NITROUS OXIDE FUEL BLEND MONOPROPELLANTS

Abstract

Compositions and methods herein provide monopropellants comprising nitrous oxide mixed with organic fuels in particular proportions creating stable, storable, monopropellants which demonstrate high ISP performance. Due to physical properties of the nitrous molecule, fuel/nitrous blends demonstrate high degrees of miscibility as well as excellent chemical stability. While the monopropellants are particularly well suited for use as propulsion propellants, they also lend themselves well to power generation in demanding situations where some specific cycle creates useable work and for providing gas pressure and/or heat for inflating deployable materials.

Description

BACKGROUND

Liquid fueled rockets have better specific impulse (I_sp) than solid rockets and are capable of being throttled, shut down and restarted. The primary performance advantage of liquid propellants is the oxidizer. Several practical liquid oxidizers (liquid oxygen, nitrogen tetroxide and hydrogen peroxide) are available that have much better I_sp than the ammonium perchlorate used in solid rocket boosters when paired with comparable fuels. However, the main difficulties with liquid propellants also are with the oxidizers. Oxidizers are generally at least moderately difficult to store and handle, either due to extreme toxicity (nitric acids), moderate cryogenicity (liquid oxygen) or both (liquid fluorine). Several oxidizers that have been proposed, for example, O₃, ClF₃, ClF₅, are unstable, energetic and toxic.

The first liquid-fuelled rocket—launched in the 1920s—used gasoline and liquid oxygen as propellants. Liquid hydrogen was used in the 1950s, and by the mid-60s, liquid hydrogen and liquid oxygen were being used. Common liquid monopropellants in use today include hydrazine and hydroxyl ammonium nitrate. Common liquid bipropellants include liquid oxygen and kerosene, liquid oxygen and liquid hydrogen, and nitrogen tetroxide and hydrazine or monomethylhydrazine. A goal of propellant design has been to develop a monopropellant having the high performance characteristics of a bipropellant. Due to the simplified system architecture of monopropellant systems, finding monopropellant chemistry that provides bipropellant-like I_sp performance has long been considered a “holy grail” in monopropellant development. Research in the field of “green” monopropellants has been ongoing to find non-toxic monopropellant alternatives to hydrazine. One such candidate is nitrous oxide. Nitrous oxide can be decomposed through the following exothermic reaction:

N₂ON₂+½O₂+Heat

Under standard conditions, this reaction generates 82 kJ/mol (515 Whr/kg) of heat per unit nitrous oxide. To liquefy the stored monopropellant requires 16.5 kJ/mol (104 Whr/kg) or approximately 20% of the enthalpy of reaction. The maximum theoretical I_sp of this reaction is 205 s. N₂O is a highly stable molecule given its high activation energy barrier ˜250 kJ/mol. As a result, thermal decomposition requires preheat temperatures >1000° C. Alternatively, catalysts can be used to significantly depress this activation energy. However, the hot (>1500° C.), highly oxidizing reaction products make catalyst bed and reaction chamber design challenging.

With the addition of hydrocarbon fuel to the reaction in the equation above, the specific energy density of liquid monopropellants can be increased up to ˜1500 Whr/kg (˜3 times the energy density of pure N₂O), and I_sp performance greater than 300 s becomes feasible. Furthermore, the hot deleterious oxygen in the exhaust stream can be consumed and the higher combustion reaction temperatures result in faster reaction kinetics as compared to pure N₂O decomposition. The faster kinetics permit rapid spark ignition. In such a case, a catalyst bed does not become the material limitation for engine design, and regeneratively cooled engine design approaches with conventional materials can be adapted for the higher I_sp performance using low cost engine fabrication techniques.

The highest I_sp chemistry ever test-fired in a rocket engine was lithium and fluorine, with hydrogen added to improve the exhaust thermodynamics. The combination delivered 542 seconds specific impulse in a vacuum. However, the impracticality of this chemistry highlights why exotic propellants, particularly bipropellants, are not used in practice. To make all three components liquids, the hydrogen must be kept below −252° C. and the lithium must be kept above 180° C. This example demonstrates dramatically a major drawback of bipropellants—they must be stored in separate tanks (and often under different temperature and/or pressure conditions), and they must be delivered to the combustion chamber at a pre-defined and specific mix ratio, typically at high pressure and high flow rates.

SUMMARY

Implementations described and claimed herein address the foregoing issues with a family of nitrous oxide fuel blend (NOFB) monopropellants comprising organic fuels mixed with nitrous oxide (N₂O). When combusted, the nitrous oxide provides both thermal decomposition energy and serves as the oxidizer to combust the fuels. Example organic fuels include ethane (C₂H₆), ethylene (C₂H₄), acetylene (C₂H₂), and mixtures thereof. Mixtures of these fuels based on oxidizer-to-fuel ratio (O/F) generate desired monopropellant characteristics including, but not limited to, I_sp, miscibility over a wide temperature and pressure range, favorable fluid handling performance, low freezing points, rapid combustion kinetics for fast engine response times, relatively high thermal decomposition limits, low mechanical shock sensitivity and impact-induced detonation, relatively high storage densities, and exhaust gas chemistries that do not produce carbon fouling or hot oxidizing environments that are difficult, if not impossible, to accommodate with combustor or reaction chamber design materials. In addition, to being a very stable oxidizer, nitrous oxide is a very good solvent with a near room temperature critical point at 36.4° C. Therefore, it is possible to dissolve fuels into the N₂O to produce a nitrous oxide fuel blend (NOFB). Care must be taken in the design of the NOFB monopropellant to ensure that the mixture is safe to handle, and that the NOFB monopropellant maintains balanced degassing of all NOFB constituents over a wide range of temperatures and tank drawdown profiles in which it may be used.

Implementations herein provide a nitrous oxide fuel blend (NOFB) monopropellant comprising nitrous oxide and an organic compound in an oxidizer-to-fuel ratio of about 2.5 to about 11.0. Preferably, the organic compound comprises, as a main component, a C2 hydrocarbon, or mixtures of C2 hydrocarbons. Specifically, implementations provide a monopropellant comprising nitrous oxide and acetylene in an oxidizer-to-fuel ratio of about 2.5 to about 11.0, or about 3.0 to about 9.0, or about 4.0 to about 8.0, or about 4.5 to about 7.5, or about 2.5 to about 6.0, or about 3.0 to about 5.0, or about 6.0 to about 11.0, or about 8.0 to about 10.0. Other implementations provide NOFB monopropellants comprising nitrous oxide and ethane in an oxidizer-to-fuel ratio of about 2.5 to about 11.0, or about 3.0 to about 9.0, or about 4.0 to about 8.0, or about 4.5 to about 7.5, or about 2.5 to about 6.0, or about 3.0 to about 5.0, or about 6.0 to about 11.0, or about 8.0 to about 10.0. Yet other implementations provide NOFB monopropellants comprising nitrous oxide and ethylene in an oxidizer-to-fuel ratio of about 2.5 to about 11.0, or about 3.0 to about 9.0, or about 4.0 to about 8.0, or about 4.5 to about 7.5, or about 2.5 to about 6.0, or about 3.0 to about 5.0, or about 6.0 to about 11.0, or about 8.0 to about 10.0. The ratios are chosen for specific uses. For instance, NOFB34 is optimized for small rocket engines (fast combustion kinetics and optimized peak Isp with frozen-at-the-throat combustion kinetics), and NOFB37 is optimized for large rocket engines (higher density monopropellant with Isp optimized for slower combustion kinetics in larger rocket diverging exhaust nozzles). In other implementations, the NOFB monopropellants may comprise other compositions or additives up to about 50% of the monopropellant, or up to about 40% of the monopropellant, or up to about 30% of the monopropellant, or up to about 20% of the monopropellant.

