chambers. The skeletal mechanism consists of 23 species and 51 reactions. The final and intermediate versions of the skeletal mechanism are compared with the parent detailed mechanism, with other reduced kinetic models and with experimental data on the ignition of methane at high pressures. This comparison shows that the developed skeletal mechanism has a better performance than other kinetic mechanisms in terms of accuracy and required computational power.
... Compared to kerosene, methane has several advantages for the application as a rocket engine fuel. Using the same engine cycle, the specific impulse of a LOX∕CH 4 motor is about 10 s higher than for a LOX/kerosene engine [1]. As a cooling fluid, methane can take off the same heat flux as kerosene while using only half of the cooling fluid mass flow rate [2]. ...
... The flow was considered to be 2-D axisymmetric. A combustion chamber ratio oxidizer to fuel (ROF) of 3.2 was chosen, which corresponds to real engine operation condition [1]. At the nozzle inlet, a reservoir-pressure boundary condition was applied, specifying the total pressure and density. ...
A numerical study is conducted to investigate the impact of different chemical reaction mechanisms on the behavior of reactive nozzle flow. Therefore, a 66-step chemical reaction mechanism for oxygen/methane combustion is implemented into German Aerospace Center’s flow solver TAU. Ignition delay simulations are conducted and compared to experimental data to demonstrate the validity of this implementation. The implemented 66-step baseline chemistry model is applied for generic nozzle flow simulations, and the results are compared to frozen nozzle flow and nozzle flow in chemical equilibrium in order to investigate the impact of the finite-rate approach. The 66-step baseline reaction mechanism is reduced to a seven-step basic configuration, which is applied to the generic nozzle flow. A good agreement of the 66-step and the seven-step model is observed. Both approaches are applied for Reynolds-averaged Navier–Stokes simulations of a dual-bell nozzle hot-gas flow. Almost no deviation between the 66-step baseline model and the reduced chemical seven-step approach is observed. The dual-bell transition behavior at different values of combustion chamber mixture ratio is investigated, applying Reynolds-averaged Navier–Stokes simulations with a reduced chemistry model. Validation data for the simulations are obtained during a hot-flow test campaign. The experimentally observed impact of the combustion chamber mixture ratio on the dual-bell transition nozzle pressure ratio is clearly reproduced by the numerical approach. A good agreement with the experimentally obtained, transition nozzle pressure ratio values is reached by the numerical simulations. A reduction of 93% of the computational cost is observed due to the reduction of the chemical reaction mechanism.
... With the development of space technology, the space propulsion field is focusing on finding easily accessible, reusable, and cost-effective fuels. Using methane as a rocket propellant provides a higher specific impulse than kerosene, has low tendency to coke and has superior cooling capacity [1][2][3], and methane is easier to store and transport than hydrogen [4]. In addition, the methane rocket engine also has the advantages of simple design, reusability and low production costs [5,6], these advantages make methane becoming a popular propellant in the aerospace propulsion field [7]. ...
Methane/Oxygen rocket engine is becoming one of the most promising rocket engines today due to its cost-effectiveness and reusability. In the design process of rocket engines, cooling system is a crucial part and film cooling is a very important method. The accurate prediction of heat transfer characteristics is crucial for the design and development of rocket engines using film cooling. In this paper, a numerical framework based on the Reynolds Averaged Navier-Stokes (RANS) method and the Eddy Dissipation Concept (EDC) reaction model is established, verified and applied to simulations of single- and multi-injector combustion chamber with film cooling. Besides, a single injector combustion experiment with film cooling is carried out to verify the numerical framework. The investigation indicates that flow and chemistry reactions near the wall coupled influence the wall heat load significantly, and the coupled wall function exploited by Direct Numerical Simulation (DNS) is modelled and embedded on the numerical frame in order to consider these coupling effects. The results of single injector chamber investigation show that by considering the chemical reactions near the wall the wall heat flux reduced 50% and agree much better to the experimental data, which indicates that coupled wall function is more effective at predicting wall heat flux than general wall functions in a chamber with film. In addition, the results also denote that the coupled wall function only acts in the near-wall region and has no effect on the main flow. Furthermore, after being verified in the single injector combustion chamber experiment, the numerical framework is applied to a multiple-injector case. The results indicate that the wall heat flux in multi-injector combustion chamber is 75% lower than the general wall functions, Afterwards, the effect of the film on the chemical enthalpy term near the wall, as well as the coupling effect of the turbulent flow and the temperature gradient near the wall are discussed. Finally, the analysis of the vorticity in the multi-injector chamber shows that the film weakens the vorticity in the front section of the combustion chamber, and subsequently affects the expansion of the flame.
... Methane can be the right candidate for next-generation fuel because of its cheaper cost, higher specific impulse and coking limits, and lesser carbon soot compared to kerosene [1]. In addition, it has attracted attention as fuel for Mars explorers, because it can be produced there. ...
SpaceX's successful development of reusable rockets and the realization of low-cost operations have significantly impacted the space industry, institutions, and companies. Price competitiveness has become a hot topic for launch vehicle development. A hydrogen-fueled rocket engine can be its solution. The developed countries are attempting to improve the performance and reliability of engines using a hydrogen-fuel. This paper summarizes the development and operation trends of hydrogen-fueled rocket engines of developed countries. It provides fundamental data for hydrogen-fueled rocket engine development, which is expected to be helpful in its future development.
... The availability of experimental data for estimation of wall temperature and code validation is even lower when new concepts based on less spread propellant combinations are considered. Among others, this is the case for the oxygen-methane propellant combination, which is catching much attention in recent years [8][9][10][11][12]. Methane is considered a valid low-cost replacement for other hydrocarbons like kerosene because of its higher specific impulse, cooling efficiency, and low level of coking and sooting, this last aspect being relevant especially for reusable launch vehicle applications. ...
The prediction of wall heat flux at the nozzle throat is of paramount importance in liquid rocket engine (LRE) design for both sizing and safety purposes. Computational fluid dynamics (CFD) simulations can aid in the prediction, provided that they can be effectively used during the design phase and that suitable modeling is employed. In this framework, this study aims at evaluating the suitability of a Reynolds-averaged Navier–Stokes-based CFD approach to predict in affordable times the nozzle wall heat flux of LREs employing the oxygen–methane propellant combination, which is nowadays attracting the attention of many developers. The interest to study the throat heat flux estimation for oxygen–methane engines comes from the known greater role played by the near-wall recombination reactions, as compared to the oxygen–hydrogen propellant pair. Nevertheless, only few indirect experimental measurements are available in the open literature for the validation of numerical tools. Recently published experimental data are used here as benchmark for the comparison of numerical simulations obtained with different assumptions. Results confirm that, for a well-designed engine, the details of injection and combustion processes have only a secondary effect on the prediction of throat heat flux.
