Liquid Propellant Research Articles

Study of Hedge Ignition and Flame Propagation Characteristics of Unsymmetrical Dimethylhydrazine and Its Metamorphosed Mixtures in a Nitrogen Tetroxide Atmosphere

Unsymmetrical dimethylhydrazine (UDMH) is a common liquid propellant widely used in rocket engines and other applications. The safety of UDMH in service is affected by its slow oxidation during long-term storage to form impurities such as dimethylamine (DMA) and formaldehyde dimethylhydrazone (FDH). How these impurities affect combustion performance is not known, and in order to assess these effects, the present experiments investigated the combustion characteristics of self-igniting fuels and carried out ignition delay time measurements and flame propagation velocity measurements of pure UDMH and its denatured mixtures in a nitrogen tetroxide (NTO) atmosphere. This experiment was carried out to measure the delay time of hedge ignition of pure UDMH and qualitative analysis of its flame propagation properties under vacuum conditions at room temperature (T = 293 K). Ignition delay time measurements and flame propagation characterization were performed under the same experimental conditions for UDMH mixed with 1%, 5% and 10% FDH, UDMH mixed with 1%, 5% and 10% H2O, UDMH mixed with 1%, 5% and 10% DMA, as well as for UDMH mixed with the same proportions of the three substances (1%, 5% and 10%). The flame propagation characteristics were analyzed. The results showed that the incorporation of DMA, H2O and FDH in different proportions could inhibit the combustion of UDMH to varying degrees and prolong its ignition delay time. It is worth noting that the introduction of FDH had the least effect on it, and the least effect was observed at a concentration of 1%. In contrast, the effect of DMA on UDMH is more obvious, and the addition of H2O has the largest increase in the ignition delay time of UDMH. In the flame propagation experiment, the flame of the experimental group adding H2O can no longer fill the whole experimental window, while the other experimental groups can still make the window full of flame. Combined with the measurements of the ignition delay time, it can be seen that the moisture content has the greatest effect on the combustion characteristics.

A wide ranging experimental and kinetic modeling study of TMEDA pyrolysis and oxidation

Traditional liquid propellants like unsymmetrical dimethylhydrazine include several issues, such as toxicity, short storage time and preservation difficulty at low temperatures. For this reason, exploring green and stable propellant fuels is imperative in the development of propellants. Tetramethyl ethylenediamine (TMEDA) is considered to be a promising green propellant fuel with high specific impulse, non-toxicity, and long-term storage stability. Thus, investigating the combustion characteristics of TMEDA and constructing a detailed chemical kinetic mechanism are of fundamental importance in describing its combustion in computational fluid simulation. For this work, ignition delay times of TMEDA were measured in a shock tube at equivalent ratios of 0.5, 1.0 and 2.0, at pressures of 2, 10, and 20 bar, and in the temperature range of 870–1500 K with fixed fuel concentration at 2 %. After the reflected shock at 990–1500 K, the pyrolysis products of 2000 and 5000 ppm TMEDA in Argon were also measured at 5 and 10 bar conditions in a single pulse shock tube. In addition, laminar flame speeds of TMEDA were measured at initial pressures of 1 bar, initial temperatures of 333 and 353 K, and equivalence ratios ranging from 0.8 to 1.5 in a constant volume reactor. The experimental results were simulated using a detailed TMEDA kinetic mechanism. Reaction path analysis and sensitivity analysis were provided to clarify the chemical kinetics of TMEDA oxidation and pyrolysis. The current work provides new experimental data and fundamental analyses for understanding the oxidation and pyrolysis characteristics of TMEDA, and sheds light on optimal directions for future models.

Design and Performance of an Electrowetting Ionic Liquid Electrospray Thruster Prototype

Porous ionic liquid electrospray thrusters are capable of meeting the propulsion requirements of micro- and nanosatellites. The liquid propellant of such a thruster is usually prefilled in the porous reservoir and emitter, then passively and uncontrollably supplied to the emitter tip, which may cause excessive propellant to submerge the emitter, thereby resulting in a short circuit of the thruster. To solve the aforementioned problems, an electrowetting ionic liquid electrospray thruster prototype is developed in this study. Electrowetting is an active control technology that can supply propellant on demand from the propellant tank to the emitter. The feasibility of electrowetting the emitter and the performance of the prototype are tested in a vacuum chamber. The results show that the porous reservoir is completely wetted by the propellant within 410 s after the electrowetting process is initiated. The prototype can work stably in the voltage range of ±2300 to ±2800 V, with a transmission efficiency of greater than 95%. The maximum specific impulse and thrust of this prototype are 5025 s and 13.99 μN, respectively, which are similar to the performance characteristics of passively fed prototypes. This study is of great significance for improving the controllability and robustness of ionic liquid electrospray thrusters.

