US 200901.33788A1 (19) United States (12) Patent Application Publication (10) Pub. No.: US 2009/0133788 A1 Mungas et al. (43) Pub. Date: May 28, 2009

(54) FUEL BLEND Related U.S. Application Data (60) Provisional application No. 60/986,991, filed on Nov. 9, 2007. (75) Inventors: Gregory Mungas, Arcadia, CA Publication Classification (US); David J. Fisher, Denver, CO (US); Christopher Mungas, (51) Int. Cl. Plymouth, CA (US); Benjamin C06B 47/04 (2006.01) Carryer, Denver, CO (US) (52) U.S. Cl...... 149/74 (57) ABSTRACT Correspondence Address: Compositions and methods herein provide monopropellants HENSLEY KIM & HOLZER, LLC comprising nitrous oxide mixed with organic fuels in particu 1660 LINCOLN STREET, SUITE 3000 lar proportions creating stable, storable, monopropellants DENVER, CO 80264 (US) which demonstrate high ISP performance. Due to physical properties of the nitrous molecule, fuel/nitrous blends dem (73) Assignee: Firestar Engineering, LLC, onstrate high degrees of miscibility as well as excellent Broomfield, CO (US) chemical stability. While the monopropellants are particu larly well Suited for use as propulsion , they also lend themselves well to power generation in demanding situ (21) Appl. No.: 12/268,266 ations where some specific cycle creates useable work and for providing gas pressure and/or heat for inflating deployable (22) Filed: Nov. 10, 2008 materials. Isp (NOFB3X at 100 psia Chamber Pressure) vs OF (To=298 K) ------arrarilla-rule 2001 F --vacuum Eq v 2001 Ed T-ch T-th

EQUBRUM { % - 4% 9%. 8. sash;94 -ko's -8%xxxx was wi:8%. g. gzge 8% is es: 8 as: 8 Patent Application Publication May 28, 2009 Sheet 1 of 17 US 2009/0133788A1

isp (NOFB3X at 100 psia Chamber Pressure) vs OF (To=298 K) --vacuum F -e-2001 F --vacuum Eq.- 2001 Ed T-ch T-th ------3500

350 d ... 3.3.xv,www. - 3OOO "

EQUBRUM -- 2500

a resis is uses - or- so so, e - e, , is a se - 330 - Ex entally Observed 2001. 2OOO Yeero -ierit." : intois's - vacuum:- - - g"." Equivalent spis r. %.

320 - are KX8: 8x8, 9%y tw: , p:% x8% arz Wee Rw8 was 888: 888 8 888. 8: 48 %. Mix %. 8%8. &. 1500

FROZEN AT THROAT

FIGURE 1 Patent Application Publication US 2009/O133788A1

Patent Application Publication May 28, 2009 Sheet 3 of 17 US 2009/0133788A1

METRIC N.H. NTOIMMH NOFB34, NOFB37 Propulsion Type Bipropellant Monopropellant (sP) 230 - 240 S 288 - 326 s >300 s - 345 s) ------

Minimum impulse bit - 15OS 1) ISP >{ 300s) ------performance SSP 2) 50-600x less impulse per valve opening ------withlower density ----- Specific Energy Density s f (Propellant) 390 Whi/kg 2049 Whi/kg 1360 - 1520 Wh/kg

Pressurized Total 56-0.65 ------internal Tank Volume (including He { -0.75 p; (C) Specific Gravity pressurant) ------.0 (G) -75 Density impulse (wsp) 1.52 NSCC {2,4} N'Sicc 1.7 Nsico (0) 31°C - 3.2} N's/CC (O) -75°C Adiabatic Flame Temp. ; - 1200°K -27OOK 3OOO-345O K Storage Pressure " . 300-400 psia (separate pressurant) 100-1000 psia (self-pressurizing) 1) -0.015%-1.6% (G1-100 psia unburned Unusable Tanked respectively 1%-5% ------Propellant Residuals O J AO 2) Last -5-20% of Propellant Load measureable wif gas pressure transducer Freezing Point 2 oC MMH: - 11°C <-77°C (insensitive to cycling) Auto Ignition of...An 400°C inside 316 Stainless Steel tube) Temperature Minimum Surface) -- Material incompatibilities not fully tested --- Auto ignition 270°C (on 650°C (after detonations on 316). May likely Temperature Maximum glass) be higher on inert surfaces. impact Sensitivity Unable to cause ignition with mechanical impact (5.5 m drop test to date) SS316 causes a minor catalytic reduction in Material Incompatibility Reacts violently with many metals, auto ignition temp. Entire trade space not fully metal oxides and porous matcrials explored yet. Exhaust Products - for Astrobiology NH3 is issue N2, CO, H2O, H., CO2 Investigations ------...W. Human DH limit 50 ppm 20 ppm (Immediately dangerous (Low Vapor (LOW Vapor Asphyxiant. Not OSHA-regulated. to life and health) Pressure) Pressure) Rapidly volatilizes into air Corrosive, Burning sensation. High doses may cause asphyxiation, low Cough, Headache. Nausea. doses may cause narcotic effects i.e. inhalation Shortness of breath. Sore throat. dizziness, headache, nausea and loss of --- Convulsions. COOrdination Ski Corrosive. MAYBE ABSORBED Rapid propellant discharge can cause Cold - | Redness. Skin burns. Pain. --- burns. Eves Corrosive. Redness. Pain. Severe High pressure discharge to eyes may cause --- cy deep burns. " | damage. ------ingestion Cogy gettps Engestion is not considered a likely route of exposure given high volatility - F GRE 3

Patent Application Publication May 28, 2009 Sheet 5 of 17 US 2009/O133788A1

NOFB37 VARATION NULLAGE GAS PHASE FTIR SPECTRA FROM -77C TO ROOMTEMPERATURE AND 114 TANK DRAWDOWN

4800 4700 4600 4500 4400 4300 WAVENUMBERS (cm-1) FIG.5A Patent Application Publication May 28, 2009 Sheet 6 of 17 US 2009/O133788A1

NOFB3x DRAWDOWN OVERVIEW EQUILIBRIUMISPI (POINTS REPRESENTAVERAGES OF TRIALS) FROZEN THROAT ISP 327292 KEYNUMBERSALONGSIDE POINTS REPRESENT 2- - - - 331/295 8.5 1.) INITIALLIQUID PHASE OF MEASUREMENT 802) INITIALULLAGEVOLUMEGASPHASE OF MEASUREMENT - 333,298 3) POST-DRAWDOWNGASPHASE OF MEASUREMENT 2ND ORDERBEST FITCALIBRATIONCURVE 334,301 70 (BASED ON2POINTSPER OF FOR NOFB 335/303 33,34,35,3637,3839) R2 = 9833. 6.5 ------/42------335.306 335/308 5 334/310 333/312 331/314 - - O NOFB35 ------327/315 A NOFB37 ------324/314 3.0 323,309 0.040 0.042 0.044 0.046 0.048 0.050 0.052 0.054 0.056 0.058 0.060 ABSORBANCE PEAKG 4430.50 cm-1 FIG5B Patent Application Publication May 28, 2009 Sheet 7 of 17 US 2009/O133788A1

