CN115038681A - Rocket propellant compositions and methods of production having enhanced cryogenic cooling, thermal stability and thrust efficiency performance - Google Patents
Rocket propellant compositions and methods of production having enhanced cryogenic cooling, thermal stability and thrust efficiency performance Download PDFInfo
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- 239000000203 mixture Substances 0.000 title claims abstract description 118
- 239000003380 propellant Substances 0.000 title claims abstract description 64
- 238000000034 method Methods 0.000 title claims description 27
- 238000001816 cooling Methods 0.000 title description 7
- 238000004519 manufacturing process Methods 0.000 title description 3
- 229930195733 hydrocarbon Natural products 0.000 claims abstract description 89
- 150000002430 hydrocarbons Chemical class 0.000 claims abstract description 88
- 239000004215 Carbon black (E152) Substances 0.000 claims abstract description 82
- MYMOFIZGZYHOMD-UHFFFAOYSA-N Dioxygen Chemical compound O=O MYMOFIZGZYHOMD-UHFFFAOYSA-N 0.000 claims abstract description 32
- 239000001257 hydrogen Substances 0.000 claims abstract description 17
- 229910052739 hydrogen Inorganic materials 0.000 claims abstract description 17
- UFHFLCQGNIYNRP-UHFFFAOYSA-N Hydrogen Chemical compound [H][H] UFHFLCQGNIYNRP-UHFFFAOYSA-N 0.000 claims abstract description 16
- NINIDFKCEFEMDL-UHFFFAOYSA-N Sulfur Chemical compound [S] NINIDFKCEFEMDL-UHFFFAOYSA-N 0.000 claims abstract description 12
- 229910052717 sulfur Inorganic materials 0.000 claims abstract description 12
- 239000011593 sulfur Substances 0.000 claims abstract description 12
- 238000002156 mixing Methods 0.000 claims description 13
- 150000001335 aliphatic alkanes Chemical class 0.000 claims description 10
- 238000009835 boiling Methods 0.000 claims description 5
- 238000004821 distillation Methods 0.000 claims description 4
- 238000007710 freezing Methods 0.000 claims description 4
- 230000008014 freezing Effects 0.000 claims description 4
- JTJMJGYZQZDUJJ-UHFFFAOYSA-N phencyclidine Chemical compound C1CCCCN1C1(C=2C=CC=CC=2)CCCCC1 JTJMJGYZQZDUJJ-UHFFFAOYSA-N 0.000 claims description 3
- 239000002760 rocket fuel Substances 0.000 claims description 3
- 150000001924 cycloalkanes Chemical class 0.000 claims description 2
- 239000007788 liquid Substances 0.000 description 12
- 239000000446 fuel Substances 0.000 description 10
- 238000002485 combustion reaction Methods 0.000 description 7
- 239000004449 solid propellant Substances 0.000 description 3
- 125000003118 aryl group Chemical group 0.000 description 2
- 239000000571 coke Substances 0.000 description 2
- 238000004939 coking Methods 0.000 description 2
- 230000000694 effects Effects 0.000 description 2
- 230000005484 gravity Effects 0.000 description 2
- 238000012986 modification Methods 0.000 description 2
- 230000004048 modification Effects 0.000 description 2
- 238000012360 testing method Methods 0.000 description 2
- 150000001336 alkenes Chemical class 0.000 description 1
- 150000001345 alkine derivatives Chemical class 0.000 description 1
- 230000015572 biosynthetic process Effects 0.000 description 1
- 239000003054 catalyst Substances 0.000 description 1
- 238000006243 chemical reaction Methods 0.000 description 1
- 125000004122 cyclic group Chemical group 0.000 description 1
- 230000001419 dependent effect Effects 0.000 description 1
- 238000009472 formulation Methods 0.000 description 1
- 238000004817 gas chromatography Methods 0.000 description 1
- 150000002431 hydrogen Chemical class 0.000 description 1
- 125000004435 hydrogen atom Chemical group [H]* 0.000 description 1
- 239000003350 kerosene Substances 0.000 description 1
- -1 liquid oxygen hydrocarbon Chemical class 0.000 description 1
