WO2014149184A2

WO2014149184A2 - Enhanced combustion rocket engine systems and methods

Info

Publication number: WO2014149184A2
Application number: PCT/US2014/013247
Authority: WO
Inventors: Andrew Stephen HOWELLS
Original assignee: Howells Andrew Stephen
Priority date: 2013-01-28
Filing date: 2014-01-27
Publication date: 2014-09-25
Also published as: WO2014149184A3

Abstract

An enhanced combustion rocket engine system (100), involving at least one primary propellant tank (110) adapted to accommodate a primary propellant (111) in at least one of a solid state, a transition melting state, and a molten state and to deliver the primary propellant (111) in the molten state, at least one secondary propellant tank adapted to accommodate a secondary propellant in a liquid state; and at least one rocket engine (140), the at least one rocket engine (140) adapted to receive the primary propellant (111) from the at least one primary propellant tank (110) in a temperature range of at least approximately above an ambient temperature, to receive the secondary propellant (121) from the at least one secondary propellant tank (120) in a temperature range of at least approximately below an ambient temperature; and exothermically react the primary and secondary propellants (111, 121) for effecting combustion to provide thrust.

Description

ENHANCED COMBUSTION ROCKET ENGINE SYSTEMS AND METHODS

CROSS-REFERENCE TO RELATED APPLICATION(S)

[0001] This document is a PCT Application which claims priority to, and the benefit of, U.S. Provisional Patent Application Serial No. 61/757,695, filed on January 28, 2013, entitled "Combustion Rocket Systems Able to Burn Higher Energy Density Propellants, Structural Fuel, and Aerodynamic Fairings," which is herein incorporated by reference in its entirety for all purposes.

TECHNICAL FIELD

[0002] The subject matter of the present disclosure technically relates to combustion rocket engine systems and methods. More specifically, the present disclosure technically relates to combustion rocket engine systems and methods, having associated external support systems therefor. Even more specifically, the present disclosure technically relates to enhanced combustion rocket engine systems and methods, having associated external support systems for use with improved fuels and improved fuel combinations.

BACKGROUND

[0003] In the related art, rockets often accelerate most of their propellant in directions of intended travel before accelerating the propellant in opposite directions to provide thrust. Therefore, current related art rocket technology faces challenges in minimizing the mass of the propellant without significantly compromising the amount of energy and thrust force available for accelerating a rocket vehicle. Yet, a decrease in required energy may decrease the cost of engineering, manufacturing, transporting, maintaining, operating, and protecting rockets and their associated support systems.

[0004] Choosing propellants which offer high energy density may minimize the mass of propellant required by rockets. Chemical potential energy is defined herein as the amount of thermal energy that can be obtained from reacting or decomposing propellants. Gravimetric energy density is the amount of chemical potential energy per unit mass of propellant. Volumetric energy density is the amount of chemical potential energy per unit volume of propellant. In the related art, application of some reactant combinations, having high energy density, has suffered from high cost due to rarity or difficulty in refining, instability, being toxic or caustic, or producing toxic or caustic exhaust, such as the tri -propellant combination of hydrogen, fluorine, and lithium. Some of these combinations may nevertheless be useful for some applications. Failure to achieve optimum pressure, as experienced by related art rockets, results in a low efficiency and a low thrust to mass ratio.

[0005] In the related art rocket engines, powdered aluminum does not burn rapidly enough for use in some compact rocket engine designs; and powdered aluminum may also produce a shower of sparks indicative of incomplete combustion. Other related art solid fueled engines are not subject to the limitation of requiring a compact engine. In such other related art rockets, a large portion of the propellant tank is able to serve as a combustion chamber. However, this related art approach may not be ideal for providing a high energy density if an implementation requires using non-liquid oxygen, or solid oxidizers, which provide an inferior oxidizing potential per unit mass. In the related art, these solid fueled engines are unsuitable for some common applications for at least their heavy casings. As such, a long-felt need exists for rockets that can burn combinations of propellants that offer high gravimetric energy density for improving stability and thrust as well as many more performance characteristics. SUMMARY

[0006] The subject matter of the present disclosure addresses many of the related art issues by increasing gravimetric energy density, thereby reducing the amount of thrust and energy required of a rocket (since some rockets must accelerate their own fuel, less fuel mass of that such rockets would accelerate in the direction of intended travel), by increasing volumetric energy density, thereby reducing the amount of thrust and energy required of a rocket (since some rockets must accelerate their own fuel tanks, engines, and structures, and there may be less mass of fuel tanks, engines and structure that such rockets would accelerate), and by improving volumetric energy density, thereby reducing the amount of power needed to force propellant into the combustion chambers. Further, the subject matter of the present disclosure also addresses other related art issues by reducing aerodynamic drag in atmospheric flights, thereby reducing the amount of thrust required. By minimizing "empty" mass, as presently disclosed, the payload fraction is maximized. Empty mass is the mass of an aircraft or spacecraft excluding the mass of the fuel or payload.

[0007] Furthermore aerodynamic drag may be reduced in atmospheric flights, which may also reduce the amount of thrust required for rockets which can burn high energy density propellant (HEDP) with oxidizers offering high oxidation potential relative to mass, such as approximately pure liquid oxygen or other liquid oxidizers, for rocket engines which can burn HEDP without diluting the reducers with inferior reducers, binders, catalysts, or other additives which decrease gravimetric energy density, for rocket engines which can burn HEDP combinations while also offering the ability to be stopped and restarted and dynamically throttled in flight, for rocket engines which are capable of burning HEDP fuels that are well suited to convenient production at remote destinations in outer space, for appropriate supporting systems to facilitate design, manufacture, transport, maintenance, operation, and protection of rockets, for rockets with low empty mass. By minimizing empty mass, payload fraction can be maximized.

[0008] The subject matter of the present disclosure is directed to enhanced rocket systems, their fabrication methods, and their operating methods, encompassing various features for facilitating burning of various propellant combinations that offer high energy density. The presently disclosed enhanced rocket engine systems facilitate ensuring stable high power density combustion and achieving consistent delivery of propellant to a combustion chamber, wherein the prescribed propellant generally comprises combinations of reactants, such as a combination of a reducer component, such as a metal component that is in a solid state at standard sea-level atmospheric temperature and pressure on Earth, e.g., silicon or calcium, combined with an oxidizer component, such as a liquid oxidizer. The presently disclosed enhanced rocket engine systems facilitate delivery of fuel, such as the propellant, to at least one rocket engine at significantly high temperatures. For example, a system using a first topology, in accordance with the present disclosure, facilitates delivery of fuel by storing the majority of the reducer component in a molten state onboard the rocket vehicle for ready use at launch time. In another embodiment, the majority of the reducer component that is stored onboard the rocket vehicle as a preheated powder or pellets mixed into a molten metal or other hot liquid reducer at launch, or, alternatively, as a hot reducer powder or pellets mixed into hot liquid reducer later, e.g., during flight. In an embodiment, the powder or pellets may be melted before injection. Powder, for example, may be stored hot and/or preheated in flight and then injected as powder.

[0009] The subject matter of the present disclosure is generally directed to an enhanced combustion rocket engine system, comprising: at least one primary propellant tank adapted to accommodate a primary propellant, for example, in at least one of a solid state, a transition melting state, and a molten state and to deliver the primary propellant in the molten state; at least one secondary propellant tank adapted to accommodate a secondary propellant, for example, in a liquid state; and at least one rocket engine, the at least one rocket engine adapted to: receive the primary propellant from the at least one primary propellant tank in a temperature range, for example, of at least approximately above an ambient temperature; receive the secondary propellant from the at least one secondary propellant tank in a temperature range, for example, of at least approximately below an ambient temperature; and exothermically react the primary and secondary propellants for effecting combustion to provide thrust, in accordance with the present disclosure.

[0010] Further, the subject matter of the present disclosure is generally directed to an enhanced combustion rocket engine vehicle comprising: at least one primary propellant tank adapted to accommodate a primary propellant, for example, in at least one of a solid state, a transition melting state, and a molten state and to deliver the primary propellant in the molten state and to accommodate at least one consumable component; at least one secondary propellant tank adapted to accommodate a secondary propellant, for example, in a liquid state; and at least one rocket engine, the at least one rocket engine adapted to: receive the primary propellant from the at least one primary propellant tank in a temperature range, for example, of at least approximately above an ambient temperature; receive the secondary propellant from the at least one secondary propellant tank in a temperature range, for example, of at least approximately below an ambient temperature; and exothermically react the primary and secondary propellants for effecting combustion to provide thrust; and at least one fairing for accommodating a payload, the fairing coupled with at least one of: the at least one primary propellant tank, the at least one secondary propellant tank, and the at least one engine, in accordance with the present disclosure.

[0011] Even further, the subject matter of the present disclosure is generally directed to a method of fabricating an enhanced combustion rocket engine system, the method comprising: providing at least one primary propellant tank adapted to accommodate a primary propellant in at least one of a solid state, a transition melting state, and a molten state and to deliver the primary propellant in the molten state; providing at least one secondary propellant tank adapted to accommodate a secondary propellant in a liquid state; and providing at least one rocket engine, the at least one rocket engine adapted to: receive the primary propellant from the at least one primary propellant tank in a temperature range of at least approximately above an ambient temperature; receive the secondary propellant from the at least one secondary propellant tank in a temperature range of at least approximately below an ambient temperature; and exothermically react the primary and secondary propellants for effecting combustion to provide thrust, in accordance with the present disclosure.

[0012] Yet even further, the subject matter of the present disclosure is generally directed to a method of using an enhanced combustion rocket engine system, the method comprising: providing an enhanced combustion rocket engine system, the rocket engine system providing comprising: providing at least one primary propellant tank adapted to accommodate a primary propellant in at least one of a solid state, a transition melting state, and a molten state and to deliver the primary propellant in the molten state; providing at least one secondary propellant tank adapted to accommodate a secondary propellant in a liquid state; and providing at least one rocket engine, the at least one rocket engine adapted to: receive the primary propellant from the at least one primary propellant tank in a temperature range of at least approximately above an ambient temperature; receive the secondary propellant from the at least one secondary propellant tank in a temperature range of at least approximately below an ambient temperature; and exothermically react the primary and secondary propellants for effecting combustion to provide thrust; charging the at least one primary propellant tank with the primary propellant; charging the at least one secondary propellant tank with the secondary propellant; preheating, heating, and delivering the primary propellant from the at least one primary propellant tank to the at least one rocket engine in a temperature range of at least approximately above an ambient temperature; delivering the secondary propellant from the at least one secondary propellant tank to the at least one rocket engine in a temperature range of at least approximately below an ambient temperature; and exothermically reacting the primary and secondary propellants for effecting combustion in the at least one rocket engine to provide thrust, in accordance with the present disclosure.

[0013] Furthermore, the subject matter of the present disclosure is generally directed to a propellant formulation for an enhanced combustion rocket engine system, comprising: a primary propellant in at least one of a solid state, a transition melting state, and a molten state and to deliver the primary propellant in the molten state, the primary propellant comprising a temperature in a range of at least approximately above an ambient temperature; and a secondary propellant in a liquid state, the secondary propellant comprising a temperature in a range of at least approximately below an ambient temperature, the primary and secondary propellants capable of exothermically reacting with each other for effecting combustion to provide thrust, in accordance with the present disclosure.

[0014] In some embodiments of the present disclosure, the rocket systems comprise an aggressive preheating system capable of melting, or, if the reducer component is already molten before launch, aggressively further preheating, most of the reducer component during flight prior to injecting the reducer component into an engine. In some embodiments of the present disclosure, the rocket systems comprise self-destruction features which facilitate self-destroying, such as burning the systems' own propellant tanks, structures, or fairings, the self-destruction features comprising at least one of: a propellant tank for accommodating hot molten fuel, such as a hot molten reducer component; features for disassembling at least one other system component, such as other structures and fairings, and features for feeding the at least one other system component into the propellant tank, wherein the at least one other system component is melted prior to being injected into the at least one engine's combustion chamber, whereby the at least one other system component is cannibalized, and wherein the propellant tank optionally comprises an injector for facilitating heating the oxidizer component by way of burning.

BRIEF DESCRIPTION OF THE DRAWINGS

[0015] The above, and other, aspects, features, and advantages of several embodiments of the present disclosure will be more apparent from the following Detailed Description as presented in conjunction with the following several figures of the Drawing.

[0016] FIG. 1 is a schematic diagram illustrating a sectional side view of an enhanced combustion rocket engine system, in accordance with an embodiment of the present disclosure.

[0017] FIG. 2A is a schematic diagram illustrating a sectional side view of an enhanced combustion rocket engine system, in accordance with an embodiment of the present disclosure.

[0018] FIG. 2B is a schematic diagram illustrating a sectional side view of the enhanced combustion rocket engine system of FIG. 2A, in accordance with an embodiment of the present disclosure.

[0019] FIG. 3 is a schematic diagram illustrating a sectional side view of an enhanced combustion rocket engine system of a rocket vehicle, in accordance with an embodiment of the present disclosure.

[0020] FIG. 4 is a schematic diagram illustrating a sectional side view of an enhanced combustion rocket engine system of a rocket vehicle, in accordance with an embodiment of the present disclosure.

[0021] FIG. 5 is a flowchart illustrating a method of fabricating an enhanced combustion rocket engine system, in accordance with an embodiment of the present disclosure.

[0022] FIG. 6 is a flowchart illustrating a method of using an enhanced combustion rocket engine system, in accordance with an embodiment of the present disclosure. [0023] FIG. 7 is a schematic diagram illustrating a semi-exploded sectional side view of a rocket vehicle powerable by an enhanced combustion rocket engine system, in accordance with an embodiment of the present disclosure.

[0023] FIG. 8 is a schematic diagram illustrating of a semi-exploded sectional side view of a rocket vehicle powerable by an enhanced combustion rocket engine system, in accordance with an embodiment of the present disclosure.

[0024] FIG. 9 is a schematic diagram illustrating a semi-exploded sectional side view of a rocket vehicle powerable by an enhanced combustion rocket engine system, in accordance with an embodiment of the present disclosure.

[0025] FIG. 10 is a schematic diagram illustrating a semi -exploded sectional side view of a rocket vehicle powerable by an enhanced combustion rocket engine system, in accordance with an embodiment of the present disclosure.

[0026] FIG. 11 is a schematic diagram illustrating a semi-exploded sectional side view of a rocket vehicle powerable by an enhanced combustion rocket engine system, in accordance with an embodiment of the present disclosure.

[0027] FIG. 12 is a schematic diagram illustrating a semi -exploded sectional side view of a rocket vehicle powerable by an enhanced combustion rocket engine system, in accordance with an embodiment of the present disclosure.

[0029] FIG. 13 is a schematic diagram illustrating a semi -exploded sectional side view of a rocket vehicle powerable by an enhanced combustion rocket engine system, in accordance with an embodiment of the present disclosure.

[0030] FIG. 14 is a schematic diagram illustrating a semi-exploded sectional side view of a rocket vehicle powerable by an enhanced combustion rocket engine system, in accordance with an embodiment of the present disclosure.

[0031] FIG. 15 is a schematic diagram illustrating a semi-exploded sectional side view of a rocket vehicle powerable by an enhanced combustion rocket engine system, in accordance with an embodiment of the present disclosure.

[0032] Corresponding reference characters indicate corresponding components throughout the several figures of the Drawing. Elements in the several figures are illustrated for simplicity and clarity and have not necessarily been drawn to scale. For example, the dimensions of some elements in the figures may be emphasized relative to other elements for facilitating understanding of the various presently disclosed embodiments. Also, common, but well- understood, elements that are useful or necessary in commercially feasible embodiments are often not depicted in order to facilitate a less obstructed view of these various embodiments of the present disclosure.

