Rocket Grade Hydrogen Peroxide Research Articles

Monoethanolamine-Based Fuels Hypergolic with Rocket-Grade Hydrogen Peroxide

The object of this work is to study the ignition performances of promising catalytically promoted amino-based green fuels to be used in combination with rocket-grade hydrogen peroxide. The main hypergolic parameter, the ignition delay time, was experimentally determined with an automated drop test setup. Additionally, the kinematic viscosity was experimentally measured, while the CEA2 software was used to determine the ideal rocket performances of the propellants. Three inorganic copper salts were selected as catalysts to be used in combination with monoethanolamine in concentrations ranging from 0.5 wt% to 20 wt%. Then, N,N-dimethylethylenediamine was introduced as part of a blend with monoethanolamine to target the high viscosity and low gravimetric specific impulse of the fuel for the pure monoethanolamine case. Due to the reduced monoethanolamine and low additive content, some formulations were observed to be characterized by gravimetric specific impulse higher than 320 s, and kinematic viscosity lower than 5 cSt while retaining ignition delay times shorter than 30 ms with 98 wt% HTP. Finally, the addition of ethanol to the blend was preliminarily investigated to improve the additive solubility. The effects on the ignition delay time were found to depend on the concentration ratio of the two amines.

Hypergolic ignition response to oxidizer droplet properties

The hypergolic ignition response to oxidizer droplet properties is an important consideration for oxidizer injector design in hybrid rocket motors. Hence, the present work focused on hypergolic ignition characterization as a function of droplet properties and impact dynamics for a promising propellant. The propellant for the present work consisted of rocket grade hydrogen peroxide (RGHP) as oxidizer, high-density polyethylene (HDPE) as fuel and sodium borohydride (NaBH4) as the hypergolic additive embedded in the HDPE matrix. This propellant was chosen for this detailed study since it showed promising results in prior studies, exhibiting low (<10 ms) ignition delays. For this work, in a drop-test setup, RGHP droplet diameters between 1.90±0.11 mm and 4.35±0.26 mm were dropped on the HDPE + NaBH4 fuel samples at velocities between 0.3 m/s to 1.93 m/s. The ensuing ignition event was studied using both visible and infrared high-speed imaging. Post-impact droplet parameters were determined to enable a deeper analysis of observed phenomena. While ignition delay was the focus of this work, flame spread areas and heat release were also studied. Higher impact velocities generally led to lower ignition delays while splashing oxidizer drops increased probability of larger flame areas and higher heat release. Additionally, the initial ignition/heat release event was found to originate in a small region at the thinnest area of the post-impact droplet. Surface temperature measurements were also performed using infrared imaging. The data suggested that the initial heat release was found to be likely due to RGHP-NaBH4 reaction with no contribution from RGHP decomposition. Regression and random forest machine learning analysis found ignition delay depended strongly on the minimum post-impact drop layer thickness. Based on the data analysis, a simple mathematical model was developed whose results matched experimentally observed trends. The study is expected to further understanding of the processes that control hypergolic ignition in addition to aiding rocket engineers design better oxidizer feed systems. Novelty and SignificanceThis work constitutes a novel in-depth study on the influence of drop impact properties on hypergolic ignition. Using innovative techniques, this work is the first to show the overwhelming importance of residual droplet layer thickness on ignition delay. Additionally, a simple mathematical model based on that finding performed well against experimental data which is of substantial significance, given the complex and multi-parametric nature of hypergolic ignition. Finally, the work establishes that operating in the splashing regime enhances heat release and flame spread. The candidate propellant for this work was hydrogen peroxide and high-density polyethylene with embedded NaBH4. However, the findings of this study are likely generalizable over a range of solid-liquid hypergolic propellants involving solid fuels with embedded energetic materials, of which there are many. The findings of this work can enhance understanding of hypergolic ignition in addition to aiding in better rocket injection system design.

