U.S. patent application number 12/268266 was filed with the patent office on 2009-05-28 for nitrous oxide fuel blend monopropellants.
This patent application is currently assigned to Firestar Engineering, LLC. Invention is credited to Benjamin Carryer, David J. Fisher, Christopher Mungas, Gregory Mungas.
Application Number | 20090133788 12/268266 |
Document ID | / |
Family ID | 40626235 |
Filed Date | 2009-05-28 |
United States Patent
Application |
20090133788 |
Kind Code |
A1 |
Mungas; Gregory ; et
al. |
May 28, 2009 |
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.
Inventors: |
Mungas; Gregory; (Arcadia,
CA) ; Fisher; David J.; (Denver, CO) ; Mungas;
Christopher; (Plymouth, CA) ; Carryer; Benjamin;
(Denver, CO) |
Correspondence
Address: |
HENSLEY KIM & HOLZER, LLC
1660 LINCOLN STREET, SUITE 3000
DENVER
CO
80264
US
|
Assignee: |
Firestar Engineering, LLC
Broomfield
CO
|
Family ID: |
40626235 |
Appl. No.: |
12/268266 |
Filed: |
November 10, 2008 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
|
|
60986991 |
Nov 9, 2007 |
|
|
|
Current U.S.
Class: |
149/74 |
Current CPC
Class: |
C10L 3/02 20130101; C06B
47/04 20130101; C06D 5/08 20130101 |
Class at
Publication: |
149/74 |
International
Class: |
C06B 47/04 20060101
C06B047/04 |
Goverment Interests
STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH OR DEVELOPMENT
[0002] This invention was supported in part by subcontract number
1265181 from the California Institute of Technology Jet Propulsion
Laboratory/NASA. The U.S. Government may have certain rights in the
invention.
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.
Description
CROSS-REFERENCE TO RELATED APPLICATIONS
[0001] This application claims benefit of U.S. Provisional
Application No. 60/986,991, entitled "Nitrous Oxide Fuel Blend and
Monopropellants" and filed on Nov. 9, 2007, which is specifically
incorporated herein by reference for all that it discloses and
teaches.
BACKGROUND
[0003] Liquid fueled rockets have better specific impulse
(I.sub.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.sub.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, O.sub.3, ClF.sub.3, ClF.sub.5, are unstable,
energetic and toxic.
[0004] 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.sub.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:
N.sub.2ON.sub.2+1/2O.sub.2+Heat
[0005] 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.sub.sp
of this reaction is 205 s. N.sub.2O is a highly stable molecule
given its high activation energy barrier .about.250 kJ/mol. As a
result, thermal decomposition requires preheat temperatures
>1000.degree. C. Alternatively, catalysts can be used to
significantly depress this activation energy. However, the hot
(>1500.degree. C.), highly oxidizing reaction products make
catalyst bed and reaction chamber design challenging.
[0006] 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 .about.1500 Whr/kg (.about.3
times the energy density of pure N.sub.2O), and I.sub.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 N.sub.2O 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.sub.sp
performance using low cost engine fabrication techniques.
[0007] The highest I.sub.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.degree. C.
and the lithium must be kept above 180.degree. 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
[0008] 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
(N.sub.2O). When combusted, the nitrous oxide provides both thermal
decomposition energy and serves as the oxidizer to combust the
fuels. Example organic fuels include ethane (C.sub.2H.sub.6),
ethylene (C.sub.2H.sub.4), acetylene (C.sub.2H.sub.2), 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.sub.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.degree. C. Therefore, it is possible to
dissolve fuels into the N.sub.2O 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.
[0009] 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.
[0010] 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.
[0011] 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.
[0012] 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
[0013] FIG. 1 is a graph of theoretical and actual I.sub.sp
performance of an NOFB monopropellant formulation.
[0014] FIG. 2 illustrates the method of making the nitrous oxide
fuel blends of the present invention.
[0015] FIG. 3 is a chart summarizing NOFB monopropellant
characteristics relative to monopropellant hydrazine and
bipropellant nitro tetroxide/monomethylhydrazine.
[0016] 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
[0017] FIG. 5A is an FTIR spectrum of NOFB monopropellant sampled
over different tank temperatures (and comparison with the remaining
gas after a 1/4 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.sub.sp performance is also shown as the blend slightly varies
during a very aggressive tank liquid expulsion (80% NOFB liquid
expulsion in .about.seconds)
[0018] FIG. 6 is a graph illustrating exemplary nozzle coefficient
values for use in vacuum equivalent I.sub.sp calculations.
[0019] 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.
[0020] 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 .about.300.degree. 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.
[0021] 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.
[0022] FIG. 10 is a graph of quenching distance based on
oxidizer-to-fuel ratios for one NOFB monopropellant
formulation.
[0023] FIG. 11 illustrates an exemplary NOFB regeneratively-cooled
thruster utilizing the high volatility of the NOFB monopropellant
to "flash-cool" the combustion chamber.
