Development and Testing of Innovative Catalytic Beds for Hydrogen Peroxide Decomposition in Space Propulsive Applications

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Introduction

1.1 Scenario

At the beginning of the space age, the environmental impact of rocket engines played virtually no role in decisions aecting their development or use. For launchers boosters, many systems evolving from the German V-2 missile used liquid oxygen and a form of kerosene as oxidizer and fuel. Eventually, engines such as M-1, J-2 and RL-10 replaced the RP-1 kerosene with liquid hydrogen, which was trickier to handle but more powerful in terms of thrust. Payload yields have been then increased adding solid fuel motors to the rocket's rst stages: storable but highly toxic and corrosive liquids have been widely used as fuels and oxidizers in some of the most successful launchers (as the Titan and Delta series). High performance upper stages and spacecraft thrusters have been almost exclusively based on hydrazine, another dicult to handle liquid fuel. Environmental eects of space propulsive systems became a factor in 1972 with the rst ight of the Space Shuttle. Being the launch site perched on the edge of a wildlife reserve, NASA was concerned that the spacecraft's anticipated high ight rates might damage the ora and the fauna that made the grounds of the Kennedy Space Centre their home. Since then, environmental issues have been taken much more seriously in the development, production and maintenance of rocket propulsion systems, with growing attention also to the impact of engine's exhausts on the atmosphere. Green fuels and oxidizers not only reduce the cost and ease of handling and assembly of the rockets they drive, but might also increase the number of potential launch sites. The greening of rocketry has spread to upper stages and satellite propulsion systems too and designers are considering alternatives to hydrazine-driven thrusters: up to present, rocket propellants as nitrogen, nitrous oxide, butane, water and xenon have been employed. But these low-end propellants are not able to attain the same performances of existing engines based on toxic and carcinogenic chemicals as hydrazine and nitrogen tetroxide, hence new propellants and propellants' combinations need to be investigated. In this framework, highly concentrated solutions of hydrogen peroxide, commonly referred to as high test peroxide (HTP) have been identied as one of the most promising solutions: in the last ten years, HTP-based propulsive systems ranging from future Single Stage To Orbit launchers to micro-satellite's thrusters have been subject of investigation. In the following sections, the advantages related to the utilization of HTP in both high-thrust and low-thrust level applications will be illustrated; then, the most critical issues related to the design of a HTP-based propulsion system will be presented.

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CHAPTER 1. INTRODUCTION 4

1.2 Analysis of Propellants for Space Propulsive Systems

Hydrogen peroxide is not a new entry in the family of chemicals used for propulsive appli-cations, as it has been widely used since the beginning of the space age: its rst application dates back to 1935, when Walter equipped the Henkel 176 with ATO (Assisted Take O) units fed with a hydrogen peroxide solution (80% weight). In 1941, the rst aircraft powered by a rocket engine, i.e. the Messerschmidt 163, made its rst ight, reaching speeds up to 1000km/h, while the most advanced military jets of that time were limited to 600-700km/h: the main engine of the Messerschmidt 163 was fed with 80% weight hydrogen peroxide as the oxidizer and a blend 30% weight hydrazine (plus liquid catalyst), 57% weight methyl alcohol and water (to lower the combustion temperature).

Germany made a massive utilization of hydrogen peroxide during World War II: beyond the above applications, it was used for underwater systems (U-Boats) and missiles (V-2). After the war, USA and UK devoted large eorts to develop their own H2O2-based propulsive systems: the Reaction Control Systems (RCS) of launchers (i.e. Centuar, Scout) and experimental vehicles (X- 1, X-15) were equipped with 90% weight peroxide.

The peroxide found its most important propulsive application in the Gamma series rocket engine, which powered the UK Black Knight and Black Arrow vehicles: this launching system (shown in Figure 1.2.1) was fed with 85% weight H2O2 and kerosene. It successfully powered the Black Knight during 22 suborbital ights from the Woomera launch site (Australia) from 1958 to 1965. In 1971, the evolution of the Black Knight (e.g. the Black Arrow) reached the altitude conventionally corresponding to a Low Earth Orbit (LEO), becoming the rst (and the last, up to now) launcher fully relying on a H2O2-based engine. After this big success, the peroxide was gradually left, due to the search of propellants able to provide higher propulsive performances, as well as to other considerations not strictly related to technical/scientic issues [1]. Black Knight programme was scrapped because of the politico-economic view that the UK would nd it cheaper to purchase US Scout launchers for future satellite missions rather than have its own independent capability [2].

In high concentrations hydrogen peroxide has a long and proven history for both propulsion and gas generation applications in Russian (former Soviet Union) rocket programmes. In fact hydrogen peroxide has been extensively used in the Soyuz Launching System both for the the attitude control system of the capsule and for the gas generator in the rst stage . In early Soyuz models the ACS was equipped with two set of hydrogen peroxide thrusters able to generate 10 N and 100 N thrust respectively. The Soyuz commander can pilot the module using a rotational hand controller that manages the ring of the eight hydrogen peroxide thrusters on the vehicle's exterior. This system was typically deactivated 15 minutes before landing, when the parachutes were deployed. Figure 1.2.2 indicates one of the hydrogen peroxide thrusters mounted on the re-entry module of the Soyuz capsule. The picture clearly shows an assembly comprising the nozzle, the feeding valve with its actuator and the presumed housing of the catalytic bed. Previous versions of Soyuz rocket were derived from the R-7 ICBM which used an RD-107 rocket engine in each strap-on booster.

The RD-107 had four main nozzles with two steering vernier engines which gimballed on one axis, the main engines did not gimbal. Each RD-107 engine consisted of four combustion chambers, feed by a single turbo-pump mounted above the chambers. The core stage used a RD-108 engine, which was the same as the RD-107, but with four steering verniers.

To start the RD-107 and RD-108 type engines, propellant valves open and propellant ows through the pumps under force of gravity. The propellant is ignited by pyrotechniques and the engine burns at what is called intermediate thrust. As the pumps are turned by the propellant ow, they also drive two small pumps that fed hydrogen peroxide into a gas or steam generator. The gas generator produced a large amount of steam that has driven the propellant pumps. This created an increased pump speed and fed more propellant into the

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Figure 1.2.1: UK's Black Arrow Satellite Launcher, 1971. (Courtesy of the Science Museum, London)

combustion chamber. The RD-117 engine is an updated version of the RD-107 capable to generate a specic impulse of 310 s powered by Lox and kerosene. The rst ight took place in 2001.

Other liquid propellants, both cryogenic (liquid hydrogen/liquid oxygen) and storable (hydrazine/nitrogen tetroxide) currently feed the largest fraction of rocket engines, for both highthrust level applications (where often solid propellants boosters are used to increase the thrust, as in the Space Shuttle) and low-thrust systems (as the hydrazine-based thruster shown in Figure 1.2.3). The development of the Shell 405 catalyst and higher purity hydrazine led to a decreased use of hydrogen peroxide for RCS applications due to the superior performance and long-term stability characteristics of hydrazine. Except for the RCS of the Russian Soyuz, all monopropellant thrusters used as propulsive sub-system for spacecraft are fuelled with hydrazine. More details about the leading catalyst for hydrazine decomposition are reported in a following section

In the last decade hydrogen peroxide has been receiving a renewed interest for propulsive application mainly driven by issues related to the cost reduction and to environmental impact during the development, production and maintenance/operation phases of a space propulsive systems.

High propulsive performances (at least comparable with the state-of-the-art propellants) are necessary to deem green propellants actual substitutes of hydrazines and nitrogen tetrox-ide. Table 1.1 shows a comparison between the specic and volumetric impulses achievable with hydrazine/nitrogen tetroxide and pure (100% weight) hydrogen peroxide/kerosene (the JP-4 fuel, widely used for fuelling aircraft engines has been considered as representative of the kerosene class).