FEATURES OF THE DEVELOPMENT OF ADDITIVE MANUFACTURING METHODS IN APPLICATION TO LIQUID PROPELLANT ROCKET ENGINES

Abstract. Modern development of additive manufacturing methods makes it necessary to seriously consider an alternative type of production for the manufacture of liquid rocket engine structures. On the one hand, additive manufacturing has a number of undeniable advantages compared to classical methods traditionally used in the rocket and space industry. On the other hand, products obtained by this method have characteristic features that must be taken into account at the early stages of design: relatively high surface roughness, the need to adapt the geometry of existing technical solutions, material properties, etc. Moreover, a full-fledged quality control system for the resulting products is currently the moment is missing, and the standardization of the production class is just emerging. It is also important to understand the direction of development of additive manufacturing, which in the future can lead to a significant increase in the characteristics of the resulting products. This paper discusses options for classifying additive manufacturing methods, provides definitions and classification of metal 3D printing methods, according to the existing ISO/ASTM 52900 standard, as well as the main features of the technology. The main advantages and disadvantages of additive manufacturing methods are considered, and a structural decomposition of the typical design of a liquid propellant rocket engine chamber is carried out. The work aimed at manufacturing structural elements for further integration into the production cycle of liquid-propellant rocket engines is considered. The work is considered, the results of which were a combination of additive technology methods, which made it possible to apply the concept of bimetallic additive production. Such products were produced without the use of complex technological equipment and belong to the product of “evolving” additive manufacturing. Based on the results of the work, it was concluded that it is possible to use additive technologies to improve the performance of new liquid-propellant rocket engine designs.

Preparation and properties of high-energy-density aluminum/boron-containing gelled fuels

Energetic nanofluid fuel has caught the attention of the field of aerospace liquid propellant for its high energy density (HED), but it suffers from the inevitable solid–liquid phase separation problem. To resolve this problem, herein we synthesized the high-Al-/B-containing (up to 30% (mass)) HED gelled fuels, with low-molecular-mass organic gellant Z, which show high net heat of combustion (NHOC), density, storage stability, and thixotropic properties. The characterizations indicate that the application of energetic particles to the gelled fuels obviously destroys their fibrous network structures but can provide the new particle–gellant gelation microstructures, resulting in the comparable stability between 1.0% (mass) Z/JP-10 + 30% (mass) Al or B and pure JP-10 gelled fuel. Moreover, the gelled fuels with high-content Al or B exhibit high shear-thinning property, recovery capability, and mechanical strength, which are favorable for their storage and utilization. Importantly, the prepared 1.0% (mass) Z/JP-10 + 30% (mass) B (or 1.0% (mass) Z/JP-10 + 30% (mass) Al) shows the density and NHOC 1.27 times (1.30) and 1.43 times (1.21) higher than pure JP-10, respectively. This work provides a facile and valid approach to the manufacturing of HED gelled fuels with high content of energetic particles for gel propellants.

A novel in-flight thrust modulation technique for SPRM: proof of concept

Thrust modulation (termination or reduction) is one of the most sophisticated topics in rocket motors. Thrust termination is employed when combustion must be terminated to ensure proper stage separation, to avoid motor explosion, to attain certain range, or for research purposes. While liquid and hybrid propellant rocket motors have the ability to extinguish thrust, thrust termination in solid propellant rocket motors (SPRMs) need the use of specialized technology. In-flight thrust termination of (SPRMs) can be achieved by sudden and sufficient increase in throat area. Increasing throat area causes high depressurization rate which consequently leads to extinguishing the rocket motor. In contrast, in-flight thrust reduction can be theoretically attained either by reducing chamber pressure to a new equilibrium level or by reducing the divergent section length. The objective of the present paper is to prove the concept of using Linear Shaped Charge (LSC) for thrust modulation; a technique novel to SPRM applications. Thrust termination concept is proved via a set of static firing tests on a standard test motor. In addition, a mathematical model is developed to predict the drop in thrust due to cutting the nozzle at any arbitrary location. The model is validated by comparing with published experiments for thrust termination and reduction. A parametric study is conducted based on model results. It is confirmed that by cutting the nozzle along its divergent section, thrust reduction less than 20% can be attained. Further reduction can be achieved by cutting the nozzle along its convergent section.