DETAILED DESCRIPTION

[0033] In accordance with the present disclosure, maximizing volumetric energy density minimizes the cost of engineering, manufacturing, transporting, maintaining, operating, and protecting rockets as well as their associated support systems, e.g., their external support systems. In some situations, related art rockets must accelerate at high rates, or deliver high thrust-to-mass ratio in order to be efficient. However, high rates of acceleration reduce the distance traveled while accelerating; and, as a consequence, the amount of energy lost to aerodynamic drag is significant in the related art. When launching heavy loads from planets or moons, e.g., bodies having strong gravity, such as Earth, rates of acceleration above 5 meters per second squared [m/s²] minimize the amount of time and energy spent by rockets on overcoming gravity during such launches, according to the present disclosure. Some types of orbital maneuvers, such as Hohmann transfer, are theoretically most efficient when maneuvering the spacecraft approximates instantaneous impulses as well as possible; and, as presently disclosed, rockets should provide a sufficient gravimetric power density, e.g., a sufficient thrust to mass ratio, for achieving such conditions. The cost and mass of adequately strong tanks and structures for rockets capable of withstanding such acceleration is reducible, as presently disclosed, if the propellant volume is reduced relative to energy. Maximizing the volumetric energy density of propellants minimizes the cost and mass of rockets by minimizing the amount of power and energy required to force propellant into combustion chambers. Furthermore, aerodynamic drag is reducible in atmospheric flights, thereby also reducing the amount of requisite thrust, as presently disclosed. Improving energy density permits, as presently disclosed, reducing the length of a rocket, thereby also reduces skin drag.

[0034] Certain propellant combinations, as presently disclosed, offer the benefits, such as high gravimetric and volumetric energy density. In addition, some of these propellants, and combinations thereof, avoid the problems experienced in the related art, such as high cost. Given some aspect considerations, some of these propellant combinations (combinations of reactants) are attractive, as presently disclosed, such as a propellant combination comprising a high energy density propellant (HEDP) and liquid oxygen (LO₂ or LOX), herein referred to as an "HEDP combination," unless otherwise noted, denoting propellant combinations which offer high energy density. As presently disclosed, several embodiments provide helpful options for ensuring efficient, fast, and stable combustion, one embodiment of which is capable of burning a propellant combination, comprising aluminum, as an HEDP component, and LOX, as an oxidizer component. In some related art rockets, reactants are stored in a mixture or a composite form before chemically reacting them. According to the present disclosure, except where otherwise noted, the term "combination," when used in relation to propellants, refers to a selection of propellant types, without necessarily indicating that the propellants are in a mixture or a composite form. By example only, the propellant combinations, as presently disclosed, comprise HEDP combinations, but are not limited to, bi-propellant combinations, comprising at least one HEDP component, such as aluminum (Al), beryllium (Be), boron (B), calcium (Ca), magnesium (Mg), silicon (Si), and titanium (Ti), and at least one oxidizer component, such as oxygen (O), nitric acid (HNO3), dinitrogen tetroxide (N2O4), and hydrogen peroxide(H202). Further, the propellant combinations, as presently disclosed, comprise HEDP combinations, include, but are not limited to, tri-propellant combinations, such as a combination of lithium (Li), fluorine (Fl), and hydrogen (H).

[0035] As presently disclosed, except where otherwise noted, the following discussion involves a preferred HEDP combination, such as aluminum with LOX; and the discussion applies to other HEDP combinations, as above listed, and to any other HEDP combinations, even if not explicitly listed herein. Developing rockets that do not require the use of fuel comprising oxidizers in a solid state, that are able to use a non-solid oxidizer being mixable with aluminum in a powder form for forming a stable composite material, and that are capable of burning the stable composite material is desirable, according the present disclosure. Such rockets have engines which accommodate propellant combinations, comprising oxidizers that provide an oxidation potential per unit mass of oxidizer, wherein the oxidation potential per unit mass comprises a range that is as high as approximately that of pure LOX or other liquid oxidizers, whereby high energy density is ensured. Furthermore, the present disclosure contemplates rockets that do not require additional fuels, catalysts, or binders and that also have a lower energy density than their primary reducer component.

[0036] In addition, for some rocket engines, achieving high pressure in the combustion chambers and in the throats is desirable, according to the present disclosure. To exploit the high energy density of HEDP combinations, the present disclosure encompasses elevating temperature and pressure in combustion chambers or expansion chambers and nozzles in significantly higher ranges than those of related art rockets. To ensure that the engine's structural material is capable of withstanding such elevated temperature and elevated pressure, some rocket engine embodiments, as presently disclosed, implement active cooling. In some rocket engine embodiments, the amount of heat capable of being absorbed by a coolant during operation is a function of whether the propellant is acting as the coolant. The amount of heat absorbed by a combustion chamber is a function of the combustion chamber's internal surface area; therefore, a practical upper bound on the combustion chamber's internal surface area exists. To minimize the mass and cost of a rocket engine, minimizing the internal surface area of the engine's hottest portions is prescribed by the present disclosure, e.g., by rapid combustion. The rate at which solid Al burns is largely a function of the surface-area-to-volume ratio of the burning Al particles. Among affordable solid forms, powdered forms for the HEDP component provide a high surface-area-to-volume ratio and, hence, burn rapidly in relation to other solid forms of the HEDP component.

[0037] To address the related art instabilities in combustion, otherwise resulting in discomfort to passengers, damage to payloads, structural damage to rockets, or inefficiency, the present disclosure prescribes developing rocket engine systems that are capable of preventing blockages or inconsistencies in feeding Al at the high flow rates into a rocket engine combustion chamber against the pre-existing high opposing pressure in the combustion chamber, as some rocket vehicles benefit from stable combustion. In achieving such function, minimizing the cost and mass of the structures of large rockets and their payloads is important, whereby the structures are engineered to be as weak as safely possible. As such, the delicate nature of such structures militates in favor of feeding fuel into the rocket engines in a manner that achieves adequate combustion stability.

[0038] While some related art solid fueled rocket engines are unable to be dynamically throttled or stopped before they have used all of their fuel, many current related art liquid fueled rocket engines do have this capability, but also have other disadvantages, such as requiring igniters. However, as presently disclosed, certain applications benefit from rocket engines, having this ability to be dynamically throttled or stopped before they have used all of their fuel, which renders such engines more cost effective. Examples of such applications include, but are not limited to, maneuvering thrusters or upper stages, wherein multiple payloads are placed into different orbits. Noteworthy is that, in the presently disclosed systems and methods, the hot fuels are hypergolic, e.g., the hot fuels ignite immediately on contact without requiring any further ignition source, such as a spark. By sufficiently preheating the relevant component(s) of the present disclosed propellant formulations, e.g., to the melting point for aluminum, hypergolic behavior is facilitated in the systems and methods of the present disclosure. For example, another hypergolic propellant, such as hydrazine, is preferred over some higher energy density propellants, such as kerosene, since hypergolic behavior facilitates many reliable and quick restarts, whereby an accumulation of unreacted propellants, or any mixtures thereof, otherwise resulting in a dangerous condition, is eliminated. However, current hypergolic propellants, such as hydrazine, are toxic, carcinogenic, and corrosive.

[0039] Although oxidizer components of the HEDP propellants may be somewhat corrosive, such oxidizer component is not nearly as toxic or carcinogenic as hydrazine in the related art. Hydrogen is a propellant used in the related art, having the best in gravimetric energy density among traditional propellants. However, storing large volumes of hydrogen for long periods is highly challenging due to its extremely low boiling point and low volumetric energy density, whereby extremely large tanks and extreme insulation would be required. Also, if a low temperature is not maintained for hydrogen, then the hydrogen must be vented, whereby a significant portion of the hydrogen is permanently lost in the related art systems and methods. Without venting, the pressure of the hydrogen in the storage tank would increase; and the tank would explode. With the presently disclosed hot propellants, e.g., hot reducers, maintaining the temperature at cryogenic temperatures at all times in storage is not a requirement. Although the reducer component of the presently disclosed propellant formulation does not, in general, require temperature maintenance in storage, the reducer component does require heating before the reducer component is transferred to the combustion chamber of the rocket engine, even in an on-demand mode, in accordance with some embodiments of the present disclosure.

[0040] Some types of HEDP materials, such as the reducers, e.g., aluminum, silicon, magnesium, and such as an oxidizer, e.g., oxygen, offer the advantage of being abundant at destinations that are interesting in terms of space exploration and exploitation. Examples of such destinations include bodies, such as planets, moons, dwarf planets, asteroids, and comets. Many such destinations are composed primarily of elements that are suitable for use as HEDP materials. For example, a regolith has high concentrations, by mass, of the following elements: oxygen, silicon, aluminum, iron, magnesium, calcium silicon, and titanium. This abundance of accessible superficial deposits potentially defrays the cost of space missions by allowing fuel for return trips to be harvested and processed locally. Such exploitation of local minerals and other natural resources is referred to as in-situ resource utilization (ISRU).

[0041] In the systems and methods of the present disclosure, Tsiolkovsky's rocket equation is applied for predicting the change in velocity, delta-v, that a rocket system can provide. The Tsiolkovsky's rocket equation is expressed as follows: Δν = v_e * ln(mo/mi), wherein v_e = exhaust velocity, mo = initial total mass including propellant, mi = final total mass, and In denotes the natural logarithm function. According to Tsiolkovsky's equation, the fuels used in rockets should be able to deliver payload fractions that are significantly higher than related art rockets actually are able to do. Herein, except where otherwise noted, the payload fraction denotes the fraction of a rocket's gross mass, including its fuel and payload, at the beginning of a flight. Tsiolkovsky's approximation assumes that rockets have an "empty" mass that is infinitesimally light. Some related art modern rockets deliver lower payload fractions for the following reasons: the engines sometimes do not reach an approximately 100% efficiency, the rocket empty mass is significant in relation to the payload mass for rockets delivering a large change in velocity, e.g., greater than 5 kilometers per second, and aerodynamic drag and gravity have significant impact on the payload fraction. Herein, empty mass of a rocket is approximated as the mass of all components of a rocket, excluding its propellant and payload.

[0042] The rockets of the present disclosure are configured to convert chemical energy to thrust by reacting propellants to generate heat and pressure, whereby the propellant is accelerated and ejected. The presently disclosed embodiments rely on preheating of propellants prior to combustion in the main combustion chambers. By preheating the propellants, the presently disclosed embodiments are able to burn fuels rapidly. By burning fuels rapidly, the presently disclosed embodiments are able to use a compact combustion chamber, while still achieving approximately complete combustion in, or adequately near, the combustion chamber to generate thrust efficiently, and while burning propellants that offer high energy density. Indeed, the presently disclosed embodiments, comprising features for sufficiently preheating propellants, efficiently generate thrust by burning aluminum and liquid oxygen as the engine's primary propellants. This propellant combination offers high energy density; and. this present approach is an improvement over some related art engines.

[0043] In some conditions, solid aluminum burns more slowly than other rocket fuels for at least the following reasons. The boiling point of aluminum is over 2000 degrees Celsius higher than that of some rocket fuels, such as hydrogen or hydrocarbons. Aluminum is solid at room temperature, and, therefore, requires the energy input of 1 or 2 more phase changes before aluminum becomes a gas. Aluminum atoms have mass of approximately 27 atomic mass units (AMU), whereas fuels for some rocket engines use the following exemplary atoms: hydrogen having a mass of 1 AMU, carbon having a mass of 12 AMU, and oxygen having a mass of mass 16 AMU. Aluminum, when burned, is assembled into larger molecules having a mass of 102 AMU with 5 atoms, whereas hydrogen and hydrocarbons form water having a mass of 18 AMU with 3 atoms, and carbon monoxide having a mass of 28 AMU with 2 atoms. Due to the greater mass and large number of atoms in each exhaust molecule, aluminum burns more slowly. Also, exhaust from burning aluminum is alumina (AI₂O₃), a refractory material having poor thermal conductivity, an extremely high melting point, and an extremely high boiling point. If aluminum powder is exposed to air, aluminum powder will oxidize, thereby forming an aluminum oxide coating which inhibits further corrosion and interferes with combustion.

[0044] Some related art solid fuel rockets use solid oxidizers which are inferior to oxygen in oxidation potential per unit mass. The presently disclosed rockets feature combustion chambers adapted to exploit liquid oxidizers. In liquid fueled rockets, the size of a combustion chamber may partly determine the amount of heat which is absorbed by a combustion chamber inner surface (with a larger chamber absorbing more heat). If a combustion chamber is too small, combustion may not nearly complete in the combustion chamber, which in some related art rockets results in poor efficiency. In accordance with the present disclosure, the minimum size needed for a combustion chamber is, in part, a function of the combustion rate, e.g., how rapidly combustion completes. In the present systems and methods, by utilizing faster combustion, a smaller combustion chamber is employed in the rocket engine, e.g., smaller size relative to the amount of thrust required, wherein the smaller combustion chamber absorbs less heat while still achieving nearly complete combustion, and whereby the rocket engine efficiency is improved.

[0045] Depending on other aspects of a rocket configuration, if "hot" fuel is stored in a main tank, then that "hot" fuel may not be able to absorb as much heat as would a coolant. In the present systems and methods, another propellant, such as an oxidizer, may be used, instead, as a coolant. Since an oxidizer has a heat absorption ceiling, for some topological configurations, an upper bound is placed on the size of the combustion chamber. Nonetheless, since combustion chambers are sometimes heavy and expensive, avoiding this upper bound is ideal in the present systems and methods and further minimizes cost, e.g., by minimizing the number propellant types used as a coolant in a rocket engine, wherein the rocket engine uses a plurality of propellant types. Some solid forms of fuel, such as powder, are poorly suited for use as coolant in some related art rockets. Thus, the present systems and methods address the need for rockets that are capable of exploiting the high energy density and other advantages of certain fuels for effecting a rapid fuel burn, e.g., by preheating fuel prior to burning the fuel. Before discussing specific recommended strategies for preheating, the following paragraph details the recommended or preferred propellants, in accordance with the present disclosure.

[0046] The following recommendations are provided for fuels that are suitable for use in conjunction with the presently disclosed systems and methods. A preferred propellant, exploiting fast combustion, is aluminum, having a low cost, a high energy density, and a low melting point. Some applications may involve rockets being accelerated many times. An example may be a rocket for a safety system which is carried on many missions, but never actually activated. Some applications may require a rocket to accelerate a payload through a large change in velocity. An example of such an application is an interstellar mission. For these applications, involving many acceleration sequences or a high change in velocity, the cost of boron or beryllium is justifiable by their respective high energy density. For use with any of the reducers, excluding lithium recommended herein, or any combination thereof, liquid oxygen is a preferred oxidizer, especially for large rockets in civilian applications which may not require a long storage period prior to being launched on short notice. For rockets which are small, for example less than 10 metric tons, or that need to be stored long term and then launched on short notice, e.g., for military applications, other oxidizers may be used which are easier to store, such as mixed oxides of nitrogen (MON), dinitrogen tetroxide, nitric acid, or hydrogen peroxide. Other fuels which may be suited to such engines comprise at least one of magnesium and silicon. These may be useful in applications that require producing fuel locally on planets other than Earth, or moons, asteroids, comets, or dwarf planets where there is little infrastructure. These elements are likely to appear as impurities in refined aluminum. To reduce cost, instead of spending energy on removing enough silicon, magnesium, and calcium to achieve high purity, simply using the silicon and magnesium deliberately as fuel as well, possibly in a mixture with the aluminum, may be less expensive. For some embodiments, magnesium is preferred over aluminum for at least that magnesium is capable of a fast burn.