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Figure 1.2.2: Hydrogen Peroxide thruster for the attitude control system of Soyuz capsule

Figure 1.2.3: EADS-ST 400N hydrazine thruster

Propellants Chemical formula Isp[s] ρIsp[kg · m−3· s−1]

Hydrazine/nitrogen tetroxide N2H4/N2O4 356 447136

JP-4/Hydrogen peroxide CH1.94/H2O2 330 466950

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CHAPTER 1. INTRODUCTION 7 Propellants Density [kg · m−3_] Hydrogen Peroxide 1500 JP-4 810 Hydrazine 1450 Nitrogen Tetroxide 1443 Table 1.2: Density of liquid propellants

Figure 1.2.4: Performance of hydrogen peroxide in comparison with hydrazine (adapted from Ventura and Mullens [3]).

Operating conditions typical for bi-propellant rockets used in both RCS and OMS (Orbital Manoeuvring 11 Systems) have been assumed: pressure levels in the combustion chamber in the range 1-10bar have been considered; the specic impulse is computed for expansion to vacuum. As summarized in Table 1.1, the specic impulse of the peroxide-based engine is lower (~7%) and the density impulse is higher (about 4%) with respect to a stoichiometric mixture of hydrazine/nitrogen tetroxide. The latter is due to the higher density of a stoichiometric mixture of the selected propellants. Densities of several propellants having a widespread use for propulsive applications are listed in Table 1.2. Such results could be advantageous in the applications with severe mass/volume constraints (these constraints grow as the dimension of the system decreases, as will be shown later ) and with signicant aerodynamic draf losses. Comparable specic impulses when hydrogen peroxide is used in bipropellant engines are also reported in Figure 1.2.5 (Ventura and Mullens [3]).

The same trends hold for a mono-propellant conguration: the comparison between hy-drazine and hydrogen peroxide is summarized in Figure 1.2.4; the peroxide guarantees a lower specic impulse (about 17%) but a higher density impulse ( about 24%).

In regard to storability issues hydrogen peroxide has been mostly considered as a very unstable compound, dicult to handle and almost impossible to store for long periods, as required in some propulsive applications. This picture is not completely true, or at least not completely updated as reported in [2]. Hydrogen peroxide can be stored for periods longer than 17 years, which means that it is as storable as hydrazine (more information on this point are summarized in Table 1.3).

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Figure 1.2.5: Comparison of hydrogen peroxide with various oxidizers and fuels (adapted from Ventura and Mullens [3]).

Propellant Storability Hydrazine ~10 years (sealed)

Hydrogen Peroxide (98%) >3 years (sealed, demonstrated 1965) >15 years (sealed, estimated ) >17 years (vented, demonstrated) Table 1.3: Storability of hydrazine and hydrogen peroxide

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CHAPTER 1. INTRODUCTION 9 Propellant Toxicity Hydrazine 0.1ppm OSHA-PEL 60mg/kg (rat) Oral LD50 570ppm (rat) Inhalation LC50 (4h) Carcinogen Mutagen Hydrogen Peroxide (98%) 1ppm OSHA-PEL

805mg/kg (rat) Oral LD50, 70% HTP

170ppm (rat) Inhalation LC50, 50%

Table 1.4: Toxicity of hydrazine and hydrogen peroxide (Ref. [4]) Total Spacecraft Mass [kg] Description

100-1000 Small spacecraft 10-100 Micro-spacecraft 1-10 Nano-spacecraft <1 Femto-spacecraft Table 1.5: AFRL proposed satellite classication standard

Toxicity level plays also a signicant role in the choice of a propellant. When addressing this issue, the advantages related to the utilization of the hydrogen peroxide become more evident. As reported in Table 1.4, the toxicity level for hydrazine is one order of magnitude larger than for hydrogen peroxide (assuming as a toxicity measurement the Personal Exposure Limit of the Occupational Safety and Health Organization, but similar conclusions are attained if dierent indicators, as the lethal dose or concentration causing the death in 50% of the subjects, are considered). Furthermore, hydrazine is a carcinogenic compound. Strong time and cost savings arise from the replacement of a toxic propellant which a much less harmful one.

Finally, a deeper insight into the miniaturization issues mentioned above is worthwhile in this section. There is an increasing demand for the ability to provide extremely low levels of thrust or impulse bit: the need for micro-propulsion arises in essentially two distinct future mission scenarios. In the rst setting, a more traditional scale type of satellite may be used in a mission which requires its orbital attitude and position to be maintained with unprece-dented levels of precision. Extremely small perturbations to the attitude and/or position of the satellite may result because of such eects as solar radiation pressure or gravitational non-uniformities. To oset the perturbations, periodic (or continuous) thrust corrections on the order of micro-Newtons are required. It is important to note that the design challenge here is to develop a propulsion system capable of delivering very low thrust levels: however, the propulsion system itself is not subject to any severe constraints on size, mass or power us-age. To this end, electric-based micro-propulsion concepts have been identied as a potential solution.

The second scenario in which micro-scale levels of thrust are required is when the size of the satellite is itself signicantly reduced. There is an overall movement toward the development of miniaturized spacecraft which can be used for a number of applications. Having masses in the range of 10ö100kg, these platforms are more commonly known as micro-spacecrafts, according to the classication proposed by the US Air Force Research Laboratory [5] and summarized in Table 1.5.

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Thruster Type Specic impulse [s] Thrust [μN] Developing agency H2O2-monopropellant 160 1ö1000 NASA GSFC

Cold gas 40ö80 500ö50000 MIT, NASA JPL Digital solid 200 10ö100000 NASA GRC, TRW, CNES Vaporizing liquid 75ö125 1ö100 NASA JPL

Ion engine 1400ö2000 0.1ö10 NASA JPL FEEP 17000 10ö200 Alta S.p.A.

Table 1.6: Survey of some micro-propulsion initiatives in academia, government and industry in recent years

achieve specic mission goals. Distributed spacecraft concept such as formation ying rep-resents a departure from traditional satellite philosophy, which is based on a single, massive multifunctional spacecraft. An excellent example of such a mission is the joint ESA/NASA's laser interferometry space antenna (LISA) program, which is intended to detect gravitational waves in the solar system.Distributed spacecraft mission architectures oer a number of advan-tages such as reduced mission cost (production and launch), increased exibility and reliability and improved data resolution. Owing to the substantially reduced size, the micro-satellites have unique propulsion requirements, including extremely low thrust levels and/or extremely low minimum impulse requirements for orbital manoeuvres and attitude control. Propulsion systems for these satellites must also satisfy additional mass, volume and power constraints: micro-electromechanical-systems (MEMS) have been identied as potential solutions.: Several strategies have been identied and are summarized in Table 1.6.

In assessing the relative merit of a given approach, it is important to recognize that each mission specication will have its own set of unique constraints: one micro-propulsion strategy might be an excellent t for one type of mission but be inappropriate for others. Electrical-based microthrusters, e.g. ion engines or FEEP (Field Emission Electrical Propulsion) engines, are capable of delivering very low thrust levels but currently are large in size and mass, and have relatively high power requirements with respect to micro-satellite concepts. Solid pro-pellant devices are limited primarily by their digital nature: once red they are not reusable. These devices therefore are constructed in arrays and necessarily have a pre-set number of rings in their lifetime. Another potential problem in micro-satellite applications is that the location of the thrust vector will vary from ring to ring as the devices are depleted: the resulting shift in the thrust location can produce additional perturbations to the attitude of a small satellite, which must be accounted for in each ring.

