TECHNICAL FIELD
The present disclosure relates to the field of rocket propellants. More specifically, it relates to a propellant composition comprising mixed oxides of nitrogen as oxidizer compound, and fuel, along with methods and rocket propulsion applications thereof. The present disclosure also evaluates the performance of said propellant for use in rocket propulsion system. In particular, the present disclosure relates to propellant composition comprising MON-10 and UH-25, method of preparing said propellant composition and applications thereof. Further, said propellant composition of the present disclosure is chemically stable and successfully operates at low temperature, for instance -20oC. The present disclosure also relates to a device for combustion of the propellant composition.

BACKGROUND AND PRIOR ART OF THE DISCLOSURE
The requirement for high performance propulsion systems for space access and satellites has been existing for decades. Propulsion systems are needed for a variety of tasks including rocket boost, orbit insertion and maintenance, attitude control systems (ACS), reaction control systems (RCS), station keeping, orbital maneuvering systems (OMS), and auxiliary power units (APU). The draw backs and consequences associated with these systems utilizing current propellants are still daunting and research and development efforts have not greatly improved the technology over the years.

The list of potential oxidizers currently available to be used in chemical propulsion is quite limited and there are significant shortcomings associated with such oxidizers. The problems associated with current oxidizers that are used for liquid and hybrid propulsion systems are varied and well known. The high operating costs result from occupational safety requirements associated with the handling of toxic, hypergolic propellants using inherently dangerous materials. The other drawbacks includes requirement of cryogenic storage (e.g., propellants using liquid oxygen, nitrogen fluoride, etc.) or additional operating complications such as the storage of the materials in a way that prevents “boil off’ prior to usage. Cryogenic storage systems also require the use of insulation which adds dry weight to both launch and space vehicles thereby reducing the vehicle’s payload weight fraction. Other limitations of current technologies are complexity of pressurants and feed systems for the propellants, state change of the propellants during prolonged storage. The current solutions available to address the storage and handling issues severely impact the performance of the propellant.

Even though less toxic and easily stored propellant oxidizers are known [e.g., nitrous oxide (N2O)], their energy (i.e., heat of formation or ?Hf) is generally too low to provide the required performance. Research to identify new liquid propellants (i.e., both fuels and oxidizers) is needed to enhance performance and minimize the above-described undesirable properties without added complexity and cost. Based on the above shortcomings, it is clear that development of new improved propellant systems would constitute a critical technological enhancement in the field of chemical propulsion. The current challenge is to attain the high energy and density goals for these propulsion systems while maintaining acceptable physical properties for the propellants. Desirable attributes for the oxidizing component of the propellant system includes high specific impulse (Isp) performance with common fuels, high density, good combustion stability and efficiency characteristics, chemical stability, nontoxicity, storable under normal conditions, low freezing point, significant reduction in cost of operations, hypergolic behavior with common fuels, ease of handling, environmental aspects (‘green propellants’) and compatibility with tank and feed system materials. While the problem is well understood, practical solutions which meet the objectives have been elusive and research has not been very fruitful to date.

The present disclosure aims at overcoming the drawbacks of the prior art as discussed above by providing an improved propellant composition having desirable properties as mentioned above.

STATEMENT OF THE INVENTION
Accordingly, the present disclosure relates to a propellant composition comprising oxidizer and fuel, wherein the oxidizer is mixed oxides of nitrogen-10 (MON-10) and the fuel is UH-25; a method of preparing the propellant composition, said method comprising acts of: a) preparing NO gas, b) dissolving NO gas with N2O4 to obtain MON-10 comprising nitric oxide (NO) at a concentration of about 10% (w/w) and nitrogen tetroxide (N2O4) at a concentration of about 90% (w/w), and c) mixing the MON-10 with UH-25 to obtain the propellant composition comprising MON-10 and UH-25; and a device for combustion of the propellant composition, said device comprising: a) a first container comprising the mixed oxide of nitrogen-10 (MON-10) oxidizer, b) a second container comprising the UH-25 fuel, c) a combustion chamber connected to the first container and optionally to the second container, wherein said combustion chamber allows combustion of the propellant composition to produce combustion gases, and d) an outlet allowing the release of combustion gases.

BRIEF DESCRIPTION OF THE ACCOMPANYING FIGURES
In order that the invention may be readily understood and put into practical effect, reference will now be made to exemplary embodiments as illustrated with reference to the accompanying figures. However, the figures are purely for the purpose of exemplifying and are non-limiting in nature. The figures together with the detailed description below, are incorporated in and form part of the specification, and serve to further illustrate the embodiments and explain various principles and advantages, in accordance with the present invention where:

Figure 1 provides experimental set up for MON-10 preparation.
Figure 2 provides predictive test performance of oxidizer-to-fuel ratio versus specific impulse for UDMH fuel and the oxidizers mentioned in the Table 3.
Figure 3 provides predictive test performance oxidizer-to-fuel ratio versus specific impulse for UH-25 fuel and the oxidizers mentioned in the Table 3.
Figure 4 provides predictive performance comparison between UDMH-25/N2O4 and UH-25/MON-10 with respect to oxidizer-to-fuel ratio and characteristic velocity.
Figure 5 represents schematic of a rocket propulsion system using the oxidizer compound of the present invention.
Figure 6 provides the specific impulse performance of the NO/N2O4 mixtures at different NO concentrations, with the data for pure N2O4, included as reference schematic of a rocket propulsion system using the oxidizer compound of the present invention.
Figure 7 provides the characteristic velocity (C*) performance of the NO/N2O4 mixtures at different NO concentrations, with the data for pure N2O4, included as reference schematic of a rocket propulsion system using the oxidizer compound of the present invention.
Figure 8a provides 125 Kgf thruster with injector for static testing.
Figure 8b shows the connection of the thruster assembly to the propellant tanks in pressure fed mode.
Figure 9 provides variation of absolute chamber pressure with time for Static Test 1 performed on the propellant composition comprising oxidizer- MON-10 and the fuel- UH-25.
Figure 10 provides comparison of oxidizer manifold pressure with chamber pressure for Static Test 1 performed on the propellant composition comprising oxidizer- MON-10 and the fuel- UH-25.
Figure 11 shows variation of thrust at sea level with time for Static Test 1 performed on the propellant composition comprising oxidizer- MON-10 and the fuel- UH-25.
Figure 12 provides variation of absolute chamber pressure with time for static Test 2 performed on the propellant composition comprising oxidizer- MON-10 and the fuel- UH-25.
Figure 13 shows variation of thrust at sea level with time for Static Test 2 performed on the propellant composition comprising oxidizer- MON-10 and the fuel- UH-25.
Figure 14 provides variation of absolute chamber pressure with time for Static Test 3 performed on the propellant composition comprising oxidizer- MON-10 and the fuel- UH-25.
Figure 15 shows variation of thrust at sea level with time for Static Test 3 performed on the propellant composition comprising oxidizer- MON-10 and the fuel- UH-25.