Microencapsulation of hydrocarbon fuels for monopropellant creation with hydrogen peroxide

Monopropellant systems are used in rocket engines when high reliability and low technical complexity is demanded. However, they usually have lower specific impulses compared to bipropellants. A mixture of oxidizer and fuel from typical bipropellant systems is generally difficult to convert into a monopropellant as they are normally immiscible and would constitute a potentially explosive mixture.Microencapsulation of one of the two components of a bipropellant could be a possible way to create an insensitive mixture of the two substances together in the same tank while creating a monopropellant. This approach would bring together the advantages of both monopropellant and bipropellant, a high specific impulse combined with low technical complexity.In this work we present a new concept of microencapsulated hydrocarbon fuels for the creation of a monopropellant with hydrogen peroxide as green oxidizer component. Hydrocarbon filled microcapsules were synthesized by interfacial polymerization and analyzed by thermogravimetry, optical and electron microscopy as well as infrared spectroscopy. Additionally, the compatibility of the polyamide capsule wall with rocket grade hydrogen peroxide (up to 98%) and vice versa was studied and was found to be suitable.

HYPERGOLIC IGNITION OF SOLID FUELS WITH REACTIVE ADDITIVES USING HYDROGEN PEROXIDE

Hypergolic ignition offers significant benefits for hybrid rockets since it eliminates the need for external ignition devices or catalyst beds. Numerous studies have demonstrated promising results when utilizing various reactive additives and oxidizers. Nevertheless, the impact of fuel temperature and particle size on ignition performance, which is critical in practical applications, is often overlooked. In this study, we explored the ignition behavior of a metal hydride-based hypergolic fuel with rocketgrade hydrogen peroxide (RGHP) as the oxidizer. The solid fuel was composed of low-density polyethylene as the binder and sodium borohydride as the reactive additive. Through image processing, we analyzed the ignition behavior by measuring the ignition delay time in droplet tests using different concentrations of the additives, particle sizes, and initial fuel surface temperatures. Three types of ignitions were observed, and all of the droplet tests were successfully ignited. To examine the initial fuel temperature and powder particle size effects on the ignition, we varied the initial temperature of the fuel pellet and composition percentage of two different particle sizes in the powder. Our experimental findings suggest that both the initial fuel surface temperature and powder particle size play a crucial role in the hypergolic ignition delay time.

Comprehensive ignition characterization of a non-toxic hypergolic hybrid rocket propellant

This paper describes a comprehensive characterization of ignition properties of a metal-hydride based non-toxic hypergolic hybrid rocket propellant. The propellant consists of Rocket Grade Hydrogen Peroxide (RGHP) as oxidizer, high-density Polyethylene (HDPE) as fuel and sodium borohydride (NaBH4) as the additive, embedded in the HDPE matrix. Ignition quality was characterized as ignition delay, ignition probability and flame spread. In a drop-test setup, ignition characteristics were determined as a function of seven parameters: RGHP concentration, additive loading, oxidizer droplet impact velocity, oxidizer droplet volume, pressure, diluent gas, and environmental exposure. The parameters encompass thermo-chemical, fluid/droplet dynamics and environmental factors affecting ignition. Ignition delays as low as 3 ms were observed, one of the lowest using non-toxic hypergolic hybrid propellants in open-air. An overwhelming majority of conditions tested yielded <10 ms ignition delays and 100% ignition success. All conditions tested affected ignition to varying degrees with RGHP concentration, NaBH4 loading and drop impact velocity significantly affecting ignition. Further, contrary to expectations, exposing sanded fuel samples to humidity for a few h enhanced ignition instead of hampering it and exposure for 24 h did not lead to ignition degradation. Tests with diluent gases other than air (at atmospheric and elevated pressures) revealed that atmospheric oxygen played a negligible role in the reaction process. This proved that oxygen for the initial ignition event was obtained from RGHP decomposition, with atmospheric oxygen playing no role in ignition performance. Aside from demonstrating excellent ignition characteristics, our results further show a need to go beyond thermo-chemical properties and to consider aspects of ignition other than ignition delay in hypergolic propellant research to enable a complete understanding of the ignition processes. The comprehensive ignition characterization demonstrates the chosen propellant's ability to overcome ignition challenges in hybrid rockets and serves as a proof of concept for its further development.

