Specific Impulse Performance Research Articles

Numerical investigation of the distributed combustion in the thermal choked solid fuel combined cycle engine

In this study, a solid fuel combined cycle engine concept based on the distributed combustion and the self-adaptive thermal choking is proposed, which can operate effectively within a wide Mach number range. By regulating the heat release through the distributed fuel injection, the thermal choking can be formed in the fixed geometry combustor. The air intake controls the flow capture and circulation area through the movement of its head cone, ensuring stable air breathing over a wide range. Numerical simulations conducted under Ma = 3.0 and 6.0 conditions using the validated Eulerian-Lagrangian method coupled with gaseous and particle reaction models demonstrate the concept's capability to operate within the Mach number range of 3.0–6.0, delivering improved mixing, combustion, and specific impulse performance.

Insights into thermodynamic performance of a hypersonic precooled air-breathing engine with a complicated multi-branch closed cycle

An advanced precooled airbreathing engine with a closed Brayton cycle is a promising solution for high-speed propulsion, of which the Synergetic Air Breathing Rocket Engine (SABRE) is a representative configuration. The performance of the latest SABRE-4 cycle was analyzed in this paper. Firstly, a relatively complete engine performance model that considers the characteristics of turbomachinery and heat exchangers was developed. Then, Sobol’ global sensitivity analysis of key performance parameters was carried out to identify the most influential design variables. Optimal specific impulses under different target specific thrusts were obtained by particle swarm optimization, of which the thermodynamic parameters corresponding to a specific thrust of 1.12 kN·s·kg−1 and a specific impulse of 3163 s were chosen as the design values. Four different control laws were analyzed in contrast, and the charge control method had the strongest ability of thrust regulation as well as maintaining a favorable specific impulse performance. Finally, working characteristics under the charge control and over a typical flight envelope were calculated, in which the average value of the maximum specific impulse was as high as 5315 s. This study would help to deepen the understanding of SABRE-4 thermodynamic characteristics and other precooled airbreathing engine cycles with similar layouts.

Combustion of Date Stone and Jojoba Solid Waste in a Hybrid Rocket-like Combustion Chamber

The performance of two solid biomass wastes, date stone and jojoba solid waste, was experimentally examined for their potential application in combustion and propulsion systems. The fuels were tested in a hybrid rocket-like combustion environment, and the test result was analyzed with combustion and propulsion parameters. The performance of both fuels was comparatively evaluated and compared with a conventional hydrocarbon fuel in a hybrid rocket, with paraffin wax serving as a baseline. A compression device was introduced to compress the solid biomass wastes into a circular-shaped fuel grain compatible with a hybrid rocket combustion chamber with a hot surface ignitor. Thermogravimetric analysis (TGA) and chemical equilibrium analysis (CEA) results revealed that the performance of the biomass fuel can be comparable to conventionally used hydrocarbon paraffin-wax-based propellant within a certain range of oxidizer-to-fuel ratio, in terms of theoretical specific impulse performance. Through experimental performance tests, it was found that the compressed biomass fuel grains were successfully ignited and produced thrust. Both biomass fuels tested in a hybrid rocket combustion chamber are expected to pave the way for further developments in biomass fuels in the waste-to-energy field for their application in combustion and propulsion systems, potentially replacing fossil fuels with renewable resources.

Numerical study on throttling characteristics of RBCC engines in ejector mode

Investigating the throttling characteristics of Rocket-Based Combined Cycle (RBCC) engines in ejector mode is crucial for adapting to varied flight mission demands. This study employs validated RANS numerical methods to investigate throttling effects on an experimental kerosene-fueled RBCC engine under two representative flight conditions: ground take-off (Mach 0) and supersonic flight (Mach 2.0). Results indicate that, at Mach 0, throttling reduces the entrained secondary stream and increases mixing efficiency, but exacerbates mixing losses, which cannot be offset by the thrust produced by the secondary stream due to its weak working capacity. Therefore, specific impulse performance deteriorates during throttling. Moreover, the optimum mixture ratio remains at 2.4, consistent with conventional rockets. Conversely, at Mach 2.0, throttling mitigates the primary stream’s squeezing effect to augment the entrained secondary stream. Simultaneously, throttling enhances mixing efficiency to improve the utilization of residual fuel. Consequently, the optimum mixture ratio shifts from 2.4 to below 1.4, which contributes to harnessing the work capacity of the secondary stream generated by ram effects, and significantly improves specific impulse performance. This study emphasizes the distinct throttling characteristics at different Mach numbers and provides pivotal insights for developing more efficient RBCC engines in ejector mode.