Chemical mono-propellants have traditionally be attractive for satellites because they oer relatively high thrust-to-weight ratios owing to the inherently high amount of stored chemical energy per unit mass. This benet is compounded in MEMS-scale devices because of the interplay of volume and surface area scaling. The mass of propellant is a volumetric quantity which decreases as the cube of the length scale; in contrast, thrust is a surface quantity which depends on the surface area of the nozzle outlet and so decreases as the square of the length scale. Thus, the thrust-to-weight ratio increases linearly with decreases in the device size. Furthermore, mono-propellant thrusters typically rely upon a catalyzed chemical decomposition of the liquid propellant as the source of energy: as will be shown later in the present work, the eciency of the catalytic decomposition increases as the surface-to-volume ratio increases. Since the surface-to-surface-to-volume ratio is inversely proportional to the size of the catalytic chamber, additional benets should be expected in MEMS devices. If the higher density impulse of hydrogen peroxide with respect to hydrazine is accounted for in this context, the advantages related to the utilization of H2O2 as propellant not only for macroscale systems but for micro-thrusters too are straightforward [1].

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Figure 1.3.1: Adiabatic decomposition temperature of H2O2 against initial concentration of the solution.

1.3 H2O2 Propulsive System: Performance and Design

Considerations

Hydrogen peroxide (HP) monopropellant engines consist in a propellant supply and feed sys-tem, a combustion chamber and an exhaust nozzle. In particular, in hydrogen peroxide mono-propellant rockets the hot gases accelerated in the nozzle are not the products of a combustion reaction but the results of a decomposition. In fact, hydrogen peroxide has the characteristic of being able to decompose exothermically into water (steam) and oxygen according to the reaction:

H2O2(l)→ H2O(g)+1₂O2(g)

The decomposition reaction usually takes place inside a catalytic bed when liquid hydro-gen peroxide comes in contact with the active surface of a suitable catalyst that enhance the reaction speed by allowing to complete the decomposition at the exit of the catalytic bed. The adiabatic decomposition temperature of hydrogen peroxide depends on the concentration as indicated in Figure 1.3.1. For a xed geometry of the thruster a complete experimental characterization of the propulsive performance in terms of specic impulse, characteristic ve-locity and thrust coecient can be achieved by the experimental measurements of the thrust, the mass ow rate and the chamber pressure.

The specic impulse summarizes the overall propulsive performance both in terms of the eectivenesses of the combustion chamber and of the nozzle. The specic impulse is dened as:

Isp=_mg_˙F

0

where F is the thrust developed by the rocket, ˙m is the propellant mass ow rate and g0

is the gravitational acceleration at the sea level. An alternative formulation of the specic im-pulse leads to an expression in terms of the characteristic velocity c∗_{, and the thrust coecient,}

CF , according to the following equation:

Isp= c

∗_C F

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The characteristic velocity primarily depends on the chemical nature of the propellant. However, the actual characteristic velocity can be aected by the operating conditions and geometry of the thrust chamber. The conventional denition for this propulsive parameter is:

c

∗

=

pcAt

˙ m

where pc and Atare the chamber pressure and the cross-sectional area of the nozzle throat

respectively. Once the chamber pressure and the propellant mass ow rate are experimentally measured, the experimental evaluation of the characteristic velocity depends on the geomet-rical parameters of the nozzle. Theoretical gas-dynamics computations with the assumptions of quasi-1-D, steady, isentropic, frictionless, perfect gas ows in variable cross-section ducts, yield to the following expression for the characteristic velocity:

c

∗

=

√ γRT c γ γ+1 2

_2(γ−1)γ+1

In order to evaluate the deviation of the real behavior of the propellant gas in the thrust chamber from the theoretical one, a conventional parameter, called c∗ _{eciency, is usually}

introduced: ηc∗ = c∗ exp c∗ theo = pexp_c At mexp r RTcexp γ γ+1 2 γ+1γ−1

When the theoretical characteristic velocity is computed from the actual values of the chamber temperature (Texp

c ) and composition (R and γ ), it measures the eciency of the

mass expulsion. In fact, this eciency relation comprises viscous eects and the other sources of losses and ineciencies that are not considered in the quasi-1-D model of ideal rocket performance. On the other hand, as it will be claried later, if the theoretical characteristic velocity refers to the best propulsive performance theoretically achievable with the selected propellant (i.e. the adiabatic complete decomposition for hydrogen peroxide), the c∗_eciency

assesses also the eectiveness of the decomposition reaction. Except for the specic heat ratio, the thrust coecient is only a function of the nozzle geometry and its matching to the actual pressure ratio. In order to experimentally compute the thrust coecient, the measured thrust must be divided by the actual chamber pressure and the cross sectional area of the nozzle throat:

CF =_p_cF_A_t

Even for this parameter, a theoretical evaluation can be performed, as follows: CF = γ s 2 γ+1 γ+1γ−1 2 γ−1 1 −pe pc γ−1γ +pe−pa pc Ae At

The theoretical formula of the CFcan be particularly useful to extrapolate the performance

of the thruster for dierent values of the expansion ratio of the nozzle and at dierent external pressures, such as in vacuum.

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1.4 Catalytic Bed Requirements

Wernimont and Muellens [6] have reported the main parameters suitable for the characteri-zation of a catalytic bed for hydrogen peroxide decomposition. The requirements of a typical mission are usually expressed in terms of operating life duration during which the catalytic bed should be capable to sustain several cycles and satisfy specic requirements in terms of transient response and steady operation. Some specic performance parameters, such as the characteristic velocity eciency, describe the eectiveness of the decomposition process which is aected in a complex way by several operational parameters the most critical of which are the mass ux and the operating pressure of the bed. In the following sections all these features will be briey analyzed.

Catalytic Bed Life Specications

The operating life of a catalytic bed can be signicantly inuenced by the quality of the propellant (i.e. uid contamination and poisoning by stabilizers and impurities), thermal cycling (the degree of the cycle and the intensity) and operating parameters. The specications of the operating life of the catalytic bed can be summarized in the following points:

Operating Life Duration: in general the largest amount of life is preferred. For modern catalyst, the typical life durations are thousands of seconds to hours with 90% H2O2.

Operating Life Cycles: the maximum number of thermal cycles is desired in particular for attitude control application. A typical goal is to achieve a total cycles on the order of greater than 5000 cycles.

Decomposition Roughness: the decomposition roughness is usually expressed in terms of oscillations of the chamber or exit pressure from the catalyst bed. A smooth decom-position is conventionally dened as a roughness of less than 5% (peak-to-peak) of the mean chamber pressure.

Transient Requirements: in rocket engines for RCS, the transient startup is crucial in order to satisfy the requirements of the attitude control system. In general, the time-evolution of a thrust prole can be separated into dierent regions: ignition delay, thrust rise time and thrust decay time. The ignition delay, called decomposition delay in monopropellant rockets, consists in the time required for the thrust to reach 1% of its nominal value as the ring valve is opened. The thrust rise time refers to the time elapsed between 1% of the nominal value to 90%. The decay time concerns with the nal transient of the thrust and it is the time between the shut-o of the valve and the reaching of 10% of thrust steady-state value. Since the behavior of the thrust in the rocket can be approximated as a rst order system with respect to an impulsive variation of the mass ow rate, the typical time constant of rst order system (dened as the time needed after the ignition delay to reach (1e−1_{) of the nominal value) can be}

eectively used to characterize the transient behaviour of the system. Catalytic Bed Performance Parameters

The typical parameters used for assessing the performance of the catalytic bed in decomposing HP for propulsive application are:

the Characteristic Velocity Eciency; the Temperature Eciency;

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The capability of a catalytic bed of eectively decomposing HP to generate thrust can be assessed by means of the characteristic velocity eciency (named also c∗ _{eciency or C-Star}

eciency) and the temperature eciency.