The other compositions include hydrocarbon fuels or mixtures thereof wherein the resulting monopropellant has the property that as the monopropellant is drawn down or the temperature changed, the balanced blend has minimal variation in liquid and ullage gas mixture-ratio chemistry as the liquid monopropellant boils-off to generate ullage gas under these conditions. The additional hydrocarbon fuels may cause a <10% variation in rocket Isp performance due to these variations in boil-off rates for the different NOFB constituents. For some applications, the addition of small amounts of detergents, emulsifiers, or other additives may be advantageous.

Additional implementations of the technology provide NOFB monopropellants comprising nitrous oxide and two or more of acetylene, ethane or ethene in an oxidizer-to-fuel ratio of about 2.5 to about 11.0, or about 3.0 to about 9.0, or about 4.0 to about 8.0, or about 4.5 to about 7.5, or about 2.5 to about 6.0, or about 3.0 to about 5.0, or about 6.0 to about 11.0, or about 8.0 to about 10.0. In other implementations, the monopropellant may comprise other compositions or additives up to about 50% of the monopropellant, or up to about 40% of the monopropellant, or up to about 30% of the monopropellant, or up to about 20% of the monopropellant.

In certain implementations, the nitrous oxide is in a gas phase when mixed with the fuel during manufacturing; in other implementations the nitrous oxide is in a liquid phase when mixed with the fuel during manufacturing; and in yet other implementations, the nitrous oxide is in a mixed gas/liquid phase when mixed with the fuel during manufacturing. The mixing is done as described in Example 1. This Summary is provided to introduce a selection of concepts in a simplified form that are further described below in the Detailed Description. This Summary is not intended to identify key or essential features of the claimed subject matter, nor is it intended to be used to limit the scope of the claimed subject matter.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a graph of theoretical and actual I_sp performance of an NOFB monopropellant formulation.

FIG. 2 illustrates the method of making the nitrous oxide fuel blends of the present invention.

FIG. 3 is a chart summarizing NOFB monopropellant characteristics relative to monopropellant hydrazine and bipropellant nitro tetroxide/monomethylhydrazine.

FIG. 4 is a graph illustrating storage characteristics (storage tank liquid and gas pressure and density versus temperature) for one NOFB monopropellant formulation (also known as a phase diagram). In FIG. 4, the NOFB monopropellant storage characteristics are also compared to pure nitrous oxide liquid and tanked hydrazine monopropellant including a typical helium pressurant load for hydrazine

FIG. 5A is an FTIR spectrum of NOFB monopropellant sampled over different tank temperatures (and comparison with the remaining gas after a ¼ tank rapid liquid expulsion) illustrating the stability of the NOFB chemical mixture to biased constituent outgassing over extreme temperature ratios. FIG. 5B similarly shows the variation in NOFB O/F ratio in the liquid and ullage gas (gas in tank with liquid) of three NOFB blends after rapid expulsion. On the right hand side of this figure, the corresponding variation in I_sp performance is also shown as the blend slightly varies during a very aggressive tank liquid expulsion (80% NOFB liquid expulsion in ˜seconds)

FIG. 6 is a graph illustrating exemplary nozzle coefficient values for use in vacuum equivalent I_sp calculations.

FIG. 7A is a graph showing thermal decomposition data for one exemplary NOFB monopropellant formulation. FIG. 7B is a summary of decomposition tests vs. NOFB pressure for an exemplary NOFB monopropellant.

FIG. 8A is a graph illustrating the specific enthalpy of vaporization of one exemplary NOFB monopropellant relative to nitrous oxide and compared to the specific energy for heating a representative quantity of bipropellant fuel from a the same temperature to ˜300° C. FIG. 8B illustrates the rapid decrease in temperature as the NOFB monopropellant is throttled or “flash-cooled” by forcing it through a pressure drop.

FIG. 9 is a graph illustrating the maximum spark propagation distance for pure nitrous oxide as a function of gas pressure at the minimum nitrous oxide spark voltage of 418 V. At this minimum voltage point (also known as Paschen curve minimum), for a given gap distance, both higher and lower gas pressure requires rapidly increasing spark voltages.

FIG. 10 is a graph of quenching distance based on oxidizer-to-fuel ratios for one NOFB monopropellant formulation.

FIG. 11 illustrates an exemplary NOFB regeneratively-cooled thruster utilizing the high volatility of the NOFB monopropellant to “flash-cool” the combustion chamber.

FIG. 12 illustrates low thrust, non-optimized engine run test data in an engine utilizing an exemplary NOFB monopropellant.

FIG. 13 is a graph illustrating a comparison of delivered payload mass of total wet mass rocket propulsion system performance of an exemplary NOFB monopropulsion system relative to hydrazine systems.

FIG. 14 summarizes the characteristics of the NOFB deployable wing spars.

DETAILED DESCRIPTION

Technology is described herein for providing a nitrous oxide fuel blend (NOFB) monopropellant comprising nitrous oxide with an organic compound, such as one or more of acetylene, ethane, or ethene resulting in a monopropellant that has a high specific impulse, low toxicity and allows for easy storage and handling in addition to other desired characteristics. The monopropellant may be used in some implementations for rocket propulsion, working fluid production, or energy or gas generation.

Before the present formulations and methods are described, it is to be understood that the invention is not limited to the particular formulations or methodologies described, as such, formulations and methods may, of course, vary. It is also to be understood that the terminology used herein is for the purpose of describing particular embodiments only, and is not intended to limit the scope of the present invention; the scope should be limited only by the appended claims. It must be noted that as used herein and in the appended claims, the singular forms “a,” “an,” and “the” include plural referents unless the context clearly dictates otherwise. Thus, for example, reference to “an agent” refers to one agent or mixtures of agents, and reference to “the method of manufacturing” includes reference to equivalent steps and methods known to those skilled in the art, and so forth.