... Methane, which is readily available, nontoxic, and low-cost, is the most promising propellant candidate for the future reusable rocket engines. 1,2 An in-depth understanding of the high-frequency combustion instabilities (HFCIs) in an O 2 /CH 4 rocket engine is of great importance for its high performance and reliability. HFCIs, also known as thermoacoustic instabilities, are a complex phenomenon widely observed in thermal systems, such as liquid rocket engines (LREs), [3][4][5] solid rocket engines, 6,7 and gas turbines. ...
Self-excited combustion instabilities of transverse modes were experimentally investigated in a rectangular multi-injector model combustor, operating with the bipropellants O2/CH4. The propellants were injected through a linear array of five oxidizer-centered shear coaxial injectors into the combustor. High-amplitude limit cycles obtained in hot-fire tests were analyzed in detail. Different combustion instability modes, including first and second width modes, were observed in cases with three different injection distribution schemes. Hence, the injection distribution strongly determined the combustion dynamics. One insight can be gained that the stable combustion could be achieved by properly designing the propellants' injection distributions.
... Several papers regarding hydrogen applications in aerospace propulsion can be found in the literature [5][6][7][8][9][10][11][12][13][14][15][16][17][18]. On the other hand, methane rose the interest of recent research and made its way to aerospace propulsion [19][20][21]. The main idea of this work is to investigate the feasibility of using methane as a fuel for a scramjet application. ...
The present work is an investigation of the reacting flow through a two-dimensional inlet-fueled scramjet configuration. Previous experimental and numerical studies showed the phenomenon of radical-farming to be responsible for the ignition of the hydrogen-air mixture. In order to reproduce the results of the previous studies, the reacting flow of a stoichiometric hydrogen-air mixture is studied using the same initial conditions. Keeping the same conditions, methane is then used as fuel instead of hydrogen. Results show that the radical-farming phenomenon is also responsible for the ignition of the methane-air mixture.
... With the rise of commercial aerospace activities, the demanding issues in terms of high operational and handling costs of cryogenic and storable propellants increased the attention for methane/oxygen in the development of future launch vehicles. Methane as a fuel can provide a higher specific impulse, together with better cooling abilities and less soot deposition than kerosene, therefore, methane/oxygen is the most promising propellant combination for the reusable rocket engine [1][2][3][4]. Compared to oxygen/hydrogen propellant combination, oxygen/methane can be considered as "space storable" and is favored by higher density, although it gives lower specific impulse [5]. ...
Methane/oxygen rocket engine is considered to be one of the most promising reusable rocket engines in future space activities. Adequate understanding and accurate prediction of heat transfer characteristics are considered key points for the development of reliable methane engines. In this paper, a methane combustion chamber with 7-elements is simulated using Reynolds averaged Navier-Stokes (RANS) method with Eddy Dissipation Concept (EDC) combustion model. The investigation reveals that the near-wall coupling effects of flow and chemistry have a significant influence on the wall heat load, and the coupled wall function developed by direct numerical simulation (DNS) is modified, validated, and incorporated in the RANS frame to consider the aforementioned coupling effects. The results show that the deviation of the wall heat load compared to experimental data is reduced from 25% to 5% for the wall of high temperature when chemistry effects are considered. The influence of the coupled wall function is limited near the wall and the properties of main flow are generally independent of the wall models adopted. The investigation also reveals that the turbulent flux of chemical enthalpy near the wall is comparable to the turbulent flux of sensible enthalpy in case of a methane combustion chamber. Finally, the effects of chemistry on the wall heat flux can be attributed to the coupling impacts of the chemical equilibrium shifting caused by the large temperature gradient near the wall and the non-uniform radial velocity brought by the powerful vortex system in the chamber.
... While there has been comprehensive research on combustion in hydrogen engines traditionally [5], [6], recent research activities have focused on developing analysis capabilities for methane engines. Methane shows good performance and cooling properties, as well as low-toxicity and space storability and is therefore an attractive option for future space transport systems [7]. ...
... Here, LH2, RP-1 and LCH4 display very different properties. Hydrogen is deemed an excellent regenerative coolant [15], and in comparison with kerosene, methane is showing superior cooling properties, higher coking limits and less soot deposition, with the latter two being especially important in the context of reusability [25]. Although methane's physical properties in the cooling channels are difficult to predict since it is operated in the trans-critical regime [26], CFD-data trained artificial neural networks have already proven to be able to quickly deliver predictions with sufficient accuracy [27]. ...
... Soot deposition and coking under fuel-rich conditions are drastically reduced compared with higher hydrocarbons, and the superior cooling properties and lower pressure drop in cooling channels make CH 4 attractive for regenerative cooling systems. A preliminary study involving an LO 2 ∕CH 4 motor concept estimated a higher specific impulse than a comparable LO 2 ∕kerosene engine [2] of about 10 s. Nevertheless, the fact that LO 2 ∕CH 4 cryogenic stages have not been employed for flight yet indicates a lack of knowledge in the field of high-pressure combustion involving hydrocarbons. ...
The design phase of a rocket combustion chamber relies on the prediction of the wall heat fluxes, in order to optimize the components to prevent thermal damage. Based on the experimental data of a subscale single-injector GCH4/GO2 combustion chamber at elevated pressure, this work thoroughly investigates the validity of a hybrid Reynolds–Averaged Navier–Stokes/Large Eddy Simulation turbulence model (namely, Improved Delayed Detached Eddy Simulation [IDDES]) combined with efficient flamelet tables. The generation of non-adiabatic flamelets is hereby optimized compared with previous studies. The effects of chemistry and enthalpy losses are first investigated on a simplified laminar test case. Then, the influence of the grid resolution on the wall thermal loads of the single-injector combustion chamber is discussed in the context of IDDES. The use of coarser meshes increases the turbulent content of the core flow, leading to an enhanced thermal load in the rear part of the chamber. Nonetheless, all investigated meshes yield a very good agreement with the experimental data, confirming the robustness of the IDDES/non-adiabatic chemistry solver. The result is confirmed by a preliminary simulation of the same combustion chamber featuring film cooling. The IDDES/non-adiabatic flamelet model is therefore recommended as a valid alternative to the widely used wall-stress-modeled large-eddy simulation.
... Methane is a kind of green fuel and easy to be produced, the propellant combination of methane and oxygen can provide high specific impulse, excellent cooling performance and low cost [1]. Therefore, LOX/Methane engines are considered to be the most promising candidate for the Reusable Launch Vehicle (RLV) propulsion system [2,3]. Several methane rocket engines, such as Raptor and BE-4, are under development for this ambitious reusable application. ...