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.

Additive-promoted hypergolic ignition of ionic liquid with hydrogen peroxide

The exploration of environmentally friendly hypergolic combinations of ionic liquid fuels and hydrogen peroxide as an oxidizer offers opportunities to replace commonly used conventional toxic, corrosive and carcinogenic hydrazine-based liquid hypergolic combinations. Various research opportunities await regarding the detailed investigations of green hypergolic ionic liquids (HILs) with rocket grade hydrogen peroxide (RGHP, >85% H2O2). In this work, the combustion of HILs 1-ethyl-3-methyl imidazolium cyanoborohydride ([EMIM][BH3CN]) and 1-allyl-3-ethyl imidazolium cyanoborohydride ([AEIM][BH3CN]), and HIL-additive mixtures with 95% H2O2 has been investigated. A new additive, 1,3-dimethyl imidazolium copper iodide ([diMIM]n[Cu2I3]n) was synthesized successfully; a structural investigation was performed through the use of single-crystal x-ray diffraction analysis through which a crystal density of 3.22 g/cm3 was revealed. The physicochemical properties (density, viscosity, and decomposition temperature) as well as performance parameters (ignition delay and specific impulse) of 2 to 15 wt% of [diMIM]n[Cu2I3]n in [EMIM][BH3CN] were determined. A 15 wt% of [diMIM]n[Cu2I3]n exhibited ignition delay time (IDT) values of 13 and 29 ms under fuel-rich and oxidizer-rich conditions, respectively; these IDT values are 100 times lower than those for [EMIM][BH3CN]. Interestingly, [diMIM]n[Cu2I3]n-[EMIM][BH3CN] combinations revealed high density (>0.98 g/cm3), good thermal (>200 °C) and chemical stability, and 5.6–6.0% higher density specific impulse than those found for unsymmetrical dimethyl hydrazine. The additive-promoted hypergolic combustion of HIL with RGHP opens a new avenue to the replacement of conventional toxic hypergolic combinations.

Hypergolic ignition delay studies of solidified ethanol fuel with hydrogen peroxide for hybrid rockets

The hypergolic interaction and ignition delay times of catalytically promoted solidified ethanol (CPSE) formulated using organic gellant, namely, methylcellulose (MC) and hydroxypropyl methylcellulose (HPMC), has been investigated using rocket grade hydrogen peroxide (90% RGHP; H2O2) as an oxidizer, for the use in hybrid rocket engines. Using iso-conversional Friedman method, the apparent activation energy (Ea) is found to be in the range of 9.06–14.01 kJ/mol for unloaded solidified ethanol samples. The hypergolic ignition delay (ID) of the CPSE fuels with H2O2 are investigated using manganese (III) acetylacetonate (Mnacac) as a catalyst using the drop-test method and found that the ID lies in the range of 49–307 ms. Five-stage events are identified based on the signals from the high-speed camera, photodiode, and microphone obtained from the drop-test. Stage I: prior to of H2O2 drop impact; Stage II: inertial spreading of H2O2 drop over CPSE pellet; Stage III: onset of exothermic decomposition reaction H2O2 droplet by Mnacac particles which leads to the formation of aerosol; Stage IV: local explosion of aerosol and ignition of CPSE pellet occur results in two types of aerosol-CPSE fuel interaction; Stage V: a weak self-sustained flame from CPSE fuel. A key finding of the study reveals that a minimum catalyst concentration ([C]Mn3+,L) is required to ignite CPSE fuel using H2O2 and ID decreases with an increase in the [C]Mn3+. Similarly, there exists an upper catalyst loading ([C]Mn3+,U) above which ID is increased due to the increase in Mnacac particle size by agglomeration. [C]Mn3+,L and [C]Mn3+,U depends on the type and concentration of the gellant and the availability of active hydroxyl (–OH) group in the gellant. Finally, SEM-EDS analysis of the residue indicates the formation of oxides of manganese with a small amount of carbon (~1.66 to 2.53 wt%), hinting complete combustion of the gellant particles.