Unless defined otherwise, all technical and scientific terms used herein have the same meaning as commonly understood by one of ordinary skill in the art to which this invention belongs. All publications mentioned herein are incorporated herein by reference for the purpose of describing and disclosing devices, formulations and methodologies that are described in the publication and that may be used in connection with the claimed invention, including related U.S. application Ser. No. 60/868,523, filed Dec. 4, 2006 entitled “Injector Head.”

Where a range of values is provided, it is understood that each intervening value, between the upper and lower limit of that range and any other stated or intervening value in that stated range is encompassed within the invention. The upper and lower limits of these smaller ranges may independently be included in the smaller ranges and are also encompassed within the invention, subject to any specifically excluded limit in the stated range. Where the stated range includes one or both of the limits, ranges excluding either or both of those included limits are also included in the invention.

The art of chemical rocket propulsion makes use of controlled release of chemically reacted or un-reacted fluids to achieve thrust in a desired direction. The thrust acts to change a body's linear or angular momentum. Similar to rocket propellants that have found application in other working fluid production and power generation applications, the claimed invention may be utilized in many alternative types of applications as well, including gas generation for inflation systems and inflatable deployments, in systems used to convert thermal energy in hot exhaust gases to mechanical and electrical power, and in high energy storage media for projectiles, munitions, and explosives. Examples where the claimed technology could be applied specifically include earth-orbiting spacecraft and missile propulsion systems; launch vehicle upper stage propulsion systems and booster stages; deep space probe propulsion and power systems; deep space spacecraft ascent and earth return stages; precision-controlled spacecraft station-keeping propulsion systems; human-rated reaction control propulsion systems; spacecraft lander descent propulsion, power, and pneumatic systems for excavation (NOFB monopropellant can be used to both provide mechanical power to run drills in extraterrestrial drilling applications and to provide gases to remove debris from the area of the cutting surfaces), spacecraft pneumatic science sample acquisition and handling systems; micro-spacecraft high performance propulsion systems; military divert and kill interceptors; high altitude aircraft engines, aircraft backup power systems; remote low temperature power systems (e.g., arctic power generators); combustion powered terrestrial tools including high temperature welding and cutting torches as well as reloadable charges for drive mechanisms (e.g., nail guns, anchor bolt guns), and the like. In terrestrial applications, NOFB monopropellants can provide power in situations where atmospheric oxygen in not in sufficient quantity to provide an oxidizer for combustion reactions (such as very high altitude aircraft powerplants or underwater equipment). Moreover, there are many derivative applications related to using combustion stored energy.

A monopropellant is a single fluid that typically is used for generating thrust, gas generation, and/or power (mechanical and/or electrical) generation. Monopropellants commonly undergo exothermic chemical reactions through a catalytic, hypergolic, or spark ignition mechanism in order to release additional heat energy (commonly providing an ideally low molar mass exhaust gas as well) in order to increase mass efficiency in generating thrust and power. Monopropellants, for example, can be used in a liquid or gas rocket engine. A common example of a monopropellant is hydrazine, often used in spacecraft propulsion for vehicle translation maneuvers (linear momentum changes) and attitude control (angular momentum changes). Another example of a monopropellant is hydroxyl ammonium nitrate (HAN) which is currently being investigated as a lower toxicity monopropellant alternative to hydrazine.

Additionally, a working fluid that has a pressure gradient between it and the surrounding environment is capable of producing mechanical work/power. This mechanical work/power can subsequently be converted into alternative energy forms (for example, electric power generation, mechanical shaft power can be used to power an electric generator or alternator to provide electric power). Pressure from either the natural vapor pressure of an NOFB monopropellant and/or through NOFB monopropellant decomposition/combustion processes in combination with NOFB monopropellant-derived working fluids can be used strategically to produce useable work beyond simple thrust. Example work extracting cycles that can implement the NOFB monopropellants may include, without limitation, gas turbine cycles (e.g., Brayton or similar cycles), constant pressure expansions of combusted monopropellant (similar to pneumatic machines), and various piston cycle engines including but not limited to spark-ignited Otto cycles, and compression-ignited Diesel cycles. The maximum energy that can be extracted from a chemical medium is related to its specific energy density (stored chemical energy per unit mass). As shown in FIG. 3, the specific energy density of NOFB liquid monopropellant (>1300 Whr/kg) is ˜3.5 to 3.9 times greater than hydrazine. For comparison, state-of-the-art lithium ion batteries store ˜145 Whr/kg. The NOFB propulsion system would require additional mass that would effectively lower the NOFB monopropellant's specific energy. Nevertheless, for many primary power applications not requiring energy recharge, the very high specific energy density of NOFB monopropellants is desirable.

For the specific case of rocket propulsion, a variety of metrics determine how efficiently a particular rocket propulsion system performs. One of the most important metrics in rocket propulsion is specific impulse (I_sp). This metric essentially measures the amount of total impulse or imparted momentum change (integrated force over time) produced by a given propulsion system divided by the total mass of propellant consumed. This result is normalized by the earth's gravitational constant (9.81 m/s²) such that I_sp has units of seconds regardless of what international system of units are being used (English or System Internationale (SI) units)). Higher I_sp values indicate greater ability to impart velocity changes to vehicles for a given amount of propellant consumed. By crude analogy, Isp performance is similar in connotation to “miles per gallon” in a combustion-powered car engine (although one caveat here is that more engine-specific characteristics go into defining “miles per gallon” for a car as compared to rocket propulsion for a spacecraft). Because 1) mass is extremely expensive to launch, and 2) there is an exponential dependency of propellant mass on I_sp performance [Propellant Mass=Spacecraft_Dry_Mass×exp[Change_in_Spacecraft_Velocity/I_sp/earth gravity]−1], high I_sp propellants are very attractive for demanding aerospace applications. In chemical propulsion systems, in order to achieve high I_sp systems, exothermic chemical reactions are generally required. Currently, a common industry standard commercial monopropellant, hydrazine, has an I_sp of around 230 s (slight deviations of this number are dependent on specific thruster design parameters). The class of NOFB (nitrous oxide fuel blend) monopropellant formulations disclosed herein can achieve engine I_sp values of up to 345 s and potentially larger I_sp values. Recent experimentally measured engine I_sp values exceed 300 s (see FIG. 1). FIG. 13 compares the dry-spacecraft-loaded-mass-ratio vs. required vehicle velocity changes for NOFB monopropellants as compared to hydrazine.

EXAMPLE 1

The mixing of fuels and oxidizer must be done in a controlled, measured manner to ensure the resulting monopropellant has desired performance characteristics.

