Rocket Engine Performance Research Articles

Studies of flow field characteristics during the impact of a gaseous jet on liquid–water column**Project supported by the National Natural Science Foundation of China (Grant No. 51305204).

Both experimental and numerical studies were presented on the flow field characteristics in the process of gaseous jet impinging on liquid–water column. The effects of the impinging process on the working performance of rocket engine were also analyzed. The experimental results showed that the liquid–water had better flame and smoke dissipation effect in the process of gaseous jet impinging on liquid–water column. However, the interaction between the gaseous jet and the liquid–water column resulted in two pressure oscillations with large amplitude appearing in the combustion chamber of the rocket engine with instantaneous pressure increased by 17.73% and 17.93%, respectively. To analyze the phenomena, a new computational method was proposed by coupling the governing equations of the MIXTURE model with the phase change equations of water and the combustion equation of propellant. Numerical simulations were carried out on the generation of gas, the accelerate gas flow, and the mutual interaction between gaseous jet and liquid–water column. Numerical simulations showed that a cavity would be formed in the liquid–water column when gaseous jet impinged on the liquid–water column. The development speed of the cavity increased obviously after each pressure oscillation. In the initial stage of impingement, the gaseous jet was blocked due to the inertia effect of high-density water, and a large amount of gas gathered in the area between the nozzle throat and the gas–liquid interface. The shock wave was formed in the nozzle expansion section. Under the dual action of the reverse pressure wave and the continuously ejected high-temperature gas upstream, the shock wave moved repeatedly in the nozzle expansion section, which led to the flow of gas in the combustion chamber being blocked, released, re-blocked, and re-released. This was also the main reason for the pressure oscillations in the combustion chamber.

Particle swarm algorithm-based damage-mitigating control law analysis and synthesis for liquid-propellant rocket engine

The damage-mitigating control is a novel technique to ameliorate the reliability and safety of liquid-propellant rocket engines by achieving an optimized trade-off level between overall dynamic performance of the liquid-propellant rocket engine and structural durability of some selected critical damageable components under the condition of no impact on the achievement of the launch and flight mission. Thus, it is needed to be solved for the damage-mitigating control that the global optimization of the best trade-off between the damage of the critical damageable components and the performance of rocket engine. The major challenge should focus on: (i) to construct model of a certain rocket engine system dynamics, critical components structural dynamics, and damage dynamics; (ii) to optimize open loop feed-forward control law based on liquid-propellant rocket engine system dynamic model, structural and damage dynamics model, by using particle swarm optimization algorithm; (iii) to synthesize an intelligent damage-mitigating control system using the optimized open loop control law. In this paper, synthesis procedure of damage mitigation is introduced; structure and damage dynamic model of damageable components are formulated. The results of the simulation computation show that the synthesized control laws are implemented and achieve the effect of damage mitigating for the liquid-propellant rocket engine. It can provide important theoretical and practical value not only for improving the safety and reliability of the liquid-propellant rocket engine, but also for the complex thermo-flow-mechanical systems such as airplane engines, automobile engines, and fossil-fueled power plant because their service life is very critical too.

Experimental research on propulsive performance of the pulse detonation rocket engine with a fluidic nozzle

To explore an effective way for the nozzle design of a pulse detonation rocket engine (PDRE), theoretical and experimental researches have been carried out. The exhaust process and optimal converging/diverging area ratios of a nozzle were theoretically analyzed. Pressure and velocity fluctuations of the exhaust flow were considered to be the major issues. A fluidic solution which employs nitrogen flows in the nozzle throat and divergent section has been tested. The fluidic nozzle is able to adjust the effective area ratios according to the incoming flow. The effective convergent area ratio varied from 2.0 to 2.2, and the divergent one varied from 5.0 to 1.8. It is proved that the fluidic nozzle effectively improved the propulsive performance of the engine. A maximum average thrust increase of 137.8% was obtained in proper conditions.

Modeling and Analysis of a LOX/Ethanol Liquid Rocket Engine

ABSTRACT 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.

Improved efficiency in rocket engine performance via a laser kinetic energy transfer chamber allowing single vehicle orbital flights

Improved efficiency in rocket engine performance via a laser kinetic energy transfer chamber allowing single vehicle orbital flights