Comparative effects of rocket-grade hydrogen peroxide solution on POM and UHMWPE: aging behaviors and tribological properties

Polymeric materials have a short history of employment as tribological materials in liquid-propellant rocket engines, but they hold promise for the future. Even those polymer materials served already in rocket engines are based on inaccurate or unsubstantiated data. For these reasons, aging tests of polyformaldehyde (POM) and ultra-high molecular weight polyethylene (UHMWPE) in rocket-grade hydrogen peroxide solution were carried out, and then tribological experiments were conducted in this work. The objective is to provide a sound foundation of further selecting or developing polymeric materials for tribological application in rocket-grade hydrogen peroxide environment. The micrographs of polymer specimen with various aging time indicate that POM has poorer aging resistance in hydrogen peroxide solution than UHMWPE. A new absorption peak assigned to hydroxyl group appears on FTIR spectra of POM after aging test due to the oxidation by hydrogen peroxide solution. It is deduced from thermoanalysis that the amorphous regions is the main area influenced by hydrogen peroxide solution. Results of tribological experiments display that the friction coefficients of both POM and UHMWPE decrease with the increase in aging time from 0 to 240 h. However, the wear rates of POM and UHMWPE present different trends with the increase of aging time; the wear rate of POM decreases, whereas the wear rate of UHMWPE increases.

Response characteristics of H2O2 monopropellant thrusters with MnO2-mixed PbO catalyst

As environmental concerns increase, rocket-grade hydrogen peroxide has been considered as one of the candidates for replacing toxic propellants. Various components for the development of monopropellant thrusters using hydrogen peroxide have been studied, including catalysts, catalytic supports, and design parameters in monopropellant thrusters. As a part of these efforts, an experimental investigation, including design parameters, for a catalyst bed is performed with MnO2-mixed PbO catalysts in a hydrogen peroxide monopropellant thruster. In addition to basic performance of the monopropellant thruster, the response characteristics of rising time and tail-off time are compared with respect to two different catalysts. From the experimental results, it is found that a catalyst with high reactivity is much more preferable for reducing not only rising time but also tail-off time. Tail-off time in case of MnO2-mixed PbO is considerably shortened compared to that of MnO2. Moreover, design parameters, such as catalyst capacity and catalyst bed loading, for designing a catalyst bed in a monopropellant thruster can be suggested from an analysis of all experimental results.

Demonstration of 500 N scale bipropellant thruster using non-toxic hypergolic fuel and hydrogen peroxide

While shedding light on green space propulsion, the development of non-toxic hypergolic bipropellants has provided alternatives to the conventional, highly toxic propellants. Non-toxic energetic hydrocarbon fuels promoted by sodium borohydride were produced, and their hypergolicity with rocket grade hydrogen peroxide was confirmed. Ignition delays ranged from 4 ms to 17 ms, depending on the oxidizer concentration. A 500 N scale thruster utilizing a non-toxic hypergolic bipropellant combination was also developed. With an oxidizer-to-fuel ratio of 5.48, the injector was designed with pentad unlike-impinging jets, one fuel jet orifice and four oxidizer jet orifices for enhancing the mix and atomization of the hypergolic propellants. The operation test was performed, and burning time was established to be 3.5 s. For the demonstrator, the rising time was approximately 158 ms, and the pressure oscillation inside the combustion chamber was approximately ±7.69%. These results indicate the feasibility of this novel concept engine for future space missions.

