Monopropellant Hydrazine Research Articles

Oxidation induced emissivity variations of the Haynes 25 alloy at high temperatures

Haynes 25 alloy is used in the design of propulsion systems for spacecrafts and it can constitute a space debris at the end of life of the mission. All the thruster combustion chambers and nozzles are manufactured with Haynes 25 alloy and the thrusters used storable monopropellant hydrazine. During the atmospheric reentry of thruster pieces at high velocity, the Haynes 25 alloy is submitted to high temperatures in an extreme environment with atomic oxygen. This is why the high temperature oxidation of Haynes 25 was studied, in air plasma conditions in the temperature range 1000–1700 K. Moreover, as emissivity is an important parameter to know for the heat transfer, the spectral directional emissivity was measured on several samples surfaces: pristine, pre-oxidized in air plasma and in situ oxidized in air. The high temperature oxidation of Haynes 25 is moderate up to 1500 K and above an important mass loss occurred up to 1700 K (around 40 mg/cm2), that is the determined melting temperature. The emissivity of pristine Haynes 25 is vey low, around 0.20 at 1350 K and it increases with oxidation up to 0.85–0.90 at 1600 K.

Numerical modeling and parametric analysis of performance of a monopropellant thruster using a single-part catalyst bed model

Monopropellant hydrazine thruster, depending on their thrust level, specific impulse, and unique functional regime, are widely used in situation control, orbital transmission, and position correction systems of satellites. In these thrusters, hydrazine decomposes by passing through the catalyst bed in a highly exothermic reaction to hot gas products. Hot gases generate thrust force by passing through a convergent-divergent nozzle. Pore scale analysis of catalytic reactions is very common in various industries and is of interest to researchers due to its accuracy. In this paper, the decomposition chamber of a monopropellant hydrazine thruster is numerically simulated with a single-part bed model at the pore-scale. The length of decomposition chamber was 2.48 cm. Then the effects of parameters such as catalyst granule diameter, catalyst bed porosity coefficient and also chamber inlet pressure on the performance of the decomposition chamber and thruster are investigated. Simulations have been performed for catalyst granules with diameters of 0.88, 1.00 and 1.15 mm in three porosity coefficients of 0.4, 0.55 and 0.65. The inlet pressure is also changed from 10 to 25 bar in four different levels. The results showed that the porosity coefficient is the most effective parameter and with its decrease, the specific impulse and temperature rise, while the thrust force and mass flow rate intensify. Also, the size of the catalyst granules affects the performance of the bed and thruster so that by increasing it (at a certain porosity coefficient), a trend similar to the effect of decreasing the porosity coefficient can be seen in the results. On the other hand, with enhancing inlet pressure, the thrust force increases significantly. In this paper, the effect of bed parameters on the thruster performance is discussed in detail, which contains helpful results for researchers that work on improving the decomposition chamber efficiency.

Development of bimetallic spinel catalysts for low-temperature decomposition of ammonium dinitramide monopropellants

Ammonium dinitramide (ADN) based liquid monopropellants have been identified as environmentally benign substitutes for hydrazine monopropellant. However, new catalysts are to be developed for making ADN monopropellants cold-start capable. In the present study, performance of Co and Ba doped CuCr2O4 nanocatalysts prepared by hydrothermal method was evaluated on the decomposition of aqueous ADN solution and ADN liquid monopropellant (LMP103X). The catalysts were characterized by PXRD (Powder X-ray Diffraction), FTIR (Fourier Transform Infrared spectroscopy), SEM (Scanning Electron Microscopy), TEM (Transmission Electron Microscopy), EDS (Energy Dispersive X-ray Spectroscopy), and XPS (X-ray Photoelectron Spectroscopy). The nanosize was confirmed by SEM and TEM, while the nanoflake morphology was confirmed by the SEM analysis. Further, we obtained the elemental composition from the EDS analysis. We investigated the catalytic activity of the catalysts by thermogravimetric (TG) analysis and the developed catalysts lowered the decomposition temperature of ADN monopropellant by about 55 °C. The XPS analysis confirmed the presence of metal ions with different chemical states. Apparently, increase in the surface area of the catalysts and the mixed active sites as well as the development of oxygen vacancy on the catalyst surface introduced by metal doping are influencing the decomposition temperature of ADN samples.