Cf VSAelAt FOR NOFB3 BLEND OF = 5

-O-EQUILIBRIUM--FROZENAT COMBUSTOR-a-FROZEN AT THROAT

1H:e. Patent Application Publication May 28, 2009 Sheet 8 of 17 US 2009/O133788A1

NOFB34THERMALDECOMP FLOW

- BLOCK TEMP - GASTEMP - SYSTEMPRESSURE 560

AUTOIGNITION 555 370 C 550 545 540 535 530 525 520 O 2 4 6 8 10 12 14 TIME(S) FIG.7A Patent Application Publication May 28, 2009 Sheet 9 of 17 US 2009/O133788A1

NOFB35 SS316 CAPILLARYTUBETHERMAL DECOMPTESTS

0 NODECOMPOSITION THERMALDECOMPOSITION

500 450 G 400 350 s 300 250 200 o 150 NOTE: THERMALDETONATION 35 100. LIMIT IS 270°C ON ACID-CLEANED 50 QUARTZAND 24°C ONOXIDZED IRON O O 100 200 300 400 500 600 GAS PRESSURE (psia) FG.7B Patent Application Publication May 28, 2009 Sheet 10 of 17 US 2009/0133788A1

NOFB34 HEAT OF WAPORIZATION VS. TEMP

-NOFB34 - N2O ------it - ENTHALPY OF RAISING 20% OF FUELTO 300°C (BIPROP. REGEN. COOLING)

-75 -50 -25 O 25 50 TEMPERATURE (C) FIG.8A Patent Application Publication May 28, 2009 Sheet 11 of 17 US 2009/0133788A1

NOFB34 TEMPERATURE (T), DENSITY (p), AND QUALITY(x) VS. THROTTLED PRESSURE 1.0 25 0.9 15 S, 0.70.8 Po,2. (TnCo-260) = 26OK is5 a 0.6 -15 Ais 05 25 5. 0.4 -35 5 0.3 -45 x(t-oo- -55 0.0' ' 2 (0.-200a, 2 (TOF R -65-75 800 700 600 500 400 300 200 100 O THROTTLED PRESSURE, P2 (psia) FIG.8B Patent Application Publication May 28, 2009 Sheet 12 of 17 US 2009/0133788A1

N2O Max Gap vs. Pressure at Paschen Curve Minimum (418 W)

E.

eart; 0.0010 (f)

f ( ). N s ( ----- 0.0001 ------...------1------, O 200 400 600 800 O Pressure (psia)

FGRE 9 Patent Application Publication May 28, 2009 Sheet 13 of 17 US 2009/0133788A1

NCFE Filter Vedia Quenching Distance as a Function of Loaded NOFB line Density aircra-a-a-iwar---ir------...-kect----e-r------reluro i: -, -334. Fai CE 1 a

NOF835 Fail : - - - - - r - ...... 'worn iv. 'wr m r - - - C se t is sak s t

s See . e

- k :

w

. i w ok w ...... w A w.w is 8 br w - - w is 4 r a w w w A w w a y

i t

5 N o 2. o E. No.2 : role------o-, -vi (5 O O.5 C2 NCFB loaded line tensity (gicc)

FBG RE 10 Patent Application Publication May 28, 2009 Sheet 14 of 17 US 2009/O133788A1

(MW88-)HEGWWHO1SnHHINZZEdÅ1010\dd Patent Application Publication May 28, 2009 Sheet 15 of 17 US 2009/0133788A1

8/13/06 ENGINE RUN 1

O 0.5 1 15 2 2.5 3 3.5 TIME (S) FIG. 12 Patent Application Publication May 28, 2009 Sheet 16 of 17 US 2009/0133788A1

DRYMASS (MINUSTANKS) TO TOTALMASS VS. AV

0.6 320S NOFBOBS 7%. TANKAGE O 5

300S NOFBOBS 0.3 13%. TANKAGE ES 0.2 N2H4 0.1 7%. TANKAGE S 0.0 O O 1000 2000 3000 4000 5000 6000 7000 8000 AV (m/s) FIG.13 Patent Application Publication May 28, 2009 Sheet 17 of 17 US 2009/0133788A1

Table 5. Summary of NOFB deployable Wing Spar Specs are - in Tubular deployable rigidizable Cross |- section centered on 4 chord. m Max Wehicle Weight 1100 kg 10,791 N Combined Tapered Wing Area 40 m ------Root chord 2.4 m ------Tip chord 1.0 m (0.4x root) ------individual wingspan 12m Max composite tensie stress 13.3 ksi mulu Spar internal gas pressure 100 psia -- run. Deployed Spar Density 12 g/cc Total Spar Plass (2 Wings) 47 kg NOFB thermal gas mass for High energy density. kg deploying both wings | Low Energy Density, 3 kg ---