Effect of Phosphate Stabilizers in Hydrogen Peroxide Decomposition on Manganese-Based Catalysts

Adverse effects of phosphate ions used as a stabilizing agent in rocket-grade hydrogen peroxide on the decomposition catalyst were investigated. For a comparative study, and were selected as the catalyst to decompose 90 wt% hydrogen peroxide containing of phosphate ions. The reactivity of the catalysts measured by reactivity tests decreased by half in the hydrogen peroxide, including phosphate ions. Endurance of the catalyst was evaluated on a 50 N monopropellant thruster by measuring the chamber pressure. The degradation of the decomposition efficiency of the catalyst was evident after the initial 5 s, during which the lead-doped catalyst revealed better performance than the pure catalyst. After 30 s of thruster operation, performance degradation was particularly severe with the lead-doped catalyst. The catalysts sampled after thruster operation were characterized by mass analysis, the Brunauer Emmet Teller model, and energy dispersive spectroscopy. A large quantity of phosphorus was detected on both catalysts tested with a high concentration of phosphate ions. The amount of phosphorus detected on the lead-doped catalyst was four times as large as the amount of phosphorus on the pure catalyst, indicating the affinity of lead and phosphate. The accumulation of phosphate on the catalyst surface inhibited the contact of hydrogen peroxide and the active material, preventing efficient decomposition.

CoMn-Based Oxides as Bulk Catalyst for Rocket-Grade Hydrogen Peroxide Decomposition

Novel bulk catalysts based on CoMn mixed oxides were prepared by a coprecipitation method for rocket-grade hydrogen peroxide decomposition. A 2 N thruster was used to evaluate and compare the catalysts in continuous and pulsed firings. The aluminum-containing mixed oxide catalyst, CoMnAl, provided the best performance, spontaneously decomposing the propellant and sustaining several thermal cycles without fractures and deactivation. Because experimental results were close to theoretical ones, this material has a good potential for peroxide decomposition in satellite and auxiliary propulsion systems.

Performance of Dicyclopentadiene/-Based Hybrid Rocket Motors with Metal Hydride Additives

An experimental performance evaluation of metal hydride fuel additives for hybrid rocket motor propulsion systems is examined in this paper. Some metal hydride additives offer improved performance, but a common issue is material aging. An accelerated aging study revealed that dicyclopentadiene can protect sodium borohydride () particles from exposure to air and water vapor much better than conventional hydroxyl-terminated polybutadiene. Static hybrid rocket motor experiments were conducted using dicyclopentadiene as the fuel. Sodium borohydride and aluminum hydride () were examined as fuel additives. Ninety percent rocket-grade hydrogen peroxide was used as the oxidizer. In this paper, the sensitivity of solid fuel regression rate and characteristic velocity efficiency to total fuel grain port mass flux and particle loading is examined. Chamber pressure histories revealed steady motor operation in most tests, with reduced ignition delays when using the fuel additive. The addition of and produces up to a 47 and 85% increase in regression rate over neat dicyclopentadiene, respectively. For all test conditions examined characteristic velocity efficiency ranges between 80 and 90%. The regression rate and characteristic velocity efficiency mass flux dependence indicate a shift toward a more diffusion controlled system with metal hydride particle addition.

Design and Fabrication of LTCC Catalyst Chambers.

The reduction in satellite size and mass presents the need to develop a proportionally smaller propulsion system for orbital station keeping. A liquid, monopropellant micropropulsion device made from Low Temperature Co-Fired Ceramics (LTCC) has been developed at Boise State University. This robust, simple design uses an embedded silver catalyst chamber to decompose a rocket-grade hydrogen peroxide monopropellant into a hot gas, which is then expelled out through a nozzle to generate thrust. Using LTCC eliminates the planar geometry fabrication constraint commonly found in silicon MEMS processing. This report presents the design and fabrication, and optimization of the hydrogen peroxide catalyst chamber used in these monopropellant microthrusters. Using the standard fabrication process for LTCC an initial prototype was developed. The design of this initial device was developed to measure the efficiency of the catalyst chamber by evaluating the ability of the device to decompose hydrogen peroxide. Catastrophic cracking within the device substrate was observed during initial testing. In order to obtain sufficient data, it was assumed that the cracking was due to thermal expansion and so a new functional design was implemented that decreased the overall cross sectional area of the device and decreased failure rates. To ensure that this assumption is correct, an investigation of device failure will be presented using an embedded resistor to simulate the catalytic reaction occurring inside the substrate. The results from this investigation will be documented. Additionally, optical microscope images will be used to document the failure investigation process. Several conclusions will be presented to improve the ability to use LTCC for high temperature applications.