[0047] Referring to FIG. 1, this schematic diagram illustrates, in a sectional side view, an enhanced combustion rocket engine system 100, in accordance with an embodiment of the present disclosure. Most, or all, of the propellant 111 that is stored in propellant tank 110 is to be supplied to at least one engine 140, e.g., a rocket engine, and is also to be preheated prior to launch. By melting most or all of this fuel, e.g., the propellant 111, in propellant tank 110 before a launch, the engine 140 receives fuel in a liquid form, thereby eliminating a need for any additional heating features that would otherwise be required by the engine 140 for rapidly melting the fuel in flight, thereby eliminating a need for any additional mechanisms for feeding solid fuel, e.g., pellets or powders, and thereby eliminating any risk of blockages in flow or inconsistencies in flow rate that would otherwise occur.

[0048] Still referring to FIG. 1, heat also slightly increases the gravimetric energy density of the fuel. Fuel, e.g., the propellant 111, in a liquid state, also comprises a higher volumetric energy density than do either powder or pellets, whereby the higher volumetric energy density facilitates minimizing the mass of propellant tanks, such as the propellant tank 110, despite the higher temperature, and facilitates at least the other above listed benefits. Another benefit applies to a scenario involving launching a rocket vehicle from planets, moons, or other locations having an ambient temperature that is above the melting point of the fuel, e.g., the propellant 111. The preheating technique of the present disclosure eliminates the need to cool the fuel, e.g., the propellant 111. As the propellant tanks, e.g., the propellant 111, must withstand high temperature, the present systems and methods involve considering the strength- to-mass ratio and the materials therefor. However, such considerations are at least partly mitigated by the high volumetric energy density of HEDP fuel, e.g., the HEDP reducer component, facilitating minimizing the mass of the tanks, e.g., the propellant 111, and by increasing the gravimetric energy density of the fuel, e.g., the propellant 111, via melting the fuel.

[0049] Still referring to FIG. 1, the presently disclosed technique is especially recommended for fuels having a low melting point, such as aluminum, magnesium, and lithium; and recommended or preferred preheating temperatures comprise approximately 700 degrees Celsius for aluminum, approximately 700 degrees Celsius magnesium, and approximately 200 degrees Celsius for lithium. Depending on the conditions, these temperatures and ranges should be sufficiently high for preventing the fuel e.g., the propellant 111, from accidentally dropping below the melting point on its way to the engine, e.g., the engine 140. Solidifying the propellant 111 en route through the fuel lines may lead to a blockage. Temperatures higher than the foregoing temperatures facilitate faster combustion and improve gravimetric energy density of the propellant 111 and are also encompassed by the present disclosure. While the higher temperatures may weaken a related art fuel tank or require a heavier or more expensive related art fuel tank, the system 100 comprises as light a tank as safely possible which is thermally, chemically, and structurally compatible with the recommended fuels for the propellant 111, e.g., that can handle such elevated temperatures. For example, the propellant 111 comprises at least one of a fuel having a high melting point, such as boron in a suspension of a powdered form, and a fuel having a lower melting point, such as aluminum a liquid form.

[0050] Still referring to FIG. 1, the enhanced rocket engine system 100 comprises a rocket engine 140, the rocket engine 140 comprising a combustion chamber 141, a throat 142, and an exhaust nozzle 143. The system 100 further comprises: a propellant tank 110, such as a reducer propellant tank, for accommodating a propellant 111, such as a reducer component of a fuel combination, e.g., a hot HEDP reducer component. The propellant tank 110 comprises a pressurizing feature (not shown) for accommodating a pressurization gas 112, the pressurizing gas 112 for pressurizing the propellant 111; and a preheating feature, such as a heater 117, for preheating the propellant 111. Thus, the propellant tank 110 is not only adapted to accommodate the propellant 111, but also to preheat the propellant 111 to elevated temperatures by way of the preheating feature and to accommodate the preheated propellant 111 at elevated temperatures. The propellant tank 110 further comprises at least one optional insulation layer, such as at least one external insulation layer 116, for maintaining the preheated propellant 111 at elevated temperatures, wherein the propellant tank 110 comprises a continuous structure, whereby the propellant tank 110 avoids continuous heating of the preheated propellant 111. In an embodiment, the reducer tank may contain molten reducer.

[0051] Still referring to FIG. 1, for further facilitating preheating and maintaining the propellant 111 at elevated temperatures, the preheating feature, e.g., the heater 117, is optionally disposed within the propellant tank 110. Alternatively, the preheating feature comprises an inductive heating feature, e.g., an inductive heating coil, disposable in at least one of externally disposed in relation to the tank 110 and internally disposed in relation to the tank 110. The at least one optional insulation layer of the propellant tank 110 alternatively comprises at least one interior insulation layer 115 for facilitating at least one of actively cooling and passively cooling the propellant tank 110 while also maintaining the propellant 111 at elevated temperatures. Further the at least one interior insulation layer 115 optionally comprises an adequately strong material, wherein the propellant tank 110 optionally comprises a grid structure instead of a continuous structure for facilitating overall weight savings of the system 100.

[0052] Still referring to FIG. 1, alternatively, the optional exterior insulation layer 116, if structurally sufficient, is disposed on or over the propellant tank 110, wherein the propellant tank 110 optionally comprises a grid structure instead of a continuous structure for facilitating overall weight savings of the system 100. This alternative disposition of the optional exterior insulation layer 116 facilitates removal of the exterior insulation layer 116 from the propellant tank 110 shortly before launch for further facilitating overall weight savings of the system 100. The heater 117 may be useful even if the optional interior insulation layer 115 is not included. The exterior insulation layer 116 may be helpful if the heater 117, or a different heater, is used to preheat the tank 110 so as to avoid problems that may be associated with rapid thermal expansion caused by pouring hot fuel into the tank, if the fuel is melted in a different crucible.

[0053] Still referring to FIG. 1, the system 100 further comprises at least one fuel line 130 and at least one valve 170. The engine 140 is adapted to receive the propellant 111 from the propellant tank 110 through the fuel line 130 when the valve 170 is open. The at least one fuel line 130 comprises at least one of: at least one heating feature 137 for heating the propellant 111, being transferred from the propellant tank 110 to the at least one engine 140, to elevated temperatures; and at least one optional insulation layer, such as at least one insulation layer 136, for maintaining the propellant 111, being transferred from the propellant tank 110 to the at least one engine 140, at elevated temperatures, whereby solidification of, and blockage by, the propellant 111 are prevented in the at least one fuel line 130. The system 100 further comprises an additional propellant tank 120, such as an oxidizer propellant tank, for accommodating an additional propellant 121, such as an oxidizer component of the total fuel. The additional propellant tank 120 is adapted to self-pressurize if the additional propellant 121 has an adequate vapor pressure, such as the additional propellant 121 comprising LOX; and, if not, the additional propellant tank 120 is also adapted to pressurize the additional propellant 121 by way of a pressurization gas 122. The additional propellant 121 further comprises an insulation layer 125 disposed in at least one of an exterior disposition and an interior disposition.

[0054] Still referring to FIG. 1, in the system 100, comprising a rocket engine 140, such as a liquid-fueled rocket engine, the exhaust nozzle 143 is adapted for cooling by the propellant 111, comprising a reducer component of the total fuel, in accordance with an alternative embodiment of the present disclosure. However, if the propellant 111 is hot, as in some of the presently disclosed embodiments, the exhaust nozzle 143 is alternatively adapted for cooling by the additional propellant 121, comprising an oxidizer component of the total fuel, in accordance with yet another alternative embodiment of the present disclosure. Due to the corrosive nature of the additional propellant 121, comprising an oxidizer component, the hottest portions of the exhaust nozzle 143 are adapted for indirect cooling by the additional propellant 121. To address many of the related art cooling problems, the system 100 further comprises a coolant loop 150 for facilitating indirect cooling of the exhaust nozzle 143 by accommodating and transferring a supplemental coolant 151, e.g., a secondary coolant. The supplemental coolant 151, such as an additional less-corrosive coolant, is disposed in the coolant loop 150, wherein the coolant loop 150 is adapted to absorb heat, and to transfer such heat to the supplemental coolant 151, from at least one component of the engine 140, such as a combustion chamber 141, a throat 142, and the exhaust nozzle 143 (some of the hottest components). The system 100 further comprises a heat exchanger 152 adapted to transfer heat from the supplemental coolant 151, disposed in the coolant loop 150, to the additional propellant 121, disposed in the additional propellant tank 120. The system 100 further comprises a pump 153 for pumping the supplemental coolant 151 through the coolant loop 150. Portions of the rocket engine 140, especially the relatively cooler portions, may also be cooled directly by the additional propellant 121 by way of the heat exchanger 154, thereby improving the cooling efficiency. The coolant loop 150 is optional; and some embodiments of the system 100 are adapted to directly cool the engine 140 by transferring heat to additional propellant 121, comprising an oxidizer component.

[0055] Still referring to FIG. 1, another embodiment of the system 100 involves preheating most of the fuel of at least one of the types, such as the propellant 111 and the additional propellant 121, to a temperature that is in a range below the melting point to be fed to a rocket engine 140 before activating the rocket engine 140. As recommended for some space missions, the system 100 comprises at least one of: a heating feature for melting or further melting the propellant 111 in flight or in transit before injecting the propellant 111 into the rocket engine 140, wherein the preheating feature, e.g., the heater 117, for preheating the propellant 111 while disposed in the propellant tank 110 and prior to activating the engine, facilitates faster melting, whereby the "in-flight" or "in-transit" heating feature experiences less demand; and whereby the in-flight" or "in-transit" heating feature comprises a lighter or less expensive material. This embodiment avoids using a more expensive or heavier tank otherwise withstanding the higher temperature of molten fuel; and this embodiment is recommended for fuels having higher melting points, such as beryllium, boron, and silicon. In an embodiment, heater 117 may be disposed within a tank. Also in an embodiment, the propellant 111, 121 may be heated to a temperature that is greater than what the main propellant tank 110 may withstand in order to further preheat the propellant 111, 121 within the tank 110 before injecting it into the engine 140. For example, aluminum propellant in the main tank may be stored at a temperature that is greater than the melting point of aluminum. [0056] Still referring to FIG. 1, the system 100 comprises a engine 140 adapted to burn its own propellant tanks, such as the propellant tank 110 and the additional propellant tank 120, in accordance with another alternative embodiment of the present disclosure. In such embodiments, a burnable propellant tank, such as the propellant tank 110 and the additional propellant tank 120, may contain fuel which is at a temperature as high as possible, without excessively weakening the tank. For example, if the primary fuel, such as the propellant 111, comprises aluminum, the propellant tank, such as the propellant tank 110, may also comprise aluminum. In this situation, heating the majority of the fuel to approximately 350 degrees Celsius before flight is recommended.

[0057] Referring to FIG. 2A, this schematic diagram illustrates, in a sectional side view, an enhanced combustion rocket engine system 200, in accordance with an embodiment of the present disclosure. When using preheated contents, e.g., the propellant 211, 213 being preheated, in a burnable tank, e.g., the propellant tank 210, the propellant 211 comprises a powder or pellet form having a low thermal conductivity, the propellant 211 disposed in at least one shell 218, such as cylindrical shells. The hottest preheated fuel, e.g. the hottest preheated propellant 211, is loadable before flight into shells 218 disposed proximate the center of the propellant tank 210, while cooler fuel, e.g. the cooler propellant 213, is loaded proximate the walls of the propellant tank 210. The shells 218 are removable before launch, or may be left in place. If the shells 218 are retained during flight, the shells 218 may comprise a burnable material, such as aluminum, and have a rocket engine 240 burn them in flight. Deformation of the shells 218 due to heat does not introduce a significant concern as would premature deformation of the tank 210; and such deformation of the shells 218 is tolerable in this embodiment of the present disclosure.

[0058] Still referring to FIG. 2A, a presently disclosed technique for reducing the mass of tanks in the system 200 involves using fuel in the form of pellets instead of powder. Pellets can be held by grid shaped walls which may be lighter or less expensive than contiguous wall. The size of the pellets may be adjusted to minimize the mass of the tank 210, which may favor larger pellets, while balancing this against ensuring that the pellets melt fast enough, which may favor smaller pellets. In-flight or in-transit melting of the fuel in the form of these pellets, instead of powder, before injecting the fuel into the combustion chamber 241, provides a fuel that more slowly burns for at least that the pellets have lower surface-area-to-volume ratio. The presently disclosed technique for using fuel in a pellet form is combinable with the presently disclosed technique for burning the tanks 210, 220. Alternatively, the grid could be made of material that could withstand higher temperature, thereby allowing the pellets to be preheated to a higher temperature, for example just below the melting point. Regardless of whether the tanks 210, 220, e.g., grid tanks, are adapted to be burned, the presently disclosed technique for using fuel in a pellet form is combinable with the presently disclosed technique for using the shells 218 to place the hottest fuel, e.g., the propellant 211, closer to the center of the tank 210, thereby preventing heat-weakening the tank 210. Regardless of whether the tanks 210, 220 comprise a grid or a continuous form, or whether or not the tanks 210, 220 are adapted to be burned, an optional insulation layer may be placed inside the tanks 210, 220 to allow raising the temperature of the fuel, e.g., the propellants 211, 213, 221 without excessively weakening the tanks 210, 220.

[0059] Still referring to FIG. 2A, the technique of atomizing a liquid, e.g., the propellant 211, into tiny droplets for facilitating injecting the liquid into the combustion chamber 241 is encompassed by the present disclosure. This atomizing technique allows stable combustion and the use of an inexpensive and simple engine 240 and may be more beneficial for some embodiments that the technique of injecting powdered forms of fuel. Atomizing is also an inexpensive way to maximize surface area of the aluminum. Melting fuel before injecting the fuel into the rocket engine 240 does not require melting the fuel before launching the rocket vehicle. In this situation, a heating feature 217 for melting the fuel in flight may be used to preheat the propellant 211. In an embodiment, heater 217 may be disposed within a tank. The present disclosure contemplates combining in-flight heating before injection into the combustion chamber 241 with preheating fuel, e.g., the propellants 211, 221, before flight. Preheating the fuel before use of the rocket engine system 200 is recommended for more greatly reducing the cost or mass of the system 200 than preheating the fuel during operation of the engine 240.

[0060] Referring to FIG. 2B, is a schematic diagram illustrates, in a cross-sectional view taken at Section A-A, of the enhanced combustion rocket engine system 200 of FIG. 2A, in accordance with an embodiment of the present disclosure. The rocket system 200 involves both using preheated powder or pellet forms and using a burner 280 through which most, or all, of the solid fuel is transferred for melting before entering the engine 240. The reducer fuel is contained within tank 210. In an embodiment, a preheating burner 280 may be used to further heat the propellant before injecting it into the engine. In an embodiment, preheating burner 280 may be disposed outside of a tank. In an embodiment, the reducer tank 210 may contain hot solid reducer or powered reducer of higher melting point suspended in a molten reducer of lower melting temperature. The fuel is divided by the vertical cylindrical partition shell 218 into the colder mass 213 outside the shell, and the hotter mass 211 inside the shell. The hotter mass is kept spaced from the tank with the shell and the colder fuel 213 between it and the tank so as to minimize reduction of the strength to mass ratio of the tank by exposure to high temperature. The cylindrical partition 218 is open at both the top and bottom so that it can be removed without excessively disturbing the fuel by lifting it vertically. In an embodiment, the tanks 110, 210 may comprise partitions 218 in order to separate warm fuel disposed, for example, near the outer portion of the tank, from hotter fuel disposed, for example, near the center of the tank.