DETAILED DESCRIPTION OF THE DISCLOSURE
The present disclosure relates to a propellant composition comprising oxidizer and fuel, wherein the oxidizer is mixed oxides of nitrogen-10 (MON-10) and the fuel is UH-25.

In an embodiment of the present disclosure, the MON-10 comprises nitric oxide (NO) at a concentration of about 10% (w/w) and nitrogen tetroxide (N2O4) at a concentration of about 90% (w/w).

In another embodiment of the present disclosure, the UH-25 comprises about 75% unsymmetrical dimethylhydrazine (UDMH) and about 25% hydrazine hydrate.

In yet another embodiment of the present disclosure, the propellant is chemically stable and operates at a temperature ranging from about 25oC to -20oC.

In still another embodiment of the present disclosure, the propellant is chemically stable and operates at -20oC.

In still another embodiment of the present disclosure, the composition is non-toxic, self-pressurizing, storable at room temperature, and operates at high altitudes and in space.

In still another embodiment of the present disclosure, the MON-10 is a non-viscous liquid.

In still another embodiment of the present disclosure, the propellant composition has a specific impulse ranging from about 221 seconds to 231 seconds, characteristic velocity ranging from about 1670 m/s to about 1700 m/s and chamber pressure ranging from about 27 KscA to 33 KscA.

In still another embodiment of the present disclosure, at -20oC, the propellant composition has a specific impulse of about 237 seconds, characteristic velocity of about 1690 m/s and chamber pressure of about 30.4 KscA.

The present disclosure relates to a method of preparing the propellant composition said method comprising acts of:
a) preparing NO gas;
b) dissolving NO gas with N2O4 to obtain MON-10 comprising nitric oxide (NO) at a concentration of about 10% (w/w) and nitrogen tetroxide (N2O4) at a concentration of about 90% (w/w); and
c) mixing the MON-10 with UH-25 to obtain the propellant composition comprising MON-10 and UH-25.

In an embodiment of the present disclosure, the NO gas is prepared by reducing aqueous sodium nitrite (NaNO2) in presence of ferrous sulphate (FeSO4).

In another embodiment of the present disclosure, the step (b) is carried out under cold condition, preferably at a temperature ranging from about 10oC to 0oC.

In yet another embodiment of the present disclosure, the preparation of NO gas in step (a) and dissolving the NO gas with N2O4 in step (b) is controlled by drop wise addition of the aqueous NaNO2 to the FeSO4 solution with constant stirring.

The present disclosure relates to a device for combustion of the propellant composition said device comprising:
a) a first container comprising the mixed oxide of nitrogen-10 (MON-10) oxidizer;
b) a second container comprising the UH-25 fuel;
c) a combustion chamber connected to the first container and optionally to the second container, wherein said combustion chamber allows combustion of the propellant composition to produce combustion gases; and
d) an outlet allowing the release of combustion gases.
In an embodiment of the present disclosure, said device is employed in rocket propulsion system of a vehicle selected from a group comprising rocket, turbojet, combined cycle propulsion system, launch vehicle propulsion system, multi-mode spacecraft propulsion system, upper stage spacecraft propulsion system, missile propulsion system, hybrid rocket, bipropellant liquid rocket, tripropellant rocket engine, gas generation system, internal combustion engine and combinations thereof.

An embodiment of the present disclosure relates to use of the propellant composition in a vehicle selected from a group comprising rocket, turbojet, combined cycle propulsion system, launch vehicle propulsion system, multi-mode spacecraft propulsion system, upper stage spacecraft propulsion system, hybrid rocket, bipropellant liquid rocket, tripropellant rocket engine, gas generation system, internal combustion engine and combinations thereof.

The present disclosure relates to propellant composition comprising mixed oxides of nitrogen (MON) as oxidizer compound along with fuel, methods and rocket propulsion applications thereof.

In an embodiment of the present disclosure, the oxidizer compound is a mixed oxide of nitrogen 10 - “MON- 10”.

In another embodiment of the present disclosure, the fuel is UH-25 [a mixture of 75% UDMH (unsymmetrical dimethylhydrazine) and 25% hydrazine hydrate].

The present disclosure relates to a propellant comprising oxidizer compound - MON-10 and fuel - UH-25.

In an embodiment of the present disclosure, performance evaluation of a propellant comprising MON-10 and fuel- UH-25 is carried out for its use in rocket propulsion system.
In another embodiment of the present disclosure, the oxidizer compound “MON-10” comprises nitric oxide (NO) at a concentration of about 10% (w/w) and nitrogen tetroxide (N2O4 also known as dinitrogen tetroxide) at a concentration of about 90% (w/w).
In another embodiment of the present disclosure, the fuel “UH-25” comprises 75% (w/w) of unsymmetrical dimethylhydrazine (UDMH) and 25% (w/w) of hydrazine hydrate.

In yet another embodiment of the present disclosure, the performance of the propellant is evaluated by carrying out static tests from ambient conditions (i.e. 25 oC to 30 oC) to about -20oC. In an exemplary embodiment of the present disclosure, the propellant comprising MON-10 and UH-25 is stable and performs efficiently at ambient conditions and also at -20 oC.

In still another embodiment of the present disclosure, the propellant is used in a variety of rocket propulsion systems such as but not limiting to those used in rockets, turbojets, internal combustion engines, combined cycle propulsion systems, launch vehicle propulsion systems, multi-mode spacecraft propulsion systems and upper stage spacecraft propulsion systems.

The oxidizer compound - MON-10 in the propellant composition of the present disclosure strikes a balance over a variety of oxidizer attributes that have traditionally been at odds with one another. More specifically, exemplary attributes balanced by the oxidizer compound of the present propellant composition are as follows:
- reduces storage problems by providing for room temperature storage thereof;
- reduces propulsion system weight since room temperature storage reduces need for storage tank insulation;
- provides for long-term storage since cryogenic boil-off is not a problem;
- has a relatively high specific impulse when compared to traditional energetic but inherently problematic oxidizers;
- has a relatively high energy density when compared to traditional energetic but inherently problematic oxidizers;
- has reduced toxicity over pure nitrogen tetroxide in its stored state; and
- produces environmentally benign exhaust products when burned in a propulsion system, thereby ensuring the reduced toxicity of the propellant composition.

In general, the oxidizer compound in the propellant composition of the present disclosure is a mixture of nitric oxide (NO) at a concentration of 10% (w/w) and nitrogen tetroxide (N2O4) at a concentration of 90% (w/w) that is homogenous and a stable liquid at room temperature.

In an embodiment of the present disclosure, the homogenous mixture of nitric oxide and nitrogen tetroxide (MON-10) is a non-viscous liquid.

In another embodiment of the present disclosure, as both MON-10 and UH-25 are hypergolic in nature, the products/exhaust are in gaseous phase.