- 238000005259 measurement Methods 0.000 description 1
- 239000007800 oxidant agent Substances 0.000 description 1
- 230000001590 oxidative effect Effects 0.000 description 1
- 229910052760 oxygen Inorganic materials 0.000 description 1
- 239000001301 oxygen Substances 0.000 description 1
- 229920006395 saturated elastomer Polymers 0.000 description 1
- 230000035945 sensitivity Effects 0.000 description 1
- 238000000926 separation method Methods 0.000 description 1
- 239000000126 substance Substances 0.000 description 1
- 238000006467 substitution reaction Methods 0.000 description 1
- 238000005979 thermal decomposition reaction Methods 0.000 description 1
- 230000008646 thermal stress Effects 0.000 description 1
- 238000012546 transfer Methods 0.000 description 1
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- C—CHEMISTRY; METALLURGY
- C06—EXPLOSIVES; MATCHES
- C06B—EXPLOSIVES OR THERMIC COMPOSITIONS; MANUFACTURE THEREOF; USE OF SINGLE SUBSTANCES AS EXPLOSIVES
- C06B47/00—Compositions in which the components are separately stored until the moment of burning or explosion, e.g. "Sprengel"-type explosives; Suspensions of solid component in a normally non-explosive liquid phase, including a thickened aqueous phase
- C06B47/02—Compositions in which the components are separately stored until the moment of burning or explosion, e.g. "Sprengel"-type explosives; Suspensions of solid component in a normally non-explosive liquid phase, including a thickened aqueous phase the components comprising a binary propellant
- C06B47/12—Compositions in which the components are separately stored until the moment of burning or explosion, e.g. "Sprengel"-type explosives; Suspensions of solid component in a normally non-explosive liquid phase, including a thickened aqueous phase the components comprising a binary propellant a component being a liquefied normally gaseous fuel
-
- C—CHEMISTRY; METALLURGY
- C06—EXPLOSIVES; MATCHES
- C06D—MEANS FOR GENERATING SMOKE OR MIST; GAS-ATTACK COMPOSITIONS; GENERATION OF GAS FOR BLASTING OR PROPULSION (CHEMICAL PART)
- C06D5/00—Generation of pressure gas, e.g. for blasting cartridges, starting cartridges, rockets
- C06D5/08—Generation of pressure gas, e.g. for blasting cartridges, starting cartridges, rockets by reaction of two or more liquids
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- Chemical & Material Sciences (AREA)
- Organic Chemistry (AREA)
- Chemical Kinetics & Catalysis (AREA)
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- Combustion & Propulsion (AREA)
- Liquid Carbonaceous Fuels (AREA)
- Lubricants (AREA)
Abstract
The present disclosure relates to a rocket propellant composition comprising (a) liquid oxygen; (b) a hydrocarbon mixture comprising: (i) a hydrogen content of from about 14.0 mass% to about 16.0 mass% based on the mass of the hydrocarbon mixture; (ii) a kinematic viscosity of about 5mm at a temperature in the range of about-35 ℃ to about-33 ℃ 2 S to about 8mm 2 S; (iii) the sulfur content is from about 0ppm to about 0.1ppm based on the mass of the hydrocarbon mixture.
Description
Technical Field
In some embodiments, the present disclosure relates to rocket propellants having improved fuel thrust and cryogenic cooling indicators and methods of making the same.
Background
Rocket propellants may be divided into solid propellants and liquid propellants. Although solid propellants may be easier to store, they have the significant disadvantage of having a lower specific impulse compared to liquid propellants, which is an indicator of propellant efficiency, resulting in lower overall performance. In addition to having a higher specific impulse, liquid propellants enable real-time throttling, shut-down and restart, whereas solid propellants, once ignited, must burn until all of the propellant is spent.