Case study on thermal design improvement of hydrazine monopropellant pipelines for satellite propulsion system

The present research investigates a case study on thermal design and improvement of hydrazine monopropellant pipelines based on actual cases, in order to highlight the critical factors for thermal design when developing an onboard propulsion system for a satellite. A simple 1-D thermal math model was developed for thermal analysis using the lumped-parameter assumption, and three different thermal design cases of hydrazine propellant pipeline layouts were considered. By increasing heater power densities and applying modified worst cold boundary conditions for thermal analysis, the thermal design performance for the propellant pipelines of an actual satellite could be improved and verified through thermal vacuum test.After the successful operation of the pipeline heaters in the real space environment, some important lessons learned on thermal design improvements of hydrazine monopropellant pipeline layout could be attained from the present case study. These lessons learned will be useful for the future development of satellite onboard propulsion systems.

LE111 HYDRAZINE BIPROPELLANT ENGINE (HBE) WITH GAS-GAS INJECTION

The LE111 thruster utilizes hydrazine and MON3 as fuel and oxidizer, and produces a thrust of 111 N at nominal operating conditions with an Isp of up to 328 s. The thruster is throttleable through the use of separate fuel and oxidizer metering valves and is able to transition to a hydrazine monopropellant mode at any point during its operation. Completed qualification testing has demonstrated a versatile operating box that includes chamber pressure - mixture ratio excursions, heated propellants up to 341K, and gaseous helium ingestion through propellant flow paths. Overall, the thruster has achieved 15,455 s of accumulated on-time and a maximum continuous burn time of 6,000 s. The thruster achieves its exceptional performance through a novel micro-coaxial gas-gas injection scheme and an innovative regeneratively cooled combustion chamber, which uses the oxidizer as the working fluid. This paper describes the thruster’s capabilities and performance recorded during the completed qualification testing, and briefly outlines the design.

111 N HYDRAZINE BIPROPELLANT ENGINE (HBE) WITH GAS-GAS INJECTION

A thruster, designated LE111, utilizes hydrazine and MON3 as fuel and oxidizer and produces a thrust of 111 N at nominal operating conditions with an Isp of up to 328 s. The thruster is throttleable through the use of separate fuel and oxidizer metering valves and is able to transition to a hydrazine monopropellant mode at any point during its operation. Completed qualification testing has demonstrated a versatile operating box that includes chamber pressure-mixture ratio excursions, heated propellants up to 341 K, and gaseous helium ingestion through propellant flow paths. Overall, the thruster has achieved 15,455 s of accumulated on-time and a maximum continuous burn time of 6,000 s. The thruster achieves its performance through a novel micro-coaxial gas-gas injection scheme and an innovative regeneratively cooled combustion chamber, which uses the oxidizer as the working fluid. The thruster's capabilities and performance recorded during the completed qualification testing are described, and its design is outlined. This paper is published with the permission of the authors granted to 3AF – Association Aéronautique et Astronautique de France (www.3AF.fr) organizer of the Space Propulsion International conference.