Effects of Injector Configuration on the Detonation Characteristics and Propulsion Performance of Rotating Detonation Engine (RDE)

Fuel injection and mixing affect the characteristics of detonation initiation and propagation, as well as the propulsion performance of rotating detonation engine (RDE). A study on the injector is carried out in the present investigation. A rectangular-shaped hole-type fuel injector (RHFI) and slit-type fuel injector (SFI) were designed and compared experimentally at equivalent conditions. The investigation of the detonation propagation modes and the analysis of propulsion performance were carried out using fast Fourier transform (FFT), short-time Fourier transform (STFT), and unwrapped image post-processing. Under 50, 75, and 100 g/s flow rate conditions at an equivalence ratio of 1.0 ± 0.05, the RHFI has relatively stable detonation propagation characteristics, higher thrust, and specific impulse performance. Additionally, the results of the experiment indicate that the number of detonation waves affects performance.

Research on the comparative analysis of thrust and specific impulse performance - different cycle types of liquid rocket engines

As one of the most crucial parts of deep space exploration and interstellar colonisation in the global space industry, the development of rocket engines plays a vital role. The subject of this study is to investigate the effect of different cycling methods on the thrust and specific impulse of liquid rocket engines. The study will cover seven typical cycle methods and hopes to contribute to the subsequent development of more efficient and reliable liquid rocket engines, provide designers and engineers with a reference for optimising design and performance, and provide the basis for other future research in related fields that will have a profound impact on the space industry.

Parametric optimization and exergy analysis of a high mach number aeroengine with an ammonia mass injection pre-compressor cooling cycle

The Ammonia mass injection pre-compressor cooling (Ammonia MIPCC) cycle provides a substantial potential for the development of a high Mach aeroengine. To evaluate performance and analyze exergy of Ammonia MIPCC aeroengine, the parameter matching modeling method and Sequential Quadratic Programming (SQP) optimization are applied. These proposed methods can realize the constraints of compressor outlet temperature, combustor outlet temperature, power balance and flow conservation, thereby optimizing the specific impulse performance. Results show that the optimal operating line (OOL) of the Ammonia MIPCC aeroengine can be divided into two paths: the temperature limit operating line and the maximum specific impulse operating line. Moreover, the OOL of the Ammonia MIPCC aeroengine is always in a high exergy utilization state. When the ammonia rate increases from 0 to 0.08, the changes in exergy reduction rate, specific impulse and specific thrust corresponding to OOL at Ma 3.0 are 18.89%–31.83%, 2347.47 to 1113.56s, 768.91–953.47 N/(kg/s), respectively. Additionally, results show that the cycle with lower inflow cooling requirements can improve economy. In summary, parameter optimization and exergy analysis help to determine the working characteristics of high Mach number Ammonia MIPCC aeroengines, enhance the engine thrust performance and extend the operating envelope.

Effect of dielectric barrier discharge methane reforming products on the combustion performance of rocket engine

Although the low-thrust liquid oxygen/methane rocket engine has broad application prospects, the low flame propagation speed and low combustion rate of methane fuel make the liquid oxygen/methane engine still face key technical challenges. Methane fuel is partially converted into hydrogen and ethane with higher combustion rate before being injected into the combustion chamber, which is positive for the use of dielectric barrier to improve the combustion performance of the engine. Therefore, this paper studies the effect of the main four products of dielectric barrier discharge reforming with methane conversion rate of 10%, on the flow field of the combustion chamber. The results show that the addition of reforming products can effectively improve the combustion efficiency of the engine. H2 in the reforming product can also improve the specific impulse performance of the engine by increasing the total pressure of the engine chamber. C2H4 will not affect the maximum temperature of the engine, However, it can expand the medium-high temperature range of engine temperature to different degrees. The addition of H2 accelerates the oxygen consumption rate, which provides a feasible way to reduce the design size of the engine and improve the combustion efficiency of the low-thrust engine.