Liquid propellants currently used include rocket propellant-1 (RP-1) and rocket propellant-2 (RP-2) (RP-2 is a low sulfur version of RP-1), which are highly reformed forms of kerosene, which have low specific impulse and are stable at room temperature. Typically, both RP-1 and RP-2 are combusted in a rocket engine with liquid oxygen as the oxidant. However, given the heavy burden on rocket propulsion efficiency due to the weight of fuels, new advances are needed to find denser and more efficient fuels, such as fuels with higher hydrogen content than existing liquid propellants. Furthermore, given that liquid propellants are typically cooled prior to mixing with liquid oxygen, there is a need for a superior liquid propellant that can be cooled to low temperatures while having a lower viscosity than existing liquid propellants in order to properly mix with liquid oxygen when needed. In addition, since existing liquid propellants tend to produce coke when consumed, which can lead to engine fouling, there is a need for liquid propellants having a reduced potential for coke formation and operating temperatures.
Disclosure of Invention
Accordingly, there is a need for improved liquid rocket propellant compositions having high hydrogen content and low kinematic viscosity at temperatures below-33 ℃. In some embodiments, the present disclosure relates to a rocket propellant composition comprising a mixture of liquid oxygen and a hydrocarbon. During combustion, the liquid oxygen and hydrocarbon mixture are mixed and then combusted to provide propulsion for the rocket.
Generally, the higher the hydrogen content value of a hydrocarbon mixture, the higher the specific impulse that the hydrocarbon mixture can provide to the rocket, which is a measure of how effectively the rocket utilizes the hydrocarbon mixture. According to ASTM D5291, the disclosed hydrocarbon mixtures have a hydrogen content in the range of from about 14.0 mass% to about 16.0 mass%, preferably from about 15.0 to about 16.0 mass%, based on the mass of the hydrocarbon mixture. According to ASTM D7042 and/or ASTM D445, disclosed hydrocarbons have a temperature range of about-35 ℃ to about-33 ℃ of about 5mm 2 S to about 8mm 2 Relatively low kinematic viscosity in/s, preferably from about 5 to about 7mm 2 And(s) in the presence of a catalyst. Further, according to ASTM D5623, the hydrocarbon mixture includes a sulfur content of about 001 ppm to about 0.1ppm by mass of the hydrocarbon mixture. Each of these attributes work in concert to provide better fuel efficiency when the rocket uses these attributes during combustion.
In some embodiments, the present disclosure relates to a method of producing a rocket propellant from a liquid oxygen and hydrocarbon mixture. The method includes providing liquid oxygen and providing a hydrocarbon mixture. According to ASTM D5291, the hydrocarbon mixture comprises a hydrogen content in the range of from about 14.0 mass% to about 16.0 mass%, preferably from about 15.0 to about 16.0 mass%; a kinematic viscosity of about 5mm according to ASTM D7042 and/or ASTM D445 at a temperature range of about-35 ℃ to about-33 ℃ 2 S to about 8mm 2 S, preferably from about 5 to about 7mm 2 S; according to ASTM D5623, the sulfur content is from about 0.01ppm to about 0.1 mass%, based on the mass of the hydrocarbon mixture. The liquid oxygen and the hydrocarbon mixture may be mixed and then combusted.
Drawings
Some embodiments of the disclosure may be understood by referring, in part, to the disclosure and the accompanying drawings, wherein:
FIG. 1 is a graph of percent change in useful payload versus percent change in beta in accordance with a specific example embodiment of the present disclosure;
FIG. 2 is a graph of specific impulse gain versus liquid oxygen to hydrocarbon mixture ratio at 350psi, according to a specific exemplary embodiment of the present disclosure;
FIG. 3 is a graph of specific impulse gain versus liquid oxygen to hydrocarbon mixture ratio at 500psi, according to a specific exemplary embodiment of the present disclosure;
FIG. 4 is a graph of specific impact gain at 750psi versus liquid oxygen to hydrocarbon mixture ratio according to a specific exemplary embodiment of the present disclosure; and
FIG. 5 is a graph of kinematic viscosity versus temperature for various disclosed rocket propellant compositions, in accordance with specific exemplary embodiments of the present disclosure.