Effects of Injector Pressure Drop on the Performance of a Monopropellant Hydrazine Thruster

نتایج مطالعه روی رفتار یک رانشگر تک‌پیشرانه هیدرازینی هنگام استفاده از انژکتورهایی با ابعاد مختلف ارائه می‌شود. انژکتورهای استفاده شده در این رانشگرها از نوع اریفیس هستند. هرچند الگوی جریان سیال در این انژکتورها ممکن است ساده به نظر برسد، رفتار رانشگر -که ناشی از اندرکنش دینامیک انژکتور با محفظه است- ویژگی‌های جالب توجهی را نشان می‌دهد. افت فشار انژکتور، در محدوده مورد مطالعه، تاثیری بر سرعت مشخصه و تاخیر اشتعال نداشت. با این حال زمان پاسخ 90% فشار و نوسان فشار محفظه به شدت تحت اثر افت فشار انژکتور هستند؛ با افزایش افت فشار، زمان پاسخ افزایش می‌یابد اما نوسان فشار کم می‌شود. نتایج ارائه شده ثابت می‌کند که آزمون سرد انژکتور برای ارزیابی و صحه‌گذاری طراحی آن کافی نبوده و اهمیت تست گرم رانشگر را نشان می‌دهد. بحث‌های این مقاله همچنین برای طراحان هر سامانه‌ای که از اریفیس برای کنترل دبی استفاده می‌کنند و به دنبال تحلیل رفتار گذرای آن سامانه‌ها هستند، قابل توجه خواهد بود.

Investigation on Performance Characteristics of Hydrazine Monopropellant Thruster according to Reaction Chamber Adiabatic Temperature

Investigation on Performance Characteristics of Hydrazine Monopropellant Thruster according to Reaction Chamber Adiabatic Temperature

A Miniaturized, Green, End-Burning, and Sandwich Hybrid Propulsion System

Conventional hybrid rocket motors with thrust levels greater than 5 N rely on forced convection within the boundary layer as the primary heat transfer mechanism for fuel vaporization. For hybrid rockets with thrust levels of less than 5 N, oxidizer mass flow levels are sufficiently small that the rate of convective heat transfer is significantly reduced; and radiative heat transfer dominates the fuel vaporization mechanism. Radiative heating is a concern when implementing traditional hybrid rocket core-burn fuel grain designs for systems with thrust levels suitable for small satellites and CubeSats. Radiative heating induces a fuel-rich burn, which leads to low combustion efficiency and nozzle clogging. This paper presents a novel idea of using radiative heating in the design of a hybrid propulsion system suitable for CubeSats and small satellites. Additionally, this paper presents the test results of two fuel grain designs: an end-burning fuel grain design and a “sandwich” fuel grain design. This paper presents the test results of a 0.5 N end-burn system that produces a vacuum of 162 s with a rise time of 1.1 s and a minimum impulse bit of . This paper also presents the test results of a 1.0 N end-burn system that produces a vacuum of 136 s with a rise time of 2.2 s and a minimum impulse bit of . In addition, the test results for a 1 N sandwich system that produces a vacuum of 170 s with a rise time of 0.3 s and a minimum impulse bit of are presented. Acrylonitrile butadiene styrene, polyvinyl chloride, Nylon-12, and an acrylic plastic known as polymethyl methacrylate were used as fuel; and gaseous oxygen was used as the oxidizer during these testing campaigns. These propellants provide several advantages, including benign handling properties, simplified plumbing, and greater burn efficiency over traditional monopropellant hydrazine.

Effects of geometrical parameters of a two-sided catalyst bed on flow and thermal fields of a monopropellant hydrazine thruster