In one specific embodiment, the present disclosure is directed to the use of an oxidizer, a liquid mixture of an oxide of nitrogen, such as nitric oxide, and nitrogen tetroxide at specific concentrations. In an exemplary embodiment, the oxidizer is MON-10. In another embodiment, MON-10 is a liquid mixture of nitric oxide (10% w/w) and nitrogen tetroxide (90% w/w). The use of said oxidizer maximizes the benefits and reduces shortcomings relative to other common/known oxidizer such as pure nitrogen tetroxide and pure nitric oxide. The said oxidizer compound presents environmentally friendly, high density (of about 1.403 gm/cm3 at 27oC) and high performance. The oxidizer is self- pressurizing and does not need to operate under deep cryogenic conditions.

The present disclosure is based upon the fact that the addition of nitric oxide to nitrogen tetroxide at specific concentrations (i.e. MON-10) will lower the freezing point and increase the vapor pressure of the latter without substantially reducing the specific impulse attainable when used with typical hydrocarbon fuels. The mixtures remain liquid over a wider temperature range than nitrogen tetroxide alone but still possess high oxidizing power and are much more stable. In an effort to improve the liquid range of nitrogen tetroxide with additives, it is identified that nitric oxide at 10% (w/w) is compatible with nitrogen tetroxide, forming one liquid phase in all proportions without evolution of oxygen, and most significantly the said oxidizer composition MON-10 is proved to have varying properties.

In an embodiment of the present disclosure, the oxidizer compound MON-10 of the present propellant composition is a homogenous, stable-liquid room temperature mixture.
According to another embodiment of the present disclosure, the specific concentration of 10% (w/w) of NO in MON-10 is significant in achieving the desired properties in the propellant since a higher percentage of NO decreases the oxidation potential of the propellant. Nitrogen tetroxide alone/itself has a freezing point of about -11.2 °C, whereas freezing point of MON-10 is -23 °C which is much lower than that of pure nitrogen tetroxide.

In an embodiment of the present disclosure, preparation of MON-10 is provided. NO gas is prepared by the reduction of sodium nitrite by ferrous sulphate and NO slowly dissolved in N2O4 under cold conditions. A special preparation cum storage tank is designed and fabricated using SS 304 material. The composition of the SS 304 material is Carbon (0.08% max.), Silicon (1.00% max.), Manganese (2.0% max.), Sulphur (0.03% max.), Phosphorus (0.045% max.), Nickel (8-12%), Chromium (18-20% max.), Molybdenum (1.00% max.) and Copper (1.0% max).
In order to confirm the percentage of NO mixed in sampling, analysis is carried out during the preparation by the standard method which is based on stoichiometric reaction of NO with oxygen gas. Schematic diagram of experimental set up for MON-10 preparation is given in FIG. 1. The specification of MON-10 and test results are provided in Table 1. The chemical analysis on MON-10 is carried out as per MIL (Military standard) specifications (MIL- REF- 26539E). The below Table 1 shows that MON-10 meets the MIL specifications.

Table 1: Specification and test results of MON-10

SNo.
Composition MON-10
Test method
Specification Test Results
1 N2O4 assay,
% wt., min. -- 89.40 Acid-Base titration
2 NO, % wt. 10.0 to 11.0 10.44 Oxidation
with O2
3 N2O4+NO
% wt, min. 99.5 99.84 Test results
1 + 2
4 Water equivalent,
% wt. 0.17 0.16 Difference between 100 and Test result 3 i.e (100-0.16)
5 Chloride content,
% wt. 0.040 Nil AgNO3
6 Particulate matter, mg/L 10 4.5 Gravimetric

In another embodiment of the present disclosure, the form of the homogenous mixture of nitric oxide and nitrogen tetroxide (MON-10) can also be a non-viscous liquid. Non-viscous liquid forms of MON-10 in the present disclosure are typically used in launch vehicle propulsion systems and multi-mode propulsion systems.

In an embodiment of the present disclosure, by applying a predictive test by using NASA SP 273, the oxidizer compound of the present disclosure (MON-10) is compared with other mixed oxides of nitrogen (MONs), pure N2O4 oxidizer, pure NO oxidizer in a propulsion system using UDMH/ UH-25 as the rocket fuel. The comparative values of specific impulse (Isp) of the oxidizer compound MON-10 at a particular oxidizer-to-fuel (O/F) ratio of 2: 1, along with typical exhaust products against the aforementioned propellants are presented in Tables 2a and 2b.

Table 2a: Comparisons of specific impulse (ISP) along with typical mole fractions of exhaust products with UDMH as fuel

Oxidiser: NO-N2O4 Mixture; Fuel: UDMH ; O/F: 2:1
%NO in N2O4 0 3 10 30 40
Isp (in seconds) 268.5 269.2 270.2 273.2 274.7
CO .18345 .18416 .18573 .18946 .19096
CO2 .05202 .05070 .04771 .04001 .03657
H .01849 .01885 .01970 .02224 .02357
HCO .00005 .00005 .00005 .00006 .00006
H2 .14533 .14776 .15345 .16987 .17814
H2O .31001 .30623 .29740 .27211 .25948
NH .00001 .00001 .00001 .00001 .00001
NH2 .00001 .00001 .00001 .00001 .00001
NH3 .00001 .00001 .00001 .00001 .00001
NO .00216 .00244 .00239 .00224 .00216
N2 .00246 .27258 .27667 .28806 .29360
O .027081 .00116 .00115 .00111 .00109
OH .01519 .01508 .01484 .01410 .01371
O2 .00100 .00097 .00089 .00071 .00064

Table 2b: Comparisons of specific impulse (ISP) along with typical mole fractions of exhaust products with UH-25 as fuel

Oxidiser: NO-N2O4 Mixture; Fuel: UH-25 ; O/F (oxidizer:fuel): 2:1
%NO in N2O4 0 3 10 30 40
Isp (in seconds) 266.9 265.0 266.2 269.5 271.2
CO .11608 .11733 .12011 .12709 .13008
CO2 .07070 .06897 .06506 .05488 .05033
H .01557 .01600 .01702 .02009 .02171
HCO .00002 .00003 .00003 .00003 .00003
H2 .08245 .08455 .08951 .10422 .11184
H2O .38112 .37760 .36932 .34522 .33297
NH .00001 .00001 .00001 .00001 .00001
NH2 - -- -- -- .00001
NO .00652 .00646 .00632 .00591 .00571
N2 .28975 .29161 .29592 .30788 .31368
O .00324 .00323 .00321 .00315 .00311
OH .02851 .02841 .02816 .02730 .02680
O2 .00600 .00578 .00531 .00417 .00370
HO2 .00001 .00001 .00001 .00001 .00001
N .00001 .00001 .00001 .00001 .00001

Table 3: Comparison of properties (specific impulse, characteristic velocity and workability at low temperature) of MON-10/UH-25 versus other propellant compositions