Experimental study on microwave-induced puffing, micro-explosion, and combustion characteristics of ammonium dinitramide-based liquid propellant droplets

Microwave ignition technology has the advantages of high ignition energy, stable ignition, and spatial multi-point ignition. These advantages make this technology promising for future application in green single-component propellants. In this paper, the ignition characteristics of ammonium dinitramide (ADN)-based liquid propellant droplets under the influence of microwaves at room temperature are investigated using experimental methods. The effects of microwave power on puffing, micro-explosion, and combustion behavior of ADN-based liquid propellant droplets were studied. The droplet and flame diameters were statistically related to time, and the microwave-assisted droplet ignition mechanism was analyzed. A new rectangular waveguide resonant cavity was designed in which the droplet is placed at the maximum electric field strength of the device. The droplet morphology and flame profile inside the resonant cavity were photographed with a high-speed camera. The experimental results showed that the microwave positively influenced the puffing, micro-explosion, and combustion behavior of droplets. When the microwave power was increased from 200 to 280 W, the total droplet evaporation time and ignition delay time were reduced by 56.5% and 35.2%, respectively. The positive effects of microwaves on combustion have been summarized as the thermal effect of microwaves on polar molecules and the promotion of fuel oxidation reactions by microwave-induced plasma. The plasma was found to control the development of the initial flame propagation front and to influence the temperature during the combustion reaction process. In this paper, we propose the mode of droplet combustion under microwave induction as a plasma discharge and several stages of the droplet combustion process. This research provides novel insight into the study of the microwave ignition mechanism of liquid fuels.

Maximum Likelihood Identification of Backflow Vortex Instability in Rocket Engine Inducers

Abstract Bayesian estimation is applied to the analysis of backflow vortex instabilities in typical three- and four-bladed liquid propellant rocket engine inducers. The flow in the impeller eye is modeled as a set of equally intense and evenly spaced 2D axial vortices, located at the same radial distance from the axis and rotating at a fraction of the impeller speed. The circle theorem is used to predict the flow pressure in terms of the vortex number, intensity, rotational speed, and radial position. The theoretical spectra so obtained are frequency broadened to mimic the dispersion of the experimental results and parametrically fitted to the measured data by maximum likelihood estimation with equal and independent Gaussian errors. The method is applied to three inducers, tested in water at room temperature and different operating conditions. It successfully characterizes backflow instabilities using the signals of a single pressure transducer flush-mounted in the impeller eye, effectively bypassing the aliasing limitations and the data acquisition/reduction complexities of traditional multiple-sensor cross-correlation methods. The identification returns the estimates of the model parameters and their standard deviations, providing the information necessary for assessing the accuracy and statistical significance of the results. The flowrate is found to be the major factor affecting the backflow vortex instability, which, on the other hand, is rather insensitive to the occurrence of cavitation. The results are consistent with the data reported in the literature, as well as with those generated by the auxiliary models specifically developed for initializing the maximum likelihood searches and supporting the identification procedure.

Combined Cycle Nuclear System Architecture for Crewed Mars Spacecraft Propulsion and Power

Nuclear thermal propulsion can potentially reduce the time of flight and spacecraft system mass needed for human spaceflight beyond cislunar space. This nuclear propulsion system has comparable thrust capability to chemically impulsive systems, which at about twice the specific impulse, can double the delta-velocity (ΔV) for the same propellant mass. However, the canonical problem for nuclear propulsion has always been that its benefits are shadowed by low technology readiness of a complex system. This paper describes a combined cycle nuclear thermal rocket (CCNTR) system architecture for propulsion and electrical power that comprises a 42-MWt-capable nuclear reactor core to provide 9.4 kN thrust on demand at a specific impulse of 940 s. The liquid hydrogen propellant flow through the rocket chamber cools the reactor during burns, thereby producing thrust while concurrently rejecting waste heat to space. The reactor also produces up to 100 kWe power for the spacecraft, eliminating the need for solar power generation and averting challenges associated with restarting a cold reactor for propulsive burns. Radiators reject the waste heat from electrical power production. Earth-to-Mars orbital transfers less than 100 days appear feasible assuming 680,000 kg of liquid hydrogen propellant and a vehicle dry mass of 83,000 kg that includes the 13,000 kg CCNTR system. Together, these results suggest that a CCNTR could be most promising to enable crewed missions to Mars.

