Liquid Bipropellant Research Articles

Effect of temperature on galvanic corrosion of Al 6061-SS 304 in nitric acid

Nitrogen tetroxide (N2O4) is commonly used as an oxidizer for bipropellant liquid, and galvanic corrosion of its storage tanks was a serious problem. The immersion and electrochemical experiments of Al 6061-SS 304 galvanic pair were carried out in 30% nitric acid at different temperatures. The accelerated galvanic corrosion of Al 6061 and the effect of different temperatures on its corrosion behavior were studied. The surface morphology, number of corrosion pits and electrochemical parameters of Al 6061 before and after corrosion were measured and analyzed by means of SEM, EDS and electrochemical methods. The results show that when Al 6061 was coupled with SS 304 at 10 °C, the driving potential difference between them reached 366 mVSCE, and Al 6061 was used as the corrosion anode. In the SEM images, three kinds of galvanic corrosion phenomena can be seen obviously after coupling. When the temperature is between 10 °C and 30 °C, the corrosion pits are mainly composed of aluminum matrix coupled with second phase particles with positive or negative volt potential relative to it. When the temperature was between 40 °C and 50 °C, the corrosion pits showed mainly intergranular corrosion at the interface of different grains. At the same time, the increase in temperature led to the decrease in galvanic potential and the change in corrosion current. There was no simple linear relationship between galvanic current and temperature, but it satisfied a certain quantitative relationship, and the theoretical ratio was highly consistent with the actual ratio by the K–S test.

Experimental Study of a Hypergolically Ignited Liquid Bipropellant Rotating Detonation Rocket Engine

A liquid bipropellant rotating detonation rocket engine was successfully test fired. The 6.5-in.-diam engine operates on a mixture of rocket-grade hydrogen peroxide and triglyme. Ignition is accomplished hypergolically, with the addition of a promoter to the fuel. Eight cases of spontaneously initiated, single-wave, rotating detonations were observed. Detonation wave speeds were near or exceeding Chapman–Jouguet calculations. Evidence of condensed-phase reactions was also seen, damaging hardware in short order. Both a rotating detonation and a bulk-slapping mode characterized the engine operation.

Numerical Analysis and Modelling of a 100 N Hypergolic Liquid Bipropellant Thruster

This study focuses on the stepwise procedure involved in the development of a numerical model of a bi-propellant hypergolic chemical propulsion system using key features and performance characteristics of existing and planned (near future) propulsion systems. The study targets specific impulse of 100 N delivery performance of thrust chambers which is suitable for primary propulsion and attitude control for spacecraft. Results from numerical models are reported and validated with the Rocket Propulsion Analysis (RPA) computation concept. In the modelling process, there was proper consideration for the essential parts of the thruster engine such as the nozzle, combustion chamber, catalyst bed, injector, and cooling jacket. This propulsion system is designed to be fabricated in our next step in advancing this idea, using a combination of additive manufacturing technology and commercial off the shelf (COTS) parts along with non-toxic propellants. The two non-toxic propellants being considered are Hydrogen Peroxide as the oxidiser and Kerosene as the fuel, thus making it a low-cost, readily available and environmentally-friendly option for future microsatellite missions.

The ignition process measurements and performance evaluations for hypergolic ionic liquid fuels: [EMIm][DCA] and [BMIm][DCA

Bipropellant liquid rocket engines are typically fueled by hydrazines that exhibit extreme toxicity and carcinogenity. Recently hypergolic ionic liquids (HILs) have been emerging as environmentally friendly propellant that holds the potential for hydrazine replacement due to their low volatility, easy and safe storage and handling. The HIL is a novel concept both for chemistry community and for combustion community, and plenty of research waits to be initiated to promote practical propulsion application, especially for comprehensive understanding on their fundamental combustion. In this work, hypergolic ignition of two HILs (1-ethyl-3-methylimidazolium dicyanamide([EMIm][DCA]) and 1-butyl-3-methyl-imidazolium dicyanamide ([BMIm][DCA]) were tested with three oxidizers (white fuming nitric acid (WFNA), red fuming nitric acid(RFNA), and nitrogen tetroxide). A drop test system was designed to control the collision event of a single droplet of fuel with an oxidizer pool. Two independent high speed cameras were spontaneously employed to capture the self-luminous images of the whole hypergolic ignition process as well as the initial impact process. The results showed that hypergolic ignition process of [EMIm][DCA]-WFNA is quite similar to that of monomethyl hydrazine-WFNA. Combined with thermo physical properties of HILs, hypergolic ignition process of [EMIm][DCA]-WFNA is proposed. To qualitatively demonstrate the overall hypergolic ignition and reactivity, the explosion delay time and the ignition delay times of various combinations of the fuel-oxidizer pairs were determined. The effects of the Weber number on the explosion delay time and the ignition delay time of [EMIm][DCA]-WFNA are quantitatively investigated. Hypergolicity of different propellant combinations is tested to investigate the influence of oxidizer’s oxidative capacity and HILs’ cation structure. Additionally, the propulsive performance of the two dicyanamide-based IL fuels was evaluated and compared with traditional hydrazine-based fuels, based on a steady rocket model under the same initial conditions. Results show that using WFNA as the oxidizer, BMIM DCA and EMIM DCA presented a comparable propulsive performance to that of UDMH.