A 7-element rocket combustion chamber using GOX/GCH4 as propellant has been modeled and simulated to get a more comprehensive knowledge for combustion and heat transfer process inside a combustion chamber. All the computational cases in our investigation use the Eddy-Dissipation Concept (EDC) combustion model for detailed chemistry and two equations RANS model for turbulence closure. The preliminary results of the base case without considerations on the reactions within the boundary layer show a good agreement with the experimental data in terms of pressure distribution, but the wall heat load is overestimated about 30%. Further investigation found that changing the turbulence model will alter the heat transfer characteristics significantly by delayed combustion process; changing the Prandtl number will tune the wall heat load slightly without changing the combustion field too much when using RANS model. It has been also found that the different reduced kinetic mechanism will slightly change the heat transfer characteristics through reducing the overall pressure level in the chamber by altering the thermodynamic properties of the combustion products. After carefully estimation, the boundary layer reactions for methane rocket chamber with multi-elements account for the most overestimation of the wall heat load.
... The effect of propellant density on propellant performance can be also understood from the fact that most hydrogen-oxygen engines use a higher mixture ratio than the ratio at which specific impulse is maximum [1]. A propellant performance comparison study on the Methane/LOX combination and Kerosene/LOX combination [3] has concluded that the effect of the increased specific impulse of the Methane/LOX combination was diminished because of its lower density and hence had identical propellant performance to that of the Kerosene/LOX combination. Sutton and Biblarz report that the specific gravity of a rocket propellant is influential on the maximum flight velocity and range of a rocket-powered vehicle flying within Earth's atmosphere [1]. ...
A strategy to predict and compare the performance of novel propellants by taking into account the effects of both specific impulse and propellant density was developed. A fixed propellant volume approach was adopted to account for the effect of propellant density. Propellant combinations consisting of two novel energetic compounds-Ditetrazolobishomocubane (DTetzBHC) and S4-2 were selected for comparison along with a baseline propellant combination. Six-degrees-of-freedom flight simulations were carried out for all selected propellants, and their performance was compared in terms of both the propulsion system and flight performance parameters, respectively. For validation purposes, the 1 st stage of the reference mission of the VEGA launch vehicle, Arianespace® was simulated and compared against the values documented in literature. Density-specific impulse was found to be a crucial performance parameter in deciding propellant performance. The propellant combination containing S4-2 was predicted to have superior performance among the cases compared.
... The propellant combination of oxygen/methane has many favorable characteristics; e.g., methane is six times as dense as hydrogen, is easier to handle, and has preferable coking temperature limits [7] and low toxicity. Furthermore, oxygen/methane offers a slightly higher specific impulse than oxygen/kerosene [8]. ...
Methane is considered being a good choice as a propellant for future reusable launch systems. However, the heat transfer prediction for supercritical methane flowing in cooling channels of a regeneratively cooled combustion chamber is challenging. Because accurate heat transfer predictions are essential to design reliable and efficient cooling systems, heat transfer modeling is a fundamental issue to address. Advanced computational fluid dynamics (CFD) calculations achieve sufficient accuracy, but the associated computational cost prevents an efficient integration in optimization loops. Surrogate models based on artificial neural networks (ANNs) offer a great speed advantage. It is shown that an ANN, trained on data extracted from samples of CFD simulations, is able to predict the maximum wall temperature along straight rocket engine cooling channels using methane with convincing precision. The combination of the ANN model with simple relations for pressure drop and enthalpy rise results in a complete reduced order model, which can be used for numerically efficient design space exploration and optimization.
... With respect to the design of reusable engines, the marginal soot emission of methane flames offers significant advantages in terms of unfavorable carbon depositions when compared to kerosene [8]. Similarly, the cooling capability of methane exceed those of kerosene due to reduced coking and lower pressure losses in the cooling channels [9]. ...
This work presents Large Eddy Simulations of a single-element GOx-GCH 4 combustion chamber at elevated pressure. A robust and efficient flamelet model accounting for heat losses is applied to represent the hydrocarbon combustion process at affordable computational effort. Three different near-wall treatments are used: wall-resolved LES, wall-modeled LES and a hybrid RANS/LES approach. The results are compared to experimental data in terms of wall heat fluxes, chamber pressure profile and optical OH ⁄ emission images. Generally, a good agreement with respect to the measurements is observed. In addition, different features of the flow field, like, e.g., the near-injector region and the axial development of the flame, are thoroughly discussed. Finally, a detailed a priori assessment of the different modeling assumptions – regarding the combustion model and relevant transport processes – is carried out to derive possible areas for future improvements.
... Among others hydrocarbons, methane characteristics are particularly promising. It offers high specific impulse (compared to kerosene for instance [1]), high density at common tank pressures (around 6 times the density of hydrogen), low pollution and low cost both for production and handling [2]. With the objective of reusable engines, the fuel cooling properties have also become one of the key parameters for the selection of hydrocarbon fuels. ...
The single element GCH4/GOx rocket combustion chamber developed at the Technische Universität München has been computed using Large Eddy Simulation. The aim of this work is to analyze the flow and combustion features at high pressure, with a particular focus on the prediction of wall heat flux, a key point for the development of reusable engines. The impact of the flow and flame, as well as of the model used, on thermal loads is investigated. Longitudinal distribution of wall heat flux, as well as chamber pressure, have been plotted against experimental data, showing a good agreement. The link between the heat released by the flame, the heat losses and the chamber pressure has been explained by performing an energetic balance of the combustion chamber. A thermally chained numerical simulation of the combustor structure has been used to validate the hypothesis used in the LES.
... D URING the last two decades, liquid oxygen (LOx) and methane (CH 4 ) experienced increasing attention as a possible propellant combination for rocket combustion due to advantages in handling, performance, and potential cost reduction [1,2]. In the United States, SpaceX and Blue Origin are developing LOx∕CH 4 engines and in Europe, the ArianeGroup is working on a reusable LOx∕CH 4 rocket motor for the future European launchers [3]. ...
Combustion models of different fidelity are applied to a seven-element gaseous methane/oxygen subscale rocket
combustion chamber. The covered region of thermo-chemical states is analyzed for two non-adiabatic flamelet-based
methods, from which one is well established and the other one has just recently been proposed for hydrogen/oxygen
combustion. Of particular interest is their applicability in situations with strong heat losses. Consequently, the
tabulated combustion models are linked with a presumed probability density function approach, and the results are
assessed using a transported probability density function method. An analytically reduced chemical reaction
mechanism with 13 species and 73 reactions is used for all models in order to allow for direct comparison of
turbulence–chemistry interaction. The results are evaluated with respect to temperature, gas composition, and wall
heat transfer and they are brought in context with their respective computational costs. Finally, the results are
compared to available experimental data in terms of wall heat flux and pressure measurements. Locally and
temporally resolved results provide additional insights to the experiment.