Figure 4 US 2009/O 133788 A1 May 28, 2009

NITROUS OXDE FUEL BLEND energy. However, the hot (>1500° C.), highly oxidizing reac MONOPROPELLANTS tion products make catalyst bed and reaction chamber design challenging. CROSS-REFERENCE TO RELATED 0006. With the addition of hydrocarbon fuel to the reaction APPLICATIONS in the equation above, the specific energy density of liquid monopropellants can be increased up to ~1500 Whr/kg (-3 0001. This application claims benefit of U.S. Provisional times the energy density of pure N2O), and I performance Application No. 60/986,991, entitled “Nitrous Oxide Fuel greater than 300 s becomes feasible. Furthermore, the hot Blend and Monopropellants' and filed on Nov. 9, 2007, deleterious oxygen in the exhaust stream can be consumed which is specifically incorporated herein by reference for all and the higher combustion reaction temperatures result in that it discloses and teaches. faster reaction kinetics as compared to pure NO decompo sition. The faster kinetics permit rapid spark ignition. In Such STATEMENT REGARDING FEDERALLY a case, a catalyst bed does not become the material limitation SPONSORED RESEARCH ORDEVELOPMENT for engine design, and regeneratively cooled engine design 0002 This invention was supported in part by subcontract approaches with conventional materials can be adapted for number 1265181 from the California Institute of Technology the higher I performance using low cost engine fabrication Jet Propulsion Laboratory/NASA. The U.S. Government techniques. may have certain rights in the invention. 10007. The highest I chemistry ever test-fired in a engine was lithium and fluorine, with hydrogen added to BACKGROUND improve the exhaust thermodynamics. The combination delivered 542 seconds specific impulse in a vacuum. How 0003 Liquid fueled have better specific impulse ever, the impracticality of this chemistry highlights why (I) than solid rockets and are capable of being throttled, shut exotic propellants, particularly bipropellants, are not used in down and restarted. The primary performance advantage of practice. To make all three components liquids, the hydrogen liquid propellants is the oxidizer. Several practical liquid must be kept below -252° C. and the lithium must be kept oxidizers (liquid oxygen, nitrogen tetroxide and hydrogen above 180° C. This example demonstrates dramatically a peroxide) are available that have much better I than the major drawback of bipropellants—they must be stored in ammonium perchlorate used in Solid rocket boosters when separate tanks (and often under different temperature and/or paired with comparable fuels. However, the main difficulties pressure conditions), and they must be delivered to the com with liquid propellants also are with the oxidizers. Oxidizers bustion chamber at a pre-defined and specific mix ratio, typi are generally at least moderately difficult to store and handle, cally at high pressure and high flow rates. either due to extreme toxicity (nitric acids), moderate cryo genicity (liquid oxygen) or both (liquid fluorine). Several SUMMARY oxidizers that have been proposed, for example, O. ClF. ClFs, are unstable, energetic and toxic. 0008 Implementations described and claimed herein 0004. The first liquid-fuelled rocket launched in the address the foregoing issues with a family of nitrous oxide 1920s—used gasoline and liquid oxygen as propellants. Liq fuel blend (NOFB) monopropellants comprising organic uid hydrogen was used in the 1950s, and by the mid-60s, fuels mixed with nitrous oxide (NO). When combusted, the liquid hydrogen and liquid oxygen were being used. Common nitrous oxide provides both thermal decomposition energy liquid monopropellants in use today include hydrazine and and serves as the oxidizer to combust the fuels. Example hydroxyl ammonium nitrate. Common liquid bipropellants organic fuels include ethane (CH), ethylene (CH4), acety include liquid oxygen and kerosene, liquid oxygen and liquid lene (CH), and mixtures thereof. Mixtures of these fuels hydrogen, and nitrogen tetroxide and hydrazine or monom based on oxidizer-to-fuel ratio (O/F) generate desired mono ethylhydrazine. A goal of propellant design has been to propellant characteristics including, but not limited to, I, develop a monopropellant having the high performance char miscibility over a wide temperature and pressure range, acteristics of a bipropellant. Due to the simplified system favorable fluid handling performance, low freezing points, architecture of monopropellant systems, finding monopro rapid combustion kinetics for fast engine response times, pellant chemistry that provides bipropellant-like I perfor relatively high thermal decomposition limits, low mechanical mance has long been considered a "holy grail” in monopro shock sensitivity and impact-induced detonation, relatively pellant development. Research in the field of “green high storage densities, and exhaust gas chemistries that do not monopropellants has been ongoing to find non-toxic mono produce carbon fouling or hot oxidizing environments that propellant alternatives to hydrazine. One Such candidate is are difficult, if not impossible, to accommodate with combus nitrous oxide. Nitrous oxide can be decomposed through the tor or reaction chamber design materials. In addition, to being following exothermic reaction: a very stable oxidizer, nitrous oxide is a very good solvent with a near room temperature critical point at 36.4°C. There fore, it is possible to dissolve fuels into the NO to produce a 0005 Under standard conditions, this reaction generates nitrous oxide fuel blend (NOFB). Care must be taken in the 82 kJ/mol (515 Whir?kg) of heat per unit nitrous oxide. To design of the NOFB monopropellant to ensure that the mix liquefy the stored monopropellant requires 16.5 kJ/mol (104 ture is safe to handle, and that the NOFB monopropellant Whr/kg) or approximately 20% of the enthalpy of reaction. maintains balanced degassing of all NOFB constituents over The maximum theoretical I of this reaction is 205s. NO is a wide range of temperatures and tank drawdown profiles in a highly stable molecule given its high activation energy which it may be used. barrier ~250 kJ/mol. As a result, thermal decomposition 0009 Implementations herein provide a nitrous oxide fuel requires preheat temperatures >1000° C. Alternatively, cata blend (NOFB) monopropellant comprising nitrous oxide and lysts can be used to significantly depress this activation an organic compound in an oxidizer-to-fuel ratio of about 2.5 US 2009/O 133788 A1 May 28, 2009

to about 11.0. Preferably, the organic compound comprises, features of the claimed subject matter, nor is it intended to be as a main component, a C2 hydrocarbon, or mixtures of C2 used to limit the scope of the claimed subject matter. hydrocarbons. Specifically, implementations provide a monopropellant comprising nitrous oxide and acetylene in an BRIEF DESCRIPTION OF THE DRAWINGS oxidizer-to-fuel ratio of about 2.5 to about 11.0, or about 3.0 (0013 FIG. 1 is a graph of theoretical and actual I perfor to about 9.0, or about 4.0 to about 8.0, or about 4.5 to about mance of an NOFB monopropellant formulation. 7.5, or about 2.5 to about 6.0, or about 3.0 to about 5.0, or 0014 FIG. 2 illustrates the method of making the nitrous about 6.0 to about 11.0, or about 8.0 to about 10.0. Other oxide fuel blends of the present invention. implementations provide NOFB monopropellants compris 0015 FIG. 3 is a chart summarizing NOFB monopropel ing nitrous oxide and ethane in an oxidizer-to-fuel ratio of lant characteristics relative to monopropellant hydrazine and about 2.5 to about 11.0, or about 3.0 to about 9.0, or about 4.0 bipropellant nitro tetroxide/monomethylhydrazine. to about 8.0, or about 4.5 to about 7.5, or about 2.5 to about 0016 FIG. 4 is a graph illustrating storage characteristics 6.0, or about 3.0 to about 5.0, or about 6.0 to about 11.0, or (storage tank liquid and gas pressure and density versus tem about 8.0 to about 10.0. Yet other implementations provide perature) for one NOFB monopropellant formulation (also NOFB monopropellants comprising nitrous oxide and ethyl known as a phase diagram). In FIG. 4, the NOFB monopro pellant storage characteristics are also compared to pure ene in an oxidizer-to-fuel ratio of about 2.5 to about 11.0, or nitrous oxide liquid and tanked hydrazine monopropellant about 3.0 to about 9.0, or about 4.0 to about 8.0, or about 4.5 including a typical helium pressurant load for hydrazine to about 7.5, or about 2.5 to about 6.0, or about 3.0 to about (0017 FIG. 5A is an FTIR spectrum of NOFB monopro 5.0, or about 6.0 to about 11.0, or about 8.0 to about 10.0. The pellant sampled over different tank temperatures (and com ratios are chosen for specific uses. For instance, NOFB34 is parison with the remaining gas after a /4 tank rapid liquid optimized for Small rocket engines (fast combustion kinetics expulsion) illustrating the stability of the NOFB chemical and optimized peak Isp with frozen-at-the-throat combustion mixture to biased constituent outgassing over extreme tem kinetics), and NOFB37 is optimized for large rocket engines perature ratios. FIG. 5B similarly shows the variation in (higher density monopropellant with Isp optimized for slower NOFBO/F ratio in the liquid and ullage gas (gas in tank with combustion kinetics in larger rocket diverging exhaust liquid) of three NOFB blends after rapid expulsion. On the nozzles). In other implementations, the NOFB monopropel right hand side of this figure, the corresponding variation in lants may comprise other compositions or additives up to I performance is also shown as the blend slightly varies about 50% of the monopropellant, or up to about 40% of the during a very aggressive tank liquid expulsion (80% NOFB monopropellant, or up to about 30% of the monopropellant, liquid expulsion in ~seconds) or up to about 20% of the monopropellant. 0018 FIG. 6 is a graph illustrating exemplary nozzle coef 0010. The other compositions include hydrocarbon fuels ficient values for use in vacuum equivalent I calculations. or mixtures thereof wherein the resulting monopropellanthas 0019 FIG. 7A is a graph showing thermal decomposition the property that as the monopropellant is drawn down or the data for one exemplary NOFB monopropellant formulation. temperature changed, the balanced blend has minimal varia FIG. 7B is a summary of decomposition tests vs. NOFB tion in liquid and ullage gas mixture-ratio chemistry as the pressure for an exemplary NOFB monopropellant. liquid monopropellant boils-off to generate ullage gas under 0020 FIG. 8A is a graph illustrating the specific enthalpy these conditions. The additional hydrocarbon fuels may cause of vaporization of one exemplary NOFB monopropellant a <10% variation in rocket Isp performance due to these relative to nitrous oxide and compared to the specific energy variations in boil-off rates for the different NOFB constitu for heating a representative quantity of bipropellant fuel from ents. For Some applications, the addition of small amounts of a the same temperature to ~300° C. FIG. 8B illustrates the detergents, emulsifiers, or other additives may be advanta rapid decrease intemperature as the NOFB monopropellant is geous. throttled or “flash-cooled' by forcing it through a pressure 0011 Additional implementations of the technology pro drop. vide NOFB monopropellants comprising nitrous oxide and 0021 FIG. 9 is a graph illustrating the maximum spark two or more of acetylene, ethane or ethene in an oxidizer-to propagation distance for pure nitrous oxide as a function of fuel ratio of about 2.5 to about 11.0, or about 3.0 to about 9.0, gas pressure at the minimum nitrous oxide spark Voltage of or about 4.0 to about 8.0, or about 4.5 to about 7.5, or about 418V. At this minimum voltage point (also known as Paschen 2.5 to about 6.0, or about 3.0 to about 5.0, or about 6.0 to about curve minimum), for a given gap distance, both higher and 11.0, or about 8.0 to about 10.0. In other implementations, the lower gas pressure requires rapidly increasing spark Voltages. monopropellant may comprise other compositions or addi 0022 FIG. 10 is a graph of quenching distance based on tives up to about 50% of the monopropellant, or up to about oxidizer-to-fuel ratios for one NOFB monopropellant formu 40% of the monopropellant, or up to about 30% of the mono lation. propellant, or up to about 20% of the monopropellant. (0023 FIG. 11 illustrates an exemplary NOFB regenera 0012. In certain implementations, the nitrous oxide is in a tively-cooled thruster utilizing the high volatility of the gas phase when mixed with the fuel during manufacturing; in NOFB monopropellant to “flash-cool the combustion cham other implementations the nitrous oxide is in a liquid phase ber. when mixed with the fuel during manufacturing; and in yet 0024 FIG. 12 illustrates low thrust, non-optimized engine other implementations, the nitrous oxide is in a mixed gas/ run test data in an engine utilizing an exemplary NOFB liquid phase when mixed with the fuel during manufacturing. monopropellant. The mixing is done as described in Example 1. This Summary 0025 FIG. 13 is a graph illustrating a comparison of deliv is provided to introduce a selection of concepts in a simplified ered payload mass of total wet mass rocket propulsion system form that are further described below in the Detailed Descrip performance of an exemplary NOFB monopropulsion system tion. This Summary is not intended to identify key or essential relative to hydrazine systems. US 2009/O 133788 A1 May 28, 2009