Properties
MON-10 + UH-25 (Present Invention) 100% N2O4 + UH-25 MON-3+ UH-25 MON-30+ UH-25 MON-40+UH-25 100% N2O4 + UDMH MON-10+ UDMH MON-10+ hydrazine
Isp (in seconds) 266.2 266.9 265 269.5 271.2 268.5 270.2 256.6
Characteristic velocity C* (in m/s) 1682 1678 1681 1711 1720 1688 1670 1624
Chemical stability and Workability/Combustion at low temperature, i.e. at -20 oC Yes No No No No No No No
Inference from Tables 2a, 2b and 3: As is readily apparent from the data/results of Tables 2a, 2b and 3, the oxidizer compound MON-10 in the propellant composition of the present disclosure provides comparable to better values of specific impulse performance, characteristic velocity and exhaust products at oxidizer-to-fuel ratio of 2:1 and even at low temperatures, when compared to that of other propellants. More importantly, the present propellant composition made up of MON-10 and UH-25 is workable at low temperatures (eg. at -20oC) while other propellants do not possess applications/workability at low temperatures. Further, the freezing points of MON-3 oxidizer and other conventional oxidizers such as 100% N2O4, are -15 oC and -11.2oC respectively and the propellant compositions (such as 100% N2O4+ UH-25, MON-3+ UH-25, MON-30+ UH-25, MON-40+UH-25, 100% N2O4+ UDMH, MON-10+ UDMH, MON-10+ hydrazine) comprising these oxidizers are not suitable for low temperature applications at -20oC.
With regard to propellant compositions comprising MON-30 and MON-40, it is observed that the solubility of NO in N2O4 is lesser when compared to MON-10, which makes MON-30 and MON-40 more viscous in nature. Further, the method of preparation of MON-30 and MON-40 is more cumbersome as compared to MON-10. On the other hand, MON-10 is a unified and homogeneous mixture owing to good solubility of NO in N2O4. Also, MON-10 is easy to prepare.
In addition to the aforementioned advantages in terms of properties/efficiency, it is also observed that the vehicle performance is improved when the propellant composition of the present disclosure is employed owing to smaller storage tank requirements. Further, since the oxidizer MON-10 of the present disclosure can be stored at room temperature, the problems associated with the cryogenic storage of liquid oxygen are also eliminated.

The nitric oxide-nitrogen tetroxide oxidizer compound (MON-10) of the present disclosure is safe to handle and produces benign exhaust products as is evidenced by the data in Tables 2a and 2b. To further illustrate the advantages of the present invention, FIG. 2 shows predictive test performance comparison of oxidizer-to-fuel ratio versus specific impulse (Isp) and characteristic velocity (C*) for UDMH in combination with various oxidizers, and FIG. 3 shows predictive test performance of oxidizer-to-fuel ratio versus specific impulse and characteristic velocity (C*) respectively for UH-25 fuel in combination with various oxidizers..

Further, the predictive performance comparison (using NASA SP 273) between MON-10 [i.e. 10% (w/w) of NO and 90% (w/w) of N2O4] -UH-25 and pure N2O4- UDMH with respect to oxidizer-to-fuel ratio and characteristic velocity is given in FIG. 4. This figure shows that the propellant comprising MON-10 and UH-25 shows better characteristic velocity when compared to propellant comprising pure N2O4 and UDMH.

Thus, the combination of oxidizer MON-10 with fuel UH-25 results in benign exhaust products, yields better specific impulse performance compared to the highly dangerous nitrogen tetroxide oxidizer while yielding greatly superior specific impulse performance when compared to an oxidizer that is 100 weight percent nitric oxide. The propellant composition comprising MON-10 and UH-25 exhibits better efficiency when compared to a propellant composition comprising 100% NO and fuel because of the following reasons:
1. N2O4 (i.e. 90% N2O4 in MON-10) has better oxidizing properties than NO.
2. NO has a lower boiling point when compared to MON-10. The boiling point of NO is - 152°C. Due to lower boiling, handling problems due to cryogenic nature occurs. Also, NO has low oxygen content and is not effective in complete combustion.

Thus, the oxidizer compound of the present disclosure solves some of the problems associated with the use of cryogenic liquid oxygen and provides better performance than liquid oxygen, while simultaneously providing performance better than the more highly energetic oxidizers without any of the handling/storage problems associated therewith.

The schematic representation of a rocket propulsion system using the propellant composition (oxidizer compound and fuel) of the present disclosure is given in FIG. 5. The specific impulse and C* performance of the mixtures of nitric oxide and nitrogen tetroxide at different nitric oxide concentrations are calculated. All calculations are carried out using UDMH and UH-25 as the fuels with a chamber pressure of 20 ksc, the nozzle area ratio of 7 and exit pressure at sea level. The results are shown in FIG. 6 and FIG. 7 for specific impulse and characteristic velocity (C*) respectively at different NO concentrations. The data for pure oxidizer, N2O4 [i.e. at 0% NO] is included for reference. It is important to note that the specific MON-10 and UH-25 combination exceeds the performance of pure N2O4 (i.e. NO at 0%). This represents a significant advantage of the propellant of the present disclosure. In applications where the use of N2O4 as an oxidizer is problematic due to its toxicity, the oxidizer of the present disclosure could thus be used as a less hazardous alternative.

In yet another embodiment of the present disclosure, the propellant comprising MON-10 and UH-25 are employed in a device which can be a liquid rocket, in which propellants are stored in liquid form. The liquid rocket can be a bipropellant or tripropellant rocket. A device which is a bipropellant or tripropellant rocket comprises a container storing the oxidizer and an additional container storing the fuel. Any or all of the containers can be in fluid or gaseous connection with each other and with a combustion chamber to which an outlet is attached, allowing expulsion of combustion gases. In one embodiment, each container is separately in connection with the combustion chamber, allowing the fuel and oxidizer to mix within the combustion chamber. Alternatively, the device can be configured such that the fuel and oxidizer are mixed prior to injection into the combustion chamber. In a related embodiment, the device is self-pressurized. In other embodiments, pressurization is achieved using other means, such as another gas or by using pumps. Various configurations of liquid rockets can be used in the invention.

The present disclosure further provides a device comprising: (a) a first container comprising oxidizer MON-10 comprising nitric oxide at 10% (w/w) and nitrogen tetroxide at 90% (w/w); (b) a second container comprising the fuel UH-25; (c) a combustion chamber in fluid or gaseous connection with the first container and/or with the second container, wherein the combustion chamber allows the combustion of a propellant comprising a mixture of the oxidizer and the fuel to produce combustion gases; and (d) an outlet allowing the release of combustion gases. In one embodiment, the oxidizer can be substantially in liquid or gaseous form and may or may not be at thermodynamic equilibrium. In another embodiment, the oxidizer can be stored in a first container which is a metal storage tank. In another embodiment, the outlet comprises a nozzle. In related embodiments, the device is a rocket device. Alternatively, the device can be a gas generator. Further, the device can be an internal combustion engine. A device of the invention can be substantially self-pressurized.

In an exemplary embodiment of the present disclosure, the device is a rocket device comprising: a) components for a propellant wherein the components include oxidizer comprising (i) nitric oxide (10% w/w), (ii) nitrogen tetroxide (90% w/w), and a fuel UH-25; b) a combustion chamber comprising an outlet; c) means for feeding the components into the combustion chamber whereby the chamber comprises a propellant; and d) means to ignite the propellant in the combustion chamber, whereby propellant is combusted and expelled through the outlet. In one embodiment, such a device can be a liquid rocket wherein the device further comprises: (i) a first container containing the oxidizer in liquid form; (ii) a second container containing the fuel; and (iii) means to mix the oxidizer and the fuel in the combustion chamber to form the propellant.