Investigation of the influence mechanism of geometric throats in variable-geometry rocket-based combined-cycle combustors

To explore the reaction mechanism a geometric throat and its influence on the performance of the combustor, this paper studies the working process of a rocket-based combined-cycle (RBCC) variable geometry combustor with geometric throats under Ma4 inflow conditions by both numerical simulation and ground direct-connect testing. The main research contents and conclusions are as follows: (1) The geometric throat can accurately control the choking position. The fuel in the variable geometry combustor can complete the concentrated heat release and bring higher combustor pressure. The thrust of the geometric throat combustor is 8.7% higher than that of the thermal throat combustor. Therefore, the geometric throat combustor provides a better combustor performance than the thermal throat combustor. (2) When the throat area of the Ma4 configuration is adopted, the thrust and specific impulse in the combustor are at their the highest, reaching 6.0% and 0.4% higher than those of the Ma5 and Ma3 throat configuration combustors, respectively. This shows that a certain amount of combustor heat release needs to match the corresponding geometric throat area to ensure the optimal performance of the combustor. (3) In the Ma4 combustor configuration, after closing the primary rocket, the thrust in the combustor decreases by approximately 10.8%, but its specific impulse performance greatly improved by 39.3%. Therefore, the geometric throat can be used in the RBCC mode transition process in combination with the secondary fuel injection strategy to avoid the thrust trap in the mode transition process.

Multi-Objective, Multi-Disciplinary Design Optimization and Multi-Attribute Evaluation of Hybrid Rocket Motors Used for Manned Lunar Lander

This paper proposed a multi-objective, multi-disciplinary design optimization and multi-attribute evaluation method for the manned lunar lander descent stage. A system design model is established considering multiple disciplines such as propulsion, structure, trajectory and cost. With the goal of minimizing mass, minimizing cost and maximizing velocity increment, the overall scheme for hybrid rocket motors was optimized by altering various grain shapes and feed systems, acting as an alternative propulsion scheme for the manned lunar lander. Under the same design conditions, the optimal scheme of hybrid rocket motors considering continuous and discrete attributes was studied and compared with that of liquid rocket engines to elucidate the characteristics of the hybrid rocket motors for deep space exploration. The results showed that the evaluation results obtained by considering only continuous attributes were different from those by considering both continuous and discrete attributes. The liquid propulsion scheme with liquid oxygen/kerosene is superior to the hybrid propulsion schemes due to its excellent cost and specific impulse performance when only continuous attributes are considered. However, hybrid rocket motors have shown good performance in terms of operability, manufacturability, safety and environmental protection. So, after introducing discrete attributes, the hybrid propulsion schemes show greater potential. In brief, based on the multi-attribute evaluation method considering comprehensive attributes, the hybrid rocket motor provided with tube grain and gas pressure feed system was considered as the optimal propulsion system for the lunar lander. Furthermore, the parametric analysis showed that the fuel grain outside diameter, the initial design thrust and the initial oxygen-to-fuel ratio had a significant influence on the performance of the hybrid rocket motor for the lunar lander, and in particular, the effect of the fuel grain outside diameter was more than 50%.

Comparison of Submillimeter Spot Ablation of Copper and Nickel by Multipulse Picosecond and Femtosecond Laser

The current transmission and reflection laser ablation micropropulsion modes have the problem of a complex working medium supply system in engineering. Therefore, we propose large-spot laser ablation with a one-dimensional supply mode. In order to verify this ablation mode, a multipulse ablation experiment of submillimeter-scale light spots was carried out on the surface of pretreated copper and nickel under the atmosphere using an ultrafast laser with a pulse width of 290 fs and 10 ps. The results show that femtosecond laser multipulse ablation (FLMA) leads to the grain refinement of copper, the crater quality of the two metals under FLMA is better, and picosecond laser multipulse ablation (PLMA) causes the crater of nickel to form a dense remelting bulge that affects laser absorption; both metals have obvious heat-affected zones after FLMA and PLMA, the heat-affected zones of nickel are 5–10% larger than those of copper, and the ablation depth of copper is deeper. Under the same conditions, the ablation mass of copper is smaller than that of nickel, and the specific impulse performance of laser ablation micropropulsion is better.