0026 FIG. 14 Summarizes the characteristics of the upper stage propulsion systems and booster NOFB deployable wing spars. stages; deep space probe propulsion and power systems; deep space spacecraft ascent and earth return stages; precision DETAILED DESCRIPTION controlled spacecraft station-keeping propulsion systems; 0027 Technology is described herein for providing a human-rated reaction control propulsion systems; spacecraft nitrous oxide fuel blend (NOFB) monopropellant comprising lander descent propulsion, power, and pneumatic systems for nitrous oxide with an organic compound. Such as one or more excavation (NOFB monopropellant can be used to both pro of acetylene, ethane, or ethene resulting in a monopropellant vide mechanical power to run drills in extraterrestrial drilling that has a high specific impulse, low toxicity and allows for applications and to provide gases to remove debris from the easy storage and handling in addition to other desired char area of the cutting Surfaces), spacecraft pneumatic science acteristics. The monopropellant may be used in Some imple sample acquisition and handling systems; micro-spacecraft mentations for rocket propulsion, working fluid production, high performance propulsion systems; military divert and kill or energy orgas generation. interceptors; high altitude aircraft engines, aircraft backup 0028 Before the present formulations and methods are power systems; remote low temperature power systems (e.g., described, it is to be understood that the invention is not arctic power generators); combustion powered terrestrial limited to the particular formulations or methodologies tools including high temperature welding and cutting torches described, as such, formulations and methods may, of course, as well as reloadable charges for drive mechanisms (e.g., nail vary. It is also to be understood that the terminology used guns, anchor bolt guns), and the like. In terrestrial applica herein is for the purpose of describing particular embodi tions, NOFB monopropellants can provide power in situa ments only, and is not intended to limit the scope of the tions where atmospheric oxygen in not in Sufficient quantity present invention; the scope should be limited only by the to provide an oxidizer for combustion reactions (such as very appended claims. It must be noted that as used herein and in high altitude aircraft powerplants or underwater equipment). the appended claims, the singular forms “a,” “an and “the Moreover, there are many derivative applications related to include plural referents unless the context clearly dictates using combustion stored energy. otherwise. Thus, for example, reference to “an agent” refers 0032. A monopropellant is a single fluid that typically is to one agent or mixtures of agents, and reference to “the used for generating thrust, gas generation, and/or power (me method of manufacturing includes reference to equivalent chanical and/or electrical) generation. Monopropellants com steps and methods known to those skilled in the art, and so monly undergo exothermic chemical reactions through a forth. catalytic, hypergolic, or spark ignition mechanism in order to 0029. Unless defined otherwise, all technical and scien release additional heat energy (commonly providing an ide tific terms used herein have the same meaning as commonly ally low molar mass exhaust gas as well) in order to increase understood by one of ordinary skill in the art to which this mass efficiency in generating thrust and power. Monopropel invention belongs. All publications mentioned herein are lants, for example, can be used in a liquid or gas rocket incorporated herein by reference for the purpose of describ engine. A common example of a monopropellant is hydra ing and disclosing devices, formulations and methodologies Zine, often used in for vehicle transla that are described in the publication and that may be used in tion maneuvers (linear momentum changes) and attitude con connection with the claimed invention, including related U.S. trol (angular momentum changes). Another example of a application Ser. No. 60/868,523, filed Dec. 4, 2006 entitled monopropellant is hydroxyl ammonium nitrate (HAN) which “Injector Head.” is currently being investigated as a lower toxicity monopro 0030. Where a range of values is provided, it is understood pellant alternative to hydrazine. that each intervening value, between the upper and lower 0033. Additionally, a working fluid that has a pressure limit of that range and any other stated or intervening value in gradient between it and the Surrounding environment is that stated range is encompassed within the invention. The capable of producing mechanical work/power. This mechani upper and lower limits of these Smaller ranges may indepen cal work/power can Subsequently be converted into alterna dently be included in the Smaller ranges and are also encom tive energy forms (for example, electric power generation, passed within the invention, Subject to any specifically mechanical shaft power can be used to power an electric excluded limit in the stated range. Where the stated range generator or alternator to provide electric power). Pressure includes one or both of the limits, ranges excluding either or from either the natural vapor pressure of an NOFB monopro both of those included limits are also included in the inven pellant and/or through NOFB monopropellant decomposi tion. tion/combustion processes in combination with NOFB 0031. The art of chemical rocket propulsion makes use of monopropellant-derived working fluids can be used strategi controlled release of chemically reacted or un-reacted fluids cally to produce useable work beyond simple thrust. Example to achieve thrust in a desired direction. The thrust acts to work extracting cycles that can implement the NOFB mono change a body's linear or angular momentum. Similar to propellants may include, without limitation, gas turbine rocket propellants that have found application in other work cycles (e.g., Brayton or similar cycles), constant pressure ing fluid production and power generation applications, the expansions of combusted monopropellant (similar to pneu claimed invention may be utilized in many alternative types matic machines), and various piston cycle engines including of applications as well, including gas generation for inflation but not limited to spark-ignited Otto cycles, and compression systems and inflatable deployments, in Systems used to con ignited Diesel cycles. The maximum energy that can be Vert thermal energy in hot exhaust gases to mechanical and extracted from a chemical medium is related to its specific electrical power, and in high energy storage media for pro energy density (stored chemical energy per unit mass). As jectiles, munitions, and explosives. Examples where the shown in FIG. 3, the specific energy density of NOFB liquid claimed technology could be applied specifically include monopropellant (>1300 Whr/kg) is ~3.5 to 3.9 times greater earth-orbiting spacecraft and missile propulsion systems; than hydrazine. For comparison, state-of-the-art lithium ion US 2009/O 133788 A1 May 28, 2009 batteries store ~145 Whr/kg. The NOFB propulsion system the system is purged of air by turning the vacuum pump on, would require additional mass that would effectively lower opening IS-3. IS-4, SW-6, and SW-8. Once an adequate the NOFB monopropellant's specific energy. Nevertheless, vacuum is accomplished, SW-8 is closed. Next, fuel(s) is/are for many primary power applications not requiring energy added to both the mixing tank and the condensing tank. To do recharge, the very high specific energy density of NOFB this, the IS-2 is opened, REG-2 is increased to the desired monopropellants is desirable. pressure, and SW-7, SW-2, and SW-4 are all opened. Depend 0034) For the specific case of rocket propulsion, a variety ing on the vacuum pulled in the previous step, purges may be of metrics determine how efficiently a particular rocket pro required. To determine the necessity of purges, the purity can pulsion system performs. One of the most important metrics be calculated by taking the ratio of fuel added absolute pres in rocket propulsion is specific impulse (I). This metric sure to the total absolute pressure in the tank. For example, if essentially measures the amount of total impulse or imparted a vacuum were pulled to 1 psia, and fuel were loaded to 100 momentum change (integrated force over time) produced by psia, the purity of the load would be 99 psia/100 psia or 99%. a given propulsion system divided by the total mass of pro A purge can increase the purity beyond the initial load if pellant consumed. This result is normalized by the earth's required. To purge, SW-4 is closed, and SW-8 is opened so as gravitational constant (9.81 m/s) such that I has units of to draw the mixture out of the system. However, when run seconds regardless of what international system of units are ning mixed combustibles through a mechanical pump system, being used (English or System Internationale (SI) units)). adequate flashback mitigation measures should be imple Higher I values indicate greater ability to impart velocity mented. Once adequate vacuum is achieved, SW-8 is closed. changes to vehicles for a given amount of propellant con The purge sequence can be repeated as required for the Sumed. By crude analogy, Isp performance is similar in con desired purity of the mixture. The new purity is calculated by notation to “miles per gallon' in a combustion-powered car multiplying the impurity levels of each load together. For engine (although one caveathere is that more engine-specific example, if another 99% pure load is added the 1% impurity characteristics go into defining "miles per gallon' for a car as is multiplied by the new 1% impurity to result in 0.01% compared to rocket propulsion for a spacecraft). Because 1) impure or 99.99% the purity level of the starting fluid. How mass is extremely expensive to launch, and 2) there is an ever, if the starting fluid is only 98% pure, no amount of exponential dependency of propellant mass on I perfor purging can increase purity levels above the initial 98% fluid mance Propellant Mass=Spacecraft Dry Massxexp purity. Once the purity and load pressure is achieved in the Change in Spacecraft Velocity/I/earth gravity-1), high system, SW-4 is closed. The fuel is then shut off and purged I propellants are very attractive for demanding aerospace from the system by closing IS-2 and opening SW-3. Adequate applications. In chemical propulsion systems, in order to time is allowed to vent the fuel from the lines and REG-2 is achieve high I systems, exothermic chemical reactions are backed out, SW-7 is closed, SW-2, and SW-3 are closed. If generally required. Currently, a common industry standard multiple fuels are used, the secondary/tertiary fuel is added at commercial monopropellant, hydrazine, has an I of around this point on top of the prior load (FIG. 2 does not show this 230 s (slight deviations of this number are dependent on option). Once the fuel blend has been achieved, nitrous oxide specific thruster design parameters). The class of NOFB (ni is added. Here, IS-1 is opened, REG-1 is increased, and SW-1 trous oxide fuel blend) monopropellant formulations dis and SW-4 are opened. Once the desired mixture is achieved, closed herein can achieve engine I values of up to 345s and SW-4 is closed. To vent the nitrous oxide, IS-1 is closed, potentially larger I values. Recent experimentally measured SW-3 is opened, the system is allowed time to drain, REG-1 engine I values exceed300s (see FIG. 1). FIG. 13 compares is backed off, and SW-1 and SW-3 are closed. At this point the the dry-spacecraft-loaded-mass-ratio vs. required vehicle correct NOFB blend has been manufactured, so the condens velocity changes for NOFB monopropellants as compared to ing tank is placed in a cold bath adequately cold to condense hydrazine. the mixture but sufficiently above the blends freezing point. One implementation uses a cold bath sustained at ~-70 C. EXAMPLE1 Sufficient time is allowed for the mixture to condense, and IS-4 and SW-6 are closed. If sufficient monopropellant is 0035. The mixing of fuels and oxidizer must be done in a manufactured with one condensation, the condensed liquid controlled, measured manner to ensure the resulting mono tank can be removed from the system (between IS-4 and propellant has desired performance characteristics. SW-6) and allowed to equilibrate back with room tempera 0036 FIG. 2 demonstrates an exemplary schematic of an ture. If multiple loads are required, the previous steps can be apparatus used to manufacture NOFB monopropellant repeated, with the exception that gases are only mixed in the blends. Thruster performance is dependent on the propellant mixing tank and condensing consists of opening SW-6 and which is combusted. For this reason it is generally important to accurately mix the monopropellant blends. A specialized IS-4. apparatus can be used to mix high vapor pressure monopro EXAMPLE 2 pellants. Essentially, the constituents can be mixed in their vapor phases and condensed in a separate container to form a 0037 Candidate fuel blends were made and tested. The high density liquid monopropellant. The method and appara most promising blends were selected based on the following tus outlined below are for exemplary purposes, and deriva criteria: combustion and theoretical engine performance; pro tions hereof may be equally acceptable manufacturing meth pellant stability; equilibrium and non-equilibrium miscibility ods. In this implementation, SW-il indicates a general on/off performance; combustion limits, flame temperature, and valve, REG-# indicates a pressure reducing regulator, and exhaust gas chemistry for engine design; propellant phase IS-is are tank isolation valves. Unless otherwise noted, all diagram properties, and combustion reaction rates. valves begin closed and regulators backed completely off. A 0038. The monopropellants of the present invention are pressure transmitter is attached to an open SW-5 valve to named in the following manner. “NOFB designates nitrous accurately monitor system pressure. To begin manufacturing, oxide fuel blend. The next number designates the place in the US 2009/O 133788 A1 May 28, 2009