Detailed Description
In some embodiments, the present disclosure relates to rocket propellant compositions and methods of making the rocket propellant compositions. The disclosed rocket propellant compositions have superior cryogenic cooling, thermal stability and thrust efficiency performance relative to existing rocket propellants. Applicants' rocket propellant compositions achieve this by varying the formulation components of the hydrocarbon mixture including hydrogen content, kinematic viscosity, and sulfur content.
Rocket propellant composition
According to some embodiments, a rocket propellant composition includes a liquid oxygen and hydrocarbon mixture that has superior cryogenic cooling, thermal stability, and thrust efficiency performance relative to existing rocket propellants. The disclosed hydrocarbon mixture comprises: the hydrocarbon mixture disclosed has a hydrogen content ranging from about 14.0 mass% to about 16.0 mass%, preferably from about 15.0 to about 16.0 mass%, based on the mass of the hydrocarbon mixture, according to ASTM D5291 and a sulfur content ranging from about 0.01ppm to about 0.1ppm, based on the mass of the hydrocarbon mixture, according to ASTM D5623. The hydrocarbon comprises about 5mm in a temperature range of about-35 ℃ to about-33 ℃ according to ASTM D7042 and/or ASTM D445 2 S to about 8mm 2 S ofKinematic viscosity, preferably from about 5 to about 7mm 2 And s. Each of these attributes work in concert to provide better fuel efficiency when the rocket uses these attributes during combustion.
As noted above, the disclosed rocket propellant includes a mixture of liquid oxygen and hydrocarbons. The liquid oxygen may be provided from any source such as cooled oxygen, compressed oxygen, or both so that it is substantially in the liquid state. The hydrocarbon mixture may be from natural sources, such as from a wellbore (wellbore) or the like, and may be synthesized, for example, by the fischer-tropsch reaction. The amounts of these components in the rocket propellant may vary. For example, the rocket propellant may have a ratio of liquid oxygen to hydrocarbon mixture in the range of about 1 to about 5. In some embodiments, the rocket propellant contains a ratio of liquid oxygen to hydrocarbon mixture of about 1, or about 1.25, or about 1.5, or about 1.75, or about 2, or about 2.25, or about 2.5, or about 2.75, or about 3, or about 3.25, or about 3.5, or about 3.75, or about 4, including about plus or minus 0.125.
As noted above, rocket propellants include hydrocarbons. Hydrocarbons may include, but are not limited to, paraffins, linear paraffins, branched paraffins, cyclic paraffins, and combinations thereof. Hydrocarbons may include alkanes, alkenes, alkynes, aromatic, saturated, unsaturated molecules, and combinations thereof. In some embodiments, the rocket propellant composition may include a hydrocarbon mixture having a hydrogen content in the range of about 14.0 mass% to about 16.0 mass%, preferably about 15.0 to about 16.0 mass%, based on the mass of the hydrocarbon mixture, according to ASTM D5291. The hydrogen content may be due to the covalent bonding of hydrogen atoms to the hydrocarbons contained in the hydrocarbon mixture. In some embodiments, the rocket propellant composition may comprise a hydrogen content of about 14.0 mass%, or about 14.1 mass%, or about 14.2 mass%, or about 14.3 mass%, or about 14.4 mass%, 14.5 mass%, or about 14.6 mass%, or about 14.7 mass%, or about 14.8 mass%, or about 14.9 mass%, or about 15.0 mass%, or about 15.1 mass%, or about 15.2 mass%, or about 15.3 mass%, or about 15.4 mass%, or about 15.5 mass%, or about 15.6 mass%, or about 15.7 mass%, or about 15.8 mass%, or about 15.9 mass%, or about 16.0 mass%, including about plus or minus 0.05 mass%, based on the mass of the hydrocarbon mixture.
