Compartmentalized rocket propellant tank layout, modular rocket propulsion unit and a rocket launch vehicle

Abstract

A rocket propellant tank and layout (15) is disclosed. The tank (15) includes a housing operationally coupled to a rocket engine (20). The housing comprises a plurality of compartments (17a, 17b, 17c, 17d, 17e, 17f) arranged longitudinally along the housing and is adapted to store a propellant. A plurality of separator walls (24, 41, 42, 32, 33, 34) are disposed adjacent to each of the plurality of compartments (17a, 17b, 17c, 17d, 17e, 17f). A set of control valves (45) are coupled to each of the plurality of compartments (17a, 17b, 17c, 17d, 17e, 17f). The set of control valves (45) are adapted to facilitate a controlled transfer of the propellant from each of the plurality of compartments (17a, 17b, 17c, 17d, 17e, 17f) to the rocket engine (20).

Description

TITLE OF THE INVENTION “COMPARTMENTALIZED ROCKET PROPELLANT TANK LAYOUT, MODULAR ROCKET PROPULSION UNIT AND A ROCKET LAUNCH VEHICLE” The following specification describes the invention and the manner in which it is to be performed.FIELD OF THE INVENTION

[0001] The present invention relates to a field of rocket technology for launching payloads into space, and more particularly, relates to a modular rocket propulsion unit and a rocket launch vehicle.BACKGROUND OF THE INVENTION

[0002] The following description provides the information that may be useful in understanding the present invention. It is not an admission that any of the information provided herein is prior art or relevant to the presently claimed invention, or that any publication specifically or implicitly referenced is prior art.

[0003] The development of advanced rocket propulsion systems and tank designs has been a cornerstone of space exploration and technology. Conventional rocket propellant tanks have traditionally been large, singular structures designed to store either fuel or oxidizer in a homogeneous volume. Conventional rocket propellant tanks are typically configured as stacked modules, integrated with a common bulkhead, or arranged as co-axial or tank-in-tank systems. This separation of fuel and oxidizer ensures safety by preventing dangerous reactions. However, such designs often lead to inefficiencies in space utilization and add significant structural mass to the rocket. As rocket dimensions increase in diameter and length to support more ambitious missions, traditional rocket tank designs encounter severe limitations in thermal management, structural integrity, and space optimization. These challenges underscore the growing need for innovative solutions that can meet the demands of modern aerospace applications.

[0004] One critical issue with traditional rocket propellant tank designs is the susceptibility of rocket propellant tanks to sloshing, which can lead to inefficient fuel usage and destabilization during flight. Efforts to address fuel sloshing often involve complex baffle designs that, while somewhat effective, add unnecessary weight and complexity. Current tank designs are generally adequate for small tomedium-diameter rockets but struggle to maintain structural stability in larger- diameter rocket applications. As rockets grow in size, homogeneous, pipe-like tank structures become increasingly prone to structural failure under high-pressure conditions, especially as the fuel depletes. This vulnerability is particularly evident during the launch, flight, and landing phases of reusable rockets. The lack of internal reinforcement in traditional rocket propellant tanks exacerbates these challenges, making conventional tank designs less viable for modem large- diameter rocket applications.

[0005] Structural integrity poses another significant challenge in conventional rocket propellant tank designs. For large-diameter rockets, traditional tank designs often lack the necessary internal support to withstand the stresses encountered during launch and flight. As the diameter of the rocket increases, the absence of reinforcement leads to weakened structures that are prone to buckling or collapse, particularly under the high pressures of operation. Adding reinforcements to counteract these weaknesses significantly increases the overall weight of the rocket, reducing payload capacity and fuel efficiency. Additionally, the large, undivided volume of traditional rocket propellant tanks worsens fuel sloshing, destabilizing the rocket and complicating fuel management. These structural and functional drawbacks render traditional rocket propellant tank layouts unattractive for modem applications, where weight optimization, fuel stability, and reliability are critical.