... Even more than Hydrazine, it combines with Oxygen from the air yielding dimethylamine and water. It reacts with CO 2 from the air to form (CH 3 ) 2 N-NH2 -CO 2 salts which are not soluble in Vehicles and Their Characterisation www.tjprc.org SCOPUS Indexed Journal editor@tjprc.org ...
Most of the rockets designed today employ chemical fuels as rocket propellant. Rocket propellant is a material utilized by a rocket as, or to create in a synthetic response, the response mass (propulsive mass) that is launched out, normally with fast, from a rocket motor to deliver push, and in this way give shuttle impetus. A decent force is unified with a fast of fumes gas discharge which infers a high burning temperature and fumes gases with little sub-atomic weights. This paper focuses on fuels of different types of rocket propulsion systems and compares their properties to understand which one of them gives the best result. Fuel's physical, chemical and thermal properties are targeted to understand their behaviour in various conditions to shortlist the best fuel for future rocket engine motors. Methane, Kerosene, Hydroxyl Ammonium Nitrate (HAN), Hydroxyl-terminate Polybutadiene (HTPB), Paraffin Wax, Hydrazine and Hydrogen are the Rocket propellants which is discussed in this paper.
... For LOX/methane thrust chambers, cryogenic methane is chosen as the coolant due to its superior cooling properties, higher coking limits, and less soot deposition. 7,8 Traditional cooling channels are milled in the copper liner and closed out by a nickel jacket through electroforming processes. 9 Quentmeyer 10 conducted experiments to explore the failure of the thrust chamber wall under cyclic work, and results indicated that a 'doghouse'-shaped failure characterized by the thinning and bulging of the inner wall appeared after several cycles. ...
To investigate the damage localization effects of the thrust chamber wall caused by combustions in LOX/methane rocket engines, a fluid-structural coupling computational methodology with a multi-channel model is developed to obtain 3-demensioanl thermal and structural responses. Heat and mechanical loads are calculated by a validated finite volume fluid-thermal coupling numerical method considering non-premixed combustion processes of propellants. The methodology is subsequently performed on an LOX/methane thrust chamber under cyclic operation. Results show that the heat loads of the thrust chamber wall are apparently non-uniform in the circumferential direction. There are noticeable disparities between different cooling channels in terms of temperature and strain distributions at the end of the hot run phase, which in turn leads to different temperature ranges, strain ranges, and residual strains during one cycle. With the work cycle proceeding, the circumferential localization effect of the residual strain would be significantly enhanced. A post-processing damage analysis reveals that the low-cycle fatigue damage accumulated in each cycle is almost unchanged, while the quasi static damage accumulated in a considered cycle declines until stabilized after several cycles. The maximum discrepancy of the predicted lives between different cooling channels is about 30%.
... LOX/Methane Liquid Rocket Engines (LRE) has aroused the interests of many countries in the past several decades. [1][2][3][4][5] Methane is a hydrocarbon which has many advantages, including low density, non-toxic, and easy to be produced. 6 The propellant combination of oxygen and methane can provide high performance, high reliability, and low cost. ...
To predict the thermal and structural responses of the thrust chamber wall under cyclic work, a 3-D fluid-structural coupling computational methodology is developed. The thermal and mechanical loads are determined by a validated 3-D finite volume fluid-thermal coupling computational method. With the specified loads, the nonlinear thermal-structural finite element analysis is applied to obtaining the 3-D thermal and structural responses. The Chaboche nonlinear kinematic hardening model calibrated by experimental data is adopted to predict the cyclic plastic behavior of the inner wall. The methodology is further applied to the thrust chamber of LOX/Methane rocket engines. The results show that both the maximum temperature at hot run phase and the maximum circumferential residual strain of the inner wall appear at the convergent part of the chamber. Structural analysis for multiple work cycles reveals that the failure of the inner wall may be controlled by the low-cycle fatigue when the Chaboche model parameter γ3 = 0, and the damage caused by the thermal-mechanical ratcheting of the inner wall cannot be ignored when γ3 > 0. The results of sensitivity analysis indicate that mechanical loads have a strong influence on the strains in the inner wall.
... In this paper, a numerical study has been conducted to examine the flow dynamics and heat transfer characteristics of a nanofluid, methane-CuO, at a supercritical pressure of 8 MPa. The base fluid, methane, which has a liquid-like density at a supercritical pressure, is chosen for the present study because it is a good alternative propellant in the reusable aerospace propulsion systems [3,23]. The present numerical study is based on a computational fluid dynamics (CFD) model that was developed and validated in our previous research works [5,13,17]. ...
... Rocket engines that employ liquid oxygen (LOX)/hydrocarbon as propellant are increasingly becoming popular as they tend to satisfy most of the aforementioned requirements in comparison to the engines that employ cryogenic propellants. 1 In such engines, LOX/methane (CH 4 ) is considered as a more appropriate option due to its high specific energy density, specific impulse and better overall performance. 2,3 Apart from the propellant, the injector technology also plays a crucial role in the development of a liquid rocket engine as it affects the performance by altering the flame and flow dynamics as well as the stability of the combustor. For example, the oxidizer injector length can lead to stable/unstable combustion behavior of a shear coaxial high-pressure combustor. ...
... Moreover, the heat transfer performance of methane is higher compared to other hydrocarbon fuels as a result of its high thermal conductivity, speci¦c heat, and low viscosity. In general, methane shows, compared to other potential candidates, better overall performance from a system point of view [5], higher speci¦c impulse [6], no risks for human health, simple extractability from natural gases, and a density 6 times higher than hydrogen when stored in liquid state at typical tank pressures. ...
... This high coking temperature prevents carbon deposits to form in the regenerative cooling channels, which are typically observed in kerosene fueled rockets. This makes liquid methane very suitable for reusable rocket engines [4]. As such, DARE wants to utilize liquid methane as fuel for the cryogenic liquid rocket engine project. ...
... The early studies on supercritical-pressure fluid flows and heat transfer focused mainly on water, carbon dioxide, hydrogen, and heavy hydrocarbon fuels [1][2][3][4][5][6][7][8][9][10][11][12][13], because of their practical/potential applications in nuclear reactors, refrigeration, aerospace propulsion and industrial power-generation systems. In recent years, because of the renewed interest in using cryogenic methane as an alternative propellant in aerospace propulsion systems [14,15], many studies on the supercritical-pressure fluid flows and heat transfer of cryogenic methane have been carried out, intended for fundamental understanding and practical applications of the regenerative engine cooling technology [16][17][18][19][20][21][22]. ...