The hydrocarbon mixture of the disclosed embodiments may have a flash point greater than about 60 ℃ at 1 ATM. The hydrocarbon mixture has a flash point at 1ATM in the range of about 60 c to about 100 c. In some embodiments, the hydrocarbon mixture may have a flash point at 1ATM of about 60 ℃, or about 65 ℃, or about 70 ℃, or about 75 ℃, or about 80 ℃, or about 85 ℃, or about 90 ℃, or about 95 ℃, or about 100 ℃, including about plus or minus 2.5 ℃.
The hydrocarbon mixture may have a distillation initial boiling point of at least about 180 ℃ at 1 ATM. According to some embodiments, the hydrocarbon mixture may have an initial boiling point for distillation at 1ATM of from about 160 ℃ to about 200 ℃, preferably 180 ℃ to 200 ℃ at 1ATM, according to ASTM D86. For example, the hydrocarbon mixture may have an initial boiling point at 1ATM of about 180 ℃, or about 182.5 ℃, or about 185 ℃, or about 185.5 ℃, or about 190 ℃, or about 192.5 ℃, or about 195 ℃, or about 195.5 ℃ or about 200 ℃, including about plus or minus 1.25 ℃.
According to ASTM D4052, the disclosed hydrocarbon mixture may have a temperature of about 0.5kg/m at about 15 deg.C 3 To about 1kg/m 3 Density within the range. In some embodiments, the hydrocarbon mixture may have a temperature of about 0.5kg/m at about 15 ℃ 3 Or about 0.6kg/m 3 Or about 0.7kg/m 3 Or about 0.8kg/m 3 Or about 0.9kg/m 3 Or about 1kg/m 3 About 0.05kg/m or less 3 。
In some embodiments, the hydrocarbon mixture can have a freezing point of less than about-45 ℃ according to ASTM D5792. In some embodiments, the hydrocarbon mixture can have a freezing point of about-45 ℃ to about-90 ℃. For example, the hydrocarbon mixture can have a freezing point of about-45 ℃, or about-50 ℃, or about-55 ℃, or about-60 ℃, or about-65 ℃, or about-70 ℃, or about-75 ℃, or about-80 ℃, or about-85 ℃, or about-90 ℃, wherein about includes plus or minus 2.5 ℃.
In addition to hydrogen, the hydrocarbon mixture may also include sulfur. In some embodiments, the disclosed hydrocarbon mixtures contain sulfur in the range of about 0.01ppm to about 0.1 ppm. The hydrocarbon mixture may have a sulfur content of about 0.01ppm, or about 0.02ppm, or about 0.03ppm, or about 0.04ppm, or about 0.05ppm, or about 0.06ppm, or about 0.07ppm, or about 0.08ppm, or about 0.09ppm or about 0.1ppm, including plus or minus 0.05ppm, based on the mass of the hydrocarbon mixture.
According to some embodiments, the rocket propellant composition may have a specific impulse value of at least about 1% gain over RP-1/RP-2 at a set operating condition. In some embodiments, the method of measuring the specific impulse is measuring the characteristic outlet velocity (C). C by standard gravity (g) and constant C f And dash (I) sp ) This is dependent on the geometry of the engine and heat transfer effects. The formula of the beta-impulse is I SP ＝(C*X C f ) (ii) in terms of/g. C may use a formula
And (4) calculating. P is the pressure of the combustion chamber, A t Is the area of the nozzle throat, and
is the fuel mass flow. In some embodiments, the disclosed rocket propellant compositions include a specific impulse of greater than about 5,700ft/s at 750psi and MR 2.7.