Canted Slit 형상의 핀틀 인젝터 로켓엔진의 특성길이와 운동량비에 따른 연소성능

본 연구에서는 케로신과 액체산소를 추진제로 사용하는 핀틀 인젝터의 설계/제작이 이루어졌으며, 연소실 특성길이와 운동량비에 따른 특성속도효율을 통해 연소성능을 확인하였다. 연소시험 결과, 핀틀 인젝터의 추진제 분무 특성으로 인한 재순환영역에 의해, 추진제 조합에 따른 특성길이 추천치보다 짧은 영역에서도 최적의 연소성능을 발휘하는 것을 확인하였다. 또한 운동량비가 연소성능에 지대한 영향을 미치나, 추천범위 1.0~1.5 내에서 특성속도효율이 일정하게 나타남을 확인하였다. In this study, a pintle injector rocket engine which uses kerosene and liquid oxygen as propellants was manufactured by collecting basic design data and establishing a design procedure. Combustion performance of the liquid rocket engine was investigated by characteristic velocity efficiency with characteristic length of the combustion chamber and total momentum ratio. As a result of hot fire tests, it showed that the engine had shorter characteristic length comparing to those of other type injectors, which was known as recommended value with the propellant combination. Also, the characteristic velocity efficiency was greatly affected by total momentum ratio and almost constant within 1.0~1.5.

Operation of a liquid-fueled and valveless pulse detonation rocket engine at high frequency

To increase the operating frequency and enhance the propulsive performance of a pulse detonation rocket engine (PDRE), a valveless mode with the purge gas based on liquid fuel was developed. Liquid gasoline, oxygen-enriched air, and nitrogen were utilized as fuel, oxidizer, and purge gas respectively. Supply of these propellants was controlled by the periodic pressure oscillations in the detonation tube. The purge gas with the highest supply pressure was filled into the detonation tube first. The fuel and oxidizer were injected after the inner pressure decaying to a relatively low level. This delay and the vaporization of the liquid fuel prevented pre-ignition of the fresh detonable mixture. The PDRE can work stably from 10Hz to 160Hz in this scheme. Fully-developed detonation waves were proved to be obtained successfully at a maximum frequency of 130Hz. Both oxygen volume fraction and supply pressure of the purge gas were tested to determine the limits for stable operation. The propulsive performance of a practical PDRE at a wide operating frequency range was also tested and the maximum thrust was obtained at 140Hz. Compared to the results without purge gas, introduction of a purge gas extends the ranges of supply pressure and operating frequency.

Design Methodology for the Performance, Weight, and Economic Assessment of Chemical Rocket Engines

AbstractA design methodology is presented that supports the design of future aerospace rocket-powered vehicles. In particular, it provides the capabilities to rapidly evaluate the performance, weig...

Impact of nozzles on a valveless pulse detonation rocket engine without the purge process

To investigate the impact of exit nozzles on the operation and the propulsive performance of a valveless pulse detonation rocket engine (PDRE) without the purge process, experiments have been carried out under wide operating frequencies. Gasoline and oxygen-enriched air were used as fuel and oxidizer. Two detonation tubes and twenty-one nozzles with different shapes were employed in this study. It was observed that both the contraction ratio and the expansion ratio have an influence on the performance of the PDRE. Nozzles with converging sections decrease the highest operating frequency of the valveless PDRE, even if it cannot operate normally with nozzles whose contraction ratios are larger than four. The expansion ratio almost has no impact on the operation of the valveless PDRE, but the thrust decreases when the expansion ratios of diverging nozzles are larger than five. Most converging nozzles increase the thrust of the valveless PDRE, especially when the fill fractions were greater than one, while most diverging nozzles increase the thrust when the fill fractions were smaller than one. A maximum thrust increase of 25% has been obtained when nozzles were utilized.

Effects of Wall-Injection Length on Spray and Combustion in a Coaxial Porous Injector

The mixing performance of an injector determines the combustion performance of a bipropellant rocket engine. In the previous study, a novel coaxial porous injector concept was developed to improve the mixing performance of a shear coaxial injector. It demonstrated a higher characteristic velocity efficiency than that of a typical shear coaxial injector. To clarify the efficiency-improving mechanism of a coaxial porous injector, the combustion flowfields of both injectors were visualized using a high-speed shadowgraph technique. The effect of geometry factors on the spray core length and the combustion performance as well as the differences in the reactive spray behavior between the coaxial porous injector and the shear coaxial injector were observed. The coaxial porous injector with a different wall-injection length showed higher characteristic velocity efficiencies, shorter spray core lengths, and thicker phase-change patterns at the spray core boundary. Changes in wall-injection length had a nonmonotonic effect on both the spray core length and the characteristic velocity efficiency. As opposed to the expected results, the coaxial porous injector with the highest combustion efficiency showed the longest spray core length. The phase-change pattern width, caused by the evaporation of liquid fuel droplets, was proportional to the efficiency.

