Abstract:
An apparatus is provided for burning a fuel and nitrous oxide. The apparatus has a combustor, a catalyst, a nitrous oxide supply passage for directing the nitrous oxide to a contact position with the catalyst, and a fuel supply passage for supplying the fuel to the combustor. The catalyst is for facilitating decomposition of the nitrous oxide, and the combustor is for burning the fuel, the decomposed nitrous oxide and/or further nitrous oxide decomposed in the reaction.

Description:
This application claims the benefit of U.S. Provisional Application No. 60/254,003, filed Dec. 7, 2000. 
    
    
     BACKGROUND OF THE INVENTION 
     Embodiments of the invention relate to burning fuels. More particularly, embodiments of the invention relate to burning fuels with nitrous oxide. 
     SUMMARY OF THE INVENTION 
     Embodiments of the invention include an apparatus for burning a fuel and nitrous oxide. The apparatus has a combustor, a catalyst, a nitrous oxide supply passage for directing the nitrous oxide to a contact position with the catalyst, and a fuel supply passage for supplying the fuel to the combustor. The catalyst is for facilitating decomposition of the nitrous oxide, and the combustor is for burning the fuel and the decomposed nitrous oxide. 
     Other embodiments of the invention include a rocket engine. The rocket engine has a nozzle, a combustor, and a propellant supply system. The propellant supply system has a catalyst, a nitrous oxide supply passage for directing nitrous oxide to a contact position with the catalyst, and a fuel supply passage for supplying a fuel to the combustor. The catalyst is for facilitating decomposition of the nitrous oxide, and the combustor is for burning the fuel and the decomposed nitrous oxide. 
     Other embodiments of the invention include a method of burning nitrous oxide and a fuel. The method includes supplying the nitrous oxide to a catalyst, facilitating decomposition of the nitrous oxide by passing the nitrous oxide over the catalyst, supplying the fuel to a combustion point, and supplying the decomposed nitrous oxide to the combustion point. 
    
    
     BRIEF DESCRIPTION OF THE DRAWINGS 
     FIG. 1 is a sectional view of an injector in accordance with an embodiment of the invention; 
     FIG. 2 is an example of a propellant system in accordance with an embodiment of the invention; 
     FIG. 3 shows a plot of thrust coefficient versus mixture ratio; and 
     FIG. 4 shows a rocket engine in accordance with an embodiment of the invention. 
    
    
     DETAILED DESCRIPTION OF PREFERRED EMBODIMENTS 
     The invention provides a nitrous oxide (N 2 O)/propane (C 3 H 8 ) rocket engine (NOP), or any other hydrocarbon fuel, that utilizes catalytic decomposition of N 2 O as an igniter system. This propellant combination is an alternative to the present space propulsion systems that use hypergolic or cryogenic liquids, or solid propellants. Various features of the invention are discussed below by way of examples. The invention is not limited to these illustrative examples and has a scope that should be clear to one skilled in the art upon reading this disclosure. 
     A serious limitation on the ability of the commercial aerospace industry to place into and keep satellites in Low-Earth-Orbit (LEO) at economical prices is the choice of propellants and propulsion technologies used for rocket boost, attitude control systems (ACS), reaction control systems (RCS), orbital maneuvering systems (OMS), and auxiliary power units (APU). Present systems are either liquid propellants that are hypergolic or cryogenic, or solid propellants that are single use only, are unthrottleable, and are explosive in nature. 
     The invention provides a solution to the problem by selecting propellants for a chemical propulsion system that are readily available, are easier to handle, non-toxic, produce relatively high performance, and provide significant reduction in cost of operations. High operating costs are a result of occupational safety requirements associated with the handling of toxic, hypergolic propellants and of added complication of operating a cryogenic propellant system. A cryogenic system also adds considerable dry weight, further reducing the payload weight fraction. 