In the c∗ _{eciency, the experimental characteristic velocity, computed from the}

measure-ments of the propellant mass ow rate, the chamber pressure and the nozzle throat cross-sectional area, is compared to the theoretical characteristic velocity relative to the ideal case of complete adiabatic decomposition:

ηc∗= c∗_exp c∗ theo = pexp_c At mexp r RTad c γ γ+1 2 γ+1γ−1

The discrepancy between experimental and theoretical characteristic velocity is due to the actual decomposition temperature (which is lower than the adiabatic decomposition temper-ature) and to the non-idealities of the real expansion in the thrust chamber (the assumptions of the quasi 1-D theory of ideal rockets are not accurately satised). The temperature e-ciency just highlights the rst cause of degradation of the characteristic velocity, expressing how close the measured chamber temperature is to the adiabatic temperature corresponding to complete decomposition of hydrogen peroxide:

η∆T =

Texp−Tamb

Tad−Tamb

Therefore, while the temperature eciency mainly indicates the eectiveness of the cat-alytic reactor, which cannot be necessarily followed by a thrust chamber, the c∗ _{eciency is}

the key propulsive parameter that takes into account both the chemical performance of the catalyst, which can reduce the decomposition temperature below its nominal value, and the non-idealities of the gas transit through the thrust chamber and the convergent part of the expansion nozzle.

In propulsive applications, another important operational parameter is the pressure drop across the catalytic bed. It obviously inuences the propellant pressurization system and it usually should be minimized in order to reduce the overall mass of the propellant feed system. However, an excessive reduction of the bed losses reduces the uid dynamic damping in the bed, exposing the chamber to the risk of development of self-sustained ow oscillations. Bed losses also represent the primary source of catalyst pressurization during the engine start-up. Their excessive reduction delays the pressurization of the catalyst, decreasing the propellant decomposition rate and increasing the length of the start-up transient. Finally, the variation of the catalyst pressure drop gives indications about the ageing of the bed and in particular of the onset of thermal fracture of the catalyst pellets. Its theoretical evaluation is quite uncertain because of its complex dependence on several operating parameters (bed porosity, pellet shape and size, bed length and load, and the pressure, temperature, density and velocity of the bed ow), most of which are only imperfectly known.

Catalytic Bed Operational Parameters

The eectiveness of the decomposition reaction inside the catalytic bed depends on both the catalytic activity of the catalyst and the operating conditions at which the catalytic bed is used. The working conditions of the catalytic bed are usually identied as follows:

Mass Flux: the mass ux (named also bed load or bed loading) is dened as the mass ow rate through the catalyst divided by its cross sectional area. The typical values of the mass ux range from a minimum of 0.07 lbm · in−2_{· s}−1 _{(or 50 kg · m}−2_{· s}−1_{) to a}

maximum of 0.6 lbm · in−2_{· s}−1 _{(or 422 kg · m}−2_{· s}−1_{). However, more typical values}

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Overall Mass: the overall mass of the catalyst depends on the catalyst material and the volume of the catalytic bed. For a specic application, the mass ux drives the catalytic bed diameter and, consequently, the volume of the bed is xed by the length of the bed. The length of the catalytic bed typically varies from 1.0 to 6.0 inches.

Catalytic Bed Pressure: historically the nominal operating pressure of catalytic beds for 90% hydrogen peroxide decomposition ranges from 100 psi to 1000 psi. The catalyst bed pressure is a crucial parameter because it directly inuences the density of the gaseous decomposition products that is linked with both the pressure drop across the catalytic bed and the eectiveness of the decomposition reaction.

Catalytic Bed Initial Temperature: the boot strapping ignition transient of a catalytic bed depends on the design of the bed and on the start conditions, such as the initial temperature of the catalytic bed and the propellant temperature. During the initial transient, the catalytic bed may ood and liquid hydrogen peroxide can pass throughout the catalytic bed and to come out of the nozzle. In general, a non-ooding ignition transient is preferred. Otherwise, if it is not possible to avoid the ooding of the bed, the boot strapping ignition transient should be minimized. In order to accomplish the start requirements, the catalytic bed may be pre-heated before the introduction of the propellant by means of pulsing or external heating elements. However, a desirable feature of the catalytic bed is the capability to start with cold initial conditions. Propellant Temperature: the propellant that enters the catalytic bed can be directly

drawn from the storage tank or may be passed through a regenerative cooling system and later introduced in the catalytic reactor. Consequently, the propellant temperature can signicantly vary. In fact, the minimum storage temperature of the hydrogen peroxide can be a few degrees above freezing (-12 °C) whereas the feed temperature after the heating in a regenerative system can be greater than 70 °C. Obviously, the inlet temperature to the catalytic bed can aect both the start transient and the steady performance of the thruster.

Degradation Phenomena of Catalytic Performance

In the operating life of a typical catalytic bed for hydrogen peroxide decomposition the main problems that can be encountered are:

Poisoning: The catalytic activity can diminish in consequence of deposition and sub-sequent reaction with preferential active sites onto the catalyst surface. Stabilizers are typically added in the commercial solutions to avoid H2O2 decomposition during

storage. Such chemical inhibitors are Sodium Stannate trihydrate Na2SnO3· 3H2O

or Na2Sn(OH)6 and Disodium dihydrogenodiphosphate Na2H2P2O7. Poisoning by

stabilizers is well documented in [7].

Flooding: the ooding of the catalytic bed takes place when the liquid hydrogen peroxide does not decompose completely passing through the bed. In this case, the temperature at the exit of the catalytic bed is lower than the boiling point of the water-hydrogen peroxide mixture at the operating pressure. This problem usually happens at the boot strapping ignition transient or when the catalyst is strongly deteriorated.

Channeling: It is a local phenomenon in which small amounts of liquid hydrogen peroxide do not directly comes in contact with the catalyst and pass through the catalytic bed without occuring a complete decomposition. In this case, the bulk temperature of the gases at the outlet of the bed is higher than the boiling point of the hydrogen peroxide mixture. The channeling yields to a deterioration of the decomposition performance

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and can generate dangerous pressure oscillations. The onset of pressure instability as a consequence of the channeling is due to the abrupt reduction of the pressure losses across the catalytic bed along the channel lled by the liquid mixture of water and hydrogen peroxide.

Clogging: the external environmental loads and the internal loads generated during HP decomposition by dierential thermal dilatations and uid pressure losses strongly stress the materials of catalytic bed. In this respect metallic screen beds usually perform better than more brittle ceramic substrates, which tend to degrade due to mechanical failures, and pellet beds, which are more prone to dusting or catalyst chipping as a consequence of ow-induced vibrations and thermal stresses. In some cases, the powdering and fracture can yield to the clogging of the catalytic bed.

Micro-scale degradation phenomena: This group includes such phenomena involving textural property evolution of the catalyst. The most important is the enlargement and/or destruction of catalyst pores for high pressure and temperature conditions which lead to loss of surface area and thus to activity reduction. Pellet based catalysts also suer from sintering of particle size whether a strong interaction of the metal with the carrier does not occur or there is no presence of elements capable to refrain the growing process from occurring. Moreover the active sites could suer from leaching or oxidation. The former is caused by the exposure to the high oxidizing environment whereas the latter is linked with catalyst mass loss by mechanical erosion.