Performance analysis of precooled turbojet engine with a low-temperature endothermic fuel

For hydrocarbon-fueled TBCC engines, precooling is a crucial technology to solve “thrust gap” by expanding the flight Mach number of the turbojet engine, which can bring a sharp reduction of the specific impulse due to insufficient heat sink of hydrocarbon fuel. A low-temperature endothermic fuel precooling scheme is a great potential way to solve this problem. Heat sink experiments and thermodynamic cycle analysis are used to verify the effectiveness of the low-temperature endothermic fuel precooling scheme in improving specific impulse performance. The experiment results indicated that low-temperature endothermic fuel has a more considerable heat sink at lower temperatures than traditional endothermic hydrocarbon fuel. The results of thermodynamic cycle analysis show that the low-temperature endothermic fuel precooling scheme can effectively improve the specific impulse. Maximum specific impulse can reach above 800s with the low-temperature endothermic fuel precooling scheme, which is below 400s with hydrocarbon fuel precooling at Ma0 = 3.0. When cracking rate equals 1.0, and precooling efficiency equals 0.85, maximum specific impulse can reach 600s, increased by approximately 1.4 times at Ma0 = 4.0. The feasibility of precooled turbojet with the low-temperature endothermic fuel precooling scheme was preliminarily verified through this work.

Design and experimental research of a sub-Newton N2O monopropellant thruster with inner-heater

Nitrous oxide (N2O) is a green propellant with excellent application prospects. A sub-Newton N2O monopropellant thruster with inner-heater and a N2O self-pressurization stable supply system with regenerative heat compensation are designed in this paper. The experimental research of the thruster is described, including measurements of preheating power, activation temperature, vacuum thrust, specific impulse, life-span and pulsed operation performance. By inserting the heater into the catalyst-bed, preheating efficiency of the heater is significantly improved compared to the thruster with outer-heater. Thus, the preheating power demand of the thruster is successfully reduced to around 10 W. The mean vacuum thrust of 322 mN is attained and the corresponding specific impulse reaches 162 s at the mass flow rate of 0.2 g/s. Successful activation temperature of 523 K is achieved, and the activation performance of the thruster is affected by the loading factor. A long term hot-firing test longer than 12000 s is attained. The pulsed operation performance of the inner-preheating thruster is also studied by measuring chamber pressure. Impulses with different magnitudes are produced by adjusting the opening duration of the control valve. A minimum impulse of 81 mN∙s is attained. Finally, the performance of the thruster is evaluated by comparison with other thrusters of the same type. The results indicate that the proposed thruster with inner-heater is superior in terms of preheating power, activation temperature and specific impulse performance.

Numerical investigation of plasma behavior in a micro DC ion thruster using the particle-in-cell/Monte Carlo collision (PIC/MCC) method

Micro direct current (DC) ion thruster has advantages of small volume, light weight and significantly improved specific impulse performance compared to conventional chemical propulsion systems, thus having a wide application prospect in the tasks of cube-satellites and micro satellite networking systems. In this paper, a two-dimensional axisymmetric particle-in-cell/Monte Carlo collision model for the discharge chamber is developed for a micro DC ion thruster, which adopt the real configuration of original MiXI thruster with ring-cusp magnetic field. This model takes the thermal electron emission including the Schottky effect, various collision processes and the uneven background gas density distribution in the cathode-anode gap into account. The transient discharge process and the influence of operating conditions on the discharge process in the ionization chamber are presented for the first time so as to provide a detailed understanding of transport characteristics of plasma associated with the DC discharge process. Our calculated properties are compared with literature experimental data under similar conditions and qualitative and quantitative agreements are reached. The leak widths for charged species are obtained and closely related to charge separation. The simulations indicate that low neutral flow rate may lead to a discharge instability accompanied with oscillation of plasma quantities. These results lay a foundation for further optimization of thruster discharge chamber design in the future.