... In fact, liquid methane as a rocket engine fuel has been recently considered as an interesting option for both space and launcher liquid rocket engines. 4,5 In such rocket engines methane is considered as the coolant and it will typically enter the cooling channels with a supercritical pressure and a subcritical temperature. As methane is heated up, due to the entering heat from hot-gas, its temperature passes through the critical value (190.56 ...
Flow modeling in regeneratively cooled rocket engines is a challenging task because of the high wall temperature gradient, the high Reynolds number, the high aspect ratio of the channel cross section and the curved geometry. If coolant is methane, a further complication is its near-critical operating condition. In this thermodynamic regime large changes of the fluid properties can greatly influence the coolant flow field and the heat transfer. In the present study numerical simulations of transcritical methane flow field in asymmetrically heated rectangular channel with high aspect ratio and strong wall temperature differences are carried out for both straight and curved channels (heated both on the convex and concave side) by means of a validated Reynolds Averaged Navier-Stokes solver for real fluids. Results are discussed in detail by comparison of transcritical and supercritical flow field, and comparison of a "full-scale" high-Reynolds number channel flow representative of the actual cooling channel geometry versus a scaled channel, low-Reynolds solution. Emphasis is given to the most critical case in terms of cooling performance, that is the curved channel heated on the convex side.
... 6,7 Experience gained during previous kerosene engine development efforts is also an important consideration for future launch systems. 8 (In the context of this paper, the term kerosene is used conventionally to describe the middle distillate range of hydrocarbons found in practical aerospace fuels, regardless of the original source.) Therefore, for subsonic and high speed flight as well as space access applications, kerosene fuels have enabled performance and operational requirements; this has occurred largely through identifying and specifying fuel formulations which meet given engine and fuel system needs. ...
Public reporting burden for this collection of information is estimated to average 1 hour per response, including the time for reviewing instructions, searching existing data sources, gathering and maintaining the data needed, and completing and reviewing this collection of information. Send comments regarding this burden estimate or any other aspect of this collection of information, including suggestions for reducing this burden to Department of Defense, Washington Headquarters Services, Directorate for Information Operations and Reports (0704-0188), 1215 Jefferson Davis Highway, Suite 1204, Arlington, VA 22202-4302. Respondents should be aware that notwithstanding any other provision of law, no person shall be subject to any penalty for failing to comply with a collection of information if it does not display a currently valid OMB control number. PLEASE DO NOT RETURN YOUR FORM TO THE ABOVE ADDRESS. Kerosene fuels possess physical and chemical properties which make them attractive for aerospace propulsion applications from operational and performance standpoints. However, variation in fuel properties and performance owing to differences in chemical makeup can be significant as operating environments and fuel composition fall outside the realm of current experience. Both circumstances are increasingly frequent, given the incorporation of new fuels in existing systems and a desire to increase vehicle performance. The Air Force Research Laboratory (AFRL) is actively engaged in deriving relationships between fuel composition, properties, and performance in realistic operating conditions. Ideally, these models will be implemented in the optimization of fuel composition to meet requirements for future systems. Moreover, current engine development activities prompt an assessment of as-supplied rocket kerosene, the set of requirements used for its specification, and the potential impacts of compositional variations on engine operability and performance. To address these needs, several lab scale RP-1 formulations were obtained which met specification requirements but were blended from chemically unique feedstocks, thereby representing the expected compositional variation for currently produced fuel. Chemical composition was characterized in terms of hydrocarbon types and was compared between the various formulations. Kerosene fuels possess physical and chemical properties which make them attractive for aerospace propulsion applications from operational and performance standpoints. However, variation in fuel properties and performance owing to differences in chemical makeup can be significant as fuel composition and operating environments fall outside the realm of current experience. Both circumstances are increasingly frequent, given the incorporation of new fuels in existing systems and a desire to increase vehicle performance. The Air Force Research Laboratory (AFRL) is engaged in deriving relationships between fuel composition, properties, and performance in realistic operating conditions. Ideally, these models will be implemented in the optimization of fuel composition to meet requirements for future systems. Moreover, current engine development activities prompt an assessment of as-supplied rocket kerosene, the set of requirements used for its specification, and the potential impacts of compositional variations on engine operability and performance. To address these needs, several lab scale RP-1 formulations were obtained which met specification requirements but were blended from chemically unique feedstocks, thereby representing the expected compositional variation for currently produced fuel. Chemical composition was characterized in terms of hydrocarbon types and was compared between the various blends. Several property measurements provided insight to compositional influence on fuel behavior; reported in this paper are composition explicit distillation curve, density, viscosity, heat of combustion, and hydrogen content. While chemical variability for RP-1 was not as extensive as that of jet fuel, the sensitivity of several properties to feedstock selection was demonstrated, even for fuels which met specification requirements.
The injectors near-field region of LRE combustion chambers is numerically investigated by means of both multi and single-injector combustor simulations. An efficient and validated uRANS numerical framework based on non-adiabatic flamelets is used to describe methane-oxygen turbulent nonpremixed combustion. Two paradigmatic configurations, representing the flame-wall and flame-flame interactions occurring on an injection plate are used to develop a 2D axy-symetric dataset. On both configurations, the confinement length of the injector is parametrically varied in order to isolate its effect on the heat flux insisting on the injector face-plate. Finally the results are compared to a fully 3D case, consisting in 37 injectors characterized by the same operating conditions and parameters as the 2D cases.
View Video Presentation: https://doi.org/10.2514/6.2022-3275.vid The prediction of wall heat flux at the nozzle throat is of paramount importance in liquid rocket engine (LRE) design both for sizing and safety purposes. Computational fluid dynamics (CFD) simulations can aid in the prediction provided that they can be effectively used during the design phase and that suitable modeling is employed. In this framework, this study aims at evaluating the suitability of a RANS-based CFD approach to predict in affordable times the nozzle wall heat flux of LREs employing the oxygen/methane propellant combination, which is nowadays attracting the attention of many developers. The interest to study the throat heat flux estimation for oxygen/methane engines comes from the known greater role played by the near-wall recombination reactions, as compared to the oxygen/hydrogen propellant pair. Nevertheless, only few indirect experimental measurements are available in the open literature for the validation of numerical tools. Recently published experimental data are used here as benchmark for the comparison of numerical simulations obtained with different assumptions. Results confirm that, for a well-designed engine, the details of injection and combustion processes have only a secondary effect on the prediction of throat heat flux.