FIG. 2 demonstrates an exemplary schematic of an apparatus used to manufacture NOFB monopropellant blends. Thruster performance is dependent on the propellant which is combusted. For this reason it is generally important to accurately mix the monopropellant blends. A specialized apparatus can be used to mix high vapor pressure monopropellants. Essentially, the constituents can be mixed in their vapor phases and condensed in a separate container to form a high density liquid monopropellant. The method and apparatus outlined below are for exemplary purposes, and derivations hereof may be equally acceptable manufacturing methods. In this implementation, SW-# indicates a general on/off valve, REG-# indicates a pressure reducing regulator, and IS-#s are tank isolation valves. Unless otherwise noted, all valves begin closed and regulators backed completely off. A pressure transmitter is attached to an open SW-5 valve to accurately monitor system pressure. To begin manufacturing, the system is purged of air by turning the vacuum pump on, opening IS-3, IS-4, SW-6, and SW-8. Once an adequate vacuum is accomplished, SW-8 is closed. Next, fuel(s) is/are added to both the mixing tank and the condensing tank. To do this, the IS-2 is opened, REG-2 is increased to the desired pressure, and SW-7, SW-2, and SW-4 are all opened. Depending on the vacuum pulled in the previous step, purges may be required. To determine the necessity of purges, the purity can be calculated by taking the ratio of fuel added absolute pressure to the total absolute pressure in the tank. For example, if a vacuum were pulled to 1 psia, and fuel were loaded to 100 psia, the purity of the load would be 99 psia/100 psia or 99%. A purge can increase the purity beyond the initial load if required. To purge, SW-4 is closed, and SW-8 is opened so as to draw the mixture out of the system. However, when running mixed combustibles through a mechanical pump system, adequate flashback mitigation measures should be implemented. Once adequate vacuum is achieved, SW-8 is closed. The purge sequence can be repeated as required for the desired purity of the mixture. The new purity is calculated by multiplying the impurity levels of each load together. For example, if another 99% pure load is added the 1% impurity is multiplied by the new 1% impurity to result in 0.01% impure or 99.99% the purity level of the starting fluid. However, if the starting fluid is only 98% pure, no amount of purging can increase purity levels above the initial 98% fluid purity. Once the purity and load pressure is achieved in the system, SW-4 is closed. The fuel is then shut off and purged from the system by closing IS-2 and opening SW-3. Adequate time is allowed to vent the fuel from the lines and REG-2 is backed out, SW-7 is closed, SW-2, and SW-3 are closed. If multiple fuels are used, the secondary/tertiary fuel is added at this point on top of the prior load (FIG. 2 does not show this option). Once the fuel blend has been achieved, nitrous oxide is added. Here, IS-1 is opened, REG-1 is increased, and SW-1 and SW-4 are opened. Once the desired mixture is achieved, SW-4 is closed. To vent the nitrous oxide, IS-1 is closed, SW-3 is opened, the system is allowed time to drain, REG-1 is backed off, and SW-1 and SW-3 are closed. At this point the correct NOFB blend has been manufactured, so the condensing tank is placed in a cold bath adequately cold to condense the mixture but sufficiently above the blend's freezing point. One implementation uses a cold bath sustained at ˜−70 C. Sufficient time is allowed for the mixture to condense, and IS-4 and SW-6 are closed. If sufficient monopropellant is manufactured with one condensation, the condensed liquid tank can be removed from the system (between IS-4 and SW-6) and allowed to equilibrate back with room temperature. If multiple loads are required, the previous steps can be repeated, with the exception that gases are only mixed in the mixing tank and condensing consists of opening SW-6 and IS-4.

EXAMPLE 2

Candidate fuel blends were made and tested. The most promising blends were selected based on the following criteria: combustion and theoretical engine performance; propellant stability; equilibrium and non-equilibrium miscibility performance; combustion limits, flame temperature, and exhaust gas chemistry for engine design; propellant phase diagram properties, and combustion reaction rates.

The monopropellants of the present invention are named in the following manner. “NOFB” designates nitrous oxide fuel blend. The next number designates the place in the C2 group; 1 is ethane, 2 is ethylene, and 3 is acetylene. The next number indicates the oxidizer to fuel ratio. Thus, “NOFB34” is nitrous oxide blended with acetylene with an oxidizer to fuel ratio of 4. Additional letters (a, b, c) after the oxidizer to fuel ratio number may be used to describe deviations in the blend. For example, an NOFB34 blend may include small amounts of specific additives to improve mixture chemistry degassing characteristics. The first discovered adaptation to this blend beyond the basic nitrous oxide and fuel chemistry would therefore be denoted NOFB34a.

FIG. 1 illustrates the theoretical I_sp performance of a nitrous oxide/acetylene (N₂O/C₂H₂) monopropellant blend as a function of oxidizer-to-fuel (O/F) mass ratio as well as showing data from recent prototype engine test results based on measuring integrated chamber pressure and propellant mass consumed during an engine run. (Additional details on the particular experimental method used for acquiring the experimental measurement are discussed in [0043] below). The experimentally measured I_sp was acquired for an O/F ratio of 4 (errors bar based on uncertainty in actual nozzle coefficient during terrestrial testing). The two sets of theoretical curves (vacuum and 200/1) are shown for two different cases, equilibrium and frozen-at-the-throat chemical kinetics. These are typical bounding scenarios for actual rocket engine performance in space applications. The vacuum condition is from an ideal exit nozzle that is infinitely long. The 200/1 nozzle is a more realistic diverging nozzle scenario where the exit plane area is 200 times larger than the minimum throat area of the nozzle. The equilibrium chemical kinetics scenario is one bounding scenario that assumes that the flow moves slowly enough to allow the hot gases to always maintain chemical equilibrium of the exhaust constituents (i.e. the exhaust gas chemistry changes to match the cooling conditions in the nozzle). The frozen-at-the-throat scenario assumes that the gases in the diverging nozzle immediately downstream of the throat cool so rapidly that the chemical kinetics “freeze” (i.e., the chemistry of the gas does not change) such that the gas constituents remain constant in the diverging nozzle of a rocket thruster downstream of the throat.

In addition, when the O/F ratio is altered, the chemical combustion performance is altered within the combustion chamber. By altering the chemical reactions taking place within the chamber, different I_sp performance is achieved. However, certain design considerations place additional constraints on optimal O/F ratios. The optimal N₂O-to-fuel mass ratio more commonly described as oxidizer-to-fuel mass ratio (O/F) for the NOFB monopropellant blends described here typically cover ranges from 2.5<O/F<11. At lower O/F ratios, carbon fouling becomes a concern. At higher O/F ratios, the hot highly oxidizing environment of the exhaust gases make combustion chamber and engine design very difficult due to the aggressive oxidation of nearly any type of material in this type of gas environment.