Insight into thermal decomposition behavior of ammonium dinitramide‐based liquid propellant through reactive molecular dynamics simulations

AbstractThe thermal decomposition characteristics and fuel oxidation mechanism of ammonium dinitramide (ADN)‐based liquid propellant are of great practical significance, but it is difficult to detect the whole intermediates and products using traditional experimental methods. In this paper, the thermal decomposition of ADN‐based liquid propellant at different temperatures (1500, 2000, 2500, and 3000 K) was simulated using the reactive molecular dynamics method. The effect of temperature on the thermal decomposition of propellant and the distribution of the main products were investigated. The results showed that the temperature played a role in promoting the decomposition of propellant. The main products of ADN‐based liquid propellant were N2, H2O, NH3, HNO3, ⋅NHO, CH2O, CO2, and NO2, while their evolution and generation pathways were also described in detail. The initial decomposition of ADN was the breakage of N−H bonds in ⋅NH4 or N−N/N−O bonds in ⋅N(NO2)2, and methanol generated free radicals by dehydrogenation. There was no direct chemical reaction between ADN and methanol, and the interaction was accomplished through the generation and consumption of free radicals. H2O provided a large amount of ⋅H and ⋅HO leading to a lower activation energy of ADN in propellant than that from reference. The overall reaction mechanism of the liquid propellant was deduced. The results would help to further investigate the reaction mechanism of ADN‐based liquid propellants.

Analytical models for critical heights of vortexing in flat and spherical bottom tanks with siphon and bottom drains

Critical height or critical submergence is liquid level at which air-core vortex extends from the free surface into drain hole when a liquid is drained from a container/tank. Extensive analytical and experimental studies have been reported on critical height of bath tub vortex, for liquid draining downward from flat bottom propellant tanks. Rockets making use of liquid propellants mostly employ spherical bottom propellant tanks as well as siphon or upward drain flow. Keeping in view of such practical applications, analytical models are developed for critical height, considering the effects of siphon drain and the shape of tank bottom. Additional design parameters influencing the behavior for each case are identified. Appropriate governing equations and a solution methodology are developed pertinent to the system considered, to predict the critical height for siphon drain and spherical bottom tank independently as well as for both combined. The results indicate that the critical height for spherical bottom tank is higher than for flat bottom tank, due to higher local flow velocity. Siphon geometry can be designed for critical height much less than normal drain from flat bottom tank. These observations are in accordance with the results published in the literature. This paper reports the analytical models and solution methodology to predict the critical height for vortexing for normal draining from spherical tank bottom, siphon draining from both flat and spherical bottom tanks.

Risk reduction of tank explosion based on passivation of unusable propellant residues

To reduce the risk of explosion of propellant tanks of expended spacecraft and launch vehicles with liquid rocket engines in orbit, as well as in case of emergency situation, for example, loss of orientation, the Inter-Agency Space Debris Coordination Committee recommends passivation measures, including the discharge of residual liquid propellant and pressurant gas. In ANSYS-Fluent program complex possible initial positions of liquid propellant residues in a spherical tank at its rotation under conditions of low gravitational fields are determined. The values of liquid propellant residues depending on their initial position in the spherical tank at opening of the drain line for discharge of gas–liquid mixture into the ambient space are determined. The concept of formation of two-phase flows of liquid propellant on the example of the spherical tank at tangential entry of compressed gas is offered. The relationship between the number of gas inlet points and the effectiveness of the developed method (expressed as the ratio of the mass of expelled liquid propellant to the mass of gas expended) is demonstrated. For instance, the use of 2 gas inlet points achieves an efficiency of up to 30 %, while employing 3 gas inlet points increases it to 89 %.