The paper focuses on an experimental unit developed for modeling combustion characteristics in a model oxygen-methane combustion chamber of a liquid rocket engine. The key components of the unit, i.e., the mixing head of the combustion chamber and the regeneratively cooled nozzle, were manufactured using advanced methods of additive manufacturing. The paper emphasizes the specific character of the combustion chamber components made with the use of additive technology and introduces hot-fire test results of the model combustion chamber as part of the experimental unit. The study shows the durability of the mixing head and combustion chamber nozzle under hot-fire test conditions, as well as the reliable operation of the experimental unit as a whole, which confirms the selected design and technological solutions. Within the study, we analyzed the cooling system of the experimental unit for the test conditions, estimated the thermal state of the nozzle, with account for the features of the additively manufactured cooling path. To increase the cooling system’s reliability and expand the combustion chamber pressure application, it is recommended to apply a heat-shielding coating on the firewall of the nozzle. Using new experimental data, we analyzed the parameters of improving the efficiency of the model combustion chamber with the additively manufactured components and corresponding in scale and consumption characteristics to the combustion chamber of the liquid rocket engine
This paper presents a wall modeling study of turbulent reacting flows of CH4/O2 mixtures towards accurately predicting the wall heat flux in combustion chambers. The study focuses on the description of flow and chemistry within inner layers and compares the accuracy of wall functions and ordinary differential equation (ODE) based wall models assuming equilibrium and frozen chemistry. Two test cases of turbulent reacting channel flow and GCH4/GOX rocket combustion demonstrate that the ODE-based frozen wall model is more accurate than other models in predicting wall shear stress and heat flux. The mixture composition within the inner layer is reproduced accurately by assuming the chemical frozen state, but assuming chemical equilibrium leads to overestimated changes in the mixture composition. The investigation of the effects of wall temperature shows that the wall function models significantly overestimate the wall shear stress and heat flux with decreasing wall temperature while the accuracy of the ODE-based frozen wall model is not affected by the wall temperature. The ODE-based frozen wall model derived from the proper description of the chemical state and thermodynamic properties of the gas mixture allow to accurately and robustly predict the wall heat flux in reacting flows, while the chemical frozen assumption maintains model simplicity.
Spark ignition aero piston engines have good prospects due to light weight and high power to weight ratio. Both gasoline and kerosene can be utilized on these engines by using either traditional port fuel injection (PFI) or the novel air-assisted fuel injection (A2FI). In this article, the effects of different fuels and injection methods on the performance of a four-cylinder opposed aero piston engine were studied. The spray performance test rig and the engine performance test rig were established. Firstly, the influence of different injection methods on engine performance were compared, which indicated that A2FI is superior to PFI in engine power and starting performance. Furthermore, the fuel performance comparison by using A2FI was conducted, which demonstrates that kerosene is inferior to gasoline in terms of spray characteristics and power performance. Finally, detailed working characteristics of A2FI system using kerosene were studied, which indicated that the stable and reliable operation of the spark-ignition operation can be realized and the kerosene's spark-ignition combustion process can be optimized similar to that of gasoline. Results shows that the use of kerosene combined with A2FI is the best technical way to achieve ideal working process of the spark ignition aero piston engine.
In the current study results from an experimental investigation on an oxygen/methane multi-injector combustion chamber are presented. They provide detailed information about the thermal loads at the hot inner walls of the combustion chamber at representative rocket engine conditions and pressure ranges up to 40 bar. The present study aims to contribute to the understanding of the thermal transfer processes and of the interaction between the injectors and the injector-wall. Furthermore, the test results are used as a test case for the validation of the in-house engineering tool Thermtest. Due to the complex flow phenomena linked to the use of cryogenic propellants, like extreme variation of flow properties and steep temperature gradients, in combination with intensive chemical reactions, the problem has been partially simplified by injecting the propellants in gaseous form.
This work studies the performance and dry mass of the under development LOX/Ethanol L75 liquid rocket engine. To this end, an object-oriented program written in C++ was developed. The program is intended to be versatile and easily extensible in order to analyze different configurations of liquid rocket engines. The UML (Unified Modeling Language) tool is used to model the architecture of the codes. UML diagrams help to visualize the code structure and the communication between objects, enabling a high degree of abstraction. The cryogenics Vulcain and HM7B engines power cycles along with the staged-combustion SSME engine perform the verification of the codes. Finally, the influence of changes in design parameters on the performance and dry mass of the L75 rocket engine is analyzed.
Film cooling is an effective technique that protects chamber walls in rocket combustion against chemical attacks and heat fluxes. This study discusses cooling effect in a multi-element GO2/CH4 splash platelet injector. Influence parameters, such as slot height, slot number, percentage of coolant, and injection position on cooling effect, were investigated. GCH4 with 298.15 K was applied as film coolant. In the first step, slot heights of 0.2 and 0.4 mm were compared by applying a constant film mass flow rate. Temperature, CH4 mole fraction distribution, and flow field structure were obtained. The effects of slot number, percentage of coolant, and injection position on wall temperature distribution were then determined. Finally, the reasons for the low cooling efficiency were analyzed. Improvement in the method is proposed to achieve improved cooling effect for splash platelet injectors.
Steady state engine cycle analysis is commonly used in pre-design phases of liquid propellants rocket engine development. In the engine development process, a high level systems analysis, which examines the engine cycle allows a preliminary design of the engine components in terms of the operational envelope, within which the engine components are required to function. This paper compiles the general methodology and component models used in DLR's cycle analysis tools. The paper describes briefly the tool's heritage. Methods used for component modeling are described in detail. Some sample calculations of rocket engines are provided as validation examples.
A numerical study was conducted at the German Aerospace Center in Lampoldshausen, to investigate the impact of various chemical models on reactive nozzle flow. Therefore, a chemical reaction mechanism for oxygen/methane combustion was implemented into DLR's flow solver TAU. Ignition delay simulations were conducted to demonstrate the validity of the implementation. The implemented baseline chemistry model was applied for generic nozzle flow simulations and the results were compared to frozen nozzle flow and nozzle flow in chemical equilibrium, in order to investigate the impact of the finite-rate approach. The baseline reaction mechanism was reduced to a basic configuration and applied to the generic nozzle flow. A good agreement with the baseline model was observed. Both approaches were applied for dual-bell nozzle flow simulations. Validation data for the simulations were obtained during a hot flow test campaign. The experiments yielded a clear impact of the combustion chamber mixture ratio on the dual-bell transition nozzle pressure ratio. RANS simulations of the dual-bell nozzle flow were conducted and almost no deviation between baseline and reduced chemical approach was observed. A reduction of 93~\% of the computational cost was reached with the reduced model. The dual-bell transition behavior at different values of combustion chamber mixture ratio was investigated, applying RANS simulations with reduced chemistry model. The impact of the mixture ratio on the transition NPR was clearly reproduced by the numerical approach. A good agreement with the experimentally obtained transition NPR values was reached.