C2 group; 1 is ethane, 2 is ethylene, and 3 is acetylene. The nitrogen tetroxide/monomethylhydrazine and significantly next number indicates the oxidizer to fuel ratio. Thus, higher that the hydrazine monopropellant. “NOFB34” is nitrous oxide blended with acetylene with an 0042 Minimum impulse bits are the minimum thrustx oxidizer to fuel ratio of 4. Additional letters (a, b, c) after the time that a propulsion system can impart. Characterizing oxidizer to fuel ratio number may be used to describe devia minimum impulse bit performance of a propulsion system is tions in the blend. For example, an NOFB34 blend may important for aerospace applications such as spacecraft pre include Small amounts of specific additives to improve mix cision attitude control and miniature vehicle maneuvers. ture chemistry degassing characteristics. The first discovered Typically, expected propulsion Isp performance decreases as adaptation to this blend beyond the basic nitrous oxide and propulsion systems try and achieve Smaller minimum fuel chemistry would therefore be denoted NOFB34a. impulse bits. Therefore, more spacecraft propellant must be I0039 FIG. 1 illustrates the theoretical I performance of a flown for missions that operate in a regime of Small impulse nitrous oxide/acetylene (NO/CH) monopropellant blend bit performance. A number of factors influence this degrada as a function of oxidizer-to-fuel (O/F) mass ratio as well as tion in performance: 1) In hydrazine systems, catalyst beds showing data from recent prototype engine test results based for decomposing the monopropellant must be brought up to on measuring integrated chamber pressure and propellant optimal operation temperatures to achieve more complete mass consumed during an engine run. (Additional details on decomposition of the monopropellant. In many cases, Small the particular experimental method used for acquiring the pulsed flows of hydrazine may not allow optimal bed tem experimental measurement are discussed in 0043 below). peratures to be achieved. NOFB monopropellants are most The experimentally measured I was acquired for an O/F commonly spark-ignited for rocket propulsion applications ratio of 4 (errors bar based on uncertainty in actual nozzle and are not performance-limited by these type of catalyst coefficient during terrestrial testing). The two sets of theoreti beds, 2) the minimum impulse bit that can be achieved is cal curves (vacuum and 200/1) are shown for two different directly associated with the minimum mass of propellant that cases, equilibrium and frozen-at-the-throat chemical kinet can be discharged. This minimum propellant Volume is asso ics. These are typical bounding scenarios for actual rocket ciated with the density of the monopropellant and the small engine performance in space applications. The vacuum con hardware volumes between a valve and through the reaction dition is from an ideal exit nozzle that is infinitely long. The chamber. NOFB monopropellants can be operated at very low 200/1 nozzle is a more realistic diverging nozzle Scenario pressures (<<100 psia) where the NOFB monopropellant gas where the exit plane area is 200 times larger than the mini has densities that are