در این تحقیق، محفظه تجزیه یک رانشگر تک مولفه‌ای بر پایه سوخت هیدرازین بصورت عددی شبیه-سازی شده است. بستر کاتالیستی مورد بررسی دو جزئی بوده و اثرات پارامترهایی مانند ضخامت ناحیه بالادستی و قطر ذرات تشکیل دهنده آن روی عملکرد محفظه تجزیه بررسی می‌شوند. در این راستا سه قطر گرانول استاندارد با عدد مش 16.5، 25 و 30 با ضخامت‌‌های ناحیه بالادستی (mm) 2.5، (mm) 5 و (mm) 7.5 در یک بستر کاتالیستی به طول (cm) 6.5 مورد مطالعه قرار گرفته‌اند. شبیه‌سازی‌ها برای دو نوع گرانول در بستر اصلی با قطرهای استاندارد (in) 8/1 و (in) 16/1 در ضرایب بارگذاری (kg/m2s) 16.5، (kg/m2s) 25 و (kg/m2s) 35 انجام گرفته است. نتایج حاصل نشان می‌دهند که میزان تاثیرگذاری بخش بالادست بستر کاتالیستی به شدت به نسبت اندازه ذرات تشکیل‌دهنده بستر اصلی به ناحیه بالادستی وابسته است. همچنین، طول ناحیه بالادستی و بارگذاری بستر پارامترهای مهمی در میزان تاثیرگذاری ناحیه بالادستی بشمار می‌روند.

Plume Contamination Measurements of an Additively Printed, Green-Propellant Hybrid Thruster

The dielectric breakdown properties of additively printed acrylonitrile butadiene styrene are leveraged to develop a green hybrid thruster family with capability for on-demand start, stop, and restart. Because acrylonitrile butadiene styrene is a new material for propulsion applications, no exhaust plume database exists, and the associated contamination potential for spacecraft surfaces is unknown. Analytical calculations of associated exhaust plume species are presented to identify predicted concentrations of primary condensable species , , and carbon (soot). When the hybrid thruster is operated near optimal equivalence ratio, the analytical studies show that combined mass concentrations of and constitute between 10 and 35% of total exhaust plume. Condensable carbon does not occur. This result is compared to 50% or greater mass concentrations of condensable species, and residual , for monopropellant hydrazine. Thus, the 3-D printed hybrid fuel is likely to burn cleaner than hydrazine. Results from ground-based and space-flight plume contamination experiments are presented. The contaminant sensor uses a light-sensitive photoresistor, mounted behind a quartz glass window to simulate spacecraft optical surfaces. Ground-based tests demonstrate approximately 12–13% optical attenuation after a 10-pulse burn sequence. Space flight data collected from a five-pulse burn sequence demonstrate optical attenuation levels approximately 2% higher.

Parameters Influencing the Characteristic Exhaust Velocity of a Nitrous Oxide/Ethene Green Propellant

Currently, several green propellants are under investigation to replace the toxic and carcinogenic monopropellant hydrazine (). Beside alternatives as ammonium dinitramide-based propellants, hydroxylammonium nitrate-based mixtures, or high concentrated hydrogen peroxide, mixtures of fuels with nitrous oxide () called HyNOx or NOFBX represent promising green propellants. Compared to classical , mixtures offer a higher performance ( and ), no or low toxicity, and low propellant costs. To investigate the handling and performance of a propellant mixture consisting of and ethene, hot gas combustion tests with an experimental combustor were conducted. This paper summarizes the results of 134 combustion tests conducted with the premixed propellant injected in gaseous state. Calculated and measured performances ( and ) depending on mixture ratio, characteristic combustion chamber length , and chamber pressure are shown and discussed. Furthermore, the residence timescales of the propellant mixture inside the combustion chamber are normalized by the chemical timescale of the combustion process. Thus, combustion efficiencies depending on dimensionless Damköhler numbers are derived.

Developing a Multi-objective Multi-Disciplinary Robust Design Optimization Framework

The purpose of this study is to provide an efficient Multi-Objective Multidisciplinary Robust Design Optimization (MOMRDO) framework. To this end, Bi-Level Integrated System Synthesis (BLISS) framework is implemented as a fast Multi-disciplinary Design Optimization (MDO) framework. Progressive Latin Hypercube Sampling (PLHS) is developed as a Design of Experiment (DOE) of the Uncertainty Analysis (UA). This systematic approach leads to a fast, adaptive and efficient framework for Robust Design Optimization (RDO) of complex systems. The accuracy and performance of the proposed algorithm have been evaluated with various tests. Finally, the RDO of a hydrazine monopropellant thruster is defined as a case study. The results show that the proposed method is a fast and efficient method for the multi-objective optimization design of complex systems, and this approach can be used for other engineering applications as well.