1.5 Purpose of the Research Activity

The scope of the present research is the development of advanced catalysts on ceramic sup-porting pellets for hydrogen peroxide decomposition and their integration on prototypes of hydrogen peroxide monopropellant and bi-propellant rockets. The research has focused on ceramic based catalysts because of their potentially favorable thermo-mechanical properties that allow to identify them as the most promising catalysts for the decomposition of 98% hy-drogen peroxide. Moreover catalyst choice has been set in agreement with the current status of the activities at Alta S.p.A. and according the current know-how in preparing and testing catalytic beds for hydrogen peroxide decomposition. Silver and metallic based catalyst beds are limited in operating temperature when the concentration of HTP solution exceeds the 98 percent. The goals of the present research for the catalyst development are reported in the following chapter whereas the purposes with regard to the design of the prototypes were:

to test the catalysts at the typical conditions of space applications;

to investigate dierent bed congurations over a wide range of operating conditions (op-erating pressure, mass ux, overall dimensions of the bed etc.);

to experimentally evaluate the propulsive performance of the thruster and the decom-position performance of the catalytic bed.

the experimental assessment of the steady state and transient performances of the pro-totypes at dierent operating conditions;

the identication of the inuence of the operating paramenters on the initial start-up transient;

Concerning the development of bi-propellant thrusters the research has been focused on the ex-perimental evaluation of the self-pressurizing capability of gaseous ethane (C2H6) at ambient

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following activities have been undertaken to gain an experimetal proof of the Self Pressurizing Technology:

Design of a suitable combustion chamber and cooling system so as to account for the high expected ame temperature (2650 K);

Experimental evaluation of the specic impulse of the peroxide-ethane propellant com-bination and propulsive performance of the breadboard.

1.6 Publications

The experimental results and the scientic contents of the present thesis work have been published in the following conference and journal papers:

Bramanti C., Cervone A., Romeo L., Torre L., d'Agostino L., Musker A. J., Saccoccia G., Experimental Characterization of Advanced Materials for the Catalytic Decomposition of Hydrogen Peroxide, AIAA Paper 2006-5238, 42nd AIAA/ASME/SAE/ASEE Joint Propulsion Conference, Sacramento, California, USA, July 2006.

Cervone A., Romeo L., Torre L., d'Agostino L., Calderazzo F., Musker A. J., Roberts G. T. and Saccoccia G., Development of Green Hydrogen Peroxide Monopropellant Rocket Engines and Testing of Advanced Catalytic Beds, 3rd International Conference on Green Propellants for Space Propulsion, Poitiers, France, September 2006.

Romeo L., Torre L., Pasini A., Cervone A., d'Agostino L., Calderazzo F., Performance of Dierent Catalysts Supported on Alumina Spheres for Hydrogen Peroxide Decompo-sition, AIAA Paper 2007-5466, 43rd AIAA/ASME/SAE/ASEE Joint Propulsion Con-ference, Cincinnati, Ohio, USA, July 2007.

Pasini A., Torre L., Romeo L., d'Agostino L., Cervone A., Musker A., Experimen-tal Characterization of a 5N Hydrogen Peroxide Monopropellant Thruster Prototype, AIAA Paper 2007-5464, 43rd AIAA/ASME/SAE/ASEE Joint Propulsion Conference, Cincinnati, Ohio, USA, July 2007.

Romeo, L., Torre, L., Pasini, A., d'Agostino, L., Calderazzo, F., Development and Testing of Pt/Al2O3 Catalysts for Hydrogen Peroxide Decomposition, 5th International Spacecraft Propulsion Conference, Heraklion, Greece, May 2008.

Torre, L., Pasini, A., Romeo, L., d'Agostino, L., Firing Performances of Advanced Hydrogen Peroxide Catalytic Beds in a Monopropellant Thruster Prototype AIAA Pa-per 2008-5575, 44th AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit, Hartford, Connecticut, USA, July 20-23 2008.

Romeo, L., Torre, L., Pasini, A., d'Agostino, L., Calderazzo, F., Comparative Char-acterization of Advanced Catalytic Beds in a Hydrogen Peroxide Thruster Prototype AIAA Paper 2008-5575, 44th AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit, Hartford, Connecticut, USA, July 20-23, 2008.

Pasini, A., Torre, L., Romeo, L., d'Agostino, L., Performance Modeling and Analysis of H2O2 Catalytic Pellet Reactors AIAA Paper 2008-5025, 44th AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit, Hartford, Connecticut, USA, July 20-23, 2008. Romeo L., Genovese C., Torre L., Pasini A., Cervone A., d'Agostino L., Centi G., Perathoner S., ``Use of Pt/CexO2-Zr1-xO2/Al2O3 as Advanced Catalyst for Hydro-gen Peroxide Thrusters'', AIAA Paper 2009-5637 45th AIAA/ASME/SAE/ASEE Joint Propulsion Conference, Denver, Colorado, USA, August 2009.

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Pasini, A., Torre, L., Romeo, L., Cervone, A., d'Agostino, L., Endurance Tests on Dierent Catalytic Beds for H2O2 Monopropellant Thrusters, AIAA Paper 2009-5472, 45th AIAA/ASME/SAE/ASEE Joint Propulsion Conference, Denver, Colorado, USA, August 2009.

Romeo, L., Bonavita, A., Torre, L., Pasini, A., Cervone, A., d'Agostino, L., Centi, G., Perathoner, S., Post-Reaction Characterization of a Pt/CexZrx-1/Al2O3 Catalyst After the Use in a HTP Thruster, ESA Space Propulsion 2010 Conference, San Sebastian, Spain, May 2010.

Torre, L., Romeo, L., Pasini, A., Cervone, A., d'Agostino, L., Green Propellant Re-search at Alta S.p.A., ESA Space Propulsion 2010 Conference, San Sebastian, Spain, May 2010.

Pasini, A., Torre, L., Romeo, L., Cervone, A., Masi, L., d'Agostino, L., High Mass Flux Tests on Catalytic Beds for H2O2 Monopropellant Thruster, ESA Space Propulsion 2010 Conference, San Sebastian, Spain, May 2010.

Pasini, A., Torre, L., Romeo, L., Cervone, A., d'Agostino, L., Thermal Stress Analy-sis of Ceramic Pellets for CatalyAnaly-sis, 46th AIAA/ASME/SAE/ASEE Joint Propulsion Conference, Nashville, Tennessee, USA, July 2010.

Pasini, A., Torre, L., Romeo, L., Cervone, A., d'Agostino, L., Testing and Character-ization of a Hydrogen Peroxide Monopropellant Thruster, Journal of Propulsion and Power, Vol.24, N°3, May-June 2008.

Torre, L., Pasini, A., Romeo, L., Cervone, A., d'Agostino, L., Performance of a Monopropellant Thruster Prototype Using Advanced Hydrogen Peroxide Catalytic Beds, AIAA Journal of Propulsion and Power, Vol. 25, No. 6, Nov-Dec 2009, pp. 1291-1299. Pasini, A., Torre, L., Romeo, L., Cervone, A., d'Agostino, L., A Reduced Order Model for H2O2 Catalytic Reactor Performance Analysis, AIAA Journal of Propulsion and Power, Vol. 26, No. 3, May-Jun 2010, pp. 446-453.

Pasini, A., Torre, L., Romeo, L., Cervone, A., d'Agostino, L., Performance Character-ization of Pellet Catalytic Beds for Hydrogen Peroxide Monopropellant Rockets, AIAA Journal of Propulsion and Power, Vol. 27, No. 2, Mar-Apr 2011.

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Chapter 2

Catalyst Development

The most challenging component for a monopropellant thruster is represented by the catalytic bed. An increase in its lifetime, decomposition eciencies along with in the thermo-mechanical strength will directly aect the propulsive performance of the overall thruster. The scope of the present work is the development of an innovative catalyst conguration for H2O2

decomposition capable to provide further improvements with respect to the actual solutions. The rst part of this chapter describes the state of art of the catalyst used in the aerospace and automotive elds till now bringing into focus the potential preparation techniques and materials capable to enhance the technological readyness level of the catalyst. According to the literature review the ceramic supporting pellets coated with platinum and a suitable layer of rare earths oxides are to be considered innovative choices for the improvement of the catalytic performance. Once the candidate conguration has been chosen a series of preparation techniques have been implemented in close collaboration with the Department of Industrial Chemistry of Pisa and Messina University. The following sections of this chaper reports the results of the experimental measurements which have been carried out to assess the activity level of the potential catalyst for the subsequent experimentation in the test rig. Finally the last two sections concern with the qualitative/quantitative characterization of the catalyst surface through Scanning Electron Microscope (SEM) and X-Ray Diractometry (XRD). The supercial analysis has been performed onto new samples to understand the chemical system resulting from the preparation and on the used catalysts to investigate the onset of morphology/chemical variations induced by H2O2decomposition reaction.