     By using nontoxic, benign propellants that are relatively safe to handle, low cost can be realized through simplified ground operations. Such a propellant combination could also benefit other systems for which safe and simple ground operations are a requirement. Rocket assisted takeoff systems (RATO) for unmanned aerial vehicles (UAV) could benefit from a system that would provide simplified ground operations since these systems may be deployed in future battlefield scenarios. Military personnel would benefit from the safe handling characteristics of benign propellants, and superior performance to other propellant combinations allowing the UAV to perform its mission with lower risk of neutralization by the enemy. 
     The invention exploits several unique properties of the propellants, propane and nitrous oxide, for a chemical rocket propulsion system. These self-pressurizing propellants have a distinct advantage over current systems that use hydrazine as a monopropellant and monomethyl-hydrazine and nitrogen tetroxide (MMH/NTO) as bipropellants. They are standard liquefied industrial gases and are classified by the U.S. Department of Transportation as simple asphyxiates, with propane as a flammable gas and nitrous oxide as a mild oxidizer. They are neither highly explosive nor hazardous to work with or handle. A chemical liquid propulsion system in accordance with the invention, using these environmentally benign propellants, may be economically advantageous to current hypergolic or cryogenic systems. They possess commercial availability at low prices and are easy to handle, thereby producing a significant reduction in operating costs. 
     A feature of nitrous oxide facilitates autoignition of propane without the use of hypergolics. Nitrous oxide can be catalytically decomposed using a wide variety of catalysts, including, for example, platinum, iridium, rhodium, tungsten carbide, copper, cobalt, and gold. This catalytic decomposition is discussed herein using a number of catalysts including, but not limited to, the iridium based Shell 405, which has a space flight heritage as a hydrazine catalyst. The decomposition process is exothermic resulting in nitrogen and oxygen at 2988° F., for complete decomposition. This hot oxidizer will autoignite propane (and most hydrocarbon fuels) on contact and will facilitate sustained combustion in a rocket combustion chamber. Using this technique, autoignition and rigorous and complete combustion can be accomplished using stable, non-toxic, storable propellants. Along the same lines nitrous oxide could be decomposed and used as a monopropellant in a similar fashion as hydrazine and hydrogen peroxide rocket systems. Although the NOP rocket concept deals specifically with nitrous oxide and propane as propellants, the greater invention is one of using nitrous oxide as an oxidizer and ignition source for use alone as a monopropellant or with a fuel as a bipropellant. 
     Nitrous oxide has added benefit as a space propellant in that it stores as a liquid and injects as a gas. This is important for attitude control (AC), since liquid injection rockets cannot provide the shorter pulse times required for an ACS mission. Liquid storage gives tremendous weight benefit since a liquid tank can hold many times its own weight in propellant, whereas the same is not true for gas storage systems. 
     Due to the versatility of nitrous oxide as both a monopropellant and oxidizer for a bipropellant system, the potential exists to set up a multi-mode propulsion system, which will improve space mission capability by reducing the dry weight overhead. A single propellant system serving all space propulsion missions from attitude control to orbital maneuvering would reduce the component count, system weight, and cost. Lower dry weight can be converted into higher payload weight fraction or ΔV. 
     Most propellants commonly used today have relatively low vapor pressure (lower than the rocket chamber pressure) and consequently require a separate expulsion system. In contrast to these systems, the NOP propellants are self-pressurizing due to their relatively high vapor pressures (higher than the rocket chamber pressure). Consequently, they do not require separate expulsion systems and the entire tank volume can be used to store propellant. The vapor pressure of nitrous oxide is approximately 750 psi and that of propane is 110 psia at ambient temperature. 
     The vapor pressure of propane is slightly low for it to be used as a true self-pressurizing propellant. For a NOP rocket system, high-pressure nitrous oxide vapor could be used as a pressurant gas for the propane by, for example, means of a bladder or diaphragm. Ethylene may also be appropriate as a propane replacement due to its higher vapor pressure. 