[0006] Propellant sloshing further complicates the performance of large-diameter rocket propellant tanks. In varying gravitational environments, such as during launch, coasting phases, or landing in reusable rockets, the liquid propellant moves unpredictably within the tank. This uncontrolled motion disrupts the balance and stability of the rocket, creating significant challenges in maintaining a consistent fuel supply to the engine. These issues not only impact flight performance but also pose safety risks, particularly in deep-space missions where precision and reliability are crucial.

[0007] Figures 1A-1G illustrate schematic vertical cross-sectional views of a typical rocket propulsion module, in accordance with a conventional techniques.

[0008] The rocket propulsion module includes three primary components, a propellant tank system, a combustion chamber, and an exhaust nozzle. The rocket propulsion module includes a host of supporting structures, such as plumbing, electronics, ignition systems, turbo pumps, and other accessories required to operate the rocket propulsion module. The combustion chamber is integrated with the exhaust nozzle, ignition system, turbo pumps, and plumbing.

[0009] Figure 1 A illustrates a spherical tank layout comprising two tanks, one for the oxidizer and one for the fuel. The plumbing is routed externally to deliver the propellants to the rocket engine's combustion chamber located below the tanks.

[0010] Figure IB depicts the propellant tank layout similar to that in Figure 1A. However, in this configuration, the tanks are cylindrical instead of spherical. The plumbing is routed externally to deliver the propellants to the rocket engine's combustion chamber located below the tanks.[Oil] Figure 1C depicts a rocket propulsion system with a propellant tank layout similar to that in Figure IB. However, in this configuration, the plumbing is routed internally to deliver the propellants to the rocket engine's combustion chamber located below the tanks.

[0012] Figure ID depicts a cylindrical tank Layout with a common bulkhead. In this configuration, the oxidizer and fuel tanks are not separate tanks but are integrated with a common bulkhead to reduce the mass of the overall propellant tank system. The plumbing is routed internally to deliver the propellants to the rocket engine's combustion chamber located below the tanks.

[0013] Figure IE depicts a co-axial tank layout with a tank inside tank. In this configuration, all tanks are laid one inside another. In other words, the tanks arecoaxially placed, one smaller diameter tank inside a bigger diameter tank. The inner tank holds one type of propellant, and the outside tank holds another type of propellant.

[0014] Figure IF depicts a coaxial tank layout with a tank-inside-a-tank configuration. In this configuration, the fuel tank is at least partially arranged within the oxidizer tank.

[0015] Figure 1G depicts a multi-tank layout, i.e., tank-next-to-tank. In this case, multiple individual tanks are welded together to form the propellant tank system. Each tank holds a specified amount of rocket propellant and is connected to a complex common feeder line system located below. This feeder line system delivers the propellants to the rocket engine.

[0016] Figure 2 illustrates a schematic vertical cross-sectional view with major components outlined for a typical rocket propulsion module 13, in accordance with a conventional technique.

[0017] The propulsion module 13 is based on the most commonly used rocket propellant tank layout 9. The rocket propellant tank layout 9 includes a tank-over- tank configuration, with an oxidizer tank 7 located below and a fuel tank 12 positioned above. The oxidizer tank 7 contains the oxidizer 6, which is liquid oxygen or unsymmetrical dimethylhydrazine (UDMH).

[0018] The fuel tank 12 stores the fuel 10, typically RP-1 or refined kerosene or hydrazine. The tanks are stacked one above the other. A pipe 5 carries the fuel from the top tank 12 to the bottom of the propulsion module 13, where the top tank 12 connects to pipe turbopump assembly 3. The assembly 3 feeds into the combustion chamber 2. Similarly, a pipe 4 turbopump assembly delivers the oxidizer from the oxidizer tank 7 to the combustion chamber 2. The combustion chamber 2 may ignite the oxidizer and fuel, producing high-temperature gases. These gases are expelled at extremely high velocity through the exhaust nozzle 1, generating thrust.