The selection of a monopropellant for mission design, propulsion system design, and actual use requires knowledge of a large number of performance metrics, as well as storage and ground handling considerations besides just the engine I_sp performance. FIG. 3 provides a comparative summary of multiple engine, storage, and ground handling performance metrics between an exemplary NOFB monopropellant formulation and hydrazine and bipropellant nitrogen tetroxide/monomethylhydrazine. Note that the I_sp performance for the NOFB blend is comparable to the bipropellant nitrogen tetroxide/monomethylhydrazine and significantly higher that the hydrazine monopropellant.

Minimum impulse bits are the minimum thrust×time that a propulsion system can impart. Characterizing minimum impulse bit performance of a propulsion system is important for aerospace applications such as spacecraft precision attitude control and miniature vehicle maneuvers. Typically, expected propulsion Isp performance decreases as propulsion systems try and achieve smaller minimum impulse bits. Therefore, more spacecraft propellant must be flown for missions that operate in a regime of small impulse bit performance. A number of factors influence this degradation in performance: 1) In hydrazine systems, catalyst beds for decomposing the monopropellant must be brought up to optimal operation temperatures to achieve more complete decomposition of the monopropellant. In many cases, small pulsed flows of hydrazine may not allow optimal bed temperatures to be achieved. NOFB monopropellants are most commonly spark-ignited for rocket propulsion applications and are not performance-limited by these type of catalyst beds, 2) the minimum impulse bit that can be achieved is directly associated with the minimum mass of propellant that can be discharged. This minimum propellant volume is associated with the density of the monopropellant and the small hardware volumes between a valve and through the reaction chamber. NOFB monopropellants can be operated at very low pressures (<<100 psia) where the NOFB monopropellant gas has densities that are < 1/100 of liquid hydrazine. We have conducted combustion experiments indicating rapid combustion is sustained at low pressures (currently tested down to ˜12 psia). The significantly higher combustion temperatures than hydrazine (see table adiabatic flame temperature) suggest that NOFB chemical kinetics reactions will be much more rapid than hydrazine particularly when considering catalytic reactions limited by surface area (surface catalysts are commonly used to decompose hydrazine). Rapid combustion kinetics without catalyst beds would ultimately allow smaller combustion chamber/reactor volumes as compared to hydrazine. Given these various attributes described above, NOFB monopropellants can be expected to have better minimum impulse bit performance than hydrazine. FIG. 3 summarizes anticipated minimum impulse bit Isp performance of NOFB monopropellants compared to hydrazine.

EXAMPLE 3

In addition to I_sp performance, a number of additional characteristics of a monopropellant are, in general, considered desirable. Hydrazine has an OSHA human fatal exposure limit of approximately 50 ppm. Low and non-toxic chemical monopropellant formulations are desired to mitigate the relatively high costs of ground handling and working with toxic monopropellant formulations. The NOFB monopropellant formulations of the claimed invention are non-toxic and classified as asphyxiants—the NOFBs are similar to gasoline in this regard, only overexposure in very high concentrations displaces breathable air resulting in suffocation or in a more minor case can cause temporary exposure symptoms such as headaches and/or confusion. In any case, the removal to a fresh air supply mitigates the symptoms to exposure. The NOFB monopropellants rapidly volatize into air so that large concentrations of liquids are easily removed from a spill. Also, where hydrazine and the bipropellant nitrotetroxide/monomethylhydrazine are corrosive and may be absorbed into the skin, the NOFB monopropellants may only cause cold burns as a result of rapid propellant discharge. In addition, where hydrazine and the bipropellant nitrotetroxide/monomethylhydrazine can be ingested, causing abdominal cramps, convulsions, unconsciousness and vomiting, and in most cases death, ingestion of the NOFB monopropellants is unlikely due to their high volatility. Also, the exhaust products of the NOFB monopropellants are N₂, CO, H₂O, H₂ and CO₂, where ammonia gas is an exhaust product of hydrazine.

For spacecraft science missions, ammonia is an undesirable byproduct because of its reactions with soils that can readily complicate and contaminate sensitive soil measurements.

Tank storage characteristics of monopropellants are important for minimizing monopropellant fluid handling hardware and tank mass relative to monopropellant mass. Ideally, storage densities of monopropellants are very high. NOFB monopropellant densities have comparable room temperature storage tank densities as hydrazine (˜0.57 g/cc) when factoring in optimized hydrazine tank designs that include internal helium reservoirs in hydrazine tanks. These helium reservoirs are used for pressurizing the hydrazine to achieve reaction chamber pressures for engine and thruster operations. NOFB monopropellants are self-pressurizing and do not require additional pressurant system hardware or unutilized tank volume for expelling the monopropellant. While monopropellant and bipropellant hydrazine systems can typically have unutilized residual propellant in a tank that are ˜1-3% of the initial load, NOFB monopropellants can be expelled down to very low pressures (where they are a pure gas) such that the unutilized monopropellant is <<1% of the initial monopropellant load. Furthermore, this residual gas phase NOFB monopropellant can be accurately monitored with a simple pressure sensing device unlike the liquid propellant alternatives. These NOFB attributes of propellant residuals are important for spacecraft with large wet masses (bulk of launched spacecraft is loaded propellant) whose primary mission life duration are defined by small propellant fractions. Many spacecraft missions' large maneuvers are conducted early in the mission and consume the bulk of the propellant—the life of the spacecraft mission is therefore defined primarily by both available residual propellant and accurate knowledge of this available propellant for planning purposes.

FIG. 4 illustrates storage characteristics of one NOFB monopropellant formulation with both monopropellant liquid and ullage gas (gas in equilibrium with liquid in tank) density and the associated monopropellant vapor pressure plotted against temperature. Each NOFB monopropellant formulation demonstrates a unique vapor pressure and density curve. These metrics are relevant because the size of a monopropellant tank as well as the proof strength of the tank will depend on the values prescribed by data such as that contained in FIG. 4. By prescribing the total spacecraft velocity change or, equivalently, the total imparted momentum required over the lifetime of a satellite and the thermal environment of the spacecraft, the entire required monopropellant storage capacity and pressure rating can be derived with information similar to that shown in FIG. 4.