inert materials (i.e. specific grades of metals). These are very event occurred. The flashback arrestor is utilized to isolate the high temperature limits, and, in fact, a regeneratively-cooled main valve and the pressure transducer in the case of an event (propellant cools combustion chamber) NOFB monopropel Such that they are not destroyed. In this implementation, the lant engine has been developed and tested (discussed below ball valve stem was electronically isolated from the rest of the and shown in FIG. 11) that takes advantage of the high exem system via nylon gears. One possible failure mode could be plary thermal decomposition limits of NOFB monopropel electric charging of a valve stem causing a spark to propagate lants in order to provide a desirable design mechanism for within the propellant stream. Utilizing this system (and slight developing long life-cycle engines. variations hereof), over 8,000 on/off cycles have been run 0050 Additionally, accidental dry spark ignition can without a single event recorded at pressures of 100 psia (com ignite environmentally-exposed solid propellants, and there mon feed system line pressures for valves). Flight valves are fore extreme care must be taken to avoid accidental spark qualified in a similar experimental configuration with the Sources and Surface charging/discharging environmental con range of anticipated NOFB fluid properties at the valve inter ditions. Unlike solid propellants, NOFB monopropellants, by face. their nature of containment, are stored in sealed metal con 0053 FIG. 6 illustrates exemplary nozzle coefficient val tainers that behave as Faraday cages (prevents buildup of ues, Cf. for use in vacuum equivalent I engine tests charge) which essentially eliminates the possibility of dry described above. Because it is not always economical or spark ignition. Care in propulsion system design still must be possible to take measurements of an engine inside a vacuum taken with devices that could disrupt the continuous Faraday chamber thrust stand, Scaling calculations can be made which cage Such as valves with insulating valve seats and plumbing estimate what the vacuum equivalent I performance would interfaces, for example. Furthermore, the NOFB monopro be based on experimental performance observed in atmo pellants have been shown to have very high breakdown volt spheric conditions with flow that achieves sonic velocities at ages (>>10's kV) at common terrestrial tank storage tempera the minimum diameter of the engine (the throat). By calcu tures and associated pressures (in fact, NO has been lating a theoretical nozzle coefficient, Cf, determined from commonly used as a high Voltage gas insulator for high Volt exhaust gas chemistry through a nozzle expansion using equi age applications). The Paschen curve minimum breakdown librium chemical analysis software such as NASA's CEA Voltage gap of NO at even a very low storage pressure of program (Gordon and McBride (1994), “Computer Program ~100 psia is <0.001 mm (see FIG. 9). This very small maxi for Calculation of Complex Chemical Equilibrium Compo mum gap distance is significantly smaller than the NOFB sitions and Applications', NASA Reference Publication quenching distance (distance through which a flame cannot 1311) (as shown in FIG. 6), a relatively quick experimentally propagate as discussed below and experimentally shown in observed I measurement can be determined within typically FIG. 10) suggesting that even if you could directly expose the tighterror bars by measuring the integrated chamber pressure stored NOFB monopropellant to high voltages, it would not and monopropellant mass consumed during an experimental be possible to easily ignite. Furthermore, these associated engine run. Basically, the dictating equation is: 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 (Nozzle Coefficient)(Throat Area) unintentional spark ignition of NOFB monopropellants is not (Time-Integrated Chamber Pressure) likely. Intentional repeated spark ignition has been demon Sp= Mass of Propellant Consumed strated (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 0054 The nozzle coefficient can also be used to determine and spark ignition system to ensure engine ignition at startup the engine thrust in vacuum from the following equation: that occurs near the Paschen curve minimum (point where Thrust=(Chamber Pressure) (Throat Area)(Nozzle minimum Voltage is required to propagate a spark through a Coefficient) gas). 0055 FIG. 7A illustrates thermal decomposition data for 0051. The realistic ignition source in the environment is one NOFB monopropellant formulation, while FIG. 7B evaluated for its potential to initiate a combustion process. As shows a summary of decomposition Go/NoGo test vs. NOFB briefly discussed above, valves impart mechanical energy pressure for a different exemplary NOFB monopropellant. into a fluid stream which could feasibly be converted into an This metric is of specific interest for regeneratively cooled electrical discharge through triboelectric charging as a valve engine designs and in defining safe temperature handling component slides across an insulting interface (i.e. valve limits. Regeneratively cooled engines use the propellant seat). To conduct preliminary experiments to determine flowed through a jacket in the combustion chamber wall as a whether valves are a realistic ignition mechanism, an auto coolant to help maintain the combustion chamber walls below mated valve cycle test in the presence of NOFB monopropel thermal failure limits. This energy acquired during wall cool lant has been implemented. Essentially, a geared DC servo ing is not lost but rather results in hotter propellant being motor was coupled to a valve with electronic triggers to both injected backinto the chamber (hence the name regenerative). count Valve cycles and control the servo motor. While most propellants have limited cooling capacity asso 0052. The thermocouple and pressure transducer were ciated with the liquid specific heat of a propellant (energy coupled into a data acquisition system and signals fed into a required to heat the liquid by a certain change in temperature), computer program which monitored the processed signals. the NOFB monopropellants, have very high vapor pressures. The thermocouple was an exposed tip /16" K type thermo By intentionally creating a pressure drop in the regenerative couple (to reduce time lag in event detection). The pressure jacket, NOFB monopropellants can be forced to “flash” or transducer was used to ensure there was not a slow leak in the vaporize and absorb Substantially more energy from the com system therefore reducing uncertainty in the case that an bustion chamber walls by going through a phase change US 2009/O 133788 A1 May 28, 2009