[0019] The limitations of current available rocket propellant tank designs highlight the need for innovative approaches to enhance safety, performance, and efficiency. An optimized rocket propellant tank layout that addresses structural weaknesses, prevents sloshing and maximizes space utilization may revolutionize large-diameter rocket designs. Such advancements can significantly improve the performance and reliability of rockets destined for deep-space missions to the Moon, Mars, and beyond, offering greater safety and efficiency in transporting humans and cargo. The development of modular rocket propulsion units and launch vehicles tailored to address these challenges represents a promising avenue for overcoming the limitations associated with conventional rocket propellant tank designs.

[0020] Therefore, there is a need for a modular rocket propulsion unit and a rocket launch vehicle that will address the limitations as mentioned above.SUMMARY OF THE INVENTION

[0021] In accordance with an embodiment, a rocket propellant tank and layout is disclosed. The rocket propellant tank includes a housing operationally coupled to a rocket engine. The housing includes a plurality of compartments arranged longitudinally along the housing. Each of the plurality of compartments is adapted to store a propellant. Further, the housing includes a plurality of separator walls disposed adjacent to each of the plurality of compartments. Each of the plurality of separator walls is adapted to define each of the plurality of compartments. Further, the housing includes a set of control valves coupled to each of the plurality of compartments. The set of control valves are adapted to facilitate a controlled transfer of the propellant from each of the plurality of compartments to the rocket engine.

[0022] In an embodiment, each of the plurality of separator walls is adapted to balance pressure of the plurality of compartments and is adapted to serve as a thermal conductor between adjacent compartments of the plurality of compartments to reduce temperature differentials.

[0023] In an embodiment, each of the plurality of compartments is integrated with a baffle structure formed by each of the plurality of separator walls.

[0024] In an embodiment, the set of control valves comprise pressure-regulated valves and reinforced feeder lines adapted to operate under varying pressure ranges from 1 bar to 20 bar.

[0025] In an embodiment, the plurality of compartments includes one or more oxidizer compartments to store liquid oxidizer and one or more fuel compartments to store liquid fuel.

[0026] In an embodiment, an outer wall of each of the plurality of compartments is cylindrical or conical frustum, and an inner wall the plurality of compartments is cylindrical, conical frustum, or any combination thereof.

[0027] In an embodiment, the diameter of the rocket propellant tank is in a range of 3 to 20 meters.

[0028] In accordance with another embodiment, a modular rocket propulsion unit is disclosed. The unit includes a plurality of tanks detachably coupled to each other via a detachable mechanism. Each of the plurality of tanks includes a housing operationally coupled to a rocket engine. The housing includes a plurality of compartments arranged longitudinally along the housing. Each of the plurality of compartments is adapted to store a propellant. Further, the housing includes a plurality of separator walls disposed adjacent to each of the plurality of compartments. Each of the plurality of separator walls is adapted to define each of the plurality of compartments. Further, the housing includes a set of control valves coupled to each of the plurality of compartments. The set of control valves are adapted to facilitate a controlled transfer of the propellant from each of the plurality of compartments to the rocket engine.

[0029] In accordance with another embodiment, a rocket launch vehicle is disclosed. The vehicle includes a payload, and a plurality of tanks detachablycoupled to the payload via a payload separator ejection mechanism. Each of the plurality of tanks includes a housing operationally coupled to a rocket engine. The housing includes a plurality of compartments arranged longitudinally along the housing. Each of the plurality of compartments is adapted to store a propellant. Further, the housing includes a plurality of separator walls disposed adjacent to each of the plurality of compartments. Each of the plurality of separator walls is adapted to define each of the plurality of compartments. Further, the housing includes a set of control valves coupled to each of the plurality of compartments. The set of control valves are adapted to facilitate a controlled transfer of the propellant from each of the plurality of compartments to the rocket engine. Further, the vehicle includes at least one combustion chamber, coupled to each of the plurality of tanks for receiving propellants from each of the plurality of compartments. Further, the vehicle includes at least one nozzle coupled to the at least one combustion chamber, for ejecting exhaust gases from the at least one combustion chamber.