Method for passivation of propellant residues in orbital stage tank of launch vehicle

The problem of removing unused liquid propellant residues from the tanks of spent spacecraft and orbital stages of Launch Vehicles (LV) leads to their explosion and the formation of space debris in orbits. To provide a solution to this problem, a method for removing liquid propellant residues from the LV tanks after the mission completion is proposed. The method is based on the gasification of liquid propellant residues in the tanks under acoustic-vacuum exposure and the discharge of the gasification products into the surrounding outer space. Experimental investigations were carried out on a Ground-based Experimental Installation (GEI) to determine the coefficient of heat transfer from the surface of an acoustic radiator to a liquid. The obtained coefficient was then used to calculate the energy costs for the gasification of kerosene. Numerical estimates are given on the example of the tank with kerosene residues from a spent second stage of the LV “Soyuz-2.1 v”. The optimal discharge rate at which kerosene does not freeze is 0.14 m3/s. Moreover, the acoustic exposure leads to an increase in the mass of evaporated kerosene over a given time by 96.1%, and the energy costs are 1756.7 kJ (approximately 50% of the remaining electrical energy).

Development of the algorithm for the movement of the measuring module of the system of automated non-destructive testing of the nozzle of a lpre

The paper considers the development of the algorithm for the movement of the measuring module of the system of automated non-destructive testing of solder joints of the nozzle of a liquid propellant rocket engine (LPRE). The object of control, the flaws in soldered joints that arise during the manufacturing process, as well as the problems of production and testing of assembly units, are described. The main factors that reduce the probability of detecting defects are determined, and solutions are proposed. The necessity of using route construction algorithms in the developed traffic algorithm is substantiated. An algorithm for obtaining the path of the transducer movement for the measuring module of the system of automated non-destructive testing of solder joints is defined to detect such defects as "dry joint", "cold lap" and "partial dry joint".

Thermal decomposition characteristics and thermal safety performance evaluation of HAN-based propellant under different external environmental stimulations

This research aims to investigate thermal characteristics and assess thermal hazards of hydroxylamine nitrate (HAN)-based liquid propellant using a combination of accelerating rate calorimeter and high-pressure differential scanning calorimetry. Adiabatic experiments revealed that the exothermic reaction initiated at 115.7 °C and a rise sharply was at 128.4 °C with a maximum self-heating rate of 164.3 °C/min and the exothermic event was accompanied by a pressure rise of 3.7 bar in 0.03s. The corrected values of adiabatic temperature rise and time to maximum rate were 605.6 °C and 1.46 min, which confirms the vulnerability of the propellant to undergo a catastrophic explosion. To prevent thermal loss prevention accidents, time to maximum rate was obtained as 24 h under 111.3 °C. The apparent activation energy calculated decreased greatly with the increase of storage temperature from 25 °C to 45 °C, and so did the thermal explosion temperature (Tb) and self-accelerating decomposition temperature (TSADT). Additionally, the linear relationship between Ea and T was E = 991.7–2.7904T. The reliability of TSADT prediction was validated through slow cook-off test. Furthermore, the propellant exhibited more violent thermal decomposition under high pressure, resulting in a higher peak power. The thermokinetic parameters related to this phenomenon were identified, specifically at pressure of 4.0 MPa.

Comparative molecular dynamics study of liquid hydrogen annular jets in subcritical and supercritical environments

In this paper, the injection characteristics of a liquid hydrogen propellant jet under subcritical and supercritical operating conditions are comparatively studied at the molecular scale using molecular dynamics methods in the context of spacecraft propulsion systems. To control the temperature and pressure within the environment and nozzle respectively, two sub-computational domains are set up for separate modeling. The sub-computational domain merging process eliminates the broken molecular bonds by using the computational domain reduction in the x-direction. The jet velocity, density and other physical property distributions at different temperatures and pressures were comparatively studied. It is found that the rate of change of the radial density of the jet under supercritical conditions is small. Based on this, the surface area and volume of the liquid phase region of the jet were extracted and calculated. Compared to result of subcritical case, the volume of the liquid hydrogen jet in the supercritical environment is reduced by 1.9%, while the surface area is increase by 6.7%. To explain the variation in the surface area due to molecular diffusion, the molecular stress results revealed that the stress in the hydrogen jet in supercritical environment is only 88.9% of that in subcritical environment.