Towards N-Alkylimidazole Borane-based Hypergolic Fuels.

Over the past few decades, toxic and highly volatile hydrazine derivatives have been the main fuel choices for liquid bipropellants, especially in traditional hypergolic rocket engines. The search for new hypergolic fuels as replacements for hydrazine derivatives is of great interest to researchers. In this study, a series of N-alkylimidazole borane compounds has been synthesized and characterized. Interestingly, these compounds display promising applications as potential hypergolic fuels owing to their excellent physiochemical properties including low melting points, high thermal stability, low viscosities, and unique hypergolic reactivity. Compared with popular hypergolic ionic liquids, the cost-effective and scaling-up advantages of these materials highlight their promising potential as high-performance fuels in liquid bipropellant formulations.

Combustion properties of gaseous CH4/O2 coaxial jet flames in a single-element combustor

Fundamental combustion properties of gaseous methane (CH4)/oxygen (O2) coaxial jet flames in a single-element combustor are experimentally evaluated as a preliminary step for subsequent studies of injection at very low temperature or using liquid O2 for CH4/O2 bipropellants, recently appearing to be the oncoming liquid bipropellant. A combustion chamber with quartz windows, a single shear coaxial injector and an exhaust nozzle on the downstream of the chamber is considered for the present study. Focusing on the measurements of the ignition and combustion stability limits of the coaxial jet flames in the chamber, flame visualization is also conducted by OH∗ chemiluminescence, schlieren imaging and direct imaging. Results show the ignition limits restricted than the combustion stability limits. Flame behaviors are largely classified into two, the stably attached flame and the oscillating, liftoff (near-blowout) flame. Due to cooling effects on the wall of the chamber, stably liftoff flame is not observed. The stability of the flame is greatest at fuel-rich condition (based on the injected amounts of CH4 and O2). From the flame visualization flame thickness is found to be smaller than the injector lip thickness and insensitive to injection conditions. The laminar-flow behavior near the injector exit due to the strong burning of pure O2 is observed even for high Reynolds numbers (Re). The flame visualization also exhibits the recirculating O2 that enhances burning in the combustor through the reaction with the outer fuel jet. The results of ignition limits, combustion stability limits and flame visualization can be used as a database for researches of modeling the gaseous CH4/O2 jet flames in the combustion chamber.

Towards Safer Rocket Fuels: Hypergolic Imidazolylidene‐Borane Compounds as Replacements for Hydrazine Derivatives

Currently, toxic and volatile hydrazine derivatives are still the main fuel choices for liquid bipropellants, especially in some traditional rocket propulsion systems. Therefore, the search for safer hypergolic fuels as replacements for hydrazine derivatives has been one of the most challenging tasks. In this study, six imidazolylidene-borane compounds with zwitterionic structure have been synthesized and characterized, and their hypergolic reactivity has been studied. As expected, these compounds exhibited fast spontaneous combustion upon contact with white fuming nitric acid (WFNA). Among them, compound 5 showed excellent integrated properties including wide liquid operating range (-70-160 °C), superior loading density (0.99 g cm(-3) ), ultrafast ignition delay times with WFNA (15 ms), and high specific impulse (303.5 s), suggesting promising application potential as safer hypergolic fuels in liquid bipropellant formulations.

Effects of Fuel Impingement-Cooling on the Combustion Flow in a Small Bipropellant Liquid Rocket Thruster

ABSTRACTIn the present study, a three-dimensional computer code, based on the computer software KIVA-3, was developed for the combustion-flow simulation of a bipropellant liquid rocket thruster. A jets-impingement model is proposed for the unlike-doublet jet impingement issue. The computer code is employed to simulate a small bipropellant liquid rocket engine installed with three unlike-doublet injectors of NTO and MMH as well as six fuel injectors injecting MMH toward the combustor wall for cooling. Effects of the fuel-injection cooling on the combustion flow, combustion efficiency, and wall temperature were investigated. The results obtained from the present study show that under the present injector configuration and a constant total fuel-flow rate, higher cooling-fuel ratios make the atomized mixing flow of NTO and MMH shift toward the combustor wall, resulting in lower combustion efficiency and chamber pressure; however, low cooling-fuel ratios are unable to keep the wall-temperature sufficiently low. To overcome this issue, the proposed three-dimensional computer code calculates/locates the optimal cooling-fuel ratio that bears high combustion efficiency for a bipropellant liquid rocket combustor while keeping the chamber wall sufficiently cool.