Theoretical and Experimental Heat of Combustion Analysis of Paraffin-Based Fuels as Preburn Characterization for Hybrid Rocket

The energy characteristics and theoretical performance of the hybrid rocket fuels are discussed in this paper. Aluminum (Al) and boron (B) metal additives were used to increase the energy density of the paraffin-based solid fuels. To predict the energy characteristics, the heat of combustion was evaluated by adiabatic bomb calorimetry. Theoretical performance parameters such as specific impulse (Isp), flame temperature, and characteristic velocity were obtained with NASA Chemical Equilibrium with Applications (CEA) code. Calorimetric test results revealed that paraffin/polyethylene/boron (P/PE/B)-based fuel formulations exhibited the highest heat of combustion among all the tested fuels. The heat of combustion value of the P/PE/B sample at 25 wt% B loading was found to be 9612 ±16 cal/g and 9293±17 cal/g for the P/PE/Al fuel formulation. The CEA results showed that the addition of Al to paraffin is noneffective in improving specific impulse performance. When B loading increased from 5 to 25 wt% in the P/PE/B, the Isp increased by 47 s compared to pure paraffin. A specific impulse increase implies the possible propellant mass saving. The reduction of the oxidizer and fuel masses may yield increased payload performance for given boundary conditions. The P/PE/B25 formulation has reported the highest value of characteristics velocity (C*) compared to other paraffin-based formulation.

Experimental research on the thermal throat of rocket based combined cycle combustor

Based on a previous one-dimensional theoretical analysis, direct-connection experiments of thermal throat generation and the factors that influence it in the Rocket Based Combined Cycle combustor are performed relative to the fuel injection position and equivalence ratio. The Mach 5.5 airflow experimental results indicate that the performance is improved by 3.0–8.3% by optimizing the fuel injection position, and at an airflow of Mach 4.0, the increase in the fuel equivalence ratio gradually shifts the generation of the thermal throat backward. When a thermal throat is generated at the end of the combustor, the thrust and specific impulse performance are optimal. The effectiveness of employing a thermal adjustment to improve engine performance is validated by the experimental studies. The radiation spectrum method is employed to measure the temperature along the flow field of the thermal throat. The temperature and pressure measurements along the flow path indicate that the Mach number variation trend calculated in the experiment is the same as that observed in the numerical simulation results in the thermal throat flow field. The generation process of the thermal throat is characterized by the acceleration of airflow from being subsonic to supersonic.

Quantitative analysis of the carbon generation characteristics during Mg/CO2 combustion: Implications for suppressing carbon deposition

Carbon deposition can significantly reduce the working stability and specific impulse performance of magnesium/carbon dioxide (Mg/CO2) powder rocket engines; however, few studies have quantitatively analyzed carbon generation characteristics during the Mg/CO2 combustion process. Therefore, this study presents a quantitative analysis of the carbon generation characteristics and degree of heterogeneous reaction during Mg/CO2 combustion in a closed high-temperature furnace. The analytical methods used include elemental analysis (EA) and inductively coupled plasma-optical emission spectrometry (ICP-OES). Experimental results showed that increasing the ambient temperature and decreasing the particle size both effectively suppressed carbon generation through a heterogeneous reaction, which is discussed from a thermodynamic perspective. When the particle size was 5 μm, the degree of heterogeneous reaction increased from 16.2% to 26.6% as the ambient temperature decreased from 1573 K to 1173 K; and when the ambient temperature was 1573 K, the degree of heterogeneous reaction increased from 16.2% to 23.2% as the particle size increased from 5 μm to 160 μm. Excessive CO concentration on the surface of Mg particles resulted in a large amount of carbon deposition, along with an insufficient Mg burnout rate. The proportion of Mg in the two pretreated powder samples participating in the heterogeneous reaction is reduced by 41% and 52%, respectively, compared with that in the source powder. Moreover, the combustion performance of pretreated Mg powder is typically excellent in CO2 due to the micro-explosion phenomenon. These findings are expected to contribute to improved working stability and specific impulse performance of Mg/CO2 powder rocket engines intended for Mars exploration.