2.1 Catalyst Development Guidelines

The starting point for the catalyst development has been set in agreement with the current status of the activities at Alta S.p.A. concerning with the catalyst conguration already tested. The present research work has taken into consideration the rupture occured to the catalytic bed in the test rig during a former experimental campaign [17]. In order to overcome this problem dierent solutions have been considered. The eectivness of the impregnation technique developed by Alta S.p.A. in collaboration with the Department of Chemistry and Industrial Chemistry of Pisa University has been experimentally proved and for this it has been implemented on new commercial catalytic substrates. The purchasing of new catalyst carriers was driven to increase the thermo-mechanical strengh of the new compounds without deeply reducing the wet surface area and simplyng the interfacing with the catalytic reactor assembly.

During operation, hydrogen peroxide decomposition catalysts are subjected to high ther-mal stresses, mostly due to severe and sometimes intermittent therther-mal shocking. It is therefore

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CHAPTER 2. CATALYST DEVELOPMENT 20

necessary for eective catalysts to combine both superior mechanical and chemical properties, which imposes a delicate trade-o between the conicting requirements of high mechanical resistance and large surface area. Similar operational conditions are experienced by hydrazine decomposition catalysts which had been widely used in monopropellant applications over the last 40 years. For this reason it seemed to be fruitful to purchase a catalyst for N2H4

de-composition to have a benchmark as regards to performance. Several research groups devoted long and intense eorts to the development of eective hydrazine decomposition beds and the challenge to prepare an alumina support in the form of spheres or pellets for this ex-treme catalytic application has been met. The standard enthalpy of reaction for hydrazine decomposition is about 336 kJ mol-1; more than threefold the 98 kJ mol-1 released during hydrogen peroxide decomposition. One of the commercially available catalysts is the Shell 405 (USA), whose substrate consists of granules of Reynolds RA-1 alumina. The outstanding properties of Shell 405 are presumably due to a more open structure and to the presence of a high fraction of eta-alumina crystals stuck together by amorphous alumina binder. The dense crystals prevent the penetration of hydrazine into the pellets and decrease the generation of catalyst nes by break-up caused by internal overpressures [8].

In addition to the Shell 405 catalyst, another catalytic system for hydrazine decomposition was developed in Europe: the KC-12-GA catalyst. It was formerly produced by Kali-Chemie Company. Subsequently the business was rst sold to Solvay and after transferred to W.C. Heraeus GmbH. In this case the substrate is produced by means of a suitable sol-gel procedure so as to create nearly spherical granules and in such a way to assure a uniform stucking in the catalytic bed and reduced pressure drops [8].

Novel alumina-based substrates have also been developed for HAN catalysts, which have to withstand even more severe conditions due to the higher temperatures reached during the decomposition (more than 1400 °C). Previous studies have revealed that alumina doped with silicon, lanthanum or barium oxides displays higher thermal stability than pure transition alumina [9].

Considering the current status of the catalyst development, the following guidelines have been set:

To investigate innovative techniques for the washcoating of alumina with low and medium surface area in order to enhance the thermo-mechanical strength;

To purchase alternative commercial substrates capable to withstand the thermal stresses induced by the reaction. It includes bimodal alumina, γ −Al2O3doped with lanthanium

and boron oxides; silicon carbide, cordierite, foams of silicon carbide and vitreous carbon; To implement a suitable impregnation technique for depose high load of platinum onto

the catalyst carrier so as to achieve a well-homogeneous distribution;

To procure and test a commercial catalyst composed by the substrate for hydrazine decomposition and platinum as active species.

It has been decided to experimentally assess the catalytic activity of the samples preliminarily at low peroxide concentration and subsequently in the sub-scale test rig where the actual conditions experienced in a monopropellant thruster can be reproduced.

2.2 Literature Survey

2.2.1 Catalysts for Hydrogen Peroxide Decomposition

During the last three years research on hydrogen peroxide-based rockets and catalysts was conducted at Korea Advanced Institute of Science and Technology. The rst Korean Satellite

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Figure 2.2.1: Preparation steps for P t/Al2O3 catalyst (Pure Al2O3 , crushed Al2O3 and

impregnated with Platinum)

was KITSAT-1 and has been launched at 1992 in collaboration with Surrey University. These small satellites for scientic programs use gravity gradient boom, magnetorquer or reaction wheel for the attitude control. Their mission ability was strongly limited because there is no an active propulsive system. To ll this gap a propulsion system should be developed to suitable thrust level, which will be challenge to pratical use of small satellites. Hydrogen peroxide represents a potential candidate owing to the cost savings associated with its low toxicity and enhanced versatility. Sungyong et al [10]in 2007 fabricated and tested an 1N-level thruster with an aluminum oxide loaded platinum catalyst in the reaction chamber. Four thrusters were integrated with the propellant feed system including the propellant tank, pressurizing gas tank, lter and solenoid valves, and a thruster module was assembled and successfull tested. Among many catalysts and its substrate, the authors selected aluminum oxide as a substrate and platinum as catalytic metal. The catalyst was prepared from a bimodal alumina (Alfa Aesar, γ − Al2O3 type) displaying a surface area of 255 m2g−1, a

total pore volume 1.14 cc/g and median pore size of 70 µm. The preparation consisted in three steps (impregnation/drying, calcination and H2reduction) using H2P tCl6as a precursor

and the wetness impregnation method. The impregnation was followed by drying (12 hr at room temperature), calcination (2 hr at 300 °C) and H2 reduction (4 hr at 300 °C). To coat

the large amount of catalyst, impregnation/calcination processes were performed three times before the reduction under hydrogen. Finally, P t/Al2O3catalyst sample was prepared and

catalyst weight fraction about 30.4 %. Alumina pellets and prepared catalyst are shown in gure 2.2.1.

The 1.8 grams of P t/Al2O3 were inserted into the reactor chamber and re test was

per-formed repeatedly near 20 bar of feeding pressure. The quantity of the hydrogen peroxide(90.1 wt%) at each test was 200 cc (about 280 gram). Performance tests were done on the same sample until the catalytic activity decreased . Each re test was carried out eight times and propellant mass owrate was about 1.9~2.4 g/s, which is about three times owrate than design condition and is equivalent to a bed load of 9 kg · m−2_s−1_{. Total operated time and}

consumed propellant was about 1100 seconds and 2240 grams. In all cases, the pressure in the reaction chamber remained stable and pressure uctuations similar to chugging instabil-ity were not observed. After the re test number 6, a decrease in reactivinstabil-ity was observed. After having carried out the eight test (at high propellant owrate), a new test (0.53g/s, at normal propellant owrate) was performed. Thruster was red for 1300 sec continuously and consumed about 500 cc of 89.6 wt% hydrogen peroxide.

The catalyst was tested in the prototype at a mass ow rate above the design point and at normal owrate. It operated for 1100 seconds (2240 gram of propellant) and 1300 seconds (700 gram of propellant) at two dierent tests respectively . The temperature of the gaseous reaction products was near the adiabatic temperature of the propellant and eciency based on the temperature was just below 100%. Despite a full cone spray injector being used to

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vaporize peroxide into droplets with a Sauter Mean Diameter of 103-135 µm, the thruster needed about 20 seconds for achieving the steady state operation.