Thrust Performance of Rotary-Valved Four-Cylinder Pulse Detonation Rocket Engine

Thrust Performance of Rotary-Valved Four-Cylinder Pulse Detonation Rocket Engine

Experimental verification of the possibility of improving the liquid propelant rocket engine performance for the “Vega” launcher upper stage

Successful flight missions demonstrated the high level of reliability and performance of the 4th stage liquidengine developed by the Yuzhnoye design bureau for the Vega launcher. The obtained specific impulseis 3104.8 m/s which is higher than it was demanded at the initial stage of the development (3094 m/s). Meanwhile,the engine characteristics like mass or specific impulse can be further improved, which was confirmed byexperimental investigations. These experimental results and further possible thrust chamber improvements suchas implementation of a new radiation-cooling nozzle extension, cooling system redesign and injectors optimizationare discussed and analyzed in the article.

Propulsive performance of a pulse detonation rocket engine without the purge process

To measure the propulsive performance of a high-frequency PDRE (pulse detonation rocket engine), an experimental facility was established. Utilizing the valveless mode, the PDRE was operated without the purge process at a maximum operating frequency of 110 Hz successfully. In this study, oxygen-enriched air instead of oxygen was utilized as oxidizer and liquid gasoline was used as fuel because its vaporization would cool the hot combustion products, which would create a buffer zone and ensure stable operation without the purge process. The thrust under a wide range of operating frequencies was obtained including over and partially filled cases. Based on the results of partially filled conditions, one empirical formula for the partial fill effect was developed. In addition, analysis was carried out which resulted in a quite similar equation. Therefore, a partial fill model for the valveless PDRE without the purge process was obtained. Finally, comparisons between the proposed model and other models developed in traditional operations were performed.

Experimental Study of Ejectors’ Effect on the Performance of Pulse Detonation Rocket Engine

Experimental studies were performed to investigate the ejector’s effect on the performance of a pulse detonation rocket engine (PDRE). The PDRE employed in the experiments utilizes gasoline as fuel and air as oxidizer. The operating frequency of the PDRE is 10 and 20 Hz respectively. Performance is quantified by thrust measurements, and the average thrust of the PDRE is obtained by integration of the thrust measured in the experiment. The experimental results show that the PDRE can increase thrust augmentation with an ejector, a maximum thrust augmentation is measured by 18.7% at 10 Hz operating frequency. In the experiment, axial placement (X/DPDRE) of the ejector is varied downstream the PDRE tube of 0.5,1 and 2 respectively. It is found that the thrust augmentation obviously increases when the relative position at 1 and 2, and also the thrust augmentation increases at higher operating frequency. The experiment result is useful for the development of a PDRE producing higher thrust.

Performances Study of Liquid Rocket Engine

The article presents the performance of the liquid rocket engine. It will study the combustion reaction of the liquid hydrogen and oxygen without thermal dissociation. The main elements studied are the influence of the exces of the oxidant on firing temperature, on exhaust velocity, on the evacutation temperature from the engine and on the ishotermal coefficient. The pressure influence from the combustion chamber on the evacuation velocity, on the temperature from the evacuation from the exhaustion section and on the isentropic index.

Ecological aspect of launch vehicles development by criterion of minimal cost

One of the topical problems in modern aerospace engineering is accordance between ecological requirements and performance of the vehicle. On the other hand, problem of economical efficiency leads to change of the main criterion of designing to the minimization of costs (instead of maximal performance). According to modern trends of “low-cost” vehicles, different concepts of the future cost-effective launch vehicles are considered. It is necessary to validate these concepts according to requirements of ecological safety for the purpose of detection of the dominant launch vehicle configuration. Typical configurations of the future 'low-cost' launch vehicle are presented by 6 conceptual groups (Koelle, 2001). Conceptual group 1 (CG1) is presented by the Ballistic “Single stage to orbit” (SSTO) reusable vehicle. All vehicles which use classical rocketry scheme of the propulsion trajectory are called “Ballistic” i.e. the ballistic vehicle is lifted to orbit under the impact of rocket engines thrust. CG1-vehicle is able to reach the low earth orbit (LEO) without stage separation reducing the number of required rocket engines. Technological feasibility of SSTO concepts is proven by numerous studies (Koelle, 2001). CG2 representatives are ballistic “Two stages to orbit” (TSTO) reusable vehicles. The difference between CG1 and CG2 consists in application of vacuum rocket engines in the second stage and, consequently, stage separation. CG2 are the most mass-effective vehicles. CG3 is presented by the winged SSTO vehicles with rocket propulsion by “Lifting body” aerodynamic scheme. Ascensional force is provided by the aerodynamic shape of the vehicle’s structure at high speeds. Winged TSTO vehicles with rocket propulsion and parallel or tandem staging form the CG4. The winged configuration provides wide landing capability for both stages. CG5 is presented by winged TSTO vehicles with airbreathing propulsion in the first stage and rocket-propelled second stage. Airbreathing jet engines provide high reusability ratio comparing with other concepts as well as the widest landing capability. Aerospace Plane with scramjet-rocket propulsion forms CG6. The vehicle is able to reach near-cosmic speed in rarefied layers of the atmosphere and then accelerate with rocket engines. The most ecologically important resemblance of represented concepts is reusability. This reduces space debris formation (due to lack of waste hardware). Reusable launch vehicles can also be used to return the spent satellites. Structural differences between the concepts form 3 criterions of comparison by ecological impact: 1) propellant toxicity; 2) safety of surface facilities (vehicle damage inside the atmosphere); 3) probability of space debris formation (vehicle damage outside the atmosphere). Comparison of the concepts by these criterions allows substantiating the most ecologically acceptable direction of research. Results of the comparison demonstrate that the most ecologically acceptable low-cost launch vehicle configuration is: Ballistic SSTO or TSTO reusable launch vehicle with “LOX+LH2” propellant. The results can be explained by following way: combustion products of the propellant “liquid oxygen + liquid hydrogen” are absolutely safe for environment. It also provides maximal performance of rocket engine (due to the highest specific impulse). Ballistic ascent scheme allows using relatively simple technologies and provides high reliability level. In combination with minimal time of atmospheric flight this provides high level of safety for surface facilities. These results may be used for substantiation of dominant research direction.