Numerical simulation of the transient process of monopropellant rocket engines

Liquid monopropellant rocket engines are widely used in space propulsion and are developing in the direction of non-toxic and pollution-free propellants. The use of non-toxic monopropellants with a low decomposition rate makes the start-up and shutdown processes of the engines longer, and the gas inlet assumptions that were applicable to the calculation of hydrazine monopropellant engines cannot be used. Based on the steady-state calculation model of hydroxyl ammonium nitrate (HAN)-based monopropellant engine, we propose a combined model that considers decomposition of liquid monopropellant and non-equilibrium heat transfer between fluid and solid catalyst bed, and uses this model and the gas inlet hypothesis to calculate the working process of an HAN-based monopropellant rocket engine of 60 N thrust. The calculated results are compared with the test curve, and the results show that the decomposition of liquid monopropellant plays a key role in the starting and stopping processes. When the gas inlet assumption is adopted without considering the decomposition of the liquid monopropellant, the calculated start-up time is much shorter than the hot-firing test, and there is no tail-off section when shutting down. The pressure-rise and pressure-decay curve calculated by the combined model agree well with the test curve, and the maximum deviation during the steady state is less than 10%. For HAN-based monopropellant engines, the pressure rises slowly at the first start because the temperature of catalyst bed is lower, which results in a lower rate of decomposition. When restarted shortly after a shutdown, the pressure rises rapidly due to the high temperature of the catalyst bed and the high rate of decomposition. The HAN-based monopropellant can penetrate a few millimeters downstream of the injector, and there is a significant tail when the engine is shut down, which is caused by the continued decomposition of the remaining propellant after the shutdown.

An Open-source Solver to Model the Catalytic Decomposition of Monopropellants for Space Thrusters

This paper presents a solver for the CFD (computational fluid dynamics) modeling of catalytic chambers in monopropellant thrusters based on the open-source OpenFOAM® (Open Field Operation and Manipulation) framework. A model was formulated and used to simulate the physical and thermochemical processes taking place inside the catalytic chambers of the monopropellant thrusters. The code integrates reacting gas flow in porous media, with mass and heat transport. The most important implemented functionalities were the separation of two reactive gases, one transient reactive gas flowing between the catalyst bed interstices (interstitial gas), and another static reactive gas on the surface of catalyst particles, or pellets (surface gas). Homogeneous and heterogeneous reactions occur on the interstitial gas and the catalyst particle, respectively. A flexible definition of porous properties, a calculation of multicomponent diffusion-flux mass, a diffusion mass coeffcient, a mass transfer coeffcient, and a heat transfer coeffcient were implemented as well. Experimental and analytical studies about hydrazine monopropellant thrusters in the literature were used to the case tests and verification of the solver. Temperature and mass fraction fields were simulated and compared. The results of the temperature profile are in agreement with experimental and theoretical studies found in the literature, and mass fraction presents some differences.

ADN and HAN‐Based Monopropellants – A Minireview on Compatibility and Chemical Stability in Aqueous Media

AbstractThis article gives a short review on compatibility and chemical stability of selected aqueous ADN and HAN‐based energetic formulations. A brief introduction will outline Energetic Ionic Liquids (EILs) as a new class of energetic materials with beneficial physical‐chemical properties which make them valuable for application in propulsion technologies. EILs combine the advantages of e. g. low toxicity, showing equal or superior propulsion power compared to the state‐of‐the‐art monopropellant hydrazine. Focus is set on open access ADN and HAN‐based monopropellant formulations that are e. g. realized in the advanced blends LMP‐103S, FLP‐106 or LGP 1845.