[0061] Still referring to FIG. 2B and referring back to FIG. 2A, the system 200 further comprises a heating feature 280 for further heating the propellant 211, e.g., the reducer component of the total fuel, during flight before injecting the propellant 211 into engines 240 may be used regardless of whether the propellant 211 is preheated before flight, whether the propellant 211 is preheated, but still primarily solid, e.g., as shown in FIG. 2A, or whether the propellant 211 is preheated primarily to the liquid state before flight, e.g., as shown in relation to the propellant 111 in FIG. 1. For fuel which has a melting point below approximately 1500 degrees Celsius, such as beryllium, aluminum, silicon, and magnesium, the present disclosure encompasses preheating the fuel, e.g., the propellant 211, to a temperature in a range that is above the fuel's melting point before injecting the fuel into an engine, e.g., the engine 240. If the fuel, such as boron, has a melting point that may be challenging to attain in a preheating burner, the present disclosure encompasses a technique comprising preheating the fuel to a temperature in a range that is above approximately 1500 degrees Celsius before injecting it into an engine's main combustion chamber 241. Alternatively, a fuel with such high melting point may be suspended in powdered form in a molten form of another fuel such as aluminum. The enhanced rocket engine system 200 further comprises at least one optional insulation layer 215 disposed inside the tank 210, at least one optional insulation layer disposed 216 outside the tank 210, and an optional heater 217. Since elevated temperature may result in weakening of the tank 210, interior insulation may be helpful if using solid fuel as the propellant 211.

[0062] Still referring to FIG. 2B and referring back to FIG. 2A, fabrication of the system 200 involves consideration of some example temperatures that are compatible with engineering the components of the system 200. By example only, the propellants 211, 213 each primarily comprise at least one of the preferred elements: aluminum and boron. As noted above, since the physical separation of the fuel into propellants 211 and 213 facilitates protection of the tank 210 from being exposed to the hottest fuel, the tank 210 may then comprise aluminum, thereby allowing the tank 210 to also be burned (self-consumed) as fuel. The tank 210 comprises at least one of a metal, a metal alloy, non-metal, and a high temperature composite material. Aluminum melts at a temperature of around 660 degrees Celsius and weakens to the point of having inadequate strength to mass ratio at a range of approximately 300 degrees Celsius to approximately 350 degrees Celsius, depending on the amount of reinforcement that the tank 210 has from other material, such as carbon fiber or stainless steel. However, ideally, the majority of the fuel, e.g., the propellant 211, would be raised to a significantly higher temperature before launch. Therefore, the hotter fuel, e.g., the propellant 211, being the majority of the fuel, comprises a temperature in a range of approximately 450 degrees Celsius to approximately 600 degrees Celsius, while the colder or cooler fuel, e.g., the propellant 213, comprises a temperature in a range of approximately 250 degrees Celsius to approximately 350 degrees Celsius.

[0063] Referring to FIG. 3, this schematic diagram illustrates, in a sectional side view, an enhanced combustion rocket engine system 300 of a rocket R, in accordance with an embodiment of the present disclosure. In the system 300, at least one consumable system component is self-consumable; and the methods of the present disclosure contemplate a technique for easily converting the at least one consumable system component, in-flight, to a form which is well suited, or suitable, for burning in the engine 340, wherein the at least one consumable system component comprises a material, such as a metal, a solid metal, an alloy, and a high temperature metal composite, and wherein the at least one consumable system component comprises at least one of a fuel propellant tank, e.g., the propellant tanks 310, 320, a partition thereof, a fairing, and any other system structure, e.g., other than the engine 340. As recommended in relation to the propellant 311, the present disclosure also encompasses melting the at least one consumable system component into a usable form of supplemental fuel before injecting the usable form of supplemental fuel into the engine. In an embodiment, propellant tank 320 is sized to fit within propellant tank 310 holds and propellant tank 320 is designed to be submerged and melted into propellant tank 310. Also in an embodiment, tank 330 may supply oxygen to burn propellant tank 320.

[0064] Still referring to FIG. 3, the fuel tank, e.g., the propellant tank 310, may not have a convenient shape, as fuel that is already in the form of pellets or powder otherwise would. As such, the system 300 further comprises a second smaller propellant tank 330, having a short ring shape or cylindrical shape and having an outer diameter that is slightly larger than that of the at least one consumable system component, such as the fuel tank, e.g., the propellant tank 310, to be melted, wherein the second smaller propellant tank 330contains molten fuel in which to submerge the at least one consumable system component, e.g., the propellant tank 310 for facilitating melting. Submerging the propellant tank 310 into the molten fuel, such as molten aluminum, in the second smaller propellant tank 330 facilitates melting by convecting for at least that molten aluminum has high heat capacity and thermal conductivity, is subject to adsorption by the propellant tank 310, and will naturally conform to the shape of the propellant tank 310 for facilitating transferring heat to the propellant tank 310. Injecting some oxidizer into the tank 330 provides additional heat for further facilitating melting.

[0065] Still referring to FIG. 3, the foregoing tank melting technique may also be used with the propellant tank 320, containing an oxidizer component, in the system 300. Oxidizer tanks, e.g., the propellant tank 320, preferably comprise aluminum. In an embodiment, the system 300 comprises a plurality of oxidizer tanks, e.g., a plurality of propellant tanks 320; and the tank melting technique is implementable by the plurality of oxidizer tanks, wherein one oxidizer tank is initially burned as a consumable while at least one remaining oxidizer tank is subsequently burned as a consumable. Alternatively, at least one propellant tank comprises a plurality of segments, wherein at least one segment is removable as a consumable to be burned while at least one remaining segment continues to accommodate a fuel, e.g., a propellant. In an embodiment, the system 300 comprises two tanks, e.g., the tanks 320, 330. In a recommended implementation, the tanks 320, 330, each tank comprise a drastically different size in relation to the other tank, wherein one tank, e.g., tank 320, being much larger than the other tank, e.g., a second smaller propellant tank 330. The propellant 321 from the larger tank, e.g., the propellant tank 320, should be consumed first, while the propellant 331 from the smaller tank, e.g., the propellant tank 330, may be consumed while burning the larger tank, e.g., the propellant tank 320, e.g., to provide sufficient reactant for the supplemental fuel provided by the burning of the large tank.

[0066] Still referring to FIG. 3, the system 300 comprises a burnable propellant tank, such as the main tank, e.g., the propellant tank 320. After its propellant 321 is depleted, the propellant tank 320 could be prepared for burning by submerging the propellant tank 320 in the main reducer tank, e.g., the propellant tank 310, whereby the propellant tank 320 is capable of melting into a liquid form for facilitating delivery of supplemental fuel to the engines 340. This technique may require actuators (not shown) for effecting venting of a pressurization gas or a gaseous oxidizer 322, for removing an optional exterior insulation layer 325 from the outer surface of the propellant tank 320 before the propellant tank 320 is submerged in the propellant tank 310, for removing a heater 317, alternatively removable before launch, so that the heater 317 does not block the path of the propellant tank 320, for opening the top of reducer propellant tank 310, and for forcing the propellant tank 320 into the liquid 311 in the reducer propellant tank 310.

[0067] Still referring to FIG. 3, the system 300 comprises a second smaller propellant tank 330 for accommodating and providing oxidizer (sufficient reactant) for (reacting with) burning the supplemental fuel provided by melting the main tank 320. The second smaller propellant tank 330 comprises an insulation layer 335, accommodates an oxidizer 331, accommodates a pressurization gas 332, and comprises a valve 371. This tank melting technique is combinable with using a preheating burner 380 for facilitating further heating the propellant 311 before the propellant 311 is injected into the main engines 340. This entire set of tanks is used to deliver fuel through the valves 370 (associated with the tank 310), 371 (associated with the tank 330), and 372 (associated with the tank 320) to the engine 340.

[0068] Still referring to FIG. 3, noted is that the main tank, e.g., the propellant tank 320, is disposed above the reducer tank, e.g., the propellant tank 310. The diameter of the propellant tank 320 is slightly smaller than the inner diameter of the propellant tank 310; and, if present, the optional exterior insulation layer 325 of the propellant tank 320 comprises a thickness to allow the main propellant tank 320 with the optional exterior insulation layer 325 (if consumable) to fit inside the reducer propellant tank 310 with an optional interior insulation layer 315. In this manner, the main propellant tank 320 can be placed into the molten reducer 311 by lowering it down vertically once the actuators have completed the above mentioned preparations. The optional exterior insulation layer 325 may comprise a consumable material, that is the same as, or similar to, the materials described in relation to the enhanced heat shields of the present disclosure, for providing supplemental fuel as well. Vertical linear guides 350 cooperating with guide rail 351 facilitate lowering the tank 320 into the tank 310. The guide 350 is supported on a vertical structural member 353.

[0069] Still referring to FIG. 3, an implementer may prefer to burn the main reducer propellant tank 310 as well. As noted above, this situation may favor using a primarily solid fuel in the main reducer propellant tank 310. In this situation, the system 300 further comprises an additional reducer tank (not shown) capable of containing a molten fuel which is hotter than that which the main reducer propellant tank 310 can withstand. This small reducer tank (not shown) facilitates melting both the main propellant tank 320 and the main reducer propellant tank 310. The technique for burning tanks, fairings, and structural components is recommended for rockets intended for low thrust- to-mass ratio and for a long burn time, especially rockets intended to operate in vacuum or near-vacuum conditions, such as exo-atmospheric space. This technique for burning consumable system components provides more time for the various actuators to perform their functions and for melting while also preventing undesirable aerodynamic effects or high thrust-to-mass ratio otherwise causing flight instability or placing undue stress on structural components. This technique for burning consumable system components may allow the use of lighter, less expensive structures, guide rails, actuators, and other components. Examples of appropriate applications include upper stages. For mitigating flight instability, the present disclosure also encompasses a technique comprising adapting the at least one engine for providing directional thrust for spinning the rocket vehicle at a low rotational speed, whereby the propellant experiences a centripetal acceleration which urges at least one molten portion of the propellant in an outboard direction for effecting better contact of the at least one molten portion of the propellant with the solid components of the tank, and whereby an additional amount, e.g., a small amount, of forward thrust is generated.

[0070] Still referring to FIG. 3, a rocket R accommodates a payload 376 disposed proximate a top portion of the rocket R, wherein the rocket R comprises the enhanced rocket engine system 300 and a payload fairing 375. The payload 376 is covered by, or disposed within, the payload fairing 375. The payload fairing 375 may be lowered into the reducer propellant tank 310 for facilitating melting the payload fairing 375, wherein supplemental fuel from melting the payload fairing 375 is subsequently burned in a manner similar to the approach used to melt the main propellant tank 320. A variety of other consumable components of the system 300 or the rocket R could be preheated or even melted for use as supplemental fuel in addition to reducer tanks, oxidizer tanks, and payload fairings. Other well suited examples (not shown) of other consumable components include, but are not limited to, guidance or stabilizer fins, inter-stage structures, other components, or even the low pressure end of a rocket nozzle if comprising an appropriate consumable material.

[0071] Still referring to FIG. 3, techniques for melting consumable structural or aerodynamic components for use as supplemental fuel, other than by submerging consumable components into a liquid fuel for facilitating melting, may also be used and are encompassed by the present disclosure. If such solid fuel is aggressively preheated before injecting the fuel into the engines 340, the following technique are recommended: preheating the fuel to a temperature range that is beyond the melting point to ensure fast stable combustion. Also, for facilitating fast melting, consumable components may comprise a relatively large number, a plurality of, small portions, wherein the plurality of small portions is at least one of removably interlocking and removably fastenable by tension features or any other fasteners or quick-disconnecting features that allow the plurality of small portions to be easily disassembled in flight. The consumable components may further comprise at least one of mechanically frangible portion and a thermally frangible portion (by melting joints) for facilitating disassembly.

[0072] Referring to FIG. 4, this schematic diagram illustrates, in a sectional side view, an enhanced combustion rocket engine system 400 of a rocket R, in accordance with an embodiment of the present disclosure. The following strategy may be used to slightly increase the energy density of oxidizers. Liquid oxygen is typically stored at its boiling point at low pressure. However, the LOX could, instead, be stored at high pressure, thereby reducing the amount of thermal energy that the LOX absorbs when expanded by boiling. Although cooling may become more complex in this situation, especially in a rocket which also uses preheated reducer fuel, high pressure storage may also increase the energy density of the oxidizer, e.g., LOX. Such benefit may be offset by the need for heavier tanks, e.g., a propellant tank 420, at the front end of a space mission in order to withstand the high pressure of the oxidizer, e.g., the propellant 421; however, in recompense, the heavier tanks, e.g., a propellant tank 420, are burnable as reducer fuel, whereby total load is significantly reduced later in the space mission, in accordance with an embodiment of the present disclosure.

[0073] Still referring to FIG. 4, the tanks themselves, e.g., a propellant tank 420, may experience some challenges in preheating much while storing oxidizer, thereby affecting the overall energy density of the reducer fuel. As such, a preferred oxidizer temperature for the preheating technique may comprise a temperature in a range that is well below ambient temperature in this embodiment. The minimum initial structural strength requirement the tanks, e.g., the propellant tank 420, may be high, providing excess unusable strength throughout the consumption of much of the latter half of the propellant. Therefore, to mitigate such circumstance, this embodiment of the present disclosure also encompasses reducing the mass of the tanks, e.g., the propellant tank 420, by beginning to consume the tanks as fuel long before the oxidizer component, e.g., the propellant 421, is expended, e.g., during consumption of the oxidizer component. One approach for effecting mitigation of excess unusable strength comprises using a plurality of tanks, such as oxidizer tanks, e.g., a plurality of propellant tanks 420, wherein the plurality of oxidizer tanks are serially disposable, thereby providing structural and aerodynamic benefits. However, this approach, alone, may not provide a preferred ratio of the initial tank mass to the oxidizer component mass, otherwise reducing the amount of reducer fuel which can be preheated, for some missions. Thus, a second approach for effecting mitigation of excess unusable strength comprises installing at least one structural reinforcement layer 481 around the tanks, e.g., the propellant tank 420, and removing and burning the at least one structural reinforcement layer 481 as the pressure of the oxidizer decreases in the tanks.

[0074] Still referring to FIG. 4, regardless of which of the two foregoing approaches is used, and regardless of whether both approaches are used in combination, the present disclosure encompasses preparing the consumable component, such as in relation to a low pressure tank, e.g., a tank that may not be necessary for the entire duration of a mission, by melting the consumable component before feeding the resulting supplemental fuel from the consumable component to any main combustion chamber 441. In an embodiment, propellant tank 420 is sized to fit within propellant tank 410. High pressure oxidizer tanks may comprise a spherical shape, presenting some aerodynamic challenges that are addressed by the present disclosure. Noted is that the strategies described in this paragraph are applied by storing reducers at high pressure instead of, or in addition to, oxidizers. Also, using smaller high pressure tanks for applications where the rocket vehicle travels through an atmosphere and/or disposing a plurality of consumable components in a series relationship minimizes drag, in accordance with the present disclosure. Alternatively, in a vacuum or near- vacuum condition g., in space, using the reinforced tank design is appropriate and encompassed by the present disclosure.

[0075] Still referring to FIG. 4, the system 400 of the rocket R comprises a burnable high pressure propellant tank, such as a main tank, e.g., the propellant tank 420, which is burnable as a consumable component. After the propellant 421 is depleted, the propellant tank 420 is prepared for burning by submerging the propellant tank 420 in the main reducer tank, e.g., the propellant tank 410, thereby melting the propellant tank 420 into a liquid form as supplemental fuel for facilitating delivery to the engines 440. Before the propellant 421 is depleted and the propellant tank 420 can be burned, consuming some of the propellant 421 may reduce the pressure in the propellant tank 420 to a point where at least one structural reinforcement layer 481 can be removed and melted in the propellant tank 410 in a manner similar to that in which the propellant tank 420 is later to be melted when depleted. This technique may require removing at least one insulation later 425 from around the at least one structural reinforcement layer 481 before removing and melting the at least one structural reinforcement layer 481. [0076] Still referring to FIG. 4, after the at least one structural reinforcement layer 481 is removed, the at least one insulation layer 425 remains disposed around, contract-able to, or ejectable from (discarded), the propellant tank 420. The procedure for melting the tank, e.g., propellant tank 420, is similar to that described in relation to the system 300, as shown in FIG. 3. The techniques of this embodiment may require actuators to vent pressurization gas or a gaseous oxidizer 422, to remove the at least one insulation layer 425 from the outer surface of the propellant tank 420 before the propellant tank 420 is submerged, to remove a heater 417, removable before launch, thereby preventing blocking of a path of the propellant tank 420 to the propellant tank 410, to open the top of reducer tank, e.g., the propellant tank 410, and to force tank, e.g., the propellant tank 420, into the liquid reducer, e.g., the propellant 411, in reducer tank, e.g., the propellant tank 410.