The characterization of disruptive combustion of organic gellant-laden ethanol fuel droplets

The future of high-performance rocket propulsion systems depends on superior burn rates, shorter ignition delay times, and eco-friendly nature. Gelled fuel propellants are such candidates which are pivotal to achieving these performance parameters. Gel fuels are multi-component fuels, and their combustion behavior is different from conventional solid and liquid propellants. To develop a fundamental understanding of the vaporization and combustion of gel fuels, single droplet experiments in a pendant mode setup are performed for ethanol-based organic gel fuels by capturing high-speed video data. Primarily, two gel fuel combinations are taken for experiments. The gel fuels are non-metalized ethanol-based fuels containing organic gellants Hydroxypropyl methylcellulose (HPMC) and methylcellulose (MC). Both fuels have shown three distinct combustion stages namely transient heat up (stage I), disruptive burning (stage II), and carbonization (stage III). The demarcation of the stages is done by visual evidence backed by high-fidelity high-speed video data. In stage I, phase separation occurs which leads to the formation of gellant shells. Organic gellants tend to form viscoelastic shells which significantly influence the jetting behavior during the disruptive burning (stage II). The combustion of gel droplets in stage II exhibits characteristic differences in bursting mechanism, for instance, oscillatory and transient bursting in HPMC and transient bursting in MC. Moreover, droplet-bursting behaviors influence jetting attributes and burn rates. Besides, the bursting of gel droplets during combustion is also dependent on the thickness of the gellant shells, the thin and weak shells are prone to bulk droplet motion due to active jetting, whereas the thick and rigid shells act as kinetic energy dampeners and thereby inhibit the bulk motion of droplets. Apart from disruptive burning, carbonization is a significant part of the combustion during which the suppressed jetting behavior is seen due to the traces of fuel trapped in the gellant shells. A significant rise in burn rate is observed due to jetting during the disruptive burning stage in contrast to the other two stages of combustion for both the gel fuels under study. Furthermore, combustion residue analysis is carried out to assess the variation in shell morphology across the stages. Overall, the current study provides detailed insights into the combustion characterization of organic gel fuels, which can help in fabricating gel fuel compositions.

Cascade nonlinear observer design for large amplitude of slosh variables in spacecraft maneuver with liquid propellant tank

Slosh dynamics in a liquid propellant tank during a spacecraft maneuver can have an effect on the system stability and potentially lead to critical problems for the mission operation. In particular, the fuel sloshing problem has drawn significant attention in aerospace applications where a large portion of the gross weight is comprised of liquid fuel. Recently, many active control methods have been introduced as alternatives to passive approaches. However, the majority of these methods basically require the full-state condition. To guarantee the availability of the full-state condition, this paper proposes a cascade nonlinear state observer based on the sliding mode theory for estimating the unmeasurable slosh variables in a planar model. Additionally, for large amplitude estimation, a complete nonlinear mathematical model with a pendulum analogy is formulated without a linearization process. The stability analysis is theoretically proven by Lyapunov method. The effectiveness is demonstrated through simulations, and the performance comparison with a different observer method is also presented in the results.

Statistical experiment analysis of wear and mechanical behaviour of abaca/sisal fiber-based hybrid composites under liquid nitrogen environment

Composite materials are increasingly replacing synthetic fiber combinations in various applications. However, certain extreme environments on Earth and in space require structures to operate under low temperatures, specifically cryogenic conditions, which can significantly affect material reactions. Therefore, the main focus of this study is to develop and evaluate hybridized biocomposites, specifically assessing their tensile, bending, and impact strengths in a controlled liquid nitrogen environment (77 K). Utilizing the Taguchi optimization method, the statistical analysis of wearing characteristics was carried out utilizing cryogenic treatment hours, load, sliding distance, and weight percentage of abaca and sisal fibers. When 20 percent abaca and sisal were mixed, tensile performance increased from 28.96 to 36.58 MPa. Likewise, the same mixture increased bending strength from 59.63 to 75.68 MPa, and impact strength improved from 59.36 to 71.25 J/m. The cryogenic treatment of composite materials for 15–30 min improved the mechanical characteristics of the materials by enhancing the binding between reinforcements and substrate. The Taguchi 27 test outcomes showed a decreased friction coefficient of 7.79 × 105 mm3/Nm in the 10th trial with 30 min of cold working, 10% hybrid fibers, 600 m slide distance, and a 4 N load combination. Frictional coefficient data indicated the lowest rate during the third experiment with 15 min of cryogenic treatment, 10% hybrid fibers, 1,500 m slide length, and a 12 N load combination. The microstructural analysis of the fractured specimen was evaluated by scanning electron microscopy. Finally, such composite materials are employed in liquid propellant tanks, satellites, spaceships, rocket constructions, aeroplane components at cruising altitudes, and other applications.

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.