[0030] In an embodiment, the vehicle further includes a supporting structure designed to attach additional propulsion modules, payload fairing to create a multi-stage rocket.

[0031] In an embodiment, the set of control valves comprise pressure-regulated valves and reinforced feeder lines adapted to operate under varying pressure ranges from 1 bar to 20 bar.

[0032] In an embodiment, the plurality of compartments includes one or more oxidizer compartments to store liquid oxidizer and one or more fuel compartments to store liquid fuel.

[0033] In accordance with another embodiment, a method for storing and carrying propellants in a rocket is disclosed. The method includes storing a liquid oxidizer in one or more oxidizer compartments of a plurality of compartments. The method further includes storing a liquid fuel in one or more fuel compartments of the plurality of compartments. The one or more oxidizer compartments and the one ormore fuel compartments are positioned adjacent to each other separated by a plurality of separator walls. Each compartment having a set of control valves coupled to each of the plurality of compartments, the set of control valves are adapted to facilitate a controlled transfer of the propellant from each of the plurality of compartments to a rocket engine.

[0034] In an embodiment, each of the plurality of separator walls is adapted to balance pressure of the plurality of compartments and is adapted to serve as a thermal conductor between adjacent compartments of the plurality of compartments to reduce temperature differentials.

[0035] In an embodiment, each of the plurality of compartments is integrated with a baffle structure formed by each of the plurality of separator walls.

[0036] In this respect, before explaining at least one object of the invention in detail, it is to be understood that the invention is not limited in its application to the details of the set of rules and to the arrangements of the various models set forth in the following description or illustrated in the drawings. The invention is capable of other objects and of being practiced and carried out in various ways, according to the needs of that industry. Also, it is to be understood that the phraseology and terminology employed herein are for the purpose of description and should not be regarded as limiting.

[0037] These together with other objects of the invention, along with the various features of novelty which characterize the invention, are pointed out with particularity in the disclosure. For a better understanding of the invention, its operating advantages and the specific objects attained by its uses, reference should be made to the accompanying drawings and descriptive matter in which there are illustrated preferred embodiments of the invention.BRIEF DESCRIPTION OF THE DRAWINGS

[0038] The invention will be better understood and objects other than those set forth above will become apparent when consideration is given to the following detailed description thereof. Such description makes reference to the annexed drawings wherein:

[0039] Figures 1A-1G illustrate schematic vertical cross-sectional views of multiple rocket propulsion modules, in accordance with conventional techniques;

[0040] Figure 2 illustrates a schematic vertical cross-sectional view with major components outlined for a typical rocket propulsion module, in accordance with a conventional technique;

[0041] Figures 3A-3B illustrate schematic vertical cross-sectional views of a rocket propulsion module, in accordance with an example embodiment of the present disclosure;

[0042] Figure 4 illustrates an isometric view of a rocket propellant tank layout depicting a plurality of compartments, in accordance with an example embodiment of the present disclosure;

[0043] Figure 5 illustrates a top view of the rocket propellant tank layout, in accordance with an example embodiment of the present disclosure;

[0044] Figure 6A illustrates an isometric view of the plurality of compartments of the rocket propellant tank layout, in accordance with an example embodiment of the present disclosure;

[0045] Figures 6B-6C illustrate top views of plurality of compartments of the rocket propellant tank layout, in accordance with an example embodiment of the present disclosure;

[0046] Figure 7 illustrates another top view of the rocket propellant tank layout depicting a set of control valves, in accordance with an example embodiment of the present disclosure;

[0047] Figure 8 illustrates an isometric view of the plurality of compartments of the rocket propellant tank layout depicting a plurality of separator walls and the set of control valves, in accordance with an example embodiment of the present disclosure;

[0048] Figure 9 illustrates a schematic view of a rocket launch vehicle, in accordance with an example embodiment of the present disclosure; and