[0024] FIG. 12 illustrates low thrust, non-optimized engine run
test data in an engine utilizing an exemplary NOFB
monopropellant.
[0025] 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.
[0026] FIG. 14 summarizes the characteristics of the NOFB
deployable wing spars.
DETAILED DESCRIPTION
[0027] 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.
[0028] 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.
[0029] 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."
[0030] 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.
[0031] 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.
[0032] 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.
[0033] 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 .about.3.5 to 3.9 times greater than hydrazine. For
comparison, state-of-the-art lithium ion batteries store .about.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.
[0034] 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.sub.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/s.sup.2)
such that I.sub.sp has units of seconds regardless of what
international system of units are being used (English or System
Internationale (SI) units)). Higher I.sub.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.sub.sp performance [Propellant
Mass=Spacecraft_Dry_Mass.times.exp[Change_in_Spacecraft_Velocity/I.sub.sp-
/earth gravity]-1], high I.sub.sp propellants are very attractive
for demanding aerospace applications. In chemical propulsion
systems, in order to achieve high I.sub.sp systems, exothermic
chemical reactions are generally required. Currently, a common
industry standard commercial monopropellant, hydrazine, has an
I.sub.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.sub.sp values of up to 345 s
and potentially larger I.sub.sp values. Recent experimentally
measured engine I.sub.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
[0035] The mixing of fuels and oxidizer must be done in a
controlled, measured manner to ensure the resulting monopropellant
has desired performance characteristics.
[0036] 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 .about.-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
[0037] 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.
[0038] 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.
[0039] FIG. 1 illustrates the theoretical I.sub.sp performance of a
nitrous oxide/acetylene (N.sub.2O/C.sub.2H.sub.2) 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.sub.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.
[0040] 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.sub.sp performance is achieved. However, certain design
considerations place additional constraints on optimal O/F ratios.
The optimal N.sub.2O-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.
[0041] 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.sub.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.sub.sp performance for the NOFB blend is comparable to the
bipropellant nitrogen tetroxide/monomethylhydrazine and
significantly higher that the hydrazine monopropellant.
[0042] Minimum impulse bits are the minimum thrust.times.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 .about.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
[0043] In addition to I.sub.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 N.sub.2, CO, H.sub.2O, H.sub.2 and
CO.sub.2, where ammonia gas is an exhaust product of hydrazine.
[0044] For spacecraft science missions, ammonia is an undesirable
byproduct because of its reactions with soils that can readily
complicate and contaminate sensitive soil measurements.
[0045] 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 (.about.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 .about.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.
[0046] 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.
[0047] For low temperature operations and storage considerations,
NOFB monopropellant densities increase significantly up to .about.1
g/cc at -75.degree. C. and freeze at <-80.degree. 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 .about.0.degree. 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.
[0048] 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 .about.2 s). The variation in
modeled I.sub.sp performance for two bounding scenarios
(frozen-at-the-throat chemistry and equilibrium chemistry
throughout nozzle) for these extreme tank expulsion cases is
.about.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.
[0049] 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
.about.10's.degree. C. Heated capillary tube testing has
demonstrated that exemplary NOFB monopropellants have thermal
ignition temperatures that are .about.400.degree. C. (FIGS. 7A and
7B) and may be as high as 650.degree. 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.
[0050] 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, N.sub.2O has been
commonly used as a high voltage gas insulator for high voltage
applications). The Paschen curve minimum breakdown voltage gap of
N.sub.2O at even a very low storage pressure of .about.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).
[0051] 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.
[0052] 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.
[0053] FIG. 6 illustrates exemplary nozzle coefficient values, Cf,
for use in vacuum equivalent I.sub.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.sub.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.sub.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 = ( Nozzle Coefficient ) ( Throat Area ) ( Time - Integrated
- Chamber Pressure ) Mass_of _Propellant _Consumed ##EQU00001##
[0054] 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)
[0055] 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.
[0056] 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 .about.300.degree. 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 .about.1600.degree.
C.
[0057] FIG. 9 illustrates exemplary Paschen curve minimum (worst
case optimum pressure.times.gap_distance conditions for propagating
a spark across two parallel surfaces) spark propagation distance
for pure N.sub.2O (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.
[0058] 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.
[0059] 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.
[0060] 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.
[0061] 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.
[0062] 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
[0063] 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
[0064] 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.
[0065] 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.
[0066] 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.
[0067] 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.sub.L and C.sub.D across wing to achieve high
overall L/D. The characteristics obtained are shown in FIG. 14.
[0068] The monopropellants of the present invention are also used
in deployment systems to provide inflatable/rigidizable
propellers.
[0069] 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.
[0070] In these applications, the NOFB rocket monopropellant used
initially for propulsive applications is also used to operate these
additional auxiliary deployment and operational modes.
[0071] 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. .sctn.112,
6.
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