Liquid Propulsion Systems Research Articles

Synthesis, Characterization and Energetic Properties of Hydroxymethyl-Bishomocubanone Derivatives.

The present work reports synthesis, characterization and theoretical insights on novel hydroxymethyl-bishomocubanone derivatives. Twelve new bishomocubanes (BHCs) were synthesized and fully characterized by various spectroscopic techniques and single crystal X-ray analysis. The densities of the title compounds were in the range of 1.30-1.59 g/cm3. Density-functional theory (DFT) based calculations at B3LYP/6-311++G(d,p) level of theory were performed on ten selected BHC based cage compounds. Propulsive and ballistic properties of newly synthesized hydroxymethyl-bishomocubanone derivatives in solid and liquid propulsion systems were calculated, and the results suggested that these compounds are superior to conventional fuel RP1 and binder HTPB. The detonation parameters revealed that these compounds are not explosive in nature and safe to use as solid propellants. Furthermore, kinetic and thermal stabilities of the title compounds were determined by HOMO-LUMO energy gap, ESP maps, impact sensitivity (h50) and bond dissociation energies (BDEs) followed by thermogravimetric analysis (TGA) and differential thermal analysis (DTA). Three compounds, a dinitroazide (Isp,vac=310.98 s), a dinitrate (Isp,vac=309.51 s), and a dinitronitrate (Isp,vac=309.20s) were found to be excellent candidates for volume limited applications.

Effect of tube bend angle on liquid propulsion system priming event pressures

Experimental work was performed to measure the priming event pressure transient levels in unrestricted liquid propulsion system tubing. An existing water hammer experimental setup was utilized, where system tubing diameter, length, and bend angle were examined. Experiments were performed under both atmospheric and low-pressure pre-test element conditions to understand any impact of tube bend angle on the mitigation of priming event pressure transients inside the liquid system. Experimental results showed pressure levels were influenced by a combination of tube diameter, length, and pre-test pressures; however, tube bend angle had minimal to no discernible change on the average peak pressure levels or rise times to peak pressure for all conditions examined.

Design optimization of a low response time thruster control valve for small satellite missions

Small satellites and spacecrafts use propulsion systems to stabilize their attitude, control orientation and deorbit it at the end of the mission life. The flow and thrust of the liquid propulsion system are controlled by the control valve, typically operated by solenoid. According to recent literature, space-qualified solenoid valves are expensive and financially non viable, especially for small satellite missions. In this paper, a detailed methodology is presented to design the solenoid valve dedicated for small satellites. This study includes design criteria focusing on geometrical configuration. The dynamic, and electro-magnetic models are used to characterize the solenoid valve. Dynamic modelling helps to determine the response time of valve, affecting the thruster minimum impulse bit. The physical dimensions of the designed solenoid valve are smaller than those available in commercial off-the shelf (COTS) solenoid valves. However, the efficacy of the proposed design is proven effective with the COTS space grade valve specifications, and it shows that the proposed valve consumes less power than the conventional valves. Finally, it is worth noting that despite their modest and compact physical dimensions, the proposed valve are ideal for small satellite missions.

Experimental study on generation and growth of bubbles in external flash-boiling spray regime

High-performance liquid propulsion systems, such as hypersonic scramjets and liquid rockets, utilize regenerative cooling by employing fuel as a coolant to shield components from thermal damage. The flash-boiling spray, resulting from superheated liquid injection and prevalent in regenerative cooling, has received considerable research interest for enhancing atomization and combustion efficiency. In this study, superheated water jets at various temperatures between 30 and 90 ℃ were injected into a vacuum chamber at different pressures above 5 kPa, and their spray patterns were acquired through diffusive backlight imaging. Images were analyzed to determine key parameters, such as bubble nucleation flux, bubble growth rate, and spray angle, in an external flash-boiling spray regime. The estimated nucleation flux agreed well with the predictions the model based on nucleation theory in the region of sufficiently high superheat. The dynamics of bubble growth was investigated and explained using a bubble growth model. The jet-expansion energy was computed by measuring the spray angle and compared with the available energy. By mapping the evolution of flash-boiling spray in relation to relevant nondimensional numbers, the study underscores the necessity of accounting for the pressure inside the injector to adequately model bubble generation inside the injector. Furthermore, this study substantiates the feasibility of visually recognizing and analyzing individual bubbles in the external flash-boiling regime, thereby enabling the generation of experimental data crucial for the validation of bubble nucleation and growth models.