A schematic of a rocket propulsion system using the oxidizer compound of the present disclosure is shown in Figure 5. One storage container at room temperature stores the homogenous and stable-liquid, nitrous oxide/nitrogen tetroxide compound MON-10 of the present disclosure. A second storage container stores the rocket fuel UH-25. A combustion chamber is coupled to both the storage containers through control valves (not shown) that control the mixing/burning of the oxidizer compound/rocket fuel with the combustion by-products being exhausted via. a nozzle to generate thrust. A variety of well-known mixing and exhausting systems can be used. Accordingly, combustion chamber and nozzle are not limitations of the present disclosure. In a related embodiment, the device is self-pressurized. In other embodiments, pressurization is achieved using other means, such as another gas or by using pumps.

In an embodiment of the present disclosure, the fuel UH-25 is highly miscible with the oxidizer MON-10 and has a low vapor pressure at storage temperature in order to minimize the possibility of a vapor phase explosion.

In another embodiment of the present disclosure as described above, UH-25 is a mixture of 75% (w/w) of unsymmetrical dimethyl hydrazine (UDMH) and 25% (w/w) of hydrazine hydrate. In an embodiment, hydrazine hydrate is a colourless oily liquid with odor of ammonia and contains 36% (w/w) mass fraction of water.

In yet another embodiment of the present disclosure, UH-25 fuel is hypergolic when combined with the mixture of nitric oxide and nitrogen tetroxide (MON-10) and can be stored as liquid at room temperature.

The rocket propellant composition of the present disclosure is prepared by mixing the various components together in the amounts required to obtain the desired composition and then subjecting the mixture to agitation by steps such as stirring, shaking etc., until a homogeneous composition is obtained. It is immaterial in what order the components are added to the container in which the mixing is affected. For example, the oxidizer may be added to the fuel or conversely the fuel may be added to the oxidizer.

The present disclosure also relates to evaluation of the performance of hydrazine and hydrazine like materials such as UDMH and UH-25 fuels with oxidizer that includes nitric oxide and nitrogen tetroxide using a bipropellant thruster in pulse mode at both ambient and -20oC temperatures. The performance parameters such as chamber pressure, characteristic velocity and specific impulse is compared with conventional oxidizer N2O4 using a bipropellant thruster in pulse mode. An embodiment the present disclosure relates to comparison of the performance evaluation of UDMH and UH-25 with MON-10 for chamber pressure, characteristic velocity and specific impulse at the same conditions mentioned above.

Some of the terms used in the present disclosure are defined below and are to be construed to have the same scope and meaning throughout the specification.

‘Characteristic velocity (C*) in the present disclosure is expressed as
: C* = PcAt/m.

where, Pc = Chamber pressure; At=Throat area and m. = mass flow rate

‘Characteristic velocity’ is used in comparing the relative performance of different chemical rocket propulsion system designs and propellants. Chamber pressure should be sufficiently high to get high characteristic velocity which determines the performance of the propellant.

‘Thrust’ developed by a given propellant depends on the ability to generate high temperature and low molecular mass gases and also the capacity to expand the gases to low values of nozzle exit pressures.
Thrust in terms of thrust coefficient is:
F = CFPcAt
where, F is Thrust; CF is Thrust coefficient; Pc is chamber pressure and At is throat area.

‘Specific impulse, Isp’ is defined as the thrust (F) per unit mass flow rate (m):
Isp = F/m.

‘Propellant tanks’: Liquid propellant rockets comprise tanks containing liquid propellants and suitable propellant feed system for supplying fuel and oxidizer to the combustion chamber at the required flow rates and pressures (i.e. propellant tank pressures).

In an embodiment of the present disclosure, the performance of the propellant (MON-10 and UH-25) of the present disclosure is investigated by operating stationary rocket motors by conducting two static tests at ambient temperature from about 25 oC to 30 oC and one static test at low temperature using UH-25 as fuel and mixtures of nitric oxide and nitrogen tetroxide (MON-10) as oxidizer.

The propellant of the present disclosure is used in propulsion applications including bipropellant liquid rocket, tripropellant rocket engines, gas generation systems, thrust augmented liquid fueled ramjets, combined cycle propulsion systems such as turborockets and internal combustion engines. Various devices are envisioned which take advantages of the propellant of the present invention. In rocket engines, oxidizer and fuel can be stored in different tanks and mixed together in the combustion chamber. Accordingly, the present disclosure contemplates systems in which the fuels and the oxidizers of this invention are held in a single tank or in separate tanks for subsequent mixing.

The technology of the instant application is further elaborated with the help of following examples. However, the examples are only illustrative in nature and should not be construed to limit the scope of the invention.

EXAMPLES
Example 1
Preparation of MON-10
MON-10 is prepared by preparation of 10% (w/w) nitric oxide (NO) and addition to 90% (w/w) of N2O4. A special preparation cum storage tank is designed and fabricated using SS 304 material. About 15 Litres (21.75 Kg) of N2O4 is transferred to the preparation cum storage tank. In order to mix NO gas in liquid N2O4 under ambient pressure, N2O4 tank is cooled between -10 and 0 ºC using ice-salt mixture.

The following methods can be employed for lab scale and industrial production of NO gas.
a. Oxidation of ammonia at 850 oC in the presence of platinum as catalyst

b. Birkeland-Eyde process: Reaction of Oxygen and Nitrogen at temperature more than 2000oC

c. Reduction of nitric acid with copper

d. Reduction of nitrous acid in the form of sodium nitrite / potassium nitrite

Any of the aforementioned methods can be employed to generate NO gas according to the present disclosure. In the present example, NO gas required for the production of MON-10 was generated by the following process:

Reduction of sodium nitrite with ferrous sulphate in sulphuric acid media yields pure NO gas. The reaction is non-exothermic, simple and free from safety hazards. Pure NO gas is generated by this method and contamination and also formation of other gases is avoided using this method.

Rate of production of NO gas and mixing with N2O4 is controlled by drop wise addition of aqueous NaNO2 to acidic FeSO4 solution with constant stirring. The duration of mixing of NO gas is optimized to obtain 10 wt % of NO in 90 wt % of N2O4. The optimization of NO gas is done by using gravimetric method following MIL-REF-26539E. The duration of mixing of NO gas depends on the batch size. NO has a very good dispersion with N2O4.

N2O4 tank is kept under low temperature (between 10 and 0oC) till the end of mixing process. In order to confirm the percentage of NO mixed in N2O4, sampling and analysis is carried out during the preparation as per standard procedure. The tank is stored at 6 bar pressure under nitrogen gas at ambient temperature.

Example 2: Static tests to evaluate the performance of propellant
The performance of the propellant of the present disclosure is investigated by operating stationary rocket motors by conducting two static tests at ambient temperature and one static test at low temperature using UH-25 as fuel and mixture of nitric oxide (10% w/w) and nitrogen tetroxide (90% w/w) [i.e. MON-10] as oxidizer.