Simulation and experimental study of the plume during the Chang’E-3 lunar landing

The ChangE-3 is the fisrt moon-landing explorer of china.A deep-throttling bipropellant liquid rocket engine was used to slow the explorer to an acceptable landing velocity.As the explorer approaches the lunar surface, the plume from the engine is reflected upward by the lunar surface and impinges on the bottom surface of the explorer and most of the landing gear.In this paper, A calculation model of the interaction between rocket plume and moon surface was built by using hybrid N-S equations/DSMC method. The plume of engine coupled with moon surface was simulated, and the plume impingement effects on the landing gear and the bottom surface of the explorer were analyzed．An experimental study of thermal protection material was performed in order to verify plume-protection design.

Direct Dynamics Simulation of the Activation and Dissociation of 1,5-Dinitrobiuret (HDNB)

Certain room-temperature ionic liquids exhibit hypergolic activity as liquid bipropellants. Understanding the chemical pathways and reaction mechanisms associated with hypergolic ignition is important for designing new fuels. It has been proposed (J. Phys. Chem. A 2008, 112, 7816) that an important ignition step for the hypergolic ionic liquid bipropellant system of dicyanamide/nitric acid is the activation and dissociation of the 1,5-dinitrobiuret anion DNB(-). For the work reported here, a quasiclassical direct dynamics simulation, at the DFT/M05-2X level of theory, was performed to model H(+) + DNB(-) association and the ensuing unimolecular decomposition of HDNB. This association step is 324 kcal/mol exothermic, and the most probable collision event is for H(+) to directly scatter off of DNB(-), without sufficient energy transfer to DNB(-) for H(+) to associate and form a highly vibrationally excited HDNB molecule. Approximately 1/3 of the trajectories do form HDNB, which decomposes by eight different reaction paths and whose unimolecular dynamics is highly nonstatistical. Some of these paths are the same as those found in a direct dynamics simulation of the high-temperature thermal decomposition of HDNB (J. Phys. Chem. A 2011, 115, 8064), for a similar total energy.

Investigation of self-excited combustion instabilities in two different combustion systems

The objective of this paper is to characterize dynamic pressure traces measured at self-excited combustion instabilities occurring in two combustion systems of different hardware. One system is a model lean premixed gas turbine combustor and the other a fullscale bipropellant liquid rocket thrust chamber. It is commonly observed in both systems that low frequency waves at around 300Hz are first excited at the onset of combustion instabilities and after a short duration, the instability mode becomes coupled to the resonant acoustic modes of the combustion chamber, the first longitudinal mode for the lean premixed combustor and the first tangential mode for the rocket thrust chamber. Low frequency waves seem to get excited at first since flame shows the higher heat release response on the lower frequency perturbations with the smaller phase differences between heat release and pressure fluctuations. Nonlinear time series analysis of pressure traces reveals that even stable combustion might have chaotic behavior with the positive maximum Lyapunov exponent. Also, pressure fluctuations under combustion instabilities reach a limit cycle or quasi-periodic oscillations at the very similar run conditions, which manifest that a self-excited high frequency instability has strong nonlinear characteristics.

Bipropellant combustion in a liquid rocket combustion chamber

A comprehensive computer simulation code is developed for the combustion-flow analysis of a bipropellant liquid rocket engine combustor. A bipropellant group combustion model is built in the computer code to facilitate the determination of various combustion modes, including the combustion of oxidizer droplets, fuel droplets and gas mixture. Several physical submodels are also adopted to account for the interaction between bipropellant droplets and the gas-phase flow. An axisymmetric cylindrical liquid rocket engine combustion chamber installed with two annular-ring injectors is selected to demonstrate the capabilities of the present computer simulation code. Two types of injection processes are analyzed. One is characterized by a non-premixed spray process and the other is a premixed spray process. Selected numerical computations show that the premixed spray results in a better utilization of combustion chamber space and gives a higher combustion efficiency in general. But finer-droplet sprays do not necessarily lead to a higher combustion efficiency as traditionally speculated.

Barrier Film Cooling Study

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Measurements of the Combustion Time Lag in a Liquid Bipropellant Rocket Motor

Current theories concerning combustion instability in rocket motors involve the concept of a time delay between the instant of injection of a propellant element into the combustion chamber and its final evolution into hot gases. In order to apply these theories to actual design problems, it is first necessary to determine experimentally the time lag and several associated parameters for each class of propellants and chamber configurations. These measurements are now being made on a bipropellant chamber at Princeton as the second phase of a program whose ultimate objective is to provide the information necessary to direct the design 01 development of rocket engines toward suppression of high-frequency combustion instability. The introductoiy phase of this program, described in Ref. (2), dealt with the development of an experimental technique and instrumentation capable of providing the necessary data. The current paper presents some preliminary results of the second phase, in which these techniques have been used to determine the combustion chamber transfer function of a particular bipropellant configuration and the deduction therefrom of the time lag and related parameters. A brief recapitulation of the conceptual background is given, together with the methods used to obtain the desired measurements. Some of the limitations of these deduced results are discussed, and current methods for reducing the limitations are described.

Servo-Stabilization of Low-Frequency Oscillations in a Liquid Bipropellant Rocket Motor

Servo-Stabilization of Low-Frequency Oscillations in a Liquid Bipropellant Rocket Motor