The catalytic bed developed at KAIST in the former work represented the starting point in a subsequent paper[11] to obtain the design data necessary for sizing the reactor chamber of a 50 Newton thruster.

Platinum was selected as a catalyst for decomposition of hydrogen peroxide. The catalyst bed was prepared from a γ-type bimodal alumina from Alfa Aesar, which displays 255 m2_g−1

of surface area, 1.14 cc/g of total pore volume, and 70 and 5000 Å of median pore size. The preparation was performed with H2P tCl6 solution as a precursor, using the wetness

impregnation method. Impregnation was followed by drying (90 °C , 12 hrs), calcination (300 °C , 4 hrs), and H2 reduction (300 °C , 4 hrs). The catalyst coating process was performed

two times. The nal P t/Al2O3 catalyst was prepared as 23 wt. % of active material based

on the total catalyst weight.

Authors designed and tested two dierent thruster prototypes: a scaled down reactor chamber of 1 cm in diameter and 4 cm in lenght and a 50 Newton thruster with a catalytic bed 3 cm in diameter and with the same length of the scaled down prototype. The catalyst bed was 3 cm x 4 cm determined from experimental data of decomposing capacity of the catalyst for the scaled down reactor. It was scaled up in the radial direction to decompose larger propellant ow rates than the scaled down version.

Maximum catalyst capacity has been experimetally assessed by scaling up peroxide mass rate owing through the scaled down test rig and by computing the c∗ _{eciency for each}

value of mass ow rate. Adiabatic temperature of 90 wt% hydrogen peroxide (about 750 °C) was obtained at region A where the propellant was fully decomposed. In region B a maximum temperature was measured at the end of the catalyst bed and observed to be below the adiabatic one at every position of the catalytic bed in region B. Figure2.2.2shows the c∗

eciency as a function of the oxidizer mass ow rate. The eciency was over 90%, which was constant at an increased propellant mass owrate below 4.0 g/s (region A). Afterwards, the eciency considerably decreased with an increase of propellant mass owrate over 4.0 g/s (region B). Maximum catalyst capacity, which was dened as allowable propellant mass owrate divided by volume of catalyst bed, was 1.27 g ·s−1_{· cm}−3_{, as observed at boundary of}

region A and B. The 50 Newton thruster used the same catalyst as the scaled down thruster. The size of the catalyst bed was enlarged for achieving the fully decomposition of the peroxide design ow rate, but having the same capacity value of the scaled down thruster. Finally a verication test of the scaled up thruster was performed using a propellant mass ow rate of 34.8 g · s−1_{(bed load equals to 50kg · m}−2_s−1_{) and 22.8 bar as feeding pressure. As results the}

c∗eciency was 98% along with 1.4 bar of pressure drop across the catalytic bed. Pressure in the reaction chamber was very stable and instability, like chugging instability which frequently occurs in a small thruster chamber, was not observed.

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Figure 2.2.2: Eciency of characteristic velocity as a function of propellant mass ow rate

Figure 2.2.3: Temperature measurements as a function of time for the50 Newton thruster [11]. A stable reactivity was because the spray injector guaranteed the uniform injection of propellant into the frontal area of the catalyst bed. Based on decomposition eciency data, nearly all of propellant was decomposed onto catalyst but even with the usage of a pressure atomizer, a thruster response time of 14 seconds was still measured (see gure2.2.3 ).

With the aim of improving the start-up transient of KAIST's thruster, Sungyong et al. have reported a new choice for the catalytic material in a subsequent paper [12] . In this work authors measured the response times of three dierent thrusters by varying the type of injector, the reactor volume and the grain size of the catalyst. MnO2 was used as the

active material for the decomposition of H2O2because of its superior activity. MnO2from a

precursor (purchased from Aldrich) was deposited on Al2O3(purchased from Alfa Aesar) by

using a suitable impregnation method. If the size of the alumina grains increases, the resistance to the diusion of a reactant (H2O2) from the alumina surface to the inner pores increases

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size should inuence the catalytic activity, which aects the response times of the thruster. Thus to nd an experimental proof, two samples of alumina with dierent particle sizes were prepared and coated with MnO2 . As shown in gure 2.2.4 one comprised (1/8)-inch pellets

and the other consisted of granules with a mesh size of 1620 (0.851.18 mm).

Figure 2.2.4: MnO2/Al2O3catalyst (1/8 pellets and granules with a mesh size of 16-20)

To determine the eect of the catalyst size on the response time, reaction tests were performed using the pellet and granular catalyst separately. The thruster assembly included a shower head injector instead of the previous full cone spray. The injector had 19 orices each with a diameter of 300 µm. When the shower-head injector was used, it was dicult to obtain a uniform distribution of droplets as in spray injection. However, useless distance at spray injection was minimized, which in turn minimized the ullage volume in the thruster.

In the case of the granular catalyst, the response times were considerably shorter than those one of the pellet catalyst. The ignition delay, pressure rise time, and tail-o time were 12, 114, and 106 ms, respectively. A decrease in the grain size of the granular catalyst had two positive eects: one on the catalytic activity and the other on the ullage volume. A catalyst grain with a large diameter gives rise to large diusion resistance. The catalytic activity can increase if the catalyst grain size decreases because the resistance to the diusion of a reactant from the catalyst surface to the inner pores decreases, which increases the eective surface area of the catalyst grains. The diusion phenomenon controls the total reaction rate in a reactor at high temperatures. Therefore, the propellant decomposition occurs mainly on the pellet surfaces rather than in the inner pores. In other words, the increase in the total eective surface area of the catalyst in the reactor contributes to an increase in the reaction rate and a decrease in the response time in the H2O2 decomposition process. In addition, the ullage volume among grains decreases with the smaller catalyst grain by stacked with high density, which is helpful in decreasing the response time, especially the pressure rise and decay times. A great work has also been done by Astro-Technology Space Oriented Higashiosaka Lead-ing Association (SOHLA) for the propulsion subsystem of their demonstrator of Panel ExTen-sion SATellite (PETSAT) [13]. It consists of standardized subsystem panels which are joined together and could be easily assembled according to mission specication. Both monopropel-lant and bipropelmonopropel-lant chemical propulsion resulted to be the strongest candidates because of its large thrust density, low power consumption, and its multiple operation capability. The choice for PETSAT propulsion system was mainly driven by safety issues. A concrete way to realize a high safety level is to reduce the risk of propellant, especially oxidizer. The risk means toxicity, causticity, explosibility in particular: Ethanol is less toxic than methanol, less caustic than hydrazine, and less explosive than acetone or HAN, and lower concentration of Hydrogen Peroxide (HP) solution is less dangerous than the higher one or H2O4, N2O.

In view of this Sahara et al. [13] reported performance estimation, catalyst test, and captive test for hydrogen peroxide monopropellant thruster. we continuously tested various catalysts for HP decomposition as shown in gure 2.2.5. NS-1A was the best catalyst but

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Figure 2.2.5: NS-1A catalyst is broken with high pressure in the test (Top). Foamed metal and TANAKA-I and TANAKA-II (Bottom)

its strength is not so high so that it was broken under the high pressure in decomposing HP. Foamed metal catalyst shown in the left side of gure 2.2.5 was the second candidate but its eciency is not as high as NS-1A, in addition, its eciency decreased for a while after several operations. So authors were seeking new catalysts and at last found catalysts made by TANAKA KIKINZOKU GROUP. They obtained catalysts for HP decomposition (TANAKA-I) and alcohol combustion (TANAKA-I(TANAKA-I), shown in the right side of gure 2.2.5. TANAKA-I and TANAKA-II are metal-supported platinum and palladium catalysts respectively, and after an experimental assessment they decided to use TANAKA-I and TANAKA-II as catalytic systems for PETSAT. They conducted captive test using a breadboard model of a HTP monopropellant thruster. The left side picture in gure 2.2.6 shows the thruster and the plume in the test whereas the right side shows pressure and temperature histories in the thrust chamber. It indicates that its temperature reached at boiling point of 60 % HP at the pressure, and wet steam was attained in the thruster chamber. Propellant mass ow rate was 1.59 g/sec, throat area was 1.57 mm2, and specic exhaust velocity is obtained as 415 m/sec. Suppose that practical nozzle coecient is 1.5, its specic impulse and thrust were estimated to be 63 sec and 989 mN respectively.