(liquid vaporizing into a gas). This is a similar concept to how through a solid into the unreacted monopropellant must also a refrigerator works and is much more effective at cooling be ultimately considered. FIG. 10 illustrates experimental combustion chamber walls. In other regeneratively-cooled data of sintered metal pore sizes sufficient for quenching an designs and applications, advanced jacket design techniques NOFB monopropellant that has been intentionally detonated that enhance heat transport into the NOFB monopropellant to produce a flashback. These quenching distances have been (particularly for the case of flowing NOFB gases) by increas incorporated into the design of an anti-flashback system using ing jacket Surface area or enhancing boundary layer tempera pores sizes that are equivalent or Smaller than the ones that ture gradients may be used to regeneratively cool the engine didn't allow flame propagation as shown in FIG. 10. without “flash-cooling’. In either scenario, the maximum 0060 Propellants in general can undergo chemical reac cooling capacity of the monopropellant is limited by the tions with storage and feed system hardware that alter the thermal decomposition limit of the monopropellant. chemistry of the propellant over time. Preliminary long dura 0056 FIG. 8A illustrates the large enthalpy of vaporiza tion testing of candidate NOFB mixtures has shown them to tion (energy absorbed during vaporization) of an NOFB be chemically stable in the presence of common aerospace monopropellant derived from the Phase Diagram shown in propulsion system materials (e.g. stainless steel, Teflon). In FIG. 4 and compared the energy absorbed in a typical coolant this case, three different monopropellant blends were that is heated from the same starting temperatures to ~300° C. exposed to Teflon and stainless steel and allowed to sit for 1.5 FIG. 8B (derived from FIG. 4) illustrates the rapid tempera years at room temperature. No chemical alteration of the ture decrease as the propellant is “Flash-cooled' started with NOFB monopropellant has been observed as indicated by different tank temperatures and associated tank densities and Fourier Transform InfraRed (FTIR) absorption spectroscopy. flowing the propellantthrough any device and/or medium that 0061 FIG. 12 illustrates exemplary low thrust, non-opti causes a pressure drop. (Note quality as shown in this figure is mized engine run test data in an engine utilizing a NOFB the percent gas by mass in a liquid/gas mix in equilibrium). (nitrous oxide fuel blend) monopropellant. This figure is FIG. 8B is also critical for evaluating feedline propellant included to demonstrate successful thruster performance uti densities that feed an engine when considering the design of lizing NOFB monopropellant blends in a flight-like configu the anti-flashback systems described below, as well as tem ration. Thrust was calculated based on nozzle coefficients for perature limits within which monopropellant feed system vacuum equivalent expansion and engine pressures. hardware must operate. FIG. 11 illustrates the successful 0062 FIG. 13 illustrates a comparison of delivered pay operation of a regeneratively-cooled NOFB thruster demon load mass (minus tankage) to total wet mass (fueled vehicle) strating the principle of an NOFB flash-cooled engine. This is versus imparted vehicle velocity change for example NOFB one important feature of the NOFB monopropellants given monopropulsion systems relative to a hydrazine system the very high combustion chamber temperatures (see FIG. 1) assuming different tankage (percentage of rocket propulsion that make even exotic high temperature combustion chamber dry mass relative to total propulsion system mass) as a func material designs typically not feasible to implement. For tion of required spacecraft changes in Velocity. comparison, monopropellant hydrazine has an exhaust gas temperature of ~1600° C. EXAMPLE 4 0063 A small 4 cylinder engine (160 cc) was modified for 0057 FIG. 9 illustrates exemplary Paschen curve mini use with the NOFB monopropellants of the present invention mum (worst case optimum pressurexgap distance conditions to test the concept of using the NOFB monopropellant for for propagating a spark across two parallel Surfaces) spark operating extremely high altitude military aircraft engines propagation distance for pure NO (main NOFB constituent) and power Supplies for launch vehicles and manned space as a function of gas pressure. At room temperature storage craft applications (NASA's Apollo 13 mission was almost lost pressures, spark gap distances must be <0.0001 mm. Such because of the lack of a back-up power Supply that could have Small associated spark Volumes are unlikely to allow inad operated from the onboard ). In order to vertent NOFB monopropellant ignition since exemplary utilize the NOFB monopropellant in this type of application, NOFB quenching distances are at least ten times greater as it was necessary to modify the injection manifold, timing, discussed below and shown in FIG. 10. spark-gap, cylinder head, and starter/ignition system relative 0058 Monopropellants can be sensitive to shock which to the original parameters used in a gasoline/air engine. The initiates a rapid chemical reaction (i.e. detonation) resulting engine was tested with the nitrous oxide rocket fuel blends of in catastrophic system failure. Impact drop testing from 5.5 the present invention comprising either ethylene oracetylene. meters has shown exemplary NOFB monopropellants to be While the engine hardware associated with these applications insensitive to impact-induced detonation. is different from the hardware identified, the 0059 Because liquid monopropellants comprise com NOFB monopropellants are still fundamentally the same as bined fuel and oxidizer, they can form a potential ignition the rocket monopropellants and the same advantages previ mechanism (a.k.a "flashback') back into their storage tank. ously identified in combustion performance, non-toxicity, Therefore, a mechanism for preventing flashback must be fluid-handling characteristics, and rapid combustion kinetics included in the engine and feed system design. A very impor relative to hydrazine, for example, apply. Hydrazine-based tant parameter for designing an engine injector and flashback engines exist for alternative applications, but, similar to the control mechanism is the quenching distance of a monopro rocket application, a major limitation to widespread use of pellant. This is the smallest flowpath dimension through hydrazine in these applications relative to NOFB monopro which a flashback flame can propagate. In practice this pellants is the much lower energy density and toxicity of dimension is affected by additional parameters such as tortu hydrazine. osity (curviness of flow path) and to a lesser extent the tem perature of the solid containing the flowpath. Smaller flow EXAMPLE 5 path sizes will quench a flame and, in general, prevent 0064. The monopropellants of the present invention can be flashback although secondary ignition through heat transfer used in deployment system architecture. This is particularly US 2009/O 133788 A1 May 28, 2009