SHORT OVERVIEW AND HIGHLIGHTS OF THE HISTORY OF THE GERMAN GEL PROPULSION RESEARCH AND TECHNOLOGY DEVELOPMENT ACTIVITIES

Gel mono- and bipropellants (i.e., gelled fuels and/or gelled oxidizers) are interesting candidates for advanced low-hazard rocket or ramjet propulsion systems and gas dynamic systems. Gel propellants combine major advantages of solid propulsion systems and liquid propulsion systems, such as easy and simple storage and handling characteristics (solid systems) and excellent throttleability (liquid systems). This article gives a short overview of the themes and the obtained results about the work conducted in Germany within the last two and a half decades, whereas a focus is set on the progress achieved in the last years.

Liquid Propulsion Systems in ISRO - Evolution and Perspective

The Liquid Propulsion Systems Centre (LPSC) which was formally constituted as a lead Centre for advanced propulsion systems development in ISRO on 1st June 1987 celebrates Silver Jubilee during 2012. LPSC has the mandate of supporting both the Launch Vehicle and Spacecraft programmes of ISRO. Starting with small Liquid Propellant Control Thrusters and Reaction Control/SITVC system development for SLV-3, today the activities of LPSC have grown many fold and encompasses integration and delivery of Liquid Propellant Booster Stages, Cryogenic Upper Stages and total Spacecraft Propulsion systems. Indigenization and production of Vikas Engine, realisation of Cryogenic Upper Stage for GSLV and the development of 200kN thrust Cryogenic Engine for GSLV-MkIII are some of the challenging assignments handled by LPSC which is also gearing up for the 2000kN thrust Lox-Kerosene Semi-Cryo Engine development. It has been a remarkable span of twenty five years of intense learning and accomplishment.
 In this article, based on the AR & DB Satish Dhawan Memorial lecture delivered in June 2012 for Aeronautical Society of India, the evolution of Liquid Propulsion Systems development in ISRO and their increasing role in the current Indian Launch Vehicle programmes as well as the Spacecraft missions viz. IRS, GSAT, Chandrayan etc is described. The new initiatives in advanced propulsion systems such as Electric Propulsion for satellites and 2000kN thrust Semi-cryo engine development for future ISRO launch vehicles are also outlined.

The approach to numerical simulation of the spatial movement of fluid with forming free gas inclusions in propellant tank at space flight conditions

The space propulsion systems ensure se veral start-ups and shutdowns of main liquid-propellant rocket engines under microgravity conditions for the spacecraft program movements and reorientation control. During the passive flight of the space stage (after its main engine shutdown), the liquid propellant in the tanks continues to move by inertia in microgravity away from the propellant management device as much as possible. In this case, the pressurization gas is displaced to the propellant management device, which creates the potential danger of gas entering the engine inlet in quantities unacceptable for the reliable engine restart. In this regard, determining the parameters of fluid movement in propellant tanks in microgravity conditions is an urgent problem that needs to be solved in the design period of liquid propulsion systems. We have developed an approach to the theoretical computation of the parameters of the motion of the ‘gas — fluid’ system in the propellant tanks of modern space stages in microgravity conditions. The approach is based on the use of the finite element method, the Volume of Fluid method and modern computer tools for finite-element analysis (Computer Aided Engineering — CAE systems). For the passive leg of the launch vehicle space flight, we performed mathematical modeling of the spatial movement of liquid propellant and forming free gas inclusions and determined the parameters of movement and shape of the free surface of the liquid in the tank as well as the location of gas inclusions. The numerical simulation of the fluid movement in an experimental sample of a spherical shape tank was performed with regard to the movement conditions in the SE Yuzhnoye Design Bureau ‘Drop tower’ for studying space object s in microgravity. The motion parameters of the ‘gas — fluid’ interface obtained as a result of mathematical modeling are in satisfactory agreement with the experimental data obtained. The use of the developed approach will significantly reduce the amount of experimental testing of the designed space stages.

2.5 kN pump-pressure pintle engine ignition experiment in different loading cases

Advanced technology of motors and batteries has stimulated the application of the emerging electric pump in the liquid propulsion systems. However, the coupling characteristics between the electric pump and the pintle injector with high momentum ratio are not clear. In this study, suitable approaches were applied to conduct the ignition experiment that the hydrogen peroxide was catalyzed in full flow with kerosene in different cases. In the experiment, the electric pumps were used to pressurize the hydrogen peroxide and kerosene respectively. The experiment results showed that, the catalytic process would be realized between 1.0 s and 2.0 s in different cases while the ignition process could be reached between 0.1 s and 0.2 s, which was mainly caused by the overshoot between 15% and 30% in the catalytic process. When the injection valve remained closed after the expected speed was reached, the outlet pressure of electric pump would be subject to vibration. However, the low-frequency vibration predominated from 12 Hz to 31 Hz. After the injection valve opened, the pressure vibration of the kerosene electric pump was between 0.15 MPa and 0.8 MPa. As the load dropped, the pressure vibration would obviously escalate. Each time when the speed increased 3000 r/min, the propellant temperature at the outlet of the hydrogen peroxide pump would increase about 1°C and that of the kerosene pump 5°C. The experiment can promote the application of the electric pump with hydrogen peroxide and kerosene in rocket propulsion systems with pintle injectors to realize a large regulating ratio.