Method for producing rocket propellant compositions
In some embodiments, a method of producing a rocket propellant composition may comprise the steps of: (a) providing liquid oxygen and (b) providing a hydrocarbon mixture having a hydrogen content of from about 14.0 mass% to about 16.0 mass%, preferably from about 15 mass% to about 16 mass%, based on the mass of the hydrocarbon mixture, according to ASTM D5291. According to some embodiments, the hydrocarbon mixture may include about 5mm in a temperature range of about-35 ℃ to about-33 ℃ according to ASTM D7042 and/or ASTM D445 2 S to about 8mm 2 Kinematic viscosity in/s, preferably about 5mm 2 S to about 7mm 2 And s. And a sulfur content of from about 0.01ppm to about 0.1 mass%, based on the mass of the hydrocarbon mixture. One method may include mixing liquid oxygen with a hydrocarbon mixture and then combusting the mixture.
As described above, a method may include mixing liquid oxygen and a hydrocarbon mixture. The mixing of the liquid oxygen and the hydrocarbon may be done at the nozzle plate (burner plate) or at a combustion point. The disclosed method may include mixing liquid oxygen and a hydrocarbon mixture in a ratio of about 1 to about 4. For example, a method can include mixing liquid oxygen and a hydrocarbon mixture in a mixing ratio of about 1, or about 1.25, or about 1.5, or about 1.75, or about 2, or about 2.25, or about 2.5, or about 2.75, or about 3, or about 3.25, or about 3.5, or about 3.75, or about 4, wherein about includes plus or minus 0.125.
Examples
The following examples illustrate some specific exemplary embodiments of the present disclosure. These embodiments represent specific ways that are found to work well in the practice of the application and, therefore, can be considered as embodiments that constitute a mode of its practice. However, those of skill in the art should, in light of the present disclosure, appreciate that many changes can be made in the specific embodiments which are disclosed and still remain within the spirit and scope of the application.
Example 1
The percent change of useful payload versus specific impulse (I) is plotted SP ) Is plotted (fig. 1). The relationship graph was made using the ideal rocket equation applied to grade 2 SpaceX falcon No. 9 launched into the near earth orbit. Standard assumptions are used, such as increasing the final velocity to account for the lack of air drag and gravitational effects in the ideal rocket equation. As shown in FIG. 1, I sp An increase of about 1% corresponds to an increase of about 5% in useful payload. This number should be considered indicative rather than precise; but emphasizes the sensitivity of the useful payload to the specific impulse indicator (driven by the chemical and combustion characteristics of the propellant identified in this application).
Examples2
One hydrocarbon mixture sample (GTL product GS190) was produced and compared by evaluating their specific impact gains at various liquid oxygen to hydrocarbon mixture ratios (figures 2-4). Three graphs are plotted plotting the effective exhaust velocity c (an experimental measurement of rocket engine exhaust velocity, directly related to Isp by the gravity constant g) versus the liquid oxygen hydrocarbon mixture ratio at three pressures, 350psi (figure 2), 500psi (figure 3) and 750psi (figure 4). Figure 3 shows the lower and upper limits (from actual test data) for sample 1at 350 psi. FIG. 4 compares sample 1 with the lower and upper limits of RP-1 (from actual test data) at 500psi, and FIG. 5 compares RP-2 with sample 1at 750 psi. These figures show how the novel rocket fuel disclosed in this application (sample 1) shows improvement in Isp performance (measured by c) compared to RP-1/RP-2, which becomes more pronounced at higher combustion chamber pressures and higher mixing ratio values.
Example 3
Kinematic viscosity data were obtained for four rocket propellant samples, sample 1, RP-1 and RP-2 (FIG. 5). Since the viscosity of sample 1 does not reach the standard viscosity of RP-1 or RP-2 (4.5cSt) until the temperature is significantly lower than RP-1 or RP-2, the data shows how the disclosed rocket propellant sample (sample 1) has superior cryogenic cooling capabilities over known rocket propellants (RP-1, RP-2).