[0049] Figure 10 illustrates a flowchart depicting a method for storing and carrying propellants in a rocket, in accordance with an example embodiment of the present disclosure.DETAILED DESCRIPTION OF THE DRAWINGS

[0050] Reference will now be made in detail to specific embodiments or features, examples of which are illustrated in the accompanying drawings. Wherever possible, corresponding, or similar reference numbers will be used throughout the drawings to refer to the same or corresponding parts. Moreover, references to various elements described herein, are made collectively or individually when there may be more than one element of the same type. However, such references are merely exemplary in nature. It may be noted that any reference to elements in the singular may also be construed to relate to the plural and vice-versa without limiting the scope of the disclosure to the exact number or type of such elements unless set forth explicitly in the appended claims.

[0051] Some embodiments of this disclosure, illustrating all its features, will now be discussed in detail. The words “comprising,” “having,” “containing,” and “including,” and other forms thereof, are intended to be equivalent in meaning and be open-ended in that an item or items following any one of these words is notmeant to be an exhaustive listing of such item or items or meant to be limited to only the listed item or items.

[0052] It must also be noted that as used herein and in the appended claims, the singular forms “a,” “an,” and “the” include plural references unless the context dictates otherwise. Although any systems and methods similar or equivalent to those described herein can be used in the practice or testing of embodiments of the present disclosure, the preferred systems, and methods are now described.

[0053] Embodiments of the present disclosure will be described more fully hereinafter with reference to the accompanying drawings in which like numerals represent like elements throughout the several figures, and in which example embodiments are shown. Embodiments of the present disclosure may, however, be embodied in alternative forms and should not be construed as being limited to the embodiments set forth herein. The examples set forth herein are non-limiting examples and are merely examples among other possible examples.

[0054] Embodiments of the present disclosure will be described below in detail with reference to the accompanying drawings.

[0055] Embodiments of the present disclosure disclose a rocket propellant tank, a modular rocket propulsion unit, a rocket launch vehicle, and a method for storing and carrying propellants in a rocket. The rocket propellant tank is a modular structure constructed from multiple compartments arranged strategically to create more compartments that are utilized for storing propellants. The present disclosure offers an optimized approach for storing propellants, i.e., fuel and oxidizer.

[0056] Figures 3A-3B illustrate schematic vertical cross-sectional view of a rocket propulsion module 14, in accordance with an example embodiment of the present disclosure.

[0057] The rocket propulsion module 14 may be referred to as a modular rocket propulsion unit. The rocket propulsion module 14 may comprise acompartmentalized rocket propellant tank layout 15. The compartmentalized rocket propellant tank layout 15 may include a plurality of compartments 17a, 17b, 17c, 17d, 17e, 17f arranged adjacent to each other surrounding central compartmentalized tanks 21a 21b. The rocket propulsion module 14 may comprise a plurality of separator walls 24 referred to as a common wall 24. The plurality of compartments 17a, 17b, 17c, 17d, 17e, 17f and the central compartmentalized tanks 21a, 21b may be adapted for storing propellant, e.g., oxidizer and fuel. In an embodiment, the common wall 24 may separate both propellants and also act as a structural reinforcement. The common wall 24 may act as baffle sheets to reduce propellant sloshing.

[0058] Further, the compartmentalized rocket propellant tank layout 15 of the rocket propulsion module 14 may be coupled to a rocket engine 20. The rocket engine 20 may include a combustion chamber 16, additional components such as the ignition system, turbo pumps, plumbing 18 and 19, and an exhaust nozzle 23. Further, the combustion chamber 16 may be coupled to the exhaust nozzle 23. In an embodiment, the fuel and oxidizer may be mixed in the combustion chamber 16 and may be ignited to produce high-temperature gases. The exhaust nozzle 23 may be adapted to expel high-temperature gases at high velocity to generate thrust.