The effect of external heat inflow to the cryogenic liquid pressurized discharge process

The launch vehicle with a liquid propulsion system requires pressurization during the propellant discharge. While the launch vehicle is flying through the air layer, a process in which frictional heat is generated between the liquid propellant tank and the air layer can be assumed. External heat inflow occurs during the liquid discharging process, and it is necessary to investigate how the external heat influx affects the liquid pressurized discharge process. A test device was constructed to determine the effect of pressurization of external heat inflow. The test device consisted of a liquid tank located in a vacuum chamber and a heating device to simulate external heat inflow. The heating device is a halogen lamp. The cryogenic liquid discharge with and without external heat inflow was compared. It was confirmed that the amount of pressurant gas used was decreased when there was external heat inflow. External heat inflow contributes to the evaporation of the liquid, and the evaporated gas increases the pressure of the ullage.

ТЕРМО-ПРОЧНОСТНОЙ АНАЛИЗ НЕОХЛАЖДАЕМОГО КОМПОЗИТНОГО СОПЛА ЖИДКОСТНОГО РАКЕТНОГО ДВИГАТЕЛЯ

In order to improve the power characteristics of liquid propulsion systems used in rocket technology by accelerating charged particles, the authors performed a thermal and strength analysis of the liquid rocket engine design using an uncooled nozzle. A model of a nozzle made with the use of composite materials was considered and the conditions of real thermal stress and force loading on the combustion chamber shell with the use of a composite nozzle were simulated.

Hydraulic similitude assessment for cryogenic cavitation in propellant lines: The case of thick orifice

Cryogenic cavitation can be easily encountered in rocket liquid propulsion systems at the very initial phase of the launch due to the propellant passing through orifices and valves. This leads to flow instability, performance degradation, and even damages to the structures. Hence, the design of the propulsion systems cannot overlook cavitation phenomena. Despite the relevance of the topic, their accurate prediction is still hampered by the difficulty in modeling the high non-equilibrium phenomena involved. From an engineering point of view, the main interest is to design such restrictions present in propellant systems and, hence, to predict the maximum flow rate they can reach. This work questions the possibility to design orifices for cryogenic applications based on purely isothermal testing. Therefore, we performed cavitation experiments with both water and liquid nitrogen by considering two orifices hydraulically similar. This implies a geometric similitude in terms of the orifice characteristic ratio (β) and its dimensionless thickness (th). The results show that the hydraulic similitude is not sufficient to predict the orifice behavior for a cryogenic flow correctly. In particular, the level of liquid subcooling (ΔTsub) upstream of the orifice plays an important role since it impacts the minimal value of pressure of the flow downstream the orifice and thus its thermodynamic state.

Impact of Cryogenics on Cavitation through an Orifice: A Review

Cryogenic cavitation affects the operation of liquid propulsion systems during the first phase of a launch. Its effects within orifices or turbopumps can range from mild instabilities to catastrophic damages to the structures, jeopardizing the launch itself. Therefore, to ensure the proper designing of propulsion systems, cavitation phenomena cannot be neglected. Although hydrodynamic cavitation has been studied for decades, the impact of the nature of the fluid has been sparsely investigated. Therefore, this review, beginning from the basic concepts of cavitation, analyzes the literature dedicated to hydrodynamic cryogenic cavitation through an orifice. Our review provides a clear vision of the state-of-the-art from experimental and modeling viewpoints, identifies the knowledge gaps in the literature, and proposes a way to further investigate cryogenic cavitation in aerospace science.