     The NOP rocket offers a non-toxic, environmentally benign propellant combination that is storable in space over long periods of time and offers comparable specific impulse to current systems. The non-toxic nature of the NOP propellants will serve to reduce operating costs due to the handling issues associated with the hypergolic propellants currently in use for space applications. The NOP propellants are benign and not highly reactive. They remain so until the nitrous oxide is catalytically decomposed and combined with the fuel. Exhaust products are mainly nitrogen, water, and carbon dioxide. Another beneficial feature of the NOP rocket propellants are that they are storable over long periods of time without degradation. For example, inadvertent decomposition is one of the main technical obstacles for hydrogen peroxide use in space propulsion systems. 
     Tests have been conducted of a NOP rocket utilizing nitrous oxide (N 2 O) as the oxidizer and propane (C 3 H 8 ) as the fuel. The chamber pressure was 150 psia for this 50 lb f  rocket, and a fuel-rich propellant combination was used to minimize combustion instabilities. According to common practice for space-based thrusters, a pressure drop between 20-30% of chamber pressure is taken across the injectors, requiring an injection pressure of approximately 180 psia in this example. 
     The injector design for a NOP rocket engine prototype, shown in FIG. 1, is a single element, coaxial (single) swirl injector, with liquid injection of propane and two-phase phase injection of nitrous oxide. Swirl tends to increase mixing and decrease the required characteristic combustor length L* and is generated by tangentially injecting the propellant off-center, with respect to the orifice through-hole. In general, swirl is better for mixing two flows and more simple than a showerhead injector design and swirl injectors operate over a wider range of conditions and are more forgiving than other types of injectors. 
     As shown in FIG. 1, the liquid C 3 H 8  fuel is injected into an inside orifice  100  and the oxidizer is injected into an outside annulus  150 , between a C 3 H 8  injector tube  120  and a sintered mesh disc  180 . Initial consideration was given to a design with the fuel on the outside, coating the hot combustor walls (made of Glid-Cu, a Cu-0.15% alumina alloy) thus protecting the inner walls from oxidation. However, injection of liquid propane in a narrow annulus is not preferred, since the liquid propane surface tension may lead to asymmetric injection. Although copper is not prone to oxidation, refractory metals such as columbium (commonly used for space applications) and tungsten are indeed vulnerable to oxidation. 
     In this example, the liquid propane injector is sized for an injector pressure drop of 30%P c , and the gaseous nitrous-oxide injector is sized for a pressure drop of 20%P c . The liquid propane injector is designed for a nominal pressure drop of 45 psid at the orifice. For a nominal flow rate of 0.0149 lb m /sec, at injection conditions of 70° F. and 195 psia, an orifice diameter of 0.032″ is used. A ¼″ S.S. tube (0.194″ I.D.) feeds into the 0.032″ propane injector, which is offset 0.043″ from the center of a 0.118″ I.D. tube, generating a swirl component. The liquid propane is then injected into the combustor from the 0.118″ I.D. tube at about 6 ft/sec. 
     The N 2 O is fed through a ⅜″ tube (0.305″ I.D.)  130 , and into a stagnation chamber  140 , where the N 2 O flow turns into the sintered mesh disc  180 . The sintered mesh in this example is stainless steel and has a one micron porosity, sized to pass the required flow rate with the required pressure drop. The 0.118″ I.D. propane tube is preferably at the center of the sintered mesh disc. Approximately 75% of the N 2 O flows through the annulus (with an area of ˜0.039 in 2 ) between the mesh center hole and the liquid propane injector tube. Approximately 25% of the N 2 O will flow through the porous sintered mesh, providing for transpiration cooling. The N 2 O injector is designed for a nominal pressure drop of 30 psid across the stainless steel sintered mesh material. The nominal N 2 O flow rate through the annulus is 0.149 lb m /sec, at −25° F. and 175 psia. In this example, the N 2 O flows over catalyst  110  before entering stagnation chamber  140 . however, the N 2 O can come in contact with a catalyst at any point before the N 2 O is mixed with the fuel (in this example, the C 3 H 8 ). For example, sintered mesh disc  180  can be made from a catalyst and would, therefore, serve dual purposes of creating a pressure drop and acting as a catalyst for the decomposition of the N 2 O. 