Design of the rocket propulsion system for conducting static tests:
The rocket engine employed in the tests has a throat area of 262.9 mm2. The ratio of the cross-sectional area of the nozzle exit to-throat cross-sectional area is 7:1. The ratio of the cross-sectional area of the combustion chamber to the cross-sectional area of the throat is 8:1. The tests are carried out using carbon phenolic lined 125 Kgf thruster. The shell is fabricated from Aluminium HE20 material. The thruster along with the injector with exit closed is subjected to 30 ksc pneumatic pressure. FIG. 8 represents 125 Kgf thruster with injector. The disclosure provides methods of operating a motor comprising a combustion chamber, the method comprising: (a) providing a fuel comprising UH-25; (b) providing MON-10 as oxidizer; (c) combusting the oxidizer in a combustion chamber to form a combustion gas; and (d) expelling the combustion gas from the motor. Thrust can be generated during the step of expelling combustion gas. In some embodiments, the oxidizer is additionally contacted with a fuel. In related embodiments, the oxidizer is contacted with a fuel prior to combustion. The motor is operated at a combustion chamber pressure of 30 KscA and an exit nozzle pressure of substantially 0.96 KscA. The fuel composition and oxidizer are fed through separate conduits from individual storage containers to the combustion chamber where the stream of fuel composition and the stream of oxidizer are contacted with each other. The fuel oxidizer ignites on contact, producing gaseous products as a result of the spontaneous combustion of the components of the two streams. The gaseous products are ejected from the combustion chamber through the throat area and then out into the atmosphere through the exit nozzle. The ejection of the reaction product gases from the combustion chamber produces a thrust which is measured by means of a load cell mounted forward of the motor. The fuel composition and the oxidizer are metered into the motor so that the amount reacting within any particular period of time is known.

Static tests 1& 2 are planned with a cumulative burn time of 7 sec and 14 sec respectively in pulsed mode at ambient temperatures ranging from about 25 oC to about 30 oC. Pulse mode is the operation/ activity of start-stop-restart of ignition/combustion of the propellant. The approximate lap time of the operation can be varied from 1 to 10 seconds or more. Pulse train is given manually for test 1 which contains first three pulses of approximate lap time 1-2 sec and a final pulse of 10 sec duration whereas for test 2 it contains three pulses long duration of 5 sec approximately.

Test 3 is planned with a cumulative burn time of 10 sec in pulsed mode at -20oC temperature. The pulse train is given manually which contains short pulses of 14 numbers approximately. The set tank pressures for all the three tests are given in Table- 4.
Table 4: Tank Pressures for the test (Static Condition)
Oxidiser Fuel
Tank Pressure (Kscg) Test-01 42.5 41.0
Test-02 43.5 43.5
Test-03 42.0 38.7

1. ‘Static test 1’ at ambient temperature (25 oC to about 30 oC) with burn time of 7 sec
The rocket motor described hereinabove is operated on an oxidizer composition consisting essentially of 90 parts by weight N2O4 and 10 parts by weight NO (MON-10). The fuel used is UH-25, at a weight ratio of fuel-to-oxidizer (i.e. weight of the fuel divided by weight of the oxidizer) of about 1.665. The pressure measured in the combustion chamber is about 30 KscA. The oxidizer tank pressure and manifold pressure is about 42.5Kscg and about 41.5 Kscg respectively, while fuel tank pressure and manifold pressure is about 41.0 Kscg and about 40.0 Kscg respectively. The densities of oxidizer and fuel are about 1403 Kg/m3 and about 855Kg/m3 and flow rates of oxidizer and fuel are about 0.293 Kg/sec and about 0.176 Kg/sec respectively. The combustor is operated for 7 seconds.

2. ‘Static test 2’ at ambient temperature (25 oC to about 30 oC) with burn time of 14 sec
The rocket motor described hereinabove is operated on an oxidizer composition consisting essentially of 90 parts by weight N2O4 and 10 parts by weight NO (MON-10). The fuel used is UH-25, at a weight ratio of fuel-to-oxidizer of about 1.665. The pressure measured in the combustion chamber is about 30 KscA. The oxidizer tank pressure and manifold pressure is about 43.5Kscg and about 41.7 Kscg respectively, while fuel tank pressure and manifold pressure is about 43.5 Kscg and about 41.9 Kscg respectively. The densities of oxidizer and fuel are about 1403 Kg/m3and about 855Kg/m3 and flow rates of oxidizer and fuel are about 0.293 Kg/sec and about 0.176 Kg/sec respectively. The combustor is operated for 14 seconds.

3. ‘Static test 3’ at -20oC with burn time of 14 sec
The rocket motor described hereinabove is operated on an oxidizer composition consisting essentially of 90 parts by weight N2O4 and 10 parts by weight NO (MON-10). The fuel used is UH-25, at a weight ratio of fuel-to-oxidizer of 1.665. The pressure measured in the combustion chamber is about 30 KscA. The oxidizer tank pressure and manifold pressure is about 43.5Kscg and about 41.7 Kscg respectively, while fuel tank pressure and manifold pressure is about 43.5 Kscg and about 41.9 Kscg respectively. The densities of oxidizer and fuel are about 1403 Kg/m3and about 855Kg/m3 and flow rates of oxidizer and fuel are about 0.293 Kg/sec and about 0.176 Kg/sec respectively. The combustor is operated for 14 seconds.

In order to maintain -20 oC temperature, propellant tanks in filled condition are kept in a box filled with ice pieces mixed with calcium chloride and sodium chloride mixture.

Figure 9 provides variation of absolute chamber pressure with time for Static Test 1. It shows the upper and lower boundaries of expected absolute chamber pressure between 33 to 27 KscA given in pink colour. The average achieved absolute chamber pressure in the pulse mode is 29 KscA given in blue colour.

Figure 10 provides comparison of oxidizer manifold pressure with chamber pressure for Static Test 1, wherein the blue colour represents the oxidizer manifold pressure and pink colour represents the chamber pressure respectively. The average achieved oxidizer manifold pressure is 37.7 Kscg and the average achieved chamber pressure is 29 KscA.

Figure 11 shows variation of thrust at sea level with time for Static Test 1. It shows the upper and lower boundaries of expected absolute chamber pressure between 127 to 105 Kgf given in pink colour. The average achieved thrust at sea level in the pulse mode is 108.5 Kgf given in blue colour.

Figure 12 provides variation of absolute chamber pressure with time for Static Test 2.The average achieved absolute chamber pressure in the pulse mode is 29.5 KscA given in blue colour.

Figure 13 shows variation of thrust at sea level with time for Static Test 2. The average achieved thrust at sea level in the pulse mode is 121.8 Kgf given in blue colour.

Figure 14 provides variation of absolute chamber pressure with time for Static Test 3. It shows the upper and lower boundaries of expected absolute chamber pressure between 33 to 27 KscA given in pink colour. The average achieved absolute chamber pressure in the pulse mode is 30.4 KscA given in blue colour.