[0077] Still referring to FIG. 4, the system 400 further comprises a second smaller tank, e.g., the propellant tank 430, that provides oxidizer for burning the supplemental fuel from melted the main tank, e.g., the propellant tank 420, and any reducer component, e.g., the propellant 411, which was left in the reducer tank, e.g., the propellant tank 410, to assist with melting the propellant tank 420. This smaller tank, e.g., the propellant tank 430, comprises at least one insulation layer 435, accommodates an oxidizer component 431, accommodates a pressurization gas 432, and comprises a valve 471. This strategy of melting tanks is combinable with a preheating burner, such as the preheating burner480, to further heat the fuel, e.g., the propellant 411, before the fuel is injected into the main engines 440. This entire set of tanks, e.g., tanks 410, 420, 430, is used to deliver fuel through their associated valves 470, 471, and 472 to the engine 440. The at least one structural reinforcement layer 481 optionally comprises a plurality of structural reinforcement layers 481 if needed.

[0078] Still referring to FIG. 4, noted is that the main tank, e.g., the propellant tank 420, and the at least one structural reinforcement layer 481 are positioned above the reducer tank, e.g., the propellant tank 410. The diameter of the propellant tank 420, including the thickness of the at least one structural reinforcement layer 481, is in a range that is slightly smaller than the inner diameter of the propellant tank 410; and, if present, an optional interior insulation layer 415 disposed inside the propellant tank 410 facilitates fitting the propellant tank 420 and at the least one structural reinforcement layer 481 inside the propellant tank 410. This way, the tank, the propellant tank 420, is disposable in the molten reducer, e.g., the propellant 411, by simply lowering the propellant tank 420 down vertically once the actuators have completed the above mentioned preparations. Vertical linear guides 450, cooperating with guide rail 451, facilitate lowering the tank 420 into the tank 410. The guide 450 is supported on a vertical structural member 453.

[0079] Still referring to FIG. 4, an implementer may prefer to burn the main reducer propellant tank 410 as well. As noted above, this situation may favor using a primarily solid fuel in the main reducer propellant tank 410. In this situation, the system 400 further comprises an additional reducer tank (not shown) capable of containing a molten fuel which is hotter than that which the main reducer propellant tank 410 can withstand. This small reducer tank (not shown) facilitates melting both the main propellant tank 420 and the main reducer propellant tank 410. The technique for burning tanks, fairings, and structural components is recommended for rockets intended for low thrust- to-mass ratio and for a long burn time, especially rockets intended to operate in vacuum or near-vacuum conditions, such as exo-atmospheric space. This technique for burning consumable system components provides more time for the various actuators to perform their functions and for melting while also preventing undesirable aerodynamic effects or high thrust-to-mass ratio otherwise causing flight instability or placing undue stress on structural components. This technique for burning consumable system components may allow the use of lighter, less expensive structures, guide rails, actuators, and other components. Examples of appropriate applications include upper stages.

[0080] Still referring to FIG. 4, a rocket R accommodates a payload 476 disposed proximate a top portion of the rocket R, wherein the rocket R comprises the enhanced rocket engine system 400 and a payload fairing 475. The payload 476 is covered by, or disposed within, the payload fairing 475. The payload fairing 475 may be lowered into the reducer propellant tank 420 for facilitating melting the payload fairing 475, wherein supplemental fuel from melting the payload fairing 475 is subsequently burned in a manner similar to the approach used to melt the main propellant tank 420. A variety of other consumable components of the system 400 or the rocket R could be preheated or even melted for use as supplemental fuel in addition to reducer tanks, oxidizer tanks, and payload fairings. Other well suited examples (not shown) of other consumable components include, but are not limited to, guidance or stabilizer fins, inter-stage structures, other components, or even the low pressure end of a rocket nozzle if comprising an appropriate consumable material. Techniques for melting consumable structural or aerodynamic components for use as supplemental fuel, other than by submerging consumable components into a liquid fuel for facilitating melting, may also be used and are encompassed by the present disclosure. If such solid fuel is aggressively preheated before injecting the fuel into the engines 440, the following technique are recommended: preheating the fuel to a temperature range that is beyond the melting point to ensure fast stable combustion. Also, for facilitating fast melting, consumable components may comprise a relatively large number, e.g., a plurality of, small portions, wherein the plurality of small portions is at least one of removably interlocking and removably fastenable by tension features or any other fasteners or quick-disconnecting features that allow the plurality of small portions to be easily disassembled in flight. The consumable components may further comprise at least one of mechanically frangible portion and a thermally frangible portion (by melting joints) for facilitating disassembly.

[0081] Referring to FIG. 5, this flowchart illustrates a method Ml of fabricating an enhanced combustion rocket engine system, in accordance with an embodiment of the present disclosure. The method Ml comprises: providing at least one primary propellant tank adapted to accommodate a primary propellant in at least one of a solid state, a transition melting state, and a molten state and to deliver the primary propellant in the molten state, as indicated by block 501; providing at least one secondary propellant tank adapted to accommodate a secondary propellant in a liquid state, as indicated by block 502; and providing at least one rocket engine, the at least one rocket engine adapted to: receive the primary propellant from the at least one primary propellant tank in a temperature range of at least approximately above an ambient temperature; receive the secondary propellant from the at least one secondary propellant tank in a temperature range of at least approximately below an ambient temperature; and exothermically react the primary and secondary propellants for effecting combustion to provide thrust, as indicated by block 503.

[0082] Referring to FIG. 6, this flowchart illustrates a method M2 of using an enhanced combustion rocket engine system, in accordance with an embodiment of the present disclosure. The method M2 comprises: providing an enhanced combustion rocket engine system, as indicated by block 600, the rocket engine system providing comprising: providing at least one primary propellant tank adapted to accommodate a primary propellant in at least one of a solid state, a transition melting state, and a molten state and to deliver the primary propellant in the molten state, as indicated by block 601; providing at least one secondary propellant tank adapted to accommodate a secondary propellant in a liquid state, as indicated by block 602; and providing at least one rocket engine, the at least one rocket engine adapted to: receive the primary propellant from the at least one primary propellant tank in a temperature range of at least approximately above an ambient temperature; receive the secondary propellant from the at least one secondary propellant tank in a temperature range of at least approximately below an ambient temperature; and exothermically react the primary and secondary propellants for effecting combustion to provide thrust, as indicated by block 603; charging the at least one primary propellant tank with the primary propellant, as indicated by block 604; charging the at least one secondary propellant tank with the secondary propellant, as indicated by block 605; preheating, heating, and delivering the primary propellant from the at least one primary propellant tank to the at least one rocket engine in a temperature range of at least approximately above an ambient temperature, as indicated by block 606; delivering the secondary propellant from the at least one secondary propellant tank to the at least one rocket engine in a temperature range of at least approximately below an ambient temperature, as indicated by block 607; and exothermically reacting the primary and secondary propellants for effecting combustion in the at least one rocket engine to provide thrust, as indicated by block 608.

[0083] Referring back to FIGS. 1-4, the rocket engines, e.g., the engines 140, 240, 340, 440, typically must run near their maximum rated throttle setting in order to achieve a near-peak efficiency. The present disclosure encompasses controlling acceleration, such as preventing excessive acceleration, of the rocket R for at least the following concerns: preventing damage to payloads, preventing injury to passengers, and prepare for docking. Thus, if the rocket engines, e.g., the engines 140, 240, 340, 440, deliver sufficient thrust early in a flight, e.g., when the mass of the rocket R is highest, later in a flight, the rocket engines, e.g., the engines 140, 240, 340, 440, may deliver excessive thrust, higher than desirable acceleration rates, or experience a need to run the rocket engines at a lower throttle setting, wherein the engines are less efficient. The foregoing concern are addressed by the presently disclosed systems and methods involving self-consumable system components, e.g., system being capable of burning their own fuel tanks, payload fairings, and other structural components. Referring also ahead to FIGS. 7-15, the present disclosure alternatively addresses the foregoing concerns in an enhanced combustion engine system, such as the systems 100, 200, 300, 400, for use with a rocket vehicle, such as a multistage rocket vehicle 700, 800, 900, 1100, 1200, 1300, 1400, 1500, comprising a plurality of engines.

[0084] Referring to FIG. 7, this schematic diagram illustrates, in a semi-exploded sectional side view, a rocket vehicle (launch vehicle), such as a multistage rocket vehicle 700, powerable by an enhanced combustion rocket engine system, such as the systems 100, 200, 300, 400, in accordance with an embodiment of the present disclosure. The multistage rocket vehicle 700 comprises a plurality of stages, e.g., the plurality of stages 710, 720, 760, 770 comprising at least one upper stage 760, 770 and at least one lower stage 710, 720. Each stage of the plurality of stages 710, 720, 760, 770 is capable of utilizing at least one enhanced combustion engine system, such as the systems 100, 200, 300, 400, wherein each stage is capable of accommodating its own engine(s) and its own propellants. The at least one enhanced combustion engine system comprises at least one of: at least one main engine 740 associated with a nozzle 740a and at least one auxiliary engine 790, 791 associated with a nozzle extension 795, the at least one auxiliary engine 790, 791 being smaller than the at least one main engine 740. By example only in this embodiment, the multistage rocket vehicle 700, comprising the plurality of stages 710, 720, 760, 770, each stage capable of accommodating at least one engine, is initially powerable by the at least one main engine 740, e.g., on take-off and early segments of the space mission. Later in the flight (in transit during the space mission), the at least one main engine 740, is capable of shut down; and the at least one lower stage 710, 720 as well as the at least one main engine 740 are capable of ejection (being jettisoned) from the at least one upper stage 760, 770, as well as any associated structure, such as a fairing, e.g., the payload fairing 775, of the multistage rocket vehicle 700 to reduce mass, wherein the at least one upper stage 760, 770 and the payload fairing 775 are subsequently powerable by the at least one auxiliary engine 790, 791. [0085] Referring to FIG. 8, this schematic diagram illustrates, in a semi-exploded sectional side view, a rocket vehicle (launch vehicle), such as a multistage rocket vehicle 800, powerable by an enhanced combustion rocket engine system, such as the systems 100, 200, 300, 400, in accordance with an embodiment of the present disclosure. The multistage rocket vehicle 800 comprises a plurality of stages, e.g., the plurality of stages 810, 820, 860, 870 comprising at least one upper stage 860, 870 and at least one lower stage 810, 820. Each stage of the plurality of stages 810, 820, 860, 870 is capable of utilizing at least one enhanced combustion engine system, such as the systems 100, 200, 300, 400, wherein each stage is capable of accommodating its own engine(s) and its own propellants. The at least one enhanced combustion engine system comprises at least one of: at least one main engine 840 associated with a nozzle 840a and at least one auxiliary engine 890, 891 associated with a nozzle extension 895, the at least one auxiliary engine 890, 891 being smaller than the at least one main engine 840. By example only in this embodiment, the multistage rocket vehicle 800, comprising the plurality of stages 810, 820, 860, 870, each stage capable of accommodating at least one engine, is initially powerable by the at least one main engine 840, e.g., on take-off and early segments of the space mission. Later in the flight (in transit during the space mission), the at least one main engine 840, is capable of shut down; and the at least one lower stage 810, 820 as well as the at least one main engine 840 are capable of ejection (being jettisoned) from the at least one upper stage 860, 870, as well as any associated structure, such as a fairing, e.g., the payload fairing 875, of the multistage rocket vehicle 800 to reduce mass, wherein the at least one upper stage 860, 870 and the payload fairing 875 are subsequently powerable by the at least one auxiliary engine 890, 891. Further, a technique involving slightly shifting the position of an engine, e.g., the at least one auxiliary engine 890, 891, relative to a central axis of the multistage rocket vehicle 800, e.g., in an outboard direction, e.g., in a direction A, as shown by the arrows in FIG. 8, thereby ensuring symmetric thrust, is also encompassed by an embodiment of the present disclosure.

[0086] Referring to FIG. 9, this schematic diagram illustrates, in a semi-exploded sectional side view, a rocket vehicle (launch vehicle), such as a multistage rocket vehicle 900, powerable by an enhanced combustion rocket engine system, such as the systems 100, 200, 300, 400, in accordance with an embodiment of the present disclosure. The multistage rocket vehicle 900 comprises a plurality of stages, e.g., the plurality of stages 910, 920, 960, 970 comprising at least one upper stage 960, 970 and at least one lower stage 910, 920. Each stage of the plurality of stages 910, 920, 960, 970 is capable of utilizing at least one enhanced combustion engine system, such as the systems 100, 200, 300, 400, wherein each stage is capable of accommodating its own engine(s) and its own propellants. The at least one enhanced combustion engine system comprises at least one of: at least one main engine 940 associated with a nozzle 940a and at least one auxiliary engine 990 associated with a nozzle extension 995, the at least one auxiliary engine 990 being smaller than the at least one main engine 940. By example only in this embodiment, the multistage rocket vehicle 900, comprising the plurality of stages 910, 920, 960, 970, each stage capable of accommodating at least one engine, is initially powerable by the at least one main engine 940, e.g., on take-off and early segments of the space mission. Later in the flight (in transit during the space mission), the at least one main engine 940, is capable of shut down; and the at least one main engine 940 as well as its associated nozzle 940a is capable of ejection (being jettisoned), e.g., in a direction B, from the at least one lower stage 910, 920 of the multistage rocket vehicle 900 to reduce mass, wherein the plurality of stages 910, 920, 960, 970, and any associated structure, such as a fairing, e.g., a payload fairing 975, are subsequently powerable by the at least one auxiliary engine 990.

[0087] Referring to FIG. 10, this schematic diagram illustrates, in a semi -exploded sectional side view, a rocket vehicle (launch vehicle), such as a multistage rocket vehicle 1000, powerable by an enhanced combustion rocket engine system, such as the systems 100, 200, 300, 400, in accordance with an embodiment of the present disclosure. The multistage rocket vehicle 1000 comprises a plurality of stages, e.g., the plurality of stages 1010, 1020, 1060, 1070, 1072, comprising at least one upper stage 1060, 1070, 1072 and at least one lower stage 1010, 1020. Each stage of the plurality of stages 1010, 1020, 1060, 1070, 1072 is capable of utilizing at least one enhanced combustion engine system, such as the systems 100, 200, 300, 400, wherein each stage is capable of accommodating its own engine(s) and its own propellants. The at least one enhanced combustion engine system comprises at least one of: at least one main engine 1040 associated with a nozzle 1040a and at least one auxiliary engine 1090 associated with a nozzle extension 1095, the at least one auxiliary engine 1090 being smaller than the at least one main engine 1040. By example only in this embodiment, the multistage rocket vehicle 1000, comprising the plurality of stages 1010, 1020, 1060, 1070, 1072, each stage capable of accommodating at least one engine, is initially powerable by the at least one main engine 1040, e.g., on take-off and early segments of the space mission. Later in the flight (in transit during the space mission), the at least one main engine 1040, is capable of shut down; and the at least one main engine 1040 as well as its associated nozzle 1040a are optionally capable of ejection (being jettisoned) from the at least one lower stage 1010, 1020 of the multistage rocket vehicle 1000 to reduce mass, wherein the plurality of stages 1010, 1020, 1060, 1070, 1072, and any associated structure, such as a fairing, e.g., a payload fairing 1075, are subsequently powerable by the at least one auxiliary engine 1090.