     Work has been performed using an existing atmospheric test stand. A new rocket test stand was designed and constructed for rocket performance and rocket ignition testing. This new rocket test stand features palletized propellant systems, improved propellant system instrumentation, an improved, more robust thrust stand, and adequate room for the rocket exhaust survey and radiometric measurement equipment. 
     The palletized propellant system is shown in FIG.  2 . It is noted that both the nitrous oxide and propane systems are similar in layout, except for minor differences in venturi size, metering valve size, and storage tank volume. Both systems will be discussed with reference to FIG.  2 . 
     The basic idea behind the propellant system layout is to provide the user with a safe, modular and self-contained process for loading and pressurizing N 2 O and C 3 H 8  for use in the ignition circuit (spark ignition or catalyst reactor) and rocket engine propellant feed-systems. The propane and nitrous oxide are first loaded into their respective run tanks  210 . Tank  200  is a nitrogen pressurization tank, that is used to further pressurize the N 2 O and C 3 H 8  run tanks  210 , to ensure that both propellants are in the liquid state at least through the venturi, thus assuring accurate mass flow rate measurements. The propellants experience a large pressure drop through the metering valve, (ΔP˜400-700 psi), which adjusts the flow to provide the required ΔP across the injector. A nitrogen purge circuit is also used to purge the lines before and after the rocket firing sequence is performed. The rocket is mounted on a thrust stand, and is connected to the propellant system by flex lines. 
     Catalyst research showed that the Shell  405  catalyst successfully decomposes nitrous oxide with moderate light-off temperatures. Shell  405  catalyst has a history of use in space as a hydrazine catalyst material. However, it is recognized that Shell  405  is adverse to repeated use with an oxidizer such as nitrous oxide, and research shows staged catalyst beds containing other elements are preferable for an engine where multiple firings are required. 
     Nitrous oxide decomposes exothermically with adiabatic decomposition temperature reaching ≈1640° C., (2984° F.). This decomposition is accelerated by a catalyst. Free oxygen available by nitrous oxide decomposition can then be combusted with a wide variety of fuels, with or without the continued supply of decomposed nitrous oxide from the catalyst, as the reaction becomes self-sustaining after initial ignition due to the continued release of heat from combustion. A preferred chemical reaction for the decomposition of nitrous oxide results in the formation of nitrogen and oxygen according to the following reaction equation. 
     
       
         N 2 O( g )→N 2 ( g )+½ O 2 ( g )+Energy  (1). 
       
     
     However, heat input is usually required to initiate the decomposition reaction. In the case of thermal decomposition, the activation energy barrier for nitrous oxide is about 250 kJ/mole. There are other intermediate chemical reactions that can lead to oxides of nitrogen, such as NO and NO 2 , that are undesirable if complete decomposition is to be achieved. 
     In order to attain homogenous reaction rates, the gas is heated above its auto-decomposition temperature, unless a heterogeneous surface such as a catalyst is incorporated. Catalysts are designed to lower the activation energy barrier, thus allowing the decomposition to occur at much lower temperatures. The principal catalytic action can originate from charge donation into the antibonding orbitals, weakening the N—O bond and thereby lowering the activation energy and thus the reaction temperature. 
     Various catalyst combinations were tested, over a range of initial pressures and reactor lengths. These catalysts include a platinum monolith, granular tungsten carbide, rhodium (0.17% granular), gold, platinum/palladium monolith, iridium (granular) and a Shell 405-Ir based catalyst bed. Each of these catalysts was preheated to various temperatures (122° F., 303° F., 398° F., 401° F. and 662° F.) using a linear temperature controller to determine the minimum light-off temperature. Catalyst activation requires a minimum initial temperature, with activation increasing as temperature increases. Instabilities occur at temperatures above a given value, which is material dependent. 
     Chemisorption experiments were conducted to assess the adsorption characteristics of various candidate catalyst materials, a property critical in a heterogeneous catalysis process, as is the case for the NOP rocket catalyst reactor. Conclusions drawn from the relative adsorption tests on Shell-405 and Co-ZSM-5 catalysts, lead to a decision to perform reactive flow studies in an experimental reactor. 