Design of Physical Configuration of a 10N Monopropellant Hydrazine Thruster

In this paper, design and physical configuration of various components of a 10N Monopropellant Hydrazine Thruster focusing on design calculations and optimization of catalytic combustion chamber. According to this design, a prototype of the thruster will be manufactured. The mentioned thruster has been designed as a three-piece modular thruster, including an injection system, catalytic combustion chamber and nozzle. Based on analyzes done for each module, the propulsion characteristics of monopropellant thruster system have been identified and used for the next module as necessary inputs. The combustion chamber dimensions are selected based on criterion of maximum decomposition of 40% ammonia and Mach number of 0.02. Also, the third module is the nozzle, designed as a simple cone. The exterior body design of these three modules and their connections to each other, based on considerations of sizing and weight limitation, as well as being dual purpose for use in the cold and hot tests, has been performed.

Satellite design verification study based on thruster plume flow impingement effects using parallel DSMC method

A major objective of the present study is to introduce a usefulness of the parallel CFD application to a real design process of a space system and its importance. To verify the suitability of a proposed satellite design which is currently under development, the present study investigated the thruster plume impingement effects on satellite components and structures by predicting the disturbance force/torque and heat load. To analyze these plume effects quantitatively, a numerical simulation with the parallel Direct Simulation Monte Carlo (DSMC) method was performed. From the present results, the impingement effects of the released plumes from four monopropellant hydrazine thrusters on satellite components and structures were examined. And finally, the suitability of the initial satellite design configuration is discussed through the analysis data of the consequential dynamic and thermal influences.

Rationale and concept for a lunar pit reconnaissance probe

ABSTRACTSpeculation on near-term scientific reasons for the exploration of lunar pits is offered alongside comments on possible longer-term human exploitation. It is proposed that in order to determine whether or not one or more of the pits offer access the large subsurface voids e.g. a non-collapsed lava tube, a preliminary reconnaissance mission solely focused on obtaining lateral images (and/or LiDAR maps) is needed. Possible concept options for such a preliminary reconnaissance mission are discussed. It is suggested that one of the best possible strategies is to employ a micro-sized probe (~0.3 m) that would hop from a nearby main landing spacecraft to the selected pit. After the surface position of the main lander is determined accurately, the probe would perform a ballistic hop, or hover-traverse, a distance of ~3 km over the lunar surface using existing propulsive and guidance technology capability. Once hovering above the pit, the probe or a separate tethered imaging unit would then be lowered into the pit to acquire the necessary subsurface void topology data. This data would then be transmitted back to Earth, directly, via the lander, or via a store-and-forward orbiting relay. Preliminary estimates indicate that a probe of ~14 kg (dry mass) is viable using a conventional hydrazine monopropellant system with a propellant mass fraction of less than ~0.2 (20%) including margins, suggesting a piggyback architecture would be feasible.

Experimental Investigation of the Flame Propagation and Flashback Behavior of a Green Propellant Consisting of N2O and C2H4

Regarding the research on alternatives for monopropellant hydrazine, several so called green propellants are currently under investigation or qualification. Aside others, the DLR Institute of Space Propulsion investigates a N2O/C2H4 premixed green propellant. During the research activities, flashback from the rocket combustion chamber into the feeding system has been identified as a major challenge when using the propellant mixture. This paper shows the results of ignition experiments conducted in a cylindrical, optical accessible ignition chamber. During the ignition and flame propagation process, pressure, temperature and high-speed video data were collected. The high speed video data were used to analyze the flame propagation speed. The obtained propagation speed was about 20 m/s at ignition, while during further propagation of the flame speeds of up to 120 m/s were measured. Additionally, two different porous materials as flame arresting elements were tested: Porous stainless steel and porous bronze material. For both materials Peclet numbers for flashback were derived. The critical Peclet number for the sintered bronze material was around 20, while for the sintered stainless steel the critical Peclet number seems to be larger than 40. Due to the test results, sintered porous materials seem to be suitable as flashback arresters.