Transient Behavior of H2O2 Thruster: Effect of Injector Type and Ullage Volume

ROCKET-GRADE hydrogen peroxide has been used as a monopropellant and a storable oxidizer. However, because of the demand for a higher specific impulse, hydrazine andN2O4 are being used as the monopropellant and storable oxidizer, respectively. Recently, due to concerns regarding propellant toxicity, there has been a renewed interest [1] in the use ofH2O2 in propulsion systems [2–10]. A monopropellant thruster is operated in either the continuous or pulse mode. The thrust force and pressure instability are important issues in the continuous mode. For generating the desired thrust, the catalytic reactor size required for completely decomposing the propellant must be determined [8]. However, in the pulse mode (the main operationmode for attitude control systems), the response characteristics of the thruster are important. The catalyst activity, thruster component design (including the injector design), manifold volume, ullage volume in the reactor, and operating pressure influence the thruster response time. Tian et al. investigated the response time when using a combination of PbO and MnO2 catalysts [11]; they found that Ir=Al2O3 is unsuitable for use as a catalyst in a H2O2 monopropellant thruster [12]. Xu et al. studied the activities of various catalysts during H2O2 decomposition [13]. El-Aiashy et al. reported that the catalyst activity ofMnO2 increased when ZnO was added [14]. Hasan et al. reported that the activity ofMnO2 increased when promoters such as Ni, Cu, Bi, and Ce were added [15]. None of the aforementioned studies have addressed the effect of thruster design parameters on response times, although a few researchers have measured the thruster response time. Optimization of the thruster design (determination of the appropriate injector and ullage volume in the reaction chamber) can also influence the response characteristics. Therefore, we investigate the response characteristics of H2O2 monopropellant thrusters for three different thruster designs andmeasure the response times by varying the injector type, reactor volume, and catalyst grain size. AMnO2=Al2O3 catalyst is used for the decomposition of concentrated H2O2 (90 wt%).

HYPERGOLIC REACTION MECHANISMS OF CATALYTICALLY PROMOTED FUELS WITH ROCKET GRADE HYDROGEN PEROXIDE

The coupling of the physical and chemical processes controlling the ignition delays (τ ign ) of hypergolic propellants complicates the direct analysis of the transient ignition processes. The transport properties of both the propellants and the ambient gas, and the heat of vaporization and reaction affect τ ign and must be considered experimentally. Furthermore, since hypergolic ignition occurs, by definition, on contact, the mechanics of initial contact and subsequent mixing can affect τ ign and must also be considered. Experimental results from a study of the hypergolic ignition of a catalytically promoted fuel with rocket grade hydrogen peroxide (RGHP) are presented. Statistical analysis of the experimental data was conducted to evaluate the relative effects of RGHP concentration, initial ambient pressure, and the ambient gas on τ ign . At RGHP concentrations of less than 92% in water, stable ignition was difficult to achieve. At sufficiently high Lewis number, an oscillatory ignition process is observed. A phenomenological model comprising mixing, RGHP decomposition, fuel heating and vaporization, and auto-ignition is presented and compared to the experimental results.

Hydrogen Peroxide Gas Generator with Dual Catalytic Beds for Nonpreheating Startup