In contrast, rocket propellants RP-1 and RP-2 are known to reach an ultimate viscosity at a temperature of about-7 ℃ as shown in FIG. 5.
Example 4
Two-dimensional gas chromatography or GC x GC data of the disclosed rocket propellant samples (GS190) were obtained. As shown in table 1, 41.92% of the rocket propellants were branched alkanes having a length of 12 carbons, and 25% of the rocket propellants were branched alkanes having a length of 11 carbons. In addition to branched alkanes, 24.57% of the rocket propellants included straight-chain alkanes. Also as shown in table 1, 25% was linear alkane, 71% was branched alkane, 4% was cycloalkane, and 1% was dicycloalkane. An important point to note for these indicators is the lack of aromatic content: the aromatics content in RP-1 (and even more so in RP-2) needs to be low to reduce the tendency of the fuel to decompose under high thermal stress (coking). The fact that the fuel consists primarily of alkanes should enhance resistance to thermal decomposition or coking, making this rocket fuel an attractive option for rocket engines that rely on pre-burning rocket fuels for cooling capability.
Table 1 rocket propellant GC X GC data.
It is to be understood that the components of each unit are listed for illustrative purposes only and are not intended to limit the scope of the present application. Particular combinations of these or other components or units can be configured in such compositions or methods for the intended use based on the teachings herein.
Various changes in the shape, size, number, separation characteristics, and/or arrangement of parts may be made by those skilled in the art without departing from the scope of the present disclosure. In accordance with some embodiments, each disclosed component, system, or process step can be associated with any other disclosed component, system, or process step and performed in any order. Where the verb "may" appears, it is intended to convey an optional and/or permissive condition, but its use does not imply lack of operability unless otherwise specified. Various modifications may be made by those skilled in the art to methods of making and using the compositions, devices, and/or systems of the present disclosure. Some embodiments of the disclosure may be practiced, where desired, exclusive of other embodiments.
Further, where ranges have been provided, the disclosed endpoints may be treated as exact and/or approximate as desired or required for the particular embodiment. Where the endpoints are approximate, the degree of flexibility may vary in proportion to the number of steps of the range. For example, in one aspect, the end point of the range of about 50 in the range of about 5 to about 50 can include 50.5 but not 52.5 or 55, and in another aspect, the end point of the range of about 50 in the range of about 0.5 to about 50 can include 55 but not 60 or 75. Furthermore, in some embodiments, it may be desirable to mix and match range endpoints. Moreover, in some implementations, each figure disclosed (e.g., in one or more of the examples, tables, and/or figures) may form the basis for a range (e.g., a delineated value of +/-about 10%, a delineated value of +/-about 50%, a delineated value of +/-about 100%) and/or a range endpoint. With respect to the former, the value 50 depicted in the examples, tables, and/or figures may form the basis of a range, for example, from about 45 to about 55, from about 25 to about 100, and/or from about 0 to about 100.
Such equivalents and substitutions as well as obvious changes and modifications are intended to be included within the scope of the present disclosure. Accordingly, the foregoing disclosure is intended to be illustrative, but not limiting, of the scope of the disclosure, which is set forth in the following claims.
The title, abstract, background, and title may be provided in accordance with the statute and/or for the convenience of the reader. They do not include an admission as to the scope or content of the prior art and do not include limitations that apply to all of the disclosed embodiments.
Claims (21)
1. A rocket propellant composition comprising:
(a) liquid oxygen; and
(b) a hydrocarbon mixture comprising:
(i) a hydrogen content of from about 14.0 mass% to about 16.0 mass% based on the mass of the hydrocarbon mixture;
(ii) a kinematic viscosity of about 5mm at a temperature in the range of about-35 ℃ to about-33 ℃ 2 S to about 8mm 2 S; and
(iii) the sulfur content is from about 001 ppm to about 0.1ppm based on the mass of the hydrocarbon mixture.