Development and Assessment of a Backward Facing Steps Planar Rocket Nozzle

Large Rocket engine nozzles performance deteriorate because of asymmetric flow separation that lead to side loads. The dual-bell nozzle is a promising two operating modes nozzle which consists of a base nozzle with low area ratio and a nozzle extension with high area ratio. Although the dual-bell nozzle improves the performance of large rocket engines, it suffers from side loads that occur during the transition between the two operating modes. This paper presents and assess a new nozzle that consist of backward facing steps to minimize/eliminate side loads. The backward facing steps geometry guarantees fixed stable local separation at the steps edges at all altitudes, thus eliminates the possibility of side loads as the rocket ascends. The assessment of the nozzle performance is carried out using CFD simulation of turbulent Navier-Stokes equations solver (fluent 14.5). The computed results for both dual-bell and backward facing steps nozzles are compared with the dual-bell nozzle experimental data.

Research of High Precision DC Constant Current Source

In this paper, we designed a new type of high precision 1 μA constant current source after deeply studying the high precision precise dc constant current source. Compared with the traditional constant current source, in this paper we adopted complex amplifier added with floating power supply outside to design high precision dc constant current source which is for China's space use , its output current is 1 μA, resolution is 0.1 nA, accuracy is 0.5. This high precise dc constant current source is mainly used for new model rocket engine performance testing in aerospace, but also can be widely applied in different fields in need of precision measurement.

Experimental investigations on effects of wall-temperature on performance of a pulse detonation rocket engine

To study the effects of wall-temperature on performance of a pulse detonation rocket engine (PDRE), long-duration experiments were conducted. An internally grooved semi-circle spiral, instead of traditional Shchelkin spiral, was utilized to facilitate the deflagration to detonation transition (DDT) process and prolong run time in the present study. Four pressure transducers and a type K thermocouple were employed to collect data during the experiments. Experiments were carried out at three different operating frequencies, such as 10Hz, 20Hz and 30Hz, until unstable operation happened. The results showed that wall-temperature increased faster at higher operating frequency. It was observed that DDT time and DDT distance both decreased with the increase of wall-temperature, and detonation pressure increased sharply with wall-temperature first and then decreased slowly when wall-temperature exceeded a certain value. Both of them were believed to be due to the impact of hot-wall on fuel vaporization and reactant temperature. Unreasonably high wall-temperature was found to cause pre-ignition of fuel–oxidizer mixture, and this was why unstable operation of the PDRE happened after long-duration runs. In addition, effects of wall-temperature on specific impulse was also discussed.

EFFECT OF FUEL DROPLET SIZE AND INJECTION TEMPERATURE ON THE PERFORMANCE OF KEROSENE-OXYGEN PULSE DETONATION ROCKET ENGINE

In this study, the effect of fuel droplet size and injection temperature on the operating performance of a kerosene-oxygen pulse detonation rocket engine was experimentally investigated. A special particle size analyzing system was utilized to measure the kerosene droplet size. In addition, another experimental test system was also established to investigate the operating performance of a kerosene-oxygen pulse detonation rocket engine at the frequency of10 Hz. Three sizes of pressure-swirl nozzles were used to the experiment. The experimental results indicated that the SMD of kerosene droplets gradually decreased as the injection temperature increased, and three temperature ranges are separately discussed. The experiment results demonstrated that higher injection temperature and smaller SMD were beneficial to improve the operating performance. The thrust and specific impulse were very sensitive to the SMD value. As the injection temperature of the PDRE increased from 293 to 493K, the thrust increased by 23.42, 14.38, and 19.53%, and the specific impulse increased by 21.62, 19.61, and 15.89%, for nozzles 1, 2, and 3, respectively.