I NTERESTS in rocket-grade hydrogen peroxide have been renewed in recent years as a new demand for a nontoxic alternative to rocket propellants arises. The use of high-concentrated hydrogen peroxide as a propellant in propulsion dates back to the 1940s. After the SecondWorldWar, it was used as amonopropellant and as an oxidizer in a bipropellant system for thrusters. Hydrogen peroxide was eventually replaced by higher-performing propellants such as hydrazine and N2O4. In the middle of the 1990s, there was renewed interest in hydrogen peroxide due to low toxicity, clean products, and enhanced versatility [1,2]. The performance of the catalyst is crucial in the design of the propulsion devices that use the decomposition of hydrogen peroxide. Silver has been the most widely used catalyst despite many shortcomings, including low melting point, nonuniform flow path, high-pressure loss, and need of preheating [2]. In addition, silver cannot withstand the high decomposition temperatures of hydrogen peroxide at concentration levels higher than 92% [3]. Alternative catalysts including manganese oxides [3– 6] and perovskites [7] have been tested with varying degree of effectiveness. Perovskite material such as La0:8Sr0:2CoO3 (hereafter referred to as LSC) had superior characteristics at elevated temperature but displayed very slow reactivity at room temperature [8]. On the other hand, permanganate-based catalysts displayed very high reactivity at room temperature but were not stable at elevated temperature [9]. In the gas generator proposed in the present study (Fig. 1), hydrogen peroxide goes through two distinct reaction phases: catalytic and thermal decomposition [10]. In the inlet region, the liquid-phase hydrogen peroxide decomposes mainly by catalytic decomposition upon contact with the catalyst surface. In this region, unreacted hydrogen peroxide vaporizes by the heat of the decomposition process. The vapor content of the partially decomposed hydrogen peroxide increases along the axial direction of the reactor bed. Down the reactor bed, the unreacted hydrogen peroxide in vapor state decomposes by further catalytic reaction. Good startup performance and catalytic reactivity are required in the inlet region, and high catalytic reactivity and thermal stability at high temperature are required in the outlet region. The LSC catalyst has good thermal stability and high reactivity at high temperature, but is hard to start up at room temperature [8]. To use the LSC catalyst for the high-temperature region, various catalysts for a vaporizer catalyst bed were selected and evaluated using a constant volume reactor and gas generator. Finally, the gas generator with a dual catalytic bedwas tested and evaluated.

Combustion of advanced non-toxic hybrid propellants

Combustion characteristics and fuel regression behaviour have been studied for two fuels utilising 98% rocket-grade hydrogen peroxide (RGHP) as oxidiser. Two test series (A and C) involved a total of 13 firings of dicyclopentadiene (DCPD) fuel grains and series B involved nine firings of DCPD grains loaded with various amounts of lithium aluminium hydride, LiAlH4 (LAH). A laboratory-scale combustion chamber with a 4.09 cm diameter was used in all firings. Regression rate data have shown a substantial increase as compared to RGHP/low-density polyethylene oxidiser/fuel combination. Results with the LAH-loaded grains showed comparable regression rates, but disappointing levels of combustion efficiency.

이원 촉매를 이용한 과산화수소 가스발생기

고농도 과산화수소는 별도의 산화제가 필요치 않은 단일추진제로써 상온에서 액상인 장점이 있어 다양한 장치의 추진제로 이용되고 있다. 본 연구에서는 상온에서의 시동 성능 보완을 위해 이원 촉매층을 이용한 실험적인 연구를 수행하였다. 상온에서의 뛰어난 성능을 바탕으로 <TEX>$K_2MnO_4$</TEX>를 도입부의 촉매로 선정하였으며 이를 위해 알루미나 졸-겔법을 이용한 새로운 코팅법을 개발하였다. 가스발생기를 이용한 반응실험을 통해 설계유량에 적합한 기화기의 길이를 결정하였고 이를 이용한 이원 촉매 가스발생기의 성능평가를 수행하였다. 성능평가 결과, <TEX>$10^{\circ}C$</TEX> 이하의 온도에서 냉시동(cold-start)에 성공하였으며 높은 분해효율을 확인하였다. The rocket grade hydrogen peroxide has been widely used as a monopropellant in propulsion systems. In the present paper, we described an experimental study of a catalytic reactor that employs two stage catalyst beds to enhance the low temperature performance of the reactor inlet. <TEX>$K_2MnO_4$</TEX> was chosen as the catalyst for the initial stage of the reactor bed for its superior behavior in the low temperature regime. Alumina sol-gel method was successfully applied for coating <TEX>$K_2MnO_4$</TEX> on a reactor bed of cordierite monolith. LSC was used for the catalyst of the second stage of the reactor. The reactor with combined catalyst beds was built and tested to exhibit superior performance in low temperature regime and high decomposition efficiency.