     Once the adsorption characteristics of the various candidate catalyst materials were understood, the candidate catalyst materials were tested inside a reactor, in order to measure the activity of the catalyst and gain the information required to build a working reactor for the NOP rocket ignition system. The results from these experimental runs suggest that the iridium based catalyst Shell-405, had the highest activity and selectivity towards nitrous oxide decomposition. Experimental results also indicate that cobalt based ZSM-5 catalysts with sodium as promoter metal produced high activity toward the thermal decomposition of nitrous oxide in the presence of a propane/propylene mix. 
     N 2 O catalytic decomposition is achievable at 400° F. for pure nitrous oxide flowing over Shell 405, and, with the use of trace amounts of a hydrocarbon (e.g. Propane or propylene), this temperature is lowered to approximately 200° F. 
     Shell-405 catalyst was loaded into the atmospheric combustion igniter, resulting in a bed length of about ¼ inch. The reactor was preheated to 148° C. at the inlet. Nitrous oxide at a gage pressure of 117 psig was then flowed through the reactor at 0.015 ACFM. A time delay of about 4 seconds was allowed before the propane/propylene mix was injected into the reactor at 103 psig pressure and a flow rate of about 1 cc/sec. At that instant a flame was observed at the exit of the reactor. The flow of the propane mix and nitrous oxide were then shut-off and the reactor purged with N 2 . The flows were then restarted and a flame was again observed at the exit. This procedure was repeated 6 times and each time the propane mix was turned on a flame was observed. This reactor was fired approximately 15 times in rapid succession, displaying robustness in the catalyst reactor operation. 
     For comparison purposes, plots of the theoretical vacuum I sp,vac, I sp , and c*, obtained from a NASA chemical equilibrium code, are presented. Initial results are promising, showing rocket performance consistent with theoretical predictions taking into account the effects of heat transfer. 
     FIG. 3 shows the variation in the thrust coefficient for a test rocket (evaluated on Test Stands 1 and 2) as a function of mixture ratio for two combustor lengths, L*=2 m and 3 m. The average thrust coefficient measured on Test Stand 1 is 1.21, compared with a theoretical value (neglecting heat losses) of 1.27, and a measured value of 1.14 on Test Stand 2, compared with a theoretical value (neglecting heat losses) of 1.26. 
     A rocket test stand facility, equipped with palletized propellant feed systems, 1000 lbs f  thrust stand, and data acquisition systems, was built to test a nitrous oxide/propane (NOP) rocket engine. The NOP rocket was tested over a range of mixture ratios (4.89&lt;M.R.&lt;8.68). An ignition concept using catalytically decomposed nitrous oxide to autoignite propane, was explored and various catalyst materials were evaluated. Shell-405 and cobalt based ZSM-5 showed promising reactivity, demonstrating sufficient decomposition of N 2 O to ignite hydrocarbon fuels. Laboratory experiments with the catalyst reactor have shown that N 2 O catalytic decomposition is achievable at 400° F. for pure nitrous oxide flowing over Shell 405, and, with the use of trace amounts of a hydrocarbon (eg. propane or propylene), this temperature is lowered to approximately 200° F. 
     In addition to developing the catalyst ignition system, NOP rocket performance was experimentally determined to match well with theoretical predictions, with proper modeling of heat losses. Radiometric measurements were also used to determine rocket exhaust temperature and plume composition and plume pitot probe measurements provided another method for verifying thrust data. 
     FIG. 4 shows a schematic example of a rocket engine  1200  in accordance with an embodiment of the invention. In FIG. 4, rocket engine  1200  has a nozzle  1210  and a propellant supply system  1220 . The propellant supply system supplies and ignites a fuel and nitrous oxide in accordance with the invention discussed above. 
     While the invention has been described with reference to particular embodiments and examples, those skilled in the art will appreciate that various modifications may be made thereto without significantly departing from the spirit and scope of the invention.