For low temperature operations and storage considerations, NOFB monopropellant densities increase significantly up to ˜1 g/cc at −75° C. and freeze at <−80° C. These temperatures are not uncommon for deep space and planetary surface missions that are further from the sun than earth and/or shielded from the sun (e.g. the Mars polar cap). While NOFB monopropellants density performance improves with lower temperature, hydrazine freezes ˜0° C. requiring additional heater hardware and spacecraft power to prevent freezing from occurring. Compared to solid propellants (most commonly incorporating premixed solid oxidizer and fuel), NOFB monopropellants typically have higher Isp performance, and are readily throttleable (i.e. can control and vary thrust output) for optimizing propellant usage in a flight trajectory; however, NOFB monopropellants tend to have lower storage densities. For deep space environments, solid propellants have to be carefully handled and insulated to avoid thermal cycling and stress cracking of the propellant grains. Structural flaws and minute cracks in solid grains can readily cause catastrophic engine failure through rapid combustion and heat-induced crack propagation during ignition. NOFB monopropellants have been shown to be very insensitive to very large temperature alteration during static and dynamic conditions including large transient tank drawdowns (FIG. 5B). Furthermore, being a liquid and gas, NOFB monopropellants are not susceptible to failure modes that are inherently associated with solid and solid composite grain structures and, therefore, can likely be thermally cycled indefinitely. As a result, NOFB monopropellants are low temperature insensitive whereas hydrazine and its derivatives, and solid propellants require additional resources to ensure relatively warm, stable thermal conditions when exposed to low temperature environments such as found in deep space missions or for missile launch applications in terrestrial environments that have large seasonal changes in temperature, for example.

FIG. 5A illustrates the minimal variation in mixture chemistry for one exemplary NOFB monopropellant blend under exposed environmental conditions. In this experiment, the ullage gas in the propellant tank was sampled as a function of propellant temperature (tank immersed in low temperature cold bath). The Fourier Transform Infrared (FTIR) Absorption Spectrum of the ullage gas was acquired as a function of different monopropellant temperatures and compared to NOFB calibration gas “fingerprints” to determine the degree of NOFB mixture alteration as a function of temperature. A similar experiment was run for complete expulsion where the tank was loaded with a known NOFB O/F ratio. After complete expulsion, the residual gas was analyzed with the same technique above to determine the degree of mixture alteration. FIG. 5B shows the variation in O/F ratio of three NOFB blends after rapid complete expulsions (80% liquid load in 75 cc tank expelled in ˜2 s). The variation in modeled I_sp performance for two bounding scenarios (frozen-at-the-throat chemistry and equilibrium chemistry throughout nozzle) for these extreme tank expulsion cases is ˜1%. These data confirm the inherent robustness of the NOFB monopropellant mixture from variations in mixture-ratio chemistry during rapid transient phenomena and exposure to wide temperature ranges.

In general, monopropellants, including solid propellants having both an oxidizer and fuel premixed, must be carefully characterized and handled with care. Upper temperature limits on propellants are required to prevent inadvertent chemical reactions from taking place including unintentional thermal ignition. In many cases, these temperature limits may be as low as ˜10's° C. Heated capillary tube testing has demonstrated that exemplary NOFB monopropellants have thermal ignition temperatures that are ˜400° C. (FIGS. 7A and 7B) and may be as high as 650° C. in the presence of inert materials (i.e. specific grades of metals). These are very high temperature limits, and, in fact, a regeneratively-cooled (propellant cools combustion chamber) NOFB monopropellant engine has been developed and tested (discussed below and shown in FIG. 11) that takes advantage of the high exemplary thermal decomposition limits of NOFB monopropellants in order to provide a desirable design mechanism for developing long life-cycle engines.

Additionally, accidental dry spark ignition can ignite environmentally-exposed solid propellants, and therefore extreme care must be taken to avoid accidental spark sources and surface charging/discharging environmental conditions. Unlike solid propellants, NOFB monopropellants, by their nature of containment, are stored in sealed metal containers that behave as Faraday cages (prevents buildup of charge) which essentially eliminates the possibility of dry spark ignition. Care in propulsion system design still must be taken with devices that could disrupt the continuous Faraday cage such as valves with insulating valve seats and plumbing interfaces, for example. Furthermore, the NOFB monopropellants have been shown to have very high breakdown voltages (>>10's kV) at common terrestrial tank storage temperatures and associated pressures (in fact, N₂O has been commonly used as a high voltage gas insulator for high voltage applications). The Paschen curve minimum breakdown voltage gap of N₂O at even a very low storage pressure of ˜100 psia is <0.001 mm (see FIG. 9). This very small maximum gap distance is significantly smaller than the NOFB quenching distance (distance through which a flame cannot propagate as discussed below and experimentally shown in FIG. 10) suggesting that even if you could directly expose the stored NOFB monopropellant to high voltages, it would not be possible to easily ignite. Furthermore, these associated ignition volumes are so small they would unlikely be able to initiate a sustained chemical reaction. These attributes of the major constituent of the NOFB monopropellant suggest that unintentional spark ignition of NOFB monopropellants is not likely. Intentional repeated spark ignition has been demonstrated (see related U.S. Ser. No. 60/868,523, filed Dec. 4, 2006 entitled “Injector Head”, which is herein incorporated by reference in its entirety) by careful design of the injector and spark ignition system to ensure engine ignition at startup that occurs near the Paschen curve minimum (point where minimum voltage is required to propagate a spark through a gas).

The realistic ignition source in the environment is evaluated for its potential to initiate a combustion process. As briefly discussed above, valves impart mechanical energy into a fluid stream which could feasibly be converted into an electrical discharge through triboelectric charging as a valve component slides across an insulting interface (i.e. valve seat). To conduct preliminary experiments to determine whether valves are a realistic ignition mechanism, an automated valve cycle test in the presence of NOFB monopropellant has been implemented. Essentially, a geared DC servo motor was coupled to a valve with electronic triggers to both count valve cycles and control the servo motor.

The thermocouple and pressure transducer were coupled into a data acquisition system and signals fed into a computer program which monitored the processed signals. The thermocouple was an exposed tip 1/16″ K type thermocouple (to reduce time lag in event detection). The pressure transducer was used to ensure there was not a slow leak in the system therefore reducing uncertainty in the case that an event occurred. The flashback arrestor is utilized to isolate the main valve and the pressure transducer in the case of an event such that they are not destroyed. In this implementation, the ball valve stem was electronically isolated from the rest of the system via nylon gears. One possible failure mode could be electric charging of a valve stem causing a spark to propagate within the propellant stream. Utilizing this system (and slight variations hereof), over 8,000 on/off cycles have been run without a single event recorded at pressures of 100 psia (common feed system line pressures for valves). Flight valves are qualified in a similar experimental configuration with the range of anticipated NOFB fluid properties at the valve interface.