During the development phase, the launchers’ needs for aerodynamic characterization are fulfilled by a hybrid approach encompassing wind tunnel testing (WTT) and computational fluid dynamics (CFD) results [1, 2]. The joint use of WTT and CFD is a powerful tool, able to provide high-quality data as input for performance evaluations as well as launcher control and sizing [1].
The Institute of Flight Propulsion at the Technische Universität München is operating a test facility to analyze different aspects of hydrocarbon combustion in a single-element rocket combustor. The research is done in cooperation with EADS Space Transportation, Ottobrunn. The test bed was designed to enable investigations of basic aspects concerning mixing and combustion of kerosene with gaseous oxygen in the preliminary stage of the development of a future reusable booster or expendable low cost engine. Main topics of research are the design and performance analysis of different injector elements and the effects of coking and sooting on heat transfer and cooling. Tests with different injector designs at several pressure levels and mixture ratios have been conducted. The data were analyzed with respect to combustion efficiency and heat transfer rate to a water cooled chamber wall. A comparison of tests with and without fuel film cooling promises great potential but also puts a number of questions to be answered. Copyright © 2005 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.
A ground test article was prepared for RS-18 engine testing using liquid oxygen (LO2) and liquid methane (LCH4) propellants under simulated altitude conditions at NASA Johnson Space Center White Sands Test Facility (WSTF). This project is part of NASA Glenn Research Center's Propulsion and Cryogenics Advanced Development (PCAD) project. "Green" propellants, such as LO2/LCH4, offer savings in both performance and safety over equivalently sized hypergolic propellant systems in lunar vehicle ascent engine applications. LO2/LCH4 testing capability at altitude conditions did not previously exist at WSTF for this size engine, and modifications were made to the Auxiliary Propulsion Systems Test Bed (APSTB) article. Altitude simulation is achieved using the WSTF Large Altitude Simulation System, which provides altitude conditions equivalent to ∼90,000 ft (∼27 km). For specific impulse calculations, engine thrust and propellant mass flow rates are measured. Propellant flow rate is measured using a coriolis-style mass-flow meter, and accuracy is compared with a serial turbine-style flow meter. Thrust is measured using three load cells in parallel. Igniter system capability is being developed to demonstrate two methods, a gaseous oxygen/methane spark torch igniter and solid propellant pyrotechnic igniter. Design, procurement and assembly are complete for the test article and test readiness is expected for hot-fires to begin pending completion of manifold buildup and system checkout. Test objectives for the RS-18 project are 1) conduct a shakedown of the test stand for LO2/LCH4 lunar ascent engines, 2) obtain nozzle kinetics data to anchor two-dimensional kinetics codes, and 3) obtain vacuum ignition data for the torch and pyrotechnic igniters.
In recent years, methane has attracted attention as a propellant for liquid rocket engines because of its various advantages compared to typical propellants such as hydrogen. When methane is used as a coolant for a regenerative cooling system, its near-critical thermodynamic and transport properties experience large variations because its critical pressure is higher than that of typical propellants; this significantly influences the flowfield and heat transfer characteristics. Therefore, adequate understanding of the flowfield and heat transfer characteristics of methane in regenerative cooling channels is a prerequisite for future engine development. In this study, conjugated coolant and heat transfer simulations were performed to investigate the flowfield and heat transfer characteristics of transcritical methane flows in a sub-scale methane-cooled thrust chamber. The computed results were validated against experimental data measured in hot firing tests. They compared well with the measured pressures and temperatures in cooling channels, and wall temperatures were within the permitted levels. Detailed flow analysis revealed peculiar flow structures in the cooling channel: a strong secondary flow induced in the concave-heated part in the channel throat section and the coexistence of two different gas phases-ideal and real-in a single cross-section in the cylindrical region. A high wall temperature appeared in the cylindrical region of the thrust chamber under the considered conditions; this was due to the heat transfer deterioration induced by an M-shaped velocity profile and a turbulent heat flux reduction. © 2012 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.
This study sought to develop a catalytic ignition advanced torch system with a unique catalyst microtube design that could serve as a low energy alternative or redundant system for the ignition of methane and oxygen rockets. Development and testing of iterations of hardware was carried out to create a system that could operate at altitude and produce a torch. A unique design was created that initiated ignition via the catalyst and then propagated into external staged ignition. This system was able to meet the goals of operating across a range of atmospheric and altitude conditions with power inputs on the order of 20 to 30 watts with chamber pressures and mass flow rates typical of comparable ignition systems for a 100 lbf engine.
Kerosene and methane are two promising candidate propellants for a future reusable booster stage. This study assesses the merits of both propellants and compares their respective performance when used in a booster stage.
The identification of the propellant properties is the starting point. An analysis of a staged combustion cycle engine for both propellants follows. The final assessment is based on the results of a performance analysis of a launch vehicle making use of these motors in reusable fly-back boosters.
Different types of reusable first stages designed for a near-term application with heavy lift launchers are investigated. The attached reference expendable space transportation system is a future Ariane 5 with cryogenic core and upper stage but skipped solid rocket boosters. The design of the so-called liquid flyback boosters is restricted to the incorporation of powerful hydrogen or hydrocarbon rocket motors already under development or in operation. The analyzed layout variants of the reusable vehicle include single- as well as dual-booster configurations. Catamaran-type double fuselage stages are investigated to evaluate the potential in reducing the unsymmetrical thrust load of one-side-mounted booster. Along with their primary use to boost heavy lift geostationary transfer-orbit missions, a second task may be covered by the same vehicle to accelerate the upper stages of small and medium launchers. The additional design requirements in such a dual-use reusable launch vehicle are studied. The investigation includes trajectory simulations and optimizations for ascent, as well as an assessment of propellant requirements and vehicle loads during return flight to the launch site. Critical flight stability aspects are evaluated by static and dynamic simulations. A comparison of size and mass is included, as well as performance data of the different liquid flyback booster configurations. The relevant rocket engine figures of performance, mass, reusability, and throttling capability are presented.
Kerosene and methane are two promising candidate propellants for a future reusable booster stage. This study assesses the merits of both propellants and compares their respective performance when used in a booster stage. The identification of the propellant properties is the starting point. An analysis of a staged combustion cycle engine for both propellants follows. The final assessment is based on the results of a performance analysis of a launch vehicle making use of these motors in reusable fly-back boosters.