beneficial when an overall NOFB monopropulsion system is What is claimed is: already required for applications associated with the deploy 1. A monopropellant comprising nitrous oxide and at least ment application. The present invention has also been studied one hydrocarbon fuel. for use in an inflatable/rigidizable pressurized propeller, for a 2. The monopropellant of claim 1 wherein the hydrocarbon wing spar and Sustaining wing gas pressure, and for an inflat fuel is selected from the group consisting of ethane, ethylene, able/rigidizable rover wheel. and acetylene. 3. The monopropellant of claim 1, wherein the oxidizer 0065. The basic system uses the liquid to combustible gas to-fuel ratio is about 2.5 to about 11.0. generator for rapid deployment, and a Sustaining gas-pressure 4. The monopropellant of claim 3, wherein the oxidizer for robust longterm deployment of wings and/or deployables. to-fuel ratio is about 4.0 to about 8.0. 0066. The exemplary lightweight rigidizable wheel is 5. The monopropellant of claim 4, wherein the oxidizer designed to provide a wheel sized at about 1.5 meters, for less to-fuel ratio is about 4.5 to about 7.5. than 1 hazard per 100 m in aggressive 25% rock abundant 6. A monopropellant comprising nitrous oxide and ethane Mars terrains and ability to navigate with 30 cm/pixel orbital in an oxidizer-to-fuel ratio of about 2.5 to about 11.0. resolution. Further, the wheel supports more than 100 kg per 7. The monopropellant of claim 6, wherein the oxidizer <10kg wheel. The wheel employs a set of inflatable shells and to-fuel ratio is about 3.0 to about 9.0. has a composite rim. 8. The monopropellant of claim 7, wherein the oxidizer to-fuel ratio is about 4.0 to about 8.0. 0067. The exemplary wing spars utilize the monopropel 9. The monopropellant of claim 8, wherein the oxidizer lants of the present invention with an inflatable/rapid rigidiz to-fuel ratio is about 4.5 to about 7.5. ing wing spar (combustion/flash-cool) for providing rela 10. A monopropellant comprising nitrous oxide and ethyl tively stiff wing to maintain stable C and C, across wing to ene in an oxidizer-to-fuel ration of about 2.5 to about 11.0. achieve high overall L/D. The characteristics obtained are 11. The monopropellant of claim 10, wherein the oxidizer shown in FIG. 14. to-fuel ratio is about 3.0 to about 9.0. 0068. The monopropellants of the present invention are 12. The monopropellant of claim 11, wherein the oxidizer also used in deployment systems to provide inflatable/rigidiz to-fuel ratio is about 4.0 to about 8.0. able propellers. 13. The monopropellant of claim 12, wherein the oxidizer to-fuel ratio is about 4.5 to about 7.5. 0069. The deployables of the present invention may also 14. A monopropellant comprising nitrous oxide and acety contain an annihilation mechanism for post-operational life. lene in an oxidizer-to-fuel ration of about 2.5 to about 11.0. This contingency option can be used for deployment deep 15. The monopropellant of claim 14, wherein the oxidizer behind enemy lines where recovery may not be an option. to-fuel ratio is about 3.0 to about 9.0. 0070. In these applications, the NOFB rocket monopro 16. The monopropellant of claim 15, wherein the oxidizer pellant used initially for propulsive applications is also used to-fuel ratio is about 4.0 to about 8.0. to operate these additional auxiliary deployment and opera 17. The monopropellant of claim 16, wherein the oxidizer tional modes. to-fuel ratio is about 4.5 to about 7.5. 0071. The present specification provides a complete 18. A monopropellant comprising nitrous oxide and two or description of compositions of matter, methodologies, sys more of acetylene, ethane or ethene in an oxidizer-to-fuel tems and/or structures and uses in example implementations ratio of about 2.5 to about 11.0. of the presently-described technology. Although various 19. The monopropellant of claim 18, wherein the oxidizer implementations of this technology have been described to-fuel ratio is about 3.0 to about 9.0. above with a certain degree of particularity, or with reference 20. The monopropellant of claim 19, wherein the oxidizer to one or more individual implementations, those skilled in to-fuel ratio is about 4.0 to about 8.0. the art could make numerous alterations to the disclosed 21. The monopropellant of claim 20, wherein the oxidizer implementations without departing from the spirit or scope of to-fuel ratio is about 4.5 to about 7.5. the technology hereof. Since many implementations can be 22. The monopropellant of claim 1, wherein additional made without departing from the spirit and scope of the constituents comprise less than about 30% of the monopro presently described technology, the appropriate Scope resides pellant. in the claims. Other implementations are therefore contem 23. The monopropellant of claim 2 comprising nitrous plated. Furthermore, it should be understood that any opera oxide and one or more of acetylene, ethane, or ethane where tions may be performed in any order, unless explicitly the fuel(s) are mixed with nitrous oxide in the gas phase claimed otherwise or a specific order is inherently necessi during the manufacturing process prior to condensing into a tated by the claim language. It is intended that all matter liquid. contained in the above description and shown in the accom 24. The monopropellant of claim 2 comprising nitrous panying drawings shall be interpreted as illustrative only of oxide and one or more of acetylene, ethane, or ethane where particular implementations and are not limiting to the the fuel(s) are mixed with nitrous oxide in the liquid phase embodiments shown. Changes in detail or structure may be during the manufacturing process made without departing from the basic elements of the 25. The monopropellant of claim 2 comprising nitrous present technology as defined in the following claims. In the oxide and one or more of acetylene, ethane, or ethane where claims of any corresponding utility application, unless the the fuel(s) are mixed with nitrous oxide in any combination of term “means' is used, none of the features or elements recited gas and liquid phases during the manufacturing process. therein should be construed as means-plus-function limita tions pursuant to 35 U.S.C. S112,6. c c c c c