Catalyzed combustion of a nanofluid fuel droplet containing polydopamine-coated metastable intermixed composite n-Al/CuO

The ignition and combustion characteristics of kerosene fuels containing polydopamine (PDA)-coated metastable intermixed composite (MIC) nano-aluminum (n-Al)/nano-copper oxide (CuO) are experimentally investigated, with a focus on the mass content of CuO in n-Al@PDA@CuO MICs. The total heat release of n-Al@PDA@CuO (2%) is found to be higher than that of the other n-Al@PDA@CuO samples. During the combustion of n-Al@PDA@CuO (2% and 5%)/kerosene droplets, a yellow flame appears around the agglomerate, which eventually evolves into a yellow light sphere with a diameter of around 5 mm after the kerosene flame extinguishes. Examination of the normalized droplet diameter curves and the morphology of combustion residues shows that the n-Al@PDA@CuO (2%)/kerosene droplet has the best all-around combustion characteristics. A mechanism is proposed for the enhancement of kerosene combustion by n-Al@PDA@CuO. Supporting data from thermogravimetry and differential scanning calorimetry (TG-DSC) measurements are presented. The findings and data in this study are expected to advance the development of new kerosene-based nanofluid fuels that meet the high energy demands of liquid propulsion systems.

Challenges of Ablatively Cooled Hybrid Rockets for Satellites or Upper Stages

Ablative-cooled hybrid rockets could potentially combine a similar versatility of a liquid propulsion system with a much simplified architecture. These characteristics make this kind of propulsion attractive, among others, for applications such as satellites and upper stages. In this paper, the use of hybrid rockets for those situations is reviewed. It is shown that, for a competitive implementation, several challenges need to be addressed, which are not the general ones often discussed in the hybrid literature. In particular, the optimal thrust to burning time ratio, which is often relatively low in liquid engines, has a deep impact on the grain geometry, that, in turn, must comply some constrains. The regression rate sometime needs to be tailored in order to avoid unreasonable grain shapes, with the consequence that the dimensional trends start to follow some sort of counter-intuitive behavior. The length to diameter ratio of the hybrid combustion chamber imposes some packaging issues in order to compact the whole propulsion system. Finally, the heat soak-back during long off phases between multiple burns could compromise the integrity of the case and of the solid fuel. Therefore, if the advantages of hybrid propulsion are to be exploited, the aspects mentioned in this paper shall be carefully considered and properly faced.

Modular Impulsive Green Monopropellant Propulsion System (MIMPS-G): For CubeSats in LEO and to the Moon

Green propellants are currently considered as enabling technology that is revolutionizing the development of high-performance space propulsion, especially for small-sized spacecraft. Modern space missions, either in LEO or interplanetary, require relatively high-thrust and impulsive capabilities to provide better control on the spacecraft, and to overcome the growing challenges, particularly related to overcrowded LEOs, and to modern space application orbital maneuver requirements. Green monopropellants are gaining momentum in the design and development of small and modular liquid propulsion systems, especially for CubeSats, due to their favorable thermophysical properties and relatively high performance when compared to gaseous propellants, and perhaps simpler management when compared to bipropellants. Accordingly, a novel high-thrust modular impulsive green monopropellant propulsion system with a micro electric pump feed cycle is proposed. MIMPS-G500mN is designed to be capable of delivering 0.5 N thrust and offers theoretical total impulse Itot from 850 to 1350 N s per 1U and >3000 N s per 2U depending on the burnt monopropellant, which makes it a candidate for various LEO satellites as well as future Moon missions. Green monopropellant ASCENT (formerly AF-M315E), as well as HAN and ADN-based alternatives (i.e., HNP225 and LMP-103S) were proposed in the preliminary design and system analysis. The article will present state-of-the-art green monopropellants in the (EIL) Energetic Ionic Liquid class and a trade-off study for proposed propellants. System analysis and design of MIMPS-G500mN will be discussed in detail, and the article will conclude with a market survey on small satellites green monopropellant propulsion systems and commercial off-the-shelf thrusters.

On the feasibility of developing a single- stage acceleration/deceleration unit based on the oxygen-hydrocarbon engine 11D58M for a super-heavy launch vehicle

The paper presents results of unsolicited exploratory design studies done by the authors into the feasibility of developing for a super-heavy launch vehicle a single-stage oxygen-hydrocarbon acceleration/deceleration unit (ADU) with two liquid-propellant rocket engines 11D58M developed by RSC Energia, intended for insertion of manned spacecraft into lunar orbit, as well as for insertion of super-heavy spacecraft into geostationary orbit (including the orbital module high-apogee transfer profile using lunar gravity assist maneuver). It demonstrates that the single-stage ADU will have a number of important advantages over both a single-stage oxygen-hydrogen ADU and a functionally similar two-stage acceleration/deceleration system of an orbital module in the form of a tandem stack of an oxygen-hydrogen acceleration stage and correction and braking stage. To assure the start-ups of the main liquid propulsion system of the ADU, it proposes a new method for inertial propellant component phase separation in the tanks in zero-gravity environment using a pre-startup pre-programmed ullage separation turn maneuver of the orbital unit about its transverse axis of inertia.