FIG. 6 illustrates exemplary nozzle coefficient values, Cf, for use in vacuum equivalent I_sp engine tests described above. Because it is not always economical or possible to take measurements of an engine inside a vacuum chamber thrust stand, scaling calculations can be made which estimate what the vacuum equivalent I_sp performance would be based on experimental performance observed in atmospheric conditions with flow that achieves sonic velocities at the minimum diameter of the engine (the throat). By calculating a theoretical nozzle coefficient, Cf, determined from exhaust gas chemistry through a nozzle expansion using equilibrium chemical analysis software such as NASA's CEA program (Gordon and McBride (1994), “Computer Program for Calculation of Complex Chemical Equilibrium Compositions and Applications”, NASA Reference Publication 1311) (as shown in FIG. 6), a relatively quick experimentally observed I_sp measurement can be determined within typically tight error bars by measuring the integrated chamber pressure and monopropellant mass consumed during an experimental engine run. Basically, the dictating equation is:

$Isp=\frac{\begin{array}{c}\left(\mathrm{Nozzle}\mathrm{Coefficient}\right)\left(\mathrm{Throat}\mathrm{Area}\right)\\ \left(\mathrm{Time}\text{-}\mathrm{Integrated}\text{-}\mathrm{Chamber}\mathrm{Pressure}\right)\end{array}}{\mathrm{Mass\_of}\mathrm{\_Propellant}\mathrm{\_Consumed}}$

The nozzle coefficient can also be used to determine the engine thrust in vacuum from the following equation:

Thrust=(Chamber Pressure)(Throat Area)(Nozzle Coefficient)

FIG. 7A illustrates thermal decomposition data for one NOFB monopropellant formulation, while FIG. 7B shows a summary of decomposition Go/NoGo test vs. NOFB pressure for a different exemplary NOFB monopropellant. This metric is of specific interest for regeneratively cooled engine designs and in defining safe temperature handling limits. Regeneratively cooled engines use the propellant flowed through a jacket in the combustion chamber wall as a coolant to help maintain the combustion chamber walls below thermal failure limits. This energy acquired during wall cooling is not lost but rather results in hotter propellant being injected back into the chamber (hence the name regenerative). While most propellants have limited cooling capacity associated with the liquid specific heat of a propellant (energy required to heat the liquid by a certain change in temperature), the NOFB monopropellants, have very high vapor pressures. By intentionally creating a pressure drop in the regenerative jacket, NOFB monopropellants can be forced to “flash” or vaporize and absorb substantially more energy from the combustion chamber walls by going through a phase change (liquid vaporizing into a gas). This is a similar concept to how a refrigerator works and is much more effective at cooling combustion chamber walls. In other regeneratively-cooled designs and applications, advanced jacket design techniques that enhance heat transport into the NOFB monopropellant (particularly for the case of flowing NOFB gases) by increasing jacket surface area or enhancing boundary layer temperature gradients may be used to regeneratively cool the engine without “flash-cooling”. In either scenario, the maximum cooling capacity of the monopropellant is limited by the thermal decomposition limit of the monopropellant.

FIG. 8A illustrates the large enthalpy of vaporization (energy absorbed during vaporization) of an NOFB monopropellant derived from the Phase Diagram shown in FIG. 4 and compared the energy absorbed in a typical coolant that is heated from the same starting temperatures to ˜300° C. FIG. 8B (derived from FIG. 4) illustrates the rapid temperature decrease as the propellant is “Flash-cooled” started with different tank temperatures and associated tank densities and flowing the propellant through any device and/or medium that causes a pressure drop. (Note quality as shown in this figure is the percent gas by mass in a liquid/gas mix in equilibrium). FIG. 8B is also critical for evaluating feedline propellant densities that feed an engine when considering the design of the anti-flashback systems described below, as well as temperature limits within which monopropellant feed system hardware must operate. FIG. 11 illustrates the successful operation of a regeneratively-cooled NOFB thruster demonstrating the principle of an NOFB flash-cooled engine. This is one important feature of the NOFB monopropellants given the very high combustion chamber temperatures (see FIG. 1) that make even exotic high temperature combustion chamber material designs typically not feasible to implement. For comparison, monopropellant hydrazine has an exhaust gas temperature of ˜1600° C.

FIG. 9 illustrates exemplary Paschen curve minimum (worst case optimum pressure×gap_distance conditions for propagating a spark across two parallel surfaces) spark propagation distance for pure N₂O (main NOFB constituent) as a function of gas pressure. At room temperature storage pressures, spark gap distances must be <0.0001 mm. Such small associated spark volumes are unlikely to allow inadvertent NOFB monopropellant ignition since exemplary NOFB quenching distances are at least ten times greater as discussed below and shown in FIG. 10.

Monopropellants can be sensitive to shock which initiates a rapid chemical reaction (i.e. detonation) resulting in catastrophic system failure. Impact drop testing from 5.5 meters has shown exemplary NOFB monopropellants to be insensitive to impact-induced detonation.

Because liquid monopropellants comprise combined fuel and oxidizer, they can form a potential ignition mechanism (a.k.a “flashback”) back into their storage tank. Therefore, a mechanism for preventing flashback must be included in the engine and feed system design. A very important parameter for designing an engine injector and flashback control mechanism is the quenching distance of a monopropellant. This is the smallest flowpath dimension through which a flashback flame can propagate. In practice this dimension is affected by additional parameters such as tortuosity (curviness of flow path) and to a lesser extent the temperature of the solid containing the flowpath. Smaller flowpath sizes will quench a flame and, in general, prevent flashback although secondary ignition through heat transfer through a solid into the unreacted monopropellant must also be ultimately considered. FIG. 10 illustrates experimental data of sintered metal pore sizes sufficient for quenching an NOFB monopropellant that has been intentionally detonated to produce a flashback. These quenching distances have been incorporated into the design of an anti-flashback system using pores sizes that are equivalent or smaller than the ones that didn't allow flame propagation as shown in FIG. 10.

Propellants in general can undergo chemical reactions with storage and feed system hardware that alter the chemistry of the propellant over time. Preliminary long duration testing of candidate NOFB mixtures has shown them to be chemically stable in the presence of common aerospace propulsion system materials (e.g. stainless steel, Teflon). In this case, three different monopropellant blends were exposed to Teflon and stainless steel and allowed to sit for 1.5 years at room temperature. No chemical alteration of the NOFB monopropellant has been observed as indicated by Fourier Transform InfraRed (FTIR) absorption spectroscopy.

FIG. 12 illustrates exemplary low thrust, non-optimized engine run test data in an engine utilizing a NOFB (nitrous oxide fuel blend) monopropellant. This figure is included to demonstrate successful thruster performance utilizing NOFB monopropellant blends in a flight-like configuration. Thrust was calculated based on nozzle coefficients for vacuum equivalent expansion and engine pressures.

FIG. 13 illustrates a comparison of delivered payload mass (minus tankage) to total wet mass (fueled vehicle) versus imparted vehicle velocity change for example NOFB monopropulsion systems relative to a hydrazine system assuming different tankage (percentage of rocket propulsion dry mass relative to total propulsion system mass) as a function of required spacecraft changes in velocity.