The German future launcher technology research program ASTRA investigates in its system study two types of partially reusable launch vehicles. This paper describes one of those concepts, a reusable first stage designed for a near term application with a heavy lift launcher. The attached reference expendable space transportation system is a future Ariane 5 with cryogenic core and upper stage, but skipped solid rocket boosters. The design of the reference liquid fly-back boosters (LFBB) is focused on LOX/LH2 propellant and a future derivative of the Vulcain rocket motor. After achieving a convergent design in the first iteration loop, a more detailed level of investigation has been started. This includes the ascent control requirements on the booster TVC system, a refinement of the aerodynamic shape, and the preliminary mechanical lay-out of body and wing structure. All major results are presented, and used in an update of the mass budget, as well as trajectory simulations and optimizations for ascent and reentry.
The present publication introduces the fundamental principles of liquid-propellant rocket engines that are required for actual design applications. After an introduction to the gas-flow properties, performance parameters, and propellant types for these rocket engines, engine requirements are set forth for such indicators as duration, weight, envelope, and thrust level. Sample calculations are given for A-1 through A-4 stage engines, and the design of thrust chambers and combustion devices is reviewed for factors such as injectors, ignition, cooling, and instability. Other elements that are discussed in detail are: gas-pressurized and turbopump-propellant feed systems; control and condition-monitoring conditions; interconnecting components and mounts; propellant tanks; and designs for specific space applications.
In this paper, worldwide investigations of the cooling characteristics of hydrocarbon fuels are reviewed and results of Chinese experimental investigations are presented. The Chinese tests were conducted at pressures of 0.5-30 MPa, flow velocities of 2-106 m/s, and heat fluxes up to 66 MW/m 2. The heat transfer characteristics of methane, propane, no. 21 high-density kerosene, aerokerosene and rocket kerosene in stainless-steel and copper tubes, and the deposit formation rates for kerosene in stainless-steel tubes were investigated. Forced convective heat transfer correlations were obtained for liquid methane, propane, and kerosene. Heat fluxes at which test tubes burnt out were also investigated. The test results indicate that heat transfer coefficients of methane and propane decrease at high wall temperatures, and that the coefficients of kerosene increase at similar conditions because of boiling. No coking was detected in methane tests; coking temperature and coking rates were determined for kerosene in stainless-steel tubes.
The advantages of methane as a rocket fuel in comparison with kerosene is discussed. The results from engine cycle scheme research in comparison with other schemes are presented considering the same engine dimensions and the concepts of various engine modules for different class launch vehicles are proposed.
The main characteristics and design of liquid rocket engines for Proton,
Zenit, and Energia launch vehicles are described. Particular attention
is given to RD-253 engine using N2O4 and UDMH propellants, and RD-170
and RD-120 engines based on Lox and Kerosene propellants. The RD-170 and
RD-120 engines are considered to be the culmination of high pressure
staged combustion and oxidizer rich turbine drive rocket engine
experience. The RD-170 was verified with significant overstress testing
to provide necessary data for the operational health monitoring life
prediction system.
Auch in Zeiten großer Fortschritte bei der Entwicklung neuer und schneller Programme zur numerischen Strömungsanalyse besteht weiterhin Bedarf an Methoden zur semi-empirischen aerodynamischen Vorauslegung, basierend auf sogenannaten Handbuchmethoden. Auf diese Weise lässt sich am schnellsten und bequemsten eine parametrische Analyse durchführungen, die Hinweise auf optimale Auslegungsbereiche liefert.
Der vorliegende Bericht enthält die Beschreibung eines Programms zur Konturauslegung von Brennkammer und Düse eines Raketenmotors, welches im Rahmen einer Studienarbeit bei der Arbeitsgruppe ART in Köln-Porz entstand. Gegenstand des Programms ist der Entwurf von wahlweise einer Parabel-, einer konischen - oder einer idealen Düse, sowohl als zweidimensionale Kontur, als auch als CAD Modell. Für letzteres können zusätzlich Kühlkanäle in verschiedenen möglichen Anordnungen automatisch generiert werden.
Auch in Zeiten großer Fortschritte bei der Entwicklung neuer und schneller Programme zur numerischen Strömungsanalyse besteht weiterhin ein Bedarf an Methoden zur semi-empirischen aerodynamischen Vorauslegung, basierend auf sogenannten Handbuchmethoden. Auf diese Weise lässt sich am schnellsten und bequemsten eine parametrische Analyse durchführen, die Hinweise auf optimale Auslegungsbereiche liefert. In diesem Bericht wird eine kurze Übersicht gegeben über den Stand des CAC-Programms (Version 1.0) und die darin enthaltenen Rechenmöglichkeiten sowie Hinweise zur Bedienung und eine Erklärung der Input-Parameter. Der zweite Teil enthält eine Dokumentation mit Beispielrechnungen verschiedener Fahrzeuge oder Projekte mit CAC. Die Ergebnisse dieser Analyse werden mit Literaturangaben verglichen.
The Vulcain engine flow separation and side-load behavior observed and measured during thrust chamber tests is discudded in detail in this paper. It is shown by the test results and by comparison with numerical flow data that the parabolic Vulcain nozzle features a transition in separation behavior from free shock separation to restricted shock separation and vice versa during both engine start-up and shut-down. These highly transient phenomena are a major cause of side-loads. In addition, the side-load activities are measured during nozzle operation with pre free shock separation or pure restricted shock separation. By using results from numerical simulations, it is shown that a specific plume pattern, the cap-shock pattern, is responsible for the observed flow transition. Finally, a comparison of the flow behavior in the Vulcain nozzle during start-up and shut-down is compared with other published data for thrust-optimized or parabolic rocket nozzles with an internal shock emanating from the throat.
The objectives of this article are to 1) define the corrosive interaction between hydrocarbon fuels and copper combustion chamber liner materials, using test methods that do not depend on direct ohmic heating; 2) identify and demonstrate protective measures against the corrosive process; 3) identify and demonstrate refurbishment methods for copper combustion chamber liner materials inadvertently corroded; and 4) establish acceptable limits for the corrosion causative agents. Our results show that the only important corrosion process is sulfur corrosion brought about by trace sulfur-containing impurities present in the fuels. As little as 1-ppm sulfur can seriously degrade cooling channel performance. This sulfur corrosion process is effectively prevented by electrodeposited gold coatings on the cooling channel walls. Also, sulfur corroded cooling channels are successfully refurbished by treatment with dilute aqueous sodium cyanide which quantitatively dissolves the cuprous sulfide Cu2S, corrosion product. However, the overall sulfur corrosion/refurbishment process leaves highly roughened walls which show significant changes in cooling channel performance. A recommended sulfur specification for hydrocarbon fuels is discussed.
Beschreibung des Dü NCC
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Gas Guzzler 18 Military Specification Propellant Kerosene
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NIST Chem-istry WebBook, NIST Standard Reference Database Number 69
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Space Launch Initiative Triggers Hydrocarbon Engine War
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Liquid Boosters for Ariane 5: A Promising Alternative Concept
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