[0088] Still referring to FIG. 10, considering that adding more engines may increase cost for manufacturing a rocket vehicle, e.g., the rocket vehicle 1000, in an alternative embodiment, any given stage, e.g., the lower stages 1010, 1020, instead of using its own engine(s), is adapted to use a smaller engine, e.g., the at least one auxiliary engine 1090, associated with a subsequent stage, e.g., the upper stages 1060, 1070, 1072, wherein the smaller engine, e.g., the at least one auxiliary engine 1090, from the subsequent stage, e.g., the upper stages 1060, 1070, 1072, is approximately sized for performance as well as manufacturing cost savings. In this embodiment, since the smaller engine, e.g., the at least one auxiliary engine 1090, from the subsequent stage, e.g., the upper stages 1060, 1070, 1072, is smaller in relation to that, e.g., the at least one main engine 1040, of the previous stage, e.g., the lower stages 1010, 1020, to maintain requisite performance when the at least one auxiliary engine 1090 from the subsequent stage, e.g., the upper stages 1060, 1070, 1072, is also used for the previous stage, e.g., the lower stages 1010, 1020, the smaller engine, e.g., the at least one auxiliary engine 1090, from the subsequent stage, e.g., the upper stages 1060, 1070, 1072, is configurable by slightly increasing its size over that of a conventional related art subsequent stage engine. The last portion of fuel from the previous stage engine, e.g., the at least one main engine 1040, is transferrable to the smaller subsequent stage engine, e.g., the at least one auxiliary engine 1090. Specifically, this last portion of fuel comprises consumable components, e.g., the fuel tanks and other structural components, from the previous stage.

[0089] Referring to FIG. 11, this schematic diagram illustrates, in a semi -exploded sectional side view, a rocket vehicle (launch vehicle), such as a multistage rocket vehicle 1100, powerable by an enhanced combustion rocket engine system, such as the systems 100, 200, 300, 400, in accordance with an embodiment of the present disclosure. The multistage rocket vehicle 1100 comprises a plurality of stages, e.g., the plurality of stages 1110, 1120, 1160, 1170, comprising at least one upper stage 1160, 1170 and at least one lower stage 1110, 1120. Each stage of the plurality of stages 1110, 1120, 1160, 1170 is capable of utilizing at least one enhanced combustion engine system, such as the systems 100, 200, 300, 400, wherein each stage is capable of accommodating its own engine(s) and its own propellants. The at least one enhanced combustion engine system comprises at least one of: at least one main engine 1140 associated with a nozzle 1140a and at least one auxiliary engine 1190 associated with a nozzle extension 1195, the at least one auxiliary engine 1190 being smaller than the at least one main engine 1140. By example only in this embodiment, the multistage rocket vehicle 1100, comprising the plurality of stages 1110, 1120, 1160, 1170, each stage capable of accommodating at least one engine, is initially powerable by the at least one main engine 1140, e.g., on take-off and early segments of the space mission. Later in the flight (in transit during the space mission), the at least one main engine 1140, is capable of shut down; and the at least one lower stage 1110, 1120 as well as the at least one main engine 1040, its associated nozzle 1140a, and the nozzle extension 1195 (now dissociated from the at least one auxiliary engine 1190) are capable of ejection (being jettisoned), e.g., in a direction C, from the at least one upper stage 1160, 1170 of the multistage rocket vehicle 1100 to reduce mass, wherein the at least one upper stage 1160, 1170, and any associated structure, such as a fairing, e.g., a payload fairing 1175, are subsequently powerable by the at least one auxiliary engine 1190. Alternatively, a small subsequent stage engine, e.g., the at least one auxiliary engine 1190, is movable, e.g., in a direction C, down to the bottom of the rocket vehicle 1100 for propelling a "stack," comprising the plurality of stages 1110, 1120, 1160, 1170, from a pusher configuration, thereby providing symmetric thrust or an even thrust distribution from a single engine, such as the newly positioned small subsequent stage engine, e.g., the at least one auxiliary engine 1190. After propelling the stack from the pusher configuration, the small subsequent stage engine, e.g., the at least one auxiliary engine 1190, is movable back up to its original or normal location, e.g., in a direction D. Another alternative embodiment that is encompassed by the present disclosure involves the subsequent stage engines being movable in relation to the subsequent stage, wherein the subsequent stage engines are laterally disposable, e.g., at the sides, in relation to the center axis for providing thrust in a nearly "tractor" configuration, and wherein the subsequent stage engines are movable for a short distance L_ss in relation to the entire rocket length Lrv.

[0090] Referring to FIG. 12, this schematic diagram illustrates, in a semi -exploded sectional side view, a rocket vehicle (launch vehicle), such as a multistage rocket vehicle 1200, powerable by an enhanced combustion rocket engine system, such as the systems 100, 200, 300, 400, in accordance with an embodiment of the present disclosure. The multistage rocket vehicle 1200 comprises a plurality of stages, e.g., the plurality of stages 1210, 1220, 1260, 1270 comprising at least one upper stage 1260, 1270 and at least one lower stage 1210, 1220. Each stage of the plurality of stages 1210, 1220, 1260, 1270 is capable of utilizing at least one enhanced combustion engine system, such as the systems 100, 200, 300, 400, wherein each stage is capable of accommodating its own engine(s) and its own propellants. The at least one enhanced combustion engine system comprises at least one of: at least one main engine 1240 associated with a nozzle 1240a and at least one auxiliary engine 1290 associated with a nozzle extension 1295, the at least one auxiliary engine 1290 being smaller than the at least one main engine 1240. By example only in this embodiment, the multistage rocket vehicle 1200, comprising the plurality of stages 1210, 1220, 1260, 1270, each stage capable of accommodating at least one engine, is initially powerable by the at least one main engine 1240, e.g., on take-off and early segments of the space mission. Later in the flight (in transit during the space mission), the at least one main engine 1240, is capable of shut down; and the at least one main engine 1240 as well as its associated nozzle 1240a is capable of ejection (being jettisoned), e.g., in a direction E, from the at least one lower stage 1210, 1220 of the multistage rocket vehicle 1200 to reduce mass, wherein the plurality of stages 1210, 1220, 1260, 1270, and any associated structure, such as a fairing, e.g., a payload fairing 1275, are subsequently powerable by at least one of: at least one remaining engine 1240 and the at least one auxiliary engine 1290. An alternative to shifting the position of an engine, such as shown in FIG. 8, is another technique encompassed by the present disclosure that involves using at least three main engines, e.g., three main engines 1240, disposed in a symmetric arrangement, for providing an even thrust distribution, notwithstanding the possibility that at least one of the three main engines might become shut down and optionally jettisoned, such as for adjusting the direction of the thrust vector, e.g., for adjusting at least one of pitch and yaw of the vehicle 1200.

[0091] Referring to FIG. 13, this schematic diagram illustrates, in a semi -exploded sectional side view, a rocket vehicle (launch vehicle), such as a multistage rocket vehicle 1300, powerable by an enhanced combustion rocket engine system, such as the systems 100, 200, 300, 400, in accordance with an embodiment of the present disclosure. The multistage rocket vehicle 1300 comprises a plurality of stages, e.g., the plurality of stages 1310, 1320, 1360, 1370 comprising at least one upper stage 1360, 1370 and at least one lower stage 1310, 1320. Each stage of the plurality of stages 1310, 1320, 1360, 1370 is capable of utilizing at least one enhanced combustion engine system, such as the systems 100, 200, 300, 400, wherein each stage is capable of accommodating its own engine(s) and its own propellants. The at least one enhanced combustion engine system comprises at least one of: at least one main engine 1340 associated with a nozzle 1340a and at least one auxiliary engine 1390 associated with a nozzle extension 1395, the at least one auxiliary engine 1390 being smaller than the at least one main engine 1340. By example only in this embodiment, the multistage rocket vehicle 1300, comprising the plurality of stages 1310, 1320, 1360, 1370, each stage capable of accommodating at least one engine, is initially powerable by the at least one main engine 1340, e.g., on take-off and early segments of the space mission. Later in the flight (in transit during the space mission), the at least one main engine 1340, is capable of shut down; and the at least one main engine 1340 as well as its associated nozzle 1340a is capable of ejection (being jettisoned), e.g., in a direction F, from the at least one lower stage 1310, 1320 of the multistage rocket vehicle 1300 to reduce mass, wherein the plurality of stages 1310, 1320, 1360, 1370, and any associated structure, such as a fairing, e.g., a payload fairing 1375, are subsequently powerable by at least one of: at least one remaining engine 1340 and the at least one auxiliary engine 1390. An alternative to shifting the position of an engine, such as shown in FIG. 8, is another technique encompassed by the present disclosure that involves using at least two main engines, e.g., two main engines 1340, disposed in a symmetric arrangement, for providing an even thrust distribution, notwithstanding the possibility that at least one of the two main engines might become shut down and optionally jettisoned, such as for adjusting the direction of the thrust vector, e.g., for adjusting at least one of pitch and yaw of the vehicle 1200.

[0092] Referring to FIG. 14, this schematic diagram illustrates, in a semi-exploded sectional side view, a rocket vehicle (launch vehicle), such as a multistage rocket vehicle 1400, powerable by an enhanced combustion rocket engine system, such as the systems 100, 200, 300, 400, in accordance with an embodiment of the present disclosure. The multistage rocket vehicle 1400 comprises a plurality of stages, e.g., the plurality of stages 1410, 1420, 1460, 1470 comprising at least one upper stage 1460, 1470 and at least one lower stage 1410, 1420. Each stage of the plurality of stages 1410, 1420, 1460, 1470 is capable of utilizing at least one enhanced combustion engine system, such as the systems 100, 200, 300, 400, wherein each stage is capable of accommodating its own engine(s) and its own propellants. The at least one enhanced combustion engine system comprises at least one of: at least one main engine 1440 associated with a nozzle 1440a and at least one auxiliary engine 1490 associated with a nozzle extension 1495, the at least one auxiliary engine 1490 being smaller than the at least one main engine 1440. By example only in this embodiment, the multistage rocket vehicle 1400, comprising the plurality of stages 1410, 1420, 1460, 1470, each stage capable of accommodating at least one engine, is initially powerable by the at least one main engine 1440, e.g., on take-off and early segments of the space mission. Later in the flight (in transit during the space mission), the at least one main engine 1440, is optionally capable of shut down; and the at least one main engine 1440 as well as its associated nozzle 1440a is capable of ejection (being jettisoned), from the at least one lower stage 1410, 1420 of the multistage rocket vehicle 1400 to reduce mass, wherein the plurality of stages 1410, 1420, 1460, 1470, and any associated structure, such as a fairing, e.g., a payload fairing 1475, are subsequently powerable by at least one of: at least one remaining engine 1440 and the at least one auxiliary engine 1490. An alternative to shifting the position of an engine, such as shown in FIG. 8, is another technique encompassed by the present disclosure that involves using at least one main engine, e.g., the main engine 1440, disposed along a center axis of the vehicle 1400 and movable in relation thereto, e.g., in a lateral direction G, for providing an even thrust distribution and for adjusting the direction of the thrust vector e.g., for adjusting at least one of pitch and yaw of the vehicle 1400.

[0093] Referring to FIG. 15, this schematic diagram illustrates, in a semi -exploded sectional side view, a rocket vehicle (launch vehicle), such as a multistage rocket vehicle 1500, powerable by an enhanced combustion rocket engine system, such as the systems 100, 200, 300, 400, in accordance with an embodiment of the present disclosure. The multistage rocket vehicle 1500 comprises a plurality of stages, e.g., the plurality of stages 1510, 1520, 1560, 1570 comprising at least one upper stage 1560, 1570 and at least one lower stage 1510, 1520. Each stage of the plurality of stages 1510, 1520, 1560, 1570 is capable of utilizing at least one enhanced combustion engine system, such as the systems 100, 200, 300, 400, wherein each stage is capable of accommodating its own engine(s) and its own propellants. The at least one enhanced combustion engine system comprises at least one of: at least one main engine 1540 associated with a nozzle 1540a and at least one auxiliary engine 1590 associated with a nozzle extension 1595, the at least one auxiliary engine 1590 being smaller than the at least one main engine 1540. By example only in this embodiment, the multistage rocket vehicle 1500, comprising the plurality of stages 1510, 1520, 1560, 1570, each stage capable of accommodating at least one engine, is initially powerable by the at least one main engine 1540, e.g., on take-off and early segments of the space mission. Later in the flight (in transit during the space mission), the at least one main engine 1540, is capable of shut down; and the at least one main engine 1540 as well as its associated nozzle 1540a is capable of ejection (being jettisoned), e.g., in a direction H, from the at least one lower stage 1510, 1520 of the multistage rocket vehicle 1500 to reduce mass, wherein the plurality of stages 1510, 1520, 1560, 1570, and any associated structure, such as a fairing, e.g., a payload fairing 1575, are subsequently powerable by at least one of: at least one remaining engine 1540 and the at least one auxiliary engine 1590. An alternative to shifting the position of an engine, such as shown in FIG. 8, is another technique encompassed by the present disclosure that involves using at least three main engines, e.g., three main engines 1540, disposed in a symmetric arrangement, for providing an even thrust distribution, notwithstanding the possibility that at least one of the three main engines might become shut down and optionally jettisoned, such as for adjusting the direction of the thrust vector, e.g., for adjusting at least one of pitch and yaw of the vehicle 1500.

[0094] Still referring back to FIGS. 1-4, the present disclosure involves various reducer components in propellant combinations. For example, boron naturally occurs with substantial concentrations of two main isotopes: one isotope having an atomic mass of approximately 10 AMU and the other isotope having an atomic mass of approximately 11 AMU. The majority of mass in naturally occurring boron comprises the heavier atoms, wherein the lighter atoms exist in adequate concentration and are separable without high energy expenditure. Due to the low atomic mass of boron, its isotopes are separable with minimal energy expenditure, despite having mass difference of only 1 AMU. For the latter stages of multistage rocket vehicles that are given a large change in velocity (significant acceleration or significant deceleration), an unnatural high concentration of the lighter boron- 10 isotope may be used, in accordance with another embodiment of the present disclosure. In an embodiment, the reducer component and fuel may comprise titanium, for example, titanium with an unnatural high concentration of titanium atoms of lower than average mass. [0095] In relation to the consumable components of a rocket vehicle, many related art heat shields are necessary for re-entry into the Earth's atmosphere. Typically, related art heat shields are fabricated from extreme high-temperature refractory ceramic materials, such as extreme high-temperature metal oxides. With a related art heat shield, the onboard power source that is available during transit is not conducive for use in an energy intensive reduction of the related art heat shield; and such onboard power source is better directly used for propulsion. In addition, solar power is not abundant at some destinations, such as one of Saturn's moon, Titan. As such, solar power for use in reducing a related art heat shield is insufficient. While such heat shield is necessary for some portions of a space mission, e.g., for other uses on the planet surface or for the return trip to orbit or beyond, and may not be a good candidate for early consumption.

[0096] To address at least the foregoing related art challenges, the present disclosure further contemplates and encompasses an enhanced heat shield structure that is more versatile than related art heat shields for at least being capable of self-consumption by the presently disclosed enhanced combustion rocket engine system. In some space missions, landers or landing vehicles, targeting Venus, Mars, or Titan, do indeed use heat shields. The enhanced heat shield structure of the present disclosure is capable of utilizing the atmospheres of some space destinations that are not oxidizing or highly oxidizing, such as Venus and Mars, each having an atmosphere comprising mostly carbon dioxide (CO₂) atmosphere, and Titan, having an atmosphere comprising mostly nitrogen (N₂), and of being self -consumable by the presently disclosed enhanced combustion rocket engine system.