2. A rocket propellant composition according to claim 1, wherein the mixing ratio of said liquid oxygen to said hydrocarbon mixture ranges from about 1.8 to about 3.2.
3. A rocket propellant composition according to claim 1, wherein the mixing ratio of said liquid oxygen to said hydrocarbon mixture ranges from about 1 to about 4.
4. A rocket propellant composition according to claim 1 wherein the hydrogen content ranges from about 15.3 to about 15.9 mass% based on the mass of the hydrocarbon mixture.
5. A rocket propellant composition according to claim 1, wherein said hydrocarbon mixture further comprises one or more of branched alkanes, linear alkanes, cycloalkanes, and combinations thereof.
6. A rocket propellant composition according to claim 1, wherein said hydrocarbons further comprise a flash point at 1ATM of at least about 60 ℃ according to ASTM D93.
7. A rocket propellant composition according to claim 1, wherein said hydrocarbon mixture further has a distillation initial boiling point of at least about 180 ℃.
8. A rocket propellant composition according to claim 1, wherein said hydrocarbon mixture further has a temperature of about 15 ℃ of about 0.7kg/m 3 To about 0.9kg/m 3 The density of (c).
9. A rocket propellant composition according to claim 1, wherein said hydrocarbon mixture further comprises a freezing point below about-45 ℃.
10. A rocket propellant composition according to claim 1, wherein said hydrocarbon mixture is a fischer-tropsch product.
11. A rocket propellant composition according to claim 1, wherein said rocket propellant composition comprises c values at least about 1% gain over RP-1/RP-2 under set operating conditions.
12. A method of producing a rocket propellant composition, the method comprising:
(a) providing liquid oxygen; and
(b) providing a hydrocarbon mixture comprising:
(i) a hydrogen content of from about 15.0 mass% to about 16.0 mass% based on the mass of the hydrocarbon mixture;
(ii) a kinematic viscosity of about 5mm at a temperature in the range of about-35 deg.C to about-33 deg.C 2 S to about 7mm 2 S; and
(iii) the sulfur content is from about 0.01ppm to about 0.1 mass% based on the mass of the hydrocarbon mixture.
13. The method of claim 12, wherein the mixing ratio of the liquid oxygen to the hydrocarbon mixture ranges from about 1 to about 4.
14. The method of claim 12, wherein the mixing ratio of the liquid oxygen to the hydrocarbon mixture is about 1.8 to about 3.2.
15. The method of claim 12, further comprising mixing the liquid oxygen and the hydrocarbon mixture at a nozzle fixing plate.
16. The method of claim 12, wherein the hydrogen content ranges from about 15.3 mass% to about 15.9 mass% based on the mass of the hydrocarbon mixture.
17. The method of claim 12, wherein the hydrocarbon further comprises a flash point of at least about 60 ℃ at 1ATM according to ASTM D93.
18. The method of claim 12, wherein the hydrocarbon mixture further comprises a distillation initial boiling point of at least about 190 ℃.
19. The method of claim 12, wherein the rocket fuel composition further comprises about 0.7kg/m at a temperature of about 15 ℃ 3 To about 0.8kg/m 3 The density of (2).
20. The process of claim 12, wherein the hydrocarbon mixture is a fischer-tropsch product.
21. The method of claim 12, wherein the rocket propellant composition comprises a c value of at least about 1% gain over RP-1/RP-2 under set operating conditions.
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| US201962954986P | 2019-12-30 | 2019-12-30 | |
| US62/954,986 | 2019-12-30 | ||
| PCT/EP2020/087928 WO2021136757A1 (en) | 2019-12-30 | 2020-12-28 | Compositions and methods of producing rocket propellants with enhanced cryogenic cooling, thermal stability, and thrust efficiency performance |
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