Figure 15 shows variation of thrust at sea level with time for Static Test 3. It shows the upper and lower boundaries of expected absolute chamber pressure between 127 to 105 Kgf given in pink colour. The average achieved thrust at sea level in the pulse mode is 118 Kgf given in blue colour.

Table 5 summarizes the instrumentation plan carried out with the achieved results for all the three tests and Table 6 shows instrumentation plan for temperature and pressure measurement and Table 7 compares the performance of UH-25 & MON-10 combination with UH-25 & N2O4 for all the three tests.

Table 5: Instrumentation plan summary
Parameter Test-01 Test-02 Test-03
Expected Achieved Expected Achieved Expected Achieved
1. Chamber Pressure (KscA) 30+3 29.0 30+3 29.5 30+3 30.4
2. Oxidiser Tank Pressure (Kscg) 42.5 42.47 43.5 42.9 43.5 41.5
3. Oxidiser manifold Pressure (Kscg) 41.5 37.7 41.7 41.4 41.7 41.0
4. Injector drop- O side (Kscg) 10.0 9.68 12.7 12.9 12.7 11.5
5. Fuel Tank Pressure (Kscg) 41.0 40.56 43.5 43.0 43.5 38.6
6. Fuel manifold Pressure (Kscg) 40 39.4 41.9 41.7 41.9 38.2
7. Injector drop - F side (Kscg) 11.3 11.41 12.9 13.18 12.9 9.1
8. Thrust at Sea level (Kgf) 116+11 108.5 116+11 121.8 116+11 118
9. Oxidiser Flow (Kg/s) 0.293 0.259 0.293 0.300 0.293 0.290
10. Fuel flow (Kg/s) 0.176 0.174 0.176 0.187 0.176 0.173
11. Maximum skin temperature
( oC)
@ Injector
@ Chamber
@ Throat

118.2
60
62.4

119.3
39.2
34.9

82
28
22
The thruster assembly is connected to the propellant tanks in pressure fed mode as shown in figure 8b. The following pressures and temperatures are measured as part of Instrumentation plan:
Table 6: Instrumentation plan for temperature and pressure measurement
S.No. Instrumentation Channels
PT1 Tank pressures 2
PT2 Manifold pressures 2
PT3 Chamber pressure 1
T1 Temperature on Injector surface 1
T2 Temperature on cylindrical surface of shell 1
T3 Temperature near throat area of shell 1

Results of the static tests:
When a flight rocket is operated employing the oxidizer compound MON-10 and fuel UH-25, satisfactory flight performance, with no detrimental explosions, is observed. The achieved chamber pressure of UH-25/MON-10 combination for tests 1, 2 & 3 are comparable with UH-25/N2O4 combination. However, the combination of MON-10/UH-25 successfully works at low temperature and is hence advantageous over UH-25/N2O4 combination. The below table 7 compares the results of UH-25 + N2O4 combination when subjected to static test at 25 oC to about 30 oC versus the individual results of UH-25 + MON-10 combination when subjected to the three static tests. ‘Expected’ is the theoretical value obtained from NASA SP 273 whereas the ‘achieved’ is the experimental static test results.

Table 7: Comparison of the performance of UH-25 & MON-10 combination with UH-25 & N2O4 for all the three static tests

S.No
Parameters
Units UH-25 & N2O4 Combination UH-25 & MON-10 Combination
Expected Achieved Expected Test-01 Test-02 Test-03
1. Chamber Pressure KscA 27+2.7 28.5 30+3 29.0 29.5 30.4
2. Thrust-Sea level Kgf 110+11 105 116+11 108.5 121.8 118
3. Specific Impulse-
Sea level S 223+5 245+5 226+5 230.36
232.14 237
4. Oxidizer mass flow Kg/s 0.304 0.284 0.293 0.259 0.300 0.293
5. Fuel mass flow Kg/s 0.180 0.166 0.176 0.174 0.187 0.173
6. Mixture ratio 1.684+0.15 1.71+0.1 1.665+0.15 1.546 1.604 1.694
7. Cumulative Burn time S 8 14 8 7 13.4 6
8. Characteristic velocity (Actual) m/s 1667 1790.1 1675.4 1710.7 1685.2 1690
9. Thrust coefficient (Actual) 1.357 1.339 1.357 1.32 1.35 1.37
10. Nozzle efficiency % 95.0 93.8 95.0 92.4 94.5 95.1

Conclusion: The achieved specific impulse and characteristic velocity of UH-25/MON-10 is compared with the expected values, which fall well within the limits of the expected values. These values when compared with UH-25/N2O4 do not show much deviation or in other words are comparable with the results of UH-25/N2O4. These values are well accepted in the bipropellant systems. However, it is important to note that the combination of UH-25/MON-10 successfully operates at low temperature which is a great advantage in terms of low temperature applications, unlike UH-25/N2O4 which does not work at low temperature (such as -20 oC) due to the fact that the freezing point of 100% N2O4 is -11.2oC and hence propellant composition comprising 100% N2O4 is not suitable for low temperature applications at -20oC.
Similarly, other propellant compositions such as MON-30 + UH-25, MON-3+ UH-25 etc. as indicated in Table 3 are also not workable at low temperature of -20 oC owing to their freezing points. The freezing point of MON-3 oxidizer is -15 oC due to which propellant composition comprising MON-3 is not suitable for low temperature applications at -20oC. Hence, the propellant composition of the present disclosure- MON-10+UH-25 is advantageous over other propellant compositions and performs efficiently by achieving desirable values of specific impulse and characteristic velocity even at low temperature of -20 oC.
From the performance evaluation of the propellant of the present disclosure, discussion and examples given hereinabove, it is seen that the propellant of the present disclosure solves some of the problems associated with the use of cryogenic liquid oxygen and provides better performance than liquid oxygen, while simultaneously providing performance comparable to the more highly energetic propellants without any of the handling/storage problems associated therewith. This propellant with the resulting benign exhaust products, yields comparable specific impulse performance to the highly dangerous nitrogen tetroxide oxidizer while yielding comparable specific impulse performance and characteristic velocity when compared to an oxidizer that is 100 weight percent nitric oxide. Also, the performance evaluation of the propellant shows that the propellant provides comparable or better specific impulse and comparable characteristic velocity when compared to UH-25/N2O4 and other propellant systems. More importantly, the present MON-10/UH-25 propellant composition performs at low temperatures which is a great advantage when compared to the known propellant compositions of the art.

Example 3:
Comparison of performance of propellant composition comprising MON-10 and UH-25 ‘versus’ the individual components MON-10 and UH-25
The component MON-10 is an oxidizer which is a chemical required by a fuel to combust/burn, but does not combust on its own either at ambient temperature or low temperature. The component UH-25 is a fuel which does not combust without an oxidizer either at ambient temperature or low temperature. Therefore, MON-10 or UH-25 individually do not exhibit rocket performance and cannot be employed in propulsion applications. However, desired level of performance in terms of specific impulse, characteristic velocity and workability at ambient temperature as well as low temperature is achieved when the present propellant composition containing - MON-10 & UH-25 is employed. Therefore, the propellant composition comprising MON-10 and UH-25 is a synergistic composition exhibiting enhanced level of rocket performance even at low temperatures which is not achieved by its individual components MON-10 or UH-25 respectively.