Hydrogen peroxide has recently been tested also for a 100-Millinewton HTP monopropel-lant microthruster [14] . In this work the microthruster demonstrated to perform a minimum impulse bit of 10−2_{N.s, which can provide a microsatellite slew at a speed of 0.1 deg/s. The}

total volume of the microthruster prototype is about 0.9 cm3_{, with a throat diameter of 0.5}

mm packing, with silver ake as the catalyst and 92% peroxide with a ow rate of 0.18 g/s. One of the major problems in a MEMS-based hydrogen peroxide microthruster is the inhi-bition of decomposition progress due to enhanced heat loss as the surface-to-volume ratio is increased in miniaturization. In addition, gas bubbles are usually observed in the microchan-nel, and eects of impeding fresh hydrogen peroxide from contacting the catalyst surface become relatively signicant as the size is reduced.

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be-CHAPTER 2. CATALYST DEVELOPMENT 26

Figure 2.2.6: Captive test of monopropellant system. The left side shows thruster and the plume under operation, the right side shows pressure and temperature histories in the thrust chamber

cause the hydrogen peroxide is decomposed on the catalyst surface and in the volume, due to homogeneous thermal decomposition. It seems that the increased surface-to-volume ratio in microthruster design may benet the surface reaction of the catalyst. Yet, in reality, with xed catalyst cell size, the catalyst surface-to-volume ratio remains the same when reducing the chamber size. In other words, the specic heat generation per unit of hydrogen peroxide by catalytic surface reaction remains the same, but the heat loss is enhanced as the size of the hydrogen peroxide microthruster is reduced. This would pose an outstanding limitation on the design of a catalytic hydrogen peroxide microthruster. The selection of durable catalyst and associated substrate is an important issue in the design of a hydrogen peroxide microthruster. Pure silver or plated silver on a nickel screen would be ideal as a practical choice. In practice, with lower chamber temperature, the catalyst bed in the microthruster system may not be as active as that in the hundreds-ofnewtons hydrogen peroxide rockets.

Considering the durability and activity, silver was selected as the catalyst for the hydrogen peroxide microthruster in this study. First, a γ − Al2O3 coating with silver in the shape of

pellets and a pure silver pellet with a 1-mm diameter were both tested. The former cannot operate for a long time because the silver will be washed out by hydrogen peroxide. The latter has poor performance because the outlet temperature does not exceed 250 °C. Finally, pure silver ake sized between 90 and 250 µm is adopted, which is easy to ll and fabricate. It also has the best performance among these catalyst types. In the microthruster, the catalyst chamber lled up with silver akes and a ne stainless steel screen is used as the bae. This screen is also used to uniformly distribute hydrogen peroxide. The total mass of silver akes is about 0.7 g. In gure 2.2.7 is shown a schematic drawing of the HTP microthruster.

The prototype of the microthruster is designed to generate a thrust on the order of 100 mN. According to previous studies, for a 100-mN thrust, a mass ux of 100 kg · m2_{· s}−1 _is

usually quoted. However, for a miniature design in which large heat loss is prevalent, the cross-sectional diameter of the 100-mN microthruster calculated from the mass ux of large thrust engines would be 1 to 2 mm. It may be too small to sustain acceptable performance and

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Figure 2.2.7: Design of hydrogen peroxide microthruster

Figure 2.2.8: Decomposition temperature and pressure vs time for a catalyst preheating tem-perature of 300k (left) and 423 K (right)

may eectively lead to the decomposition instability in such a small chamber. In this study, a catalyst-bed diameter of 5 mm is used for a catalyst load factor equal to 9.2 kg · m2_{· s}−1

. The calculated length of the catalyst chamber is 5.5 mm. The packed weight of the silver catalyst is 0.6 g. The overall hydrogen peroxide microthruster, including the distribution plate, catalyst chamber, and nozzle, weighs 5.8 g in total.

The catalyst is resistively heated with a power between 1015 W. Once the catalyst at-tains the destined preheating temperature, the valve will open and the external power will be switched o immediately. Without preheating the catalyst bed, the hydrogen peroxide microthruster requires about 15 s to reach the steady-state pressure of 565 kN · m−2 _after

the solenoid valve is opened. The reactor temperature rapidly reaches 420 K after the valve opens and stays at 420 K for about 12 s. It then sharply increases to 900 K and drops a little to the steady state, at approximately 820 K. When the catalyst bed is preheated to 423 K, the pressure rises to 689 kN · m−2 _{in 50 ms after the valve is opened, whereas the}

catalyst-bed temperature increases to about 850 K in 1 s. Moreover, it decreases slightly to the steady-state pressure of 579 kN · m−2 _{and temperature of 840 K. In gure the time trace}

of the chamber temperature and pressure for the cases of preheating temperatures of 300 K and 423 K, respectively are shown.

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Figure 2.2.9: Ignition delay time of the microthruster versus preheating temperatures at ambient pressure

A thruster for attitude control of microspacecraft, the ignition delay time, is an important index of thruster performance. In this study, the ignition delay time is dened as the time required to attain 95% steady pressure after the valve is opened. Note that the response time of the solenoid valve, which is estimated to be 6 to 20 ms, is included in the ignition delay. The ignition delay time is about 14 s without preheating, which is not suitable for the requirement of attitude control. For preheating in the temperature range between 323 and 363 K, the ignition delay times are more than 10 s. With the preheating temperature increased to 403 K, the ignition delay time may be reduced to 150 ms, further to 50 ms at 423 K, and 35 ms at 453 K (see gure 2.2.9 ).

To coincide with the thruster duty cycle in a typical microsatellite mission, a total of 20 tests are performed to investigate the catalyst durability of the hydrogen peroxide mi-crothruster. Unlike the rocket propulsion system, the micropropulsion system is mainly used for altitude control. In the operation of a micropropulsion system, being limited by the pro-pellant tank and propro-pellant mass, the ring duration is usually less than 12 s for most altitude control commands. Therefore, in the study, each test lasts continuously for 30 s. The total run time for catalyst durability is 600 s.

The c∗_{eciencies calculated by the average pressure for these 20 tests is illustrated in}

Fig. 2.2.10. It shows that during the tests, the c∗ _{eciency constantly reaches around 90%}

with no obvious decay. The average c∗ _{eciency of the 20 tests is 0.89, with the standard}

deviation of 0.028. The average thermal eciency of the 20 tests is 0.69, with the standard deviation of 0.042. Several researchers proposed that the aging of silver is caused by the formation of silver oxide or coating of the hydrogen peroxide stabilizer on the surface of silver. The role of silver oxide in the catalytic decomposition of hydrogen peroxide is still debatable. Both promoting and prohibiting the decomposition process were proposed in the literature. Nevertheless, silver oxide is unstable and will reduce to silver above 160 °C. We nd that the silver ake becomes pale in color after testing, unlike the shining silver before using it. It may suggest that the silver surface becomes rougher with increased surface area after tests. Interestingly, the used silver bed seems to have a slightly higher performance than fresh silver. The durability test for HTP catalysts reported by Kuan et al. [14] is capable to very eectively depict the trend of the catalytic performance by allowing to detect the onset of