[0097] For some missions, e.g., not involving necessitating a heat shield for reentering Earth's atmosphere, corrosion resistance is not as significant a requirement; and the enhanced heat shield structure comprises an enhanced material, such as a metal, a metal mixture, a metal alloy, non-metal, and/or a semi-metal, e.g., a graphite (C) and any other carbon (C) allotrope, capable of withstanding a high temperature (incapable of being easily melted, without using the extreme high-temperature refractory metal oxide ceramic materials of related art heat shields. The enhanced material comprises a material that is not easily meltable, but is still meltable at a higher temperature in a range that is less than approximately the reducing temperatures for extreme high-temperatures refractory ceramic materials. While the enhanced heat shield may comprise a metal, a metal mixture, and a metal alloy, generally having a high thermal conductivity, whereby their insulation capability may be less than that of extreme high-temperatures refractory ceramic materials, the enhanced heat shield may also comprise a material that is more robust, such as a metal particle, a metal powder, a metal fiber, a metal tape, a woven metal, a metal wool, a metal foam, a porous metallic material, a semi-metal particle, a semi-metal powder, a semi-metal fiber, a semi-metal tape, a woven semi-metal, a semi-metal wool, a semi-metal foam, and a porous semi-metallic material, wherein the metal comprises a material, such as silicon, and wherein the semi-metal comprises a material, such as carbon, graphite, and any carbon allotrope, e.g., a fullerene. The graphite comprises carbon in at least one stack of graphene sheets having linked hexagonal rings. A fullerene is any molecule comprising carbon in the form of a hollow structure, e.g., a sphere, an ellipsoid, a tube, and any other hollow shape, including, but not limited to, spherical fullerenes or "buckyballs" and cylindrical fullerenes (carbon nanotubes) or "buckytubes." While fullerenes may comprise carbon in at least one stack of graphene sheets having linked hexagonal rings, fullerenes may also comprises pentagonal, and/or sometimes heptagonal, rings.

[0098] The foregoing materials for the enhanced heat shield may be combined in various configurations for effecting sufficient thermal insulation as well as providing consumability as a supplemental fuel source. By example only, a structure, such as a metal or semi-metal foam, or a porous metal or semi-metal structure, or a metal or semi-metal powder, may be sandwiched between another structure, such as sheets of solid metal or semi-metal material, thereby providing an effective thermal insulation, wherein at least the plurality of voids in the sandwich structure effectively provide long and convoluted pathways through the matrix material for reducing thermal conduction therethrough. Alternatively, the present disclosure encompasses a configuration, wherein a metal or semi-metal powder is disposed between metal sheets, and wherein the metal or semi-metal powder facilitates processing.

[0099] In addition, the enhanced heat shield comprises lithium and/or alloys thereof for some applications, in accordance with an embodiment of the present disclosure. In this embodiment, lithium (Li) is capable of alloying with aluminum. While metals, such as lithium, aluminum, and magnesium, may not withstand extreme high temperatures, substantial creep is tolerable, and possibly even mitigated, during reentry of the space vehicle into an atmosphere by alloying these metals, wherein some parts of the enhanced heat shield are exposed to elevated temperature, but not extreme high temperatures. Therefore, in another embodiment of the present disclosure, the enhanced heat shield may comprise a material, such as aluminum, magnesium, and/or alloys thereof, for the cooler portions; and a material, such as titanium, silicon, and/or alloys thereof, for the hotter portions, whereby sufficient thermal insulation and consumability as a supplemental fuel source are provided.

[0100] With respect to titanium, this metal withstands temperatures in a range that is approximately 200 degrees Celsius higher than does aluminum. The gravimetric energy density of titanium is competitive with that of hydrogen and is higher than that of some related art fuels. Titanium also has a high volumetric energy density as well as a high strength-to-mass ratio. Silicon may be more brittle than aluminum, but silicon can also withstand much higher temperatures than aluminum or magnesium. Silicon also has a higher energy density than magnesium or titanium and a lower thermal conductivity and melting point than titanium. As such, in the various embodiments of the enhanced heat shield, combinations of these metals and their alloys, or non-metals, enhance performance of a rocket vehicle for certain space missions. Alternatively, the present disclosure contemplates and encompasses systems and methods comprising onboard, and possibly deployable, refining equipment for reducing oxides in situ. While mining or harvesting metal oxides is a locally available option, exploiting the purity of the material contained in the enhanced heat shield itself (by refining the enhanced heat shield's material) may be a more valuable option for the enhanced combustion rocket engine system, e.g., if such refining equipment is sufficiently reliable. Another alternative embodiment of the present systems and methods involves forgoing refinement of the enhanced heat shield and locally mining or harvesting appropriate proportions of metal, e.g., silicon, and either locally gathering oxygen, e.g., by electrolysis of local water or by heating local carbon dioxide, or transporting a supplemental onboard oxidizer, for melting and burning the enhanced heat shield.

[0101] A rocket comprises: a first propellant tank, substantially filled with hot propellant, the hot propellant comprising mostly types of material which are typically solid near standard sea-level atmospheric temperature and pressure on Earth, much of the hot propellant being in the molten state; a second propellant tank containing cold propellant; a rocket engine configured to accept the hot propellant at temperatures near to or higher than the elevated temperature at which it is stored in the first propellant tank and the cold propellant to react the propellants in an exothermic reaction, such that pressure is created by the heat of the exothermic reaction, which expels the exhaust to create thrust. The rocket engine may be configured such that when delivering nearly the maximum amount of thrust that it can, the rocket engine derives the majority of the energy used for expelling propellant from combustion of propellant. The rocket may be configured such that the majority of the propellant stored on the rocket of at least one of the types configured to be delivered to the engine is stored primarily in the molten state and delivered to the engine at temperature above its melting temperature. The rocket may be configured to accept molten aluminum as the majority of the hot propellant. The rocket may be configured to accept oxygen as the majority of the cold propellant. The oxygen may be primarily in the liquid state when stored in the second propellant tank. The rocket may be configured to accept oxygen as the cold propellant, the oxygen may be primarily in the liquid state when stored in the cold propellant tank, and the engine may be configured to use oxygen as coolant. The rocket may be configured to accept molten beryllium as the majority of the hot propellant. The rocket may comprise: the first propellant tank filled to more than half of its maximum capacity with the hot propellant; the second propellant tank filled to more than half of its maximum capacity with the cold propellant; and the rocket may further comprise a combustion powered preheater distinct from the rocket engine configured to be able to accept the majority of the hot propellant and provide it a certain average temperature increase while supplying it to the engine. The rocket engine may be configured to accept fuel from the first propellant tank by way of the combustion powered preheater and the second propellant tank and burn the propellant to produce thrust, the thermal energy applied to the propellant by the combustion powered preheater need not be used to supply the majority of the energy used to force the propellant into the rocket engine. At least one of the types of fuels powering the combustion powered preheater may be a type of fuel different from that contained in either the first propellant tank or the second propellant tank. The combustion powered preheater may be configured to be able to provide the hot propellant from the first propellant tank an average temperature increase of more than 300 degrees Celsius, in such a way that the majority of the thermal energy added to the hot propellant by the preheater would be propagated through the combustion reaction of the rocket engine to its exhaust. The hot propellant may be primarily aluminum.

[0102] A rocket vehicle may comprise: a first propellant tank, the tank configured to contain hot solid propellant; a second propellant tank, the tank able to contain propellant; a set of one or more rocket engines configured to accept hot propellant from the first propellant tank and propellant from the second propellant tank and react them to produce thrust. The majority of the mass of the hot solid propellant may be comprised primarily of powder or pellets, wherein the majority of the mass is comprised of particles of less than 10 millimeters in length. The first propellant may be primarily comprised of a combination of aluminum and magnesium. The hot solid propellant may be primarily comprised of one or more of the following elements: aluminum, magnesium, boron, silicon, or beryllium, calcium, or titanium. The hot solid propellant may be primarily comprised of a combination of aluminum and magnesium. The hot solid propellant may be primarily comprised of aluminum. The hot solid propellant may be primarily comprised of boron. The hot solid propellant may be primarily comprised of beryllium. The majority of the mass of the hot solid propellant may be above 350 degrees Celsius when the rocket vehicle is filled with approximately the maximum amount of propellant which it can contain. The majority of the mass of the hot solid propellant may be above 350 degrees Celsius when the rocket vehicle is filled with approximately the maximum amount of propellant which it can contain. The majority of the mass of the hot solid propellant may be above 500 degrees Celsius when the rocket vehicle is filled with approximately the maximum amount of propellant which it can contain. The rocket may further comprise: a set of one or more partitions approximately conformal to the shape of the first propellant tank, smaller than the first propellant tank, and oriented inside of it in such a way that an outer layer of hot solid propellant can be positioned between the set of one or more partitions and the first propellant tank, such that the outer layer of hot solid propellant approximately conforms nearly to the entire outer surface of the set of one or more partitions excepting at the ullage space of the first propellant tank, and an inner layer of hot solid propellant may line the inner surface of the set of one or more partitions, wherein the inner layer of hot solid propellant is substantially hotter than the outer layer of hot solid propellant. The rocket may further comprise: a combustion powered preheater configured to be able to accept the majority of the hot solid propellant and provide it a certain average temperature increase while supplying it to the rocket engine. The rocket engine may be configured to accept hot propellant from the first propellant tank by way of the combustion powered preheater and propellant from the second propellant tank and burn the propellant to produce thrust, the thermal energy from combustion powered preheater need not be used to supply the majority of the energy used to force propellant into the set of one or more engines. The average temperature increase which the combustion powered preheater is able to provide may be greater than 350 degrees Celsius. The average temperature increase which the combustion powered preheater is able to provide may be greater than 600 degrees Celsius. The rocket vehicle may be filled with propellant to at least three quarters of its maximum capacity of propellant and the majority of the reducer fuel aboard the rocket vehicle may be solid. The rocket vehicle may be filled with reducer propellant to at least three quarters of its maximum capacity of propellant and the majority of the reducer fuel aboard the rocket vehicle is one or more of the following: aluminum, magnesium, silicon, boron, beryllium, calcium, or titanium. The average temperature increase which the combustion powered preheater is able to provide may be greater than 600 degrees Celsius.

[0103] A rocket vehicle may comprise: a set of one or more reducer propellant tanks wherein the majority of the reducer fuel aboard the rocket is contained when the rocket vehicle is filled with reducer propellant to at least three quarters of its maximum capacity of propellant; a propellant tank such as an oxidizer propellant tank; a set of one or more rocket engines; a combustion powered preheater configured to be able to accept the majority of the reducer fuel and provide it a certain average temperature increase while supplying it to the set of one or more engines. The set of one or more rocket engines may be configured to accept fuel from the set of one or more reducer tanks by way of the combustion powered preheater and the propellant tank and burn the propellant to produce thrust, the thermal energy applied to the propellant by the combustion powered preheater is not used to supply the majority of the energy used to force propellant into the set of one or more rocket engines. The average temperature increase which the combustion powered preheater is able to provide may be greater than 350 degrees Celsius. The average temperature increase which the combustion powered preheater is able to provide may be greater than 600 degrees Celsius. The majority of the reducer fuel aboard the rocket vehicle when the rocket vehicle is filled to at least three quarters of its maximum capacity of propellant may be solid. The majority of the reducer fuel aboard the rocket vehicle when the rocket vehicle is filled with reducer propellant to at least three quarters of its maximum capacity of propellant may be one or more of the following: aluminum, magnesium, silicon, boron, beryllium, calcium, or titanium. The majority of the reducer fuel aboard the rocket vehicle when the rocket vehicle is filled with reducer propellant to at least three quarters of its maximum capacity of propellant may be liquid. The average temperature increase which the combustion powered preheater is able to provide may be greater than 600 degrees Celsius.

[0104] A rocket vehicle may comprise: a first propellant tank primarily comprising flammable structural material, the tank configured to contain propellant; a second propellant tank, the tank able to contain propellant; a set of one or more rocket engines configured to accept propellant from the first propellant tank and the second propellant tank; a tank preparation system, capable of altering the first propellant tank and supplying the majority of the flammable structural material of which it is composed to the set of one or more rocket engines in such a way that the rocket engines can produce thrust from burning the majority of the flammable structural material, wherein the set of one or more rocket engines is configured to produce thrust by burning propellant from the first propellant tank until the tank is nearly depleted of its contents, then produce thrust by burning the majority of the mass of the flammable structural material in combination with propellant from the second propellant tank, relying on the tank preparation system to supply the flammable structure material. The majority of the mass of the propellant contained within the first propellant tank when the tank is more than half full may be liquid oxygen; and the majority of the mass of the flammable structural material may be a combination of aluminum and magnesium. The tank preparation system may comprise: a molten reducer tank, containing a reducer fuel which primarily comprises a substance which is typically solid at standard sea level atmospheric temperature and pressure on Earth, and which exists primarily in the molten state at a temperature above the melting temperature of the majority of the material of which the flammable structural material is comprised; wherein the manner in which the tank preparation system alters the first propellant tank is by melting it, and the tank preparation system supplies the first propellant tank to the set of one or more rocket engines primarily in the molten state. The rocket may further comprise: a first reducer tank, wherein the set of one or more rocket engines burns propellant from the first reducer tank by chemical reaction with the contents of the first propellant tank depleting the first reducer tank and ejecting it from the rocket before the majority of the flammable structural material has been burned.

[0105] A rocket vehicle may comprise: a set of one or more payload fairings comprising flammable structural material; a set of one or more rocket engines; a fairing preparation system, capable of altering the set of one or more payload fairings and supplying the majority of the flammable structural material of which it is composed to the set of one or more rocket engines in such a way that the rocket engines can produce thrust from burning the majority of the flammable structural material; a control system, able to activate the fairing preparation system a short time after the rocket vehicle passes from a condition wherein its speed and altitude are such that aerodynamic effects may have an undesirable effect on a payload located near the top of the rocket to a condition where its altitude is high enough relative to its speed to avoid such harmful effects, wherein the set of one or more rocket engines is configured to produce thrust by burning the majority of the mass of the flammable structural material, relying on the fairing preparation system to supply the flammable structural material. The majority of the mass of the flammable structural material may be a combination of aluminum and magnesium.

[0106] A rocket vehicle may comprise: a first propellant tank primarily comprising flammable structural material, the tank configured to contain high pressure liquid propellant, the flammable structural material comprising one or more layers of structural reinforcement material able to protect the tank from damage due to extreme pressure difference between its contents and its exterior environment, the one or more layers configured to be sequentially removed without causing damage to the first propellant tank as long as interior pressure adequately decreases before the removal of each layer; a set of one or more rocket engines configured to accept propellant from the first propellant tank; a tank unwrapping system, capable of removing the layers of structural reinforcement material from the first propellant tank and supplying the majority of the flammable structural material of which they are composed to the set of one or more rocket engines in such a way that the rocket engines can produce thrust from burning the majority of the flammable structural material; wherein the set of one or more rocket engines is configured to produce thrust by burning propellant from the first propellant tank until the tank is nearly depleted of its contents and to produce thrust by burning the majority of the mass of the flammable structural material relying on the tank unwrapping system to supply the flammable structural material. The high pressure propellant may be liquid oxygen above 110 degrees Kelvin. The flammable structural material may be comprised substantially of aluminum.