FEASIBILITY ANALYSIS OF LIQUID PROPELLANT SETTLING IN THE TANKS OF A SPACE PROPULSION SYSTEM AFTER FLYING IN ZERO GRAVITY USING A SEPARATING TURN MANUEVER OF THE ORBITAL UNIT

The paper presents results of preliminary design feasibility studies of using inertial settling of liquid propellant in tanks of space liquid rocket propulsion systems and separating from the propellant the ullage gases formed in zero gravity by means of centrifugal forces generated when the orbital unit performs the programmed turn maneuver proposed by the author. It shows that performing propellant settling and ullage gas separation by means of the separating turn maneuver makes it possible to significantly reduce propellant consumption in propulsion systems for attitude control and ullage engines. Key words: space liquid rocket propulsion system, main engine, separation of ullage gases from liquid propellant, separating turn maneuver of the orbital unit.

Feasibility analysis of liquid propellant settling in the tanks of a space propulsion system after flying in zero gravity using a separating turn maneuver of the orbital unit

The paper presents results of preliminary design feasibility studies of using inertial settling of liquid propellant in tanks of space liquid rocket propulsion systems and separating from the propellant the ullage gases formed in zero gravity by means of centrifugal forces generated when the orbital unit performs the programmed turn maneuver proposed by the author. It shows that performing propellant settling and ullage gas separation by means of the separating turn maneuver makes it possible to significantly reduce propellant consumption in propulsion systems for attitude control and ullage engines. Key words: space liquid rocket propulsion system, main engine, separation of ullage gases from liquid propellant, separating turn maneuver of the orbital unit.

Review of State-of-the-Art Green Monopropellants: For Propulsion Systems Analysts and Designers

Current research trends have advanced the use of “green propellants” on a wide scale for spacecraft in various space missions; mainly for environmental sustainability and safety concerns. Small satellites, particularly micro and nanosatellites, evolved from passive planetary-orbiting to being able to perform active orbital operations that may require high-thrust impulsive capabilities. Thus, onboard primary and auxiliary propulsion systems capable of performing such orbital operations are required. Novelty in primary propulsion systems design calls for specific attention to miniaturization, which can be achieved, along the above-mentioned orbital transfer capabilities, by utilizing green monopropellants due to their relative high performance together with simplicity, and better storability when compared to gaseous and bi-propellants, especially for miniaturized systems. Owing to the ongoing rapid research activities in the green-propulsion field, it was necessary to extensively study and collect various data of green monopropellants properties and performance that would further assist analysts and designers in the research and development of liquid propulsion systems. This review traces the history and origins of green monopropellants and after intensive study of physicochemical properties of such propellants it was possible to classify green monopropellants to three main classes: Energetic Ionic Liquids (EILs), Liquid NOx Monopropellants, and Hydrogen Peroxide Aqueous Solutions (HPAS). Further, the tabulated data and performance comparisons will provide substantial assistance in using analysis tools—such as: Rocket Propulsion Analysis (RPA) and NASA CEA—for engineers and scientists dealing with chemical propulsion systems analysis and design. Some applications of green monopropellants were discussed through different propulsion systems configurations such as: multi-mode, dual mode, and combined chemical–electric propulsion. Although the in-space demonstrated EILs (i.e., AF-M315E and LMP-103S) are widely proposed and utilized in many space applications, the investigation transpired that NOx fuel blends possess the highest performance, while HPAS yield the lowest performance even compared to hydrazine.

Influence of propellant leakage from pump area into turbine area on turbo-pump operation stability

There is an increasing trend to liquid-propellant rocket engines which run on eco-friendly storable propellant. This trend is mostly dictated by the refusal to use traditional toxic storable propellant in many countries. The most widespread eco-friendly storable propellant is hydrogen peroxide with kerosene. Though, this propellant has a lower specific impulse in comparison with traditional liquid oxygen with kerosene. To compensate the loss of specific impulse, there is a reason to design a staged combustion engine. Evidently, the turbopump is the most complicated system in the staged combustion propulsion system. This fact makes research devoted to turbo-pumps a top priority. The paper aims to determine the influence of propellant leakage from the pump area into the turbine area and create recommendations which would allow organizing the stable operation of turbopump. As a result of turbopump staged combustion cycle testing, a conclusion had been made that leakage, which opens during the test, significantly influences the stability of turbopump operation. Depending on the amount of leakage, the turbine generated power drop was between 20 and 45%, which led to a decrease in rotation speed and outlet pressure of the pump. During the R&D process, a way of leakage influence elimination had been offered. Formulated recommendations may be used during the design process of the turbopump for staged combustion liquid propulsion systems.