EXAMPLE 4

A small 4 cylinder engine (160 cc) was modified for use with the NOFB monopropellants of the present invention to test the concept of using the NOFB monopropellant for operating extremely high altitude military aircraft engines and power supplies for launch vehicles and manned spacecraft applications (NASA's Apollo 13 mission was almost lost because of the lack of a back-up power supply that could have operated from the onboard rocket propellant). In order to utilize the NOFB monopropellant in this type of application, it was necessary to modify the injection manifold, timing, spark-gap, cylinder head, and starter/ignition system relative to the original parameters used in a gasoline/air engine. The engine was tested with the nitrous oxide rocket fuel blends of the present invention comprising either ethylene or acetylene. While the engine hardware associated with these applications is different from the rocket engine hardware identified, the NOFB monopropellants are still fundamentally the same as the rocket monopropellants and the same advantages previously identified in combustion performance, non-toxicity, fluid-handling characteristics, and rapid combustion kinetics relative to hydrazine, for example, apply. Hydrazine-based engines exist for alternative applications, but, similar to the rocket application, a major limitation to widespread use of hydrazine in these applications relative to NOFB monopropellants is the much lower energy density and toxicity of hydrazine.

EXAMPLE 5

The monopropellants of the present invention can be used in deployment system architecture. This is particularly beneficial when an overall NOFB monopropulsion system is already required for applications associated with the deployment application. The present invention has also been studied for use in an inflatable/rigidizable pressurized propeller, for a wing spar and sustaining wing gas pressure, and for an inflatable/rigidizable rover wheel.

The basic system uses the liquid to combustible gas generator for rapid deployment, and a sustaining gas-pressure for robust long term deployment of wings and/or deployables.

The exemplary lightweight rigidizable wheel is designed to provide a wheel sized at about 1.5 meters, for less than 1 hazard per 100 m in aggressive 25% rock abundant Mars terrains and ability to navigate with 30 cm/pixel orbital resolution. Further, the wheel supports more than 100 kg per <10 kg wheel. The wheel employs a set of inflatable shells and has a composite rim.

The exemplary wing spars utilize the monopropellants of the present invention with an inflatable/rapid rigidizing wing spar (combustion/flash-cool) for providing relatively stiff wing to maintain stable C_L and C_D across wing to achieve high overall L/D. The characteristics obtained are shown in FIG. 14.

The monopropellants of the present invention are also used in deployment systems to provide inflatable/rigidizable propellers.

The deployables of the present invention may also contain an annihilation mechanism for post-operational life. This contingency option can be used for deployment deep behind enemy lines where recovery may not be an option.

In these applications, the NOFB rocket monopropellant used initially for propulsive applications is also used to operate these additional auxiliary deployment and operational modes.

The present specification provides a complete description of compositions of matter, methodologies, systems and/or structures and uses in example implementations of the presently-described technology. Although various implementations of this technology have been described above with a certain degree of particularity, or with reference to one or more individual implementations, those skilled in the art could make numerous alterations to the disclosed implementations without departing from the spirit or scope of the technology hereof. Since many implementations can be made without departing from the spirit and scope of the presently described technology, the appropriate scope resides in the claims. Other implementations are therefore contemplated. Furthermore, it should be understood that any operations may be performed in any order, unless explicitly claimed otherwise or a specific order is inherently necessitated by the claim language. It is intended that all matter contained in the above description and shown in the accompanying drawings shall be interpreted as illustrative only of particular implementations and are not limiting to the embodiments shown. Changes in detail or structure may be made without departing from the basic elements of the present technology as defined in the following claims. In the claims of any corresponding utility application, unless the term “means” is used, none of the features or elements recited therein should be construed as means-plus-function limitations pursuant to 35 U.S.C. §112, ¶6.

Claims

1. A monopropellant comprising nitrous oxide and at least one hydrocarbon fuel.

2. The monopropellant of claim 1 wherein the hydrocarbon fuel is selected from the group consisting of ethane, ethylene, and acetylene.

3. The monopropellant of claim 1, wherein the oxidizer-to-fuel ratio is about 2.5 to about 11.0.

4. The monopropellant of claim 3, wherein the oxidizer-to-fuel ratio is about 4.0 to about 8.0.

5. The monopropellant of claim 4, wherein the oxidizer-to-fuel ratio is about 4.5 to about 7.5.

6. A monopropellant comprising nitrous oxide and ethane in an oxidizer-to-fuel ratio of about 2.5 to about 11.0.

7. The monopropellant of claim 6, wherein the oxidizer-to-fuel ratio is about 3.0 to about 9.0.

8. The monopropellant of claim 7, wherein the oxidizer-to-fuel ratio is about 4.0 to about 8.0.

9. The monopropellant of claim 8, wherein the oxidizer-to-fuel ratio is about 4.5 to about 7.5.

10. A monopropellant comprising nitrous oxide and ethylene in an oxidizer-to-fuel ration of about 2.5 to about 11.0.

11. The monopropellant of claim 10, wherein the oxidizer-to-fuel ratio is about 3.0 to about 9.0.

12. The monopropellant of claim 11, wherein the oxidizer-to-fuel ratio is about 4.0 to about 8.0.

13. The monopropellant of claim 12, wherein the oxidizer-to-fuel ratio is about 4.5 to about 7.5.

14. A monopropellant comprising nitrous oxide and acetylene in an oxidizer-to-fuel ration of about 2.5 to about 11.0.

15. The monopropellant of claim 14, wherein the oxidizer-to-fuel ratio is about 3.0 to about 9.0.

16. The monopropellant of claim 15, wherein the oxidizer-to-fuel ratio is about 4.0 to about 8.0.

17. The monopropellant of claim 16, wherein the oxidizer-to-fuel ratio is about 4.5 to about 7.5.

18. A monopropellant comprising nitrous oxide and two or more of acetylene, ethane or ethene in an oxidizer-to-fuel ratio of about 2.5 to about 11.0.

19. The monopropellant of claim 18, wherein the oxidizer-to-fuel ratio is about 3.0 to about 9.0.

20. The monopropellant of claim 19, wherein the oxidizer-to-fuel ratio is about 4.0 to about 8.0.

21. The monopropellant of claim 20, wherein the oxidizer-to-fuel ratio is about 4.5 to about 7.5.

22. The monopropellant of claim 1, wherein additional constituents comprise less than about 30% of the monopropellant.

23. The monopropellant of claim 2 comprising nitrous oxide and one or more of acetylene, ethane, or ethane where the fuel(s) are mixed with nitrous oxide in the gas phase during the manufacturing process prior to condensing into a liquid.

24. The monopropellant of claim 2 comprising nitrous oxide and one or more of acetylene, ethane, or ethane where the fuel(s) are mixed with nitrous oxide in the liquid phase during the manufacturing process

25. The monopropellant of claim 2 comprising nitrous oxide and one or more of acetylene, ethane, or ethane where the fuel(s) are mixed with nitrous oxide in any combination of gas and liquid phases during the manufacturing process.