NON-LIMITING ADVANTAGES OF THE PROPELLANT COMPOSITION (MON-10 and UH-25) OF THE PRESENT DISCLOSURE:
1. The attributes for the oxidizing component of the propellant system includes high specific impulse (Isp) performance, high density, good combustion stability and efficiency characteristics, chemically stable, nontoxic, storable under normal conditions, low vapor pressure for pump fed systems, low freezing point, significant reduction in cost of operations, hypergolic behavior with the fuel, ease of handling, environmental aspects (‘green propellant’) and compatibility with tank and feed system materials.
2. The propellant can be used in a variety of rocket propulsion systems such as those used in rockets, turbojets, internal combustion engines, combined cycle propulsion systems, launch vehicle propulsion systems, multi-mode spacecraft propulsion systems and upper stage spacecraft propulsion systems.
3. The propellant of the present disclosure has the following attributes:
a) reduces storage problems by providing for room temperature storage thereof;
b) reduces propulsion system weight since room temperature storage reduces need for storage tank insulation;
c) provides for long-term storage since cryogenic boil-off is not a problem;
d) has a relatively high specific impulse when compared to traditional energetic but inherently problematic oxidizers;
e) has a relatively high energy density when compared to traditional energetic but inherently problematic oxidizers;
f) has reduced toxicity over pure nitrogen tetroxide in its stored state; and
g) produces environmentally benign exhaust products when burned in a propulsion system.
4. The oxidizer compound of the present propellant is self pressurizing and does not need to operate under deep cryogenic conditions as it can be stored at room temperature.
5. Mixture of 10% (w/w) NO and 90% (w/w) N2O4 can additionally present significant safety advantages compared to the use of liquid oxygen due to reduced cryogenic and fire hazards. The safety advantages of the oxidizers of the invention would thus have a substantial effect in reducing the development and operational costs associated with propulsion systems. Also, the present oxidizer MON-10 has a lower freezing point compared to pure N2O4. This feature is particularly useful in space applications, due to the low temperatures at high altitudes and in space.
6. Applications of the propellant of the present disclosure comprising NO/N2O4 oxidizer and UH-25 include hybrid rockets, bipropellant liquid rocket, tripropellant rocket engines, gas generation systems and internal combustion engines. Various devices are envisioned which take advantages of the oxidizer of the invention. In rocket engines, oxidizer and fuel can be stored in different tanks and mixed together in the combustion chamber.
7. The propellant of the present disclosure is highly stable and has the unique ability to be used for low temperature (such as -20oC) applications.

Therefore, the present disclosure demonstrates propellant composition comprising UH-25 as fuel and mixture of nitric oxide and nitrogen tetroxide with NO at a concentration of 10% (w/w) and N2O4 at a concentration of 90% (w/w) as oxidizer. Also, the present disclosure demonstrates the performance evaluation of the said propellant with respect to specific impulse, characteristic velocity and other technical parameters to establish the improved performance/properties of the present propellant even at low temperatures when compared to conventional propellants (such as propellants having pure nitrogen tetroxide or pure nitric oxide as oxidizer).

CLIAMS:1. A propellant composition comprising oxidizer and fuel, wherein the oxidizer is mixed oxides of nitrogen-10 (MON-10) and the fuel is UH-25.
2. The propellant composition as claimed in claim 1, wherein the MON-10 comprises nitric oxide (NO) at a concentration of about 10% (w/w) and nitrogen tetroxide (N2O4) at a concentration of about 90% (w/w).
3. The propellant composition as claimed in claim 1, wherein the UH-25 comprises about 75% unsymmetrical dimethylhydrazine (UDMH) and about 25% hydrazine hydrate.
4. The propellant composition as claimed in claim 1, wherein the propellant is chemically stable and operates at a temperature ranging from about 25oC to -20oC.
5. The propellant composition as claimed in claim 1, wherein the propellant is chemically stable and operates at -20oC.
6. The propellant composition as claimed in claim 1, wherein the composition is non-toxic, self-pressurizing, storable at room temperature, and operates at high altitudes and in space.
7. The propellant composition as claimed in claim 1, wherein the MON-10 is a non-viscous liquid.
8. The propellant as claimed in claim 1, wherein the propellant composition has a specific impulse ranging from about 221 seconds to 231 seconds, characteristic velocity ranging from about 1670 m/s to about 1700 m/s and chamber pressure ranging from about 27 KscA to 33 KscA.
9. The propellant as claimed in claim 1, wherein at -20oC, the propellant composition has a specific impulse of about 237 seconds, characteristic velocity of about 1690 m/s and chamber pressure of about 30.4 KscA.
10. A method of preparing the propellant composition as claimed in claim 1, said method comprising acts of:
a) preparing NO gas;
b) dissolving NO gas with N2O4 to obtain MON-10 comprising nitric oxide (NO) at a concentration of about 10% (w/w) and nitrogen tetroxide (N2O4) at a concentration of about 90% (w/w); and
c) mixing the MON-10 with UH-25 to obtain the propellant composition comprising MON-10 and UH-25.
11. The method as claimed in claim 10, wherein the NO gas is prepared by reducing aqueous sodium nitrite (NaNO2) in presence of ferrous sulphate (FeSO4).
12. The method as claimed in claim 10, wherein the step (b) is carried out under cold condition, preferably at a temperature ranging from about 10oC to 0oC.
13. The method as claimed in claim 10 or claim 11, wherein the preparation of NO gas in step (a) and dissolving the NO gas with N2O4in step (b) is controlled by drop wise addition of the aqueous NaNO2 to the FeSO4 solution with constant stirring.
14. A device for combustion of the propellant composition as claimed in claim 1, said device comprising:
a) a first container comprising the mixed oxide of nitrogen-10 (MON-10) oxidizer;
b) a second container comprising the UH-25 fuel;
c) a combustion chamber connected to the first container and optionally to the second container, wherein said combustion chamber allows combustion of the propellant composition as claimed in claim 1 to produce combustion gases; and
d) an outlet allowing the release of combustion gases.
15. The device as claimed in claim 14, wherein said device is employed in rocket propulsion system of a vehicle selected from a group comprising rocket, turbojet, combined cycle propulsion system, launch vehicle propulsion system, multi-mode spacecraft propulsion system, upper stage spacecraft propulsion system, missile propulsion system, hybrid rocket, bipropellant liquid rocket, tripropellant rocket engine, gas generation system, internal combustion engine and combinations thereof.
16. The propellant composition as claimed in claim 1 for use in a vehicle selected from a group comprising rocket, turbojet, combined cycle propulsion system, launch vehicle propulsion system, multi-mode spacecraft propulsion system, upper stage spacecraft propulsion system, hybrid rocket, bipropellant liquid rocket, tripropellant rocket engine, gas generation system, internal combustion engine and combinations thereof.