[0107] A rocket vehicle may comprise: a set of one or more stabilizer fins comprising flammable structural material; a set of one or more rocket engines; a fin preparation system, capable of altering the set of one or more stabilizer fins and supplying the majority of the flammable structural material of which it is composed to the set of one or more rocket engines in such a way that the rocket engines can produce thrust from burning the majority of the flammable structural material; a control system, able to activate the fin preparation system a short time after the rocket vehicle passes from a condition wherein its speed and altitude are such that aerodynamic benefits of using the fins for guidance or stabilization of the rocket vehicle require their use to a condition where its altitude is high enough relative to its speed to avoid requiring the fins. The set of one or more rocket engines may be configured to produce thrust by burning the majority of the mass of the flammable structural material, relying on the fin preparation system to supply the flammable structural material. The majority of the mass of the flammable structural material may be a combination of aluminum and magnesium.

[0108] Information as herein shown and described in detail is fully capable of attaining the above-described object of the present disclosure, the presently preferred embodiment of the present disclosure, and is, thus, representative of the subject matter which is broadly contemplated by the present disclosure. The scope of the present disclosure fully encompasses other embodiments which may become obvious to those skilled in the art, and is to be limited, accordingly, by nothing other than the appended claims, wherein any reference to an element being made in the singular is not intended to mean "one and only one" unless explicitly so stated, but rather "one or more." All structural and functional equivalents to the elements of the above-described preferred embodiment and additional embodiments as regarded by those of ordinary skill in the art are hereby expressly incorporated by reference and are intended to be encompassed by the present claims.

[0109] Moreover, no requirement exists for a system or method to address each and every problem sought to be resolved by the present disclosure, for such to be encompassed by the present claims. Furthermore, no element, component, or method step in the present disclosure is intended to be dedicated to the public regardless of whether the element, component, or method step is explicitly recited in the claims. However, that various changes and modifications in forms, material, work-piece, and fabrication material detail may be made, without departing from the spirit and scope of the present disclosure, as set forth in the appended claims, as may be apparent to those of ordinary skill in the art, are also encompassed by the present disclosure.

INDUSTRIAL APPLICABILITY

[0118] The subject matter of the present disclosure industrially applies to combustion rocket engine systems and methods. More specifically, the present disclosure industrially applies to combustion rocket engine systems and methods, having associated external support systems therefor. Even more specifically, the present disclosure industrially applies to enhanced combustion rocket engine systems and methods, having associated external support systems for use with improved fuels and improved fuel combinations.

Claims

CLAIMS What is claimed:

1. An enhanced combustion rocket engine system, comprising:

at least one primary propellant tank adapted to accommodate a primary propellant in at least one of a solid state, a transition melting state, and a molten state and to deliver the primary propellant in the molten state;

at least one secondary propellant tank adapted to accommodate a secondary propellant in a liquid state; and at least one rocket engine, the at least one rocket engine adapted to:

receive the primary propellant from the at least one primary propellant tank in a temperature range of at least approximately above an ambient temperature;

receive the secondary propellant from the at least one secondary propellant tank in a temperature range of at least approximately below an ambient temperature; and

exothermically react the primary and secondary propellants for effecting combustion to provide thrust.

2. The system of claim 1, wherein the at least one rocket engine is further adapted to, when delivering nearly the maximum possible amount of thrust, derive a majority of energy used for exhausting from the combustion.

3. The system of claim 1, further comprising a fuel transfer feature capable of delivering the primary propellant from the at least one primary propellant tank in the molten state to the at least one rocket engine at a temperature in a range of approximately above the melting point of the primary propellant.

4. The system of claim 1,

wherein the primary propellant comprises a hot propellant, a majority of the hot propellant comprising at least one reducer component of aluminum, beryllium, boron, calcium, magnesium, silicon, and titanium, and

wherein the secondary propellant comprises a cold propellant, a majority of the cold propellant comprising at least one oxidizer component of oxygen, nitric acid, dinitrogen tetroxide, and hydrogen peroxide.

5. The system of claim 1,

wherein the primary propellant comprises at least one form of a solid, a powder, and a pellet, and wherein the primary propellant comprises a particle size in a range of less than approximately 10 millimeters.

6. The system of claim 4,

wherein the hot propellant comprises a molten state when stored in the at least one primary propellant tank, and

wherein the cold propellant comprises a liquid state when stored in the at least one secondary propellant tank.

7. The system of claim 6, further comprising a cooling feature for cooling the at least one rocket engine, the cooling feature utilizing the cold propellant as a coolant.

8. The system of claim 3, wherein the fuel transfer feature comprises:

at least one fuel line; and

at least one combustion powered preheater coupled with the at least one fuel line, the at least one combustion powered preheater adapted for further preheating the primary propellant during delivery of the primary propellant to the at least one rocket engine.

9. The system of claim 8,

wherein the at least one primary propellant tank is filled to more than approximately one half of the primary propellant tank's maximum capacity with the hot propellant,

wherein the at least one secondary propellant tank is filled to more than approximately one half of the secondary propellant tank's maximum capacity with the cold propellant, and

wherein the at least one combustion powered preheater is adapted to apply energy to the primary propellant, the energy applied by the combustion powered preheater is distinct from the majority of energy used for forcing the primary propellant into the at least one rocket engine.

10. The system of claim 7,

wherein the combustion powered preheater is operable by way of a fuel having a type distinct from that of the at least one of the first propellant and the second propellant, and

wherein a majority of thermal energy, applied to the first propellant by the combustion powered preheater, is capable of propagating through a combustion reaction, and to an exhaust, of the at least one rocket engine.

11. The system of claim 9, wherein, when the at least one primary propellant tank is filled to approximately the at least one primary propellant tank's maximum capacity with the at least one primary propellant tank, a majority of the primary propellant comprises a temperature in a range of at least approximately 350 degrees Celsius.

12. The system of claim 11, wherein, when the at least one primary propellant tank is filled to approximately the at least one primary propellant tank's maximum capacity with the at least one primary propellant tank, a majority of the primary propellant comprises a temperature in a range of at least approximately 500 degrees Celsius.

13. The system of claim 8, wherein the combustion powered preheater is operable for preheating the first propellant to an average temperature increase range of at least one of at least approximately 300 degrees Celsius.

14. The system of claim 4, wherein the at least one primary propellant tank is filled to at least three quarters of the at least one primary propellant tank's maximum capacity with primary propellant.

15. The system of claim 5, wherein the at least one primary propellant tank is filled to at least three quarters of the at least one primary propellant tank's maximum capacity with primary propellant.

16. The system of claim 1, further comprising a tank preparation feature capable of altering the state of at least one consumable component into a supplemental fuel source and facilitating delivering the supplemental fuel source to the at least one rocket engine,

wherein the supplemental fuel source is burnable and contributes to producing thrust,

wherein the at least one primary propellant tank is capable of accommodating the at least one secondary propellant tank,

wherein the at least one secondary propellant tank comprises a plurality of secondary propellant tanks, wherein the at least one consumable component comprises a consumable structural material being at least one of meltable and flammable;

wherein the at least one consumable component comprises one secondary propellant tank of the plurality of secondary propellant tanks,

wherein the at least one rocket engine is adapted to produce thrust by:

burning the primary propellant delivered from the at least one primary propellant tank until the at least one primary propellant tank is nearly depleted of the primary propellant; and

subsequently burning the supplemental fuel source in combination with the secondary propellant from a remaining secondary propellant tank of the plurality of secondary propellant tanks.

17. The system of claim 16,

wherein a majority of the secondary propellant is disposed within the one secondary propellant tank of the plurality of secondary propellant tanks when the one secondary propellant tank is filled to more than one half of the one secondary propellant tank's maximum capacity is liquid oxygen, and

wherein a majority of the supplemental fuel source comprises a combination of aluminum and magnesium.

18. The system of claim 16,

wherein the tank preparation feature comprises a feature for delivering the at least one consumable component to the at least one primary propellant tank and a feature for melting the at least one consumable component in the at least one primary propellant tank, and

wherein the at least one primary propellant tank contains a reducer component in a molten state at a temperature range greater than approximately the melting point of the majority of the material comprising the at least one consumable component.

19. The system of claim 16, wherein the at least one consumable component is ejectable when the at least one consumable component has no further function during a mission.

20. The system of claim 16, further comprising:

at least one structural reinforcement layer for reinforcing, by wrapping, at least one of the at least one primary propellant tank and the at least one secondary propellant tank, the at least one structural reinforcement layer, at least one of the at least one primary propellant tank and the at least one secondary propellant tank being protected from damage from an extreme pressure difference between at least one of the primary propellant and the secondary propellant and an exterior environment; and a tank unwrapping feature capable of sequentially removing the at least one structural reinforcement layer from at least one of the at least one primary propellant tank and the at least one secondary propellant tank and altering the state of at least one consumable component into a supplemental fuel source and facilitating delivering the supplemental fuel source to the at least one rocket engine,

wherein the at least one primary propellant tank is capable of accommodating the at least one structural reinforcement layer,

wherein the at least one consumable component is wrapped with the at least one structural reinforcement layer.

21. The system of claim 20,

wherein the secondary propellant comprises a high pressure propellant,

wherein the high pressure propellant comprises a liquid oxygen, and

wherein the liquid oxygen comprises a temperature range of at least approximately 110 degrees Kelvin.

22. The system of claim 20, wherein the at least one structural reinforcement layer comprises aluminum.

23. The system of claim 1,

wherein the primary propellant comprises a solid form,

wherein the at least one primary propellant tank comprises at least one partition having a shape that is approximately conformal to that of the at least one primary propellant tank and a size that is smaller than the at least one primary propellant tank, the at least one partition facilitating disposition of the primary propellant and providing thermal protection to the at least one primary propellant tank,

wherein a portion of the primary propellant is disposed inboard of the partition,

wherein a portion of the primary propellant is disposed outboard of the at least one partition, and wherein the inboard portion of the primary propellant is substantially hotter than the outboard portion of the primary propellant.

24. The system of claim 23, further comprising at least one combustion powered preheater adapted for preheating the primary propellant for facilitating delivery of the primary propellant to the at least one rocket engine.

25. An enhanced combustion rocket engine vehicle comprising:

at least one primary propellant tank adapted to accommodate a primary propellant in at least one of a solid state, a transition melting state, and a molten state and to deliver the primary propellant in the molten state and to accommodate at least one consumable component;

exothermically react the primary and secondary propellants for effecting combustion to provide thrust; and at least one fairing for accommodating a payload, the fairing coupled with at least one of: the at least one primary propellant tank, the at least one secondary propellant tank, and the at least one engine.

26. The vehicle of claim 25, further comprising:

a fairing preparation feature capable of altering the state of at least one consumable component into a supplemental fuel source and facilitating delivering the supplemental fuel source to the at least one rocket engine; and

a control system operable for activating the fairing preparation feature,

wherein the at least one primary propellant tank is capable of accommodating the at least one fairing, wherein the at least one secondary propellant tank comprises a plurality of secondary propellant tanks, wherein the at least one consumable component comprises a consumable structural material being at least one of meltable and flammable,

wherein the at least one consumable component comprises the at least one fairing,

wherein the at least one fairing comprises a consumable material, and

wherein the at least one rocket engine is adapted to produce thrust by:

27. The vehicle of claim 26, wherein a majority of the at least one fairing comprises a combination of aluminum and magnesium.

28. The vehicle of claim 25, further comprising:

at least one stabilizer fin comprising flammable structural material;

a fin preparation feature capable of altering the state of at least one consumable component into a supplemental fuel source and facilitating delivering the supplemental fuel source to the at least one rocket engine; and

a control system for activating the fin preparation,

wherein the at least one primary propellant tank is capable of accommodating the at least one stabilizer fin, wherein the at least one secondary propellant tank comprises a plurality of secondary propellant tanks, wherein the at least one consumable component comprises a consumable structural material being at least one of meltable and flammable,

wherein the at least one consumable component comprises the at least one stabilizer fin,

wherein the at least one fairing comprises a consumable material, and

wherein the at least one rocket engine is adapted to produce thrust by:

burning the primary propellant delivered from the at least one primary propellant tank until the at least one primary propellant tank is nearly depleted of the primary propellant; and subsequently burning the supplemental fuel source in combination with the secondary propellant from a remaining secondary propellant tank of the plurality of secondary propellant tanks.

90

29. The vehicle of claim 28, wherein a majority of the at least one stabilizer fin comprises a combination of aluminum and magnesium.

30. The vehicle of claim 25, wherein the consumable component comprises at least one of at least one 95 secondary propellant tank, a fairing, a portion of a rocket engine nozzle, and a frame structure of a rocket.

31. A method of fabricating an enhanced combustion rocket engine system, the method comprising:

providing at least one primary propellant tank adapted to accommodate a primary propellant in at least one of a solid state, a transition melting state, and a molten state and to deliver the primary propellant in the molten state; 100 providing at least one secondary propellant tank adapted to accommodate a secondary propellant in a liquid state; and

providing at least one rocket engine, the at least one rocket engine adapted to:

105 receive the secondary propellant from the at least one secondary propellant tank in a temperature range of at least approximately below an ambient temperature; and

32. A method of using an enhanced combustion rocket engine system, the method comprising:

110 providing an enhanced combustion rocket engine system, the rocket engine system providing comprising:

providing at least one primary propellant tank adapted to accommodate a primary propellant in at least one of a solid state, a transition melting state, and a molten state and to deliver the primary propellant in the molten state;

providing at least one secondary propellant tank adapted to accommodate a secondary propellant 115 in a liquid state; and

receive the secondary propellant from the at least one secondary propellant tank in a temperature 120 range of at least approximately below an ambient temperature; and

exothermically react the primary and secondary propellants for effecting combustion to provide thrust;

charging the at least one primary propellant tank with the primary propellant;

charging the at least one secondary propellant tank with the secondary propellant;

125 preheating, heating, and delivering the primary propellant from the at least one primary propellant tank to the at least one rocket engine in a temperature range of at least approximately above an ambient temperature;

delivering the secondary propellant from the at least one secondary propellant tank to the at least one rocket engine in a temperature range of at least approximately below an ambient temperature; and exothermically reacting the primary and secondary propellants for effecting combustion in the at least one 130 rocket engine to provide thrust.

33. A propellant formulation for an enhanced combustion rocket engine system, comprising:

a primary propellant in at least one of a solid state, a transition melting state, and a molten state and to deliver the primary propellant in the molten state, the primary propellant comprising a temperature in a range of at 135 least approximately above an ambient temperature,; and

a secondary propellant in a liquid state, the secondary propellant comprising a temperature in a range of at least approximately below an ambient temperature,

the primary and secondary propellants capable of exothermically reacting with each other for effecting combustion to provide thrust.

140

34. The propellant formulation of claim 33,

wherein the primary propellant comprises a hot propellant, a majority of the hot propellant comprising at least one reducer component of aluminum, beryllium, boron, calcium, lithium, magnesium, silicon, and titanium, and

145 wherein the secondary propellant comprises a cold propellant, a majority of the cold propellant comprising at least one oxidizer component of oxygen, nitric acid, dinitrogen tetroxide, hydrogen, and hydrogen peroxide.

35. The propellant formulation of claim 34, further comprising a tertiary propellant capable of exothermically reacting with at least one of the primary propellant and the secondary propellant for further effecting combustion to

150 provide thrust,

wherein the primary propellant comprises lithium,

wherein the secondary propellant comprises hydrogen, and

wherein the tertiary propellant comprises fluorine.

The propellant formulation of claim 33,

wherein the primary propellant comprises at least one form of a solid, a powder, and a pellet, and wherein the primary propellant comprises a particle in a range of less than approximately 10 millimeters.

37. The propellant formulation of claim 34,

160 wherein the hot propellant comprises a molten state for facilitating storage in at least one primary propellant tank, and

wherein the cold propellant comprises a liquid state for facilitating storage in at least one secondary propellant tank.

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