Numerical analysis of direct-current microdischarge for space propulsion applications using the particle-in-cell/Monte Carlo collision (PIC/MCC) method

Microdischarges are an important type of plasma discharge that possess several unique characteristics, such as the presence of a stable glow discharge, high plasma density and intense excimer radiation, leading to several potential applications. The intense and controllable gas heating within the extremely small dimensions of microdischarges has been exploited in micro-thruster technologies by incorporating a micro-nozzle to generate the thrust. This kind of micro-thruster has a significantly improved specific impulse performance compared to conventional cold gas thrusters, and can meet the requirements arising from the emerging development and application of micro-spacecraft. In this paper, we performed a self-consistent 2D particle-in-cell simulation, with a Monte Carlo collision model, of a microdischarge operating in a prototype micro-plasma thruster with a hollow cylinder geometry and a divergent micro-nozzle. The model takes into account the thermionic electron emission including the Schottky effect, the secondary electron emission due to cathode bombardment by the plasma ions, several different collision processes, and a non-uniform argon background gas density in the cathode–anode gap. Results in the high-pressure (several hundreds of Torr), high-current (mA) operating regime showing the behavior of the plasma density, potential distribution, and energy flux towards the hollow cathode and anode are presented and discussed. In addition, the results of simulations showing the effect of different argon gas pressures, cathode material work function and discharge voltage on the operation of the microdischarge thruster are presented. Our calculated properties are compared with experimental data under similar conditions and qualitative and quantitative agreements are reached.

다단 하이브리드 로켓에서 AP 첨가 추진제의 연료과농 연소

본 연구는 다단 하이브리드 로켓의 낮은 비추력 성능을 향상시키는 방법으로 AP 첨가 추진제를 제안하고 있다. 추진제에 첨가된 AP 첨가비율 변화에 따라 다단 하이브리드 로켓의 비추력 성능 변화와 연료과농 연소특성 변화를 살펴보았으며, 이때 AP 첨가비율은 하이브리드 로켓의 연소 특징을 유지하기 위해 최대 15 wt%로 제한하였다. 결과에 의하면, AP 15 wt% 추진제는 AP 0 wt% 추진제와 비교하여 비추력 성능이 약 3% 향상되었다. 또한, 동일한 연소온도를 유지함에도 불구하고, AP 첨가비율을 증가시키면 산화제 유입량, O/F비 변화량, 그리고 연소압력은 감소하며 반경반향 온도 분포가 좋아지는 등 다단 하이브리드 로켓의 성능향상에 긍정적인 효과가 나타났다. 그러나 오직 AP를 추진제에 첨가하는 것만으로 다단 하이브리드 로켓의 비추력 성능을 일반 화학로켓의 수준으로 향상시키는 것이 매우 어려운 일임을 고려할 때, 추가적으로 금속입자 첨가를 통해 비추력 성능을 향상시킬 계획이다. In this study, AP added propellant has been proposed as a method of enhancing the low specific impulse performance found for staged hybrid rocket engine. Experimental tests were carried out to analyze and evaluate the effect of AP added propellant on specific impulse performance as well as fuel-rich combustion characteristics in a staged hybrid rocket engine. Upper limit of AP content in propellant was set to be 15 wt% to maintain the hybrid rocket engine advantages. As a result, 15 wt% AP added propellant showed 3% higher specific impulse performance compared to 0 wt% AP added propellant. Moreover, AP addition proved to offer less injected oxidizer mass flow, less O/F variation, and less combustion pressure while producing fuel-rich gas of the same combustion temperature. Future studies will carry out more combustion tests with metal additives to further enhance specific impulse.

Regression rate and combustion performance investigation of aluminum metallized HTPB/98HP hybrid rocket motor with numerical simulation

This paper attempts to research and predict the effect of the convention aluminum particle additives on the performance of the HTPB/98HP hybrid rocket motor with the numerical approach. The aluminum particle is injected to the flow field in form of droplet from the grain surface, and then simulated by the discrete phase model. A CFD code has been developed to solve the coupling mass, momentum and energy equations of gas and solid at the grain surface, as well as the initial parameters of the aluminum droplet. A good agreement exists between the calculated regression rate from numerical model and the experimental data, indicating that the accuracy of applying numerical approach to predict the characteristics of aluminum metallized HRM is acceptable. Series of static numerical simulation have been conducted to determine the effect of aluminum particle property on the combustion and specific impulse performance of the HRM. The regression rate increases with the weight fraction of convention aluminum particle, but shows no distinguished correlation with the particle diameter. The characteristic velocity and vacuum specific impulse display a strong dependence on the particle diameter. The results illustrate that the proposed numerical approach is potentially an effective tool for predicting and investigating the enhancement of the regression rate and performance of the HRM with the addition of aluminum particle.