[0059] In an embodiment, the plurality of compartments 17a, 17b, 17c, 17d, 17e, 17f may store oxidizers like Liquid Oxygen or UDMH, and the central compartmentalized tanks 21a, 21b may store RP1, Propane, Ethane, Methane, or Hydrazine. Further, each of the plurality of compartments 17a, 17b, 17c, 17d, 17e, 17f may have provisions 22, along with a set of control valves 45, as shown in Figures 7-8, on the bottom / top of each compartment to receive fuel or oxidizer from an upper compartment and eventually delivering the fuel and oxidizer to the rocket engine final feeder line 18, 19.

[0060] In an embodiment, the central compartmentalized tanks 21a, 21b may have a length equivalent to the entire length of the rocket propulsion module 14. The plurality of compartments 17a, 17b, 17c, 17d, 17e, 17f may surround the centralcompartmentalized tanks 21a, 21b. In an embodiment, the number of compartmentalized tanks may vary based on specific application and design requirements. In an exemplary embodiment, larger rocket diameters and greater lengths result in a higher number of compartments.

[0061] In an embodiment, the segregation of the compartmentalized rocket propellant tank layout 15 into plurality of compartments 17a, 17b, 17c, 17d, 17e, 17f configuration may reduce weight of the rocket propulsion module 14 and enhances structural integrity.

[0062] In an exemplary embodiment, the fuel such as UDMH and hydrazine have very similar freezing temperatures, and there is no requirement for insulation on the common walls 24 and the provisions 22. In an embodiment, the rocket propulsion module 14 may significantly reduce wall thickness by elimination of insulation. In another embodiment, other similar propellant combination with matching freezing / melting temperatures include oxygen-methane, oxygen-ethane, oxygen-propane.

[0063] Figure 4 illustrates an isometric view of the rocket propellant tank layout 15 depicting a plurality of compartments, in accordance with an example embodiment of the present disclosure. Figure 5 illustrates a top view of the rocket propellant tank layout 15, in accordance with an example embodiment of the present disclosure.

[0064] The compartmentalized rocket propellant tank layout 15 may also be referred to the rocket propellant tank layout 15, rocket propellant tank and may be used interchangeably.

[0065] In an embodiment, the rocket propellant tank layout 15 may include at least four propellant tanks 27, 28, 29, and 30 with a central tank 29. The central tank 29 is formed by arranging the at least three propellant tanks 27, 28, 30. Further, the plurality of separator walls or common walls 32, 33, 34 may act as separator for fluids between the at least three propellant tanks 27, 28, 30. Further,the outer wall of the tank 29 may be the common wall separating tank 29 with rest of the tanks 27, 28 and 30.

[0066] Figure 6A illustrates an isometric view of the plurality of compartments 36,37, 38, 39, 40, 36B, 37B, 38B, 39B, and 40B of the rocket propellant tank layout 15, in accordance with an example embodiment of the present disclosure. Figures 6B-6C illustrate top views of plurality of compartments 36, 37, 38, 39, and 40 of the rocket propellant tank layout, in accordance with an example embodiment of the present disclosure.

[0067] In the illustrated embodiment, each of the plurality of compartments 36, 37,38, 39, 40, 36B, 37B, 38B, 39B, and 40B may be selectively and strategically arranged. The arrangement of the each of the plurality of compartments 36, 37, 38,39, 40, 36B, 37B, 38B, and 39B may form central compartments 40, 40B arranged longitudinally. For instance, at least four compartments 36, 37, 38, 39 may be arranged longitudinally, and a fifth tank 40 may be formed at the center the at least four compartments 36, 37, 38, 39.

[0068] Figure 7 illustrates another top view of the rocket propellant tank layout 15 depicting the set of control valves 45, in accordance with an example embodiment of the present disclosure. Figure 8 illustrates an isometric view of wherein compartment 38 is sliced out or removed to show plurality of compartments 36,37, 39, 36B, 37B, 38B, 39B, and 40B of the rocket propellant tank layout 15 depicting the plurality of separator walls 41, 42 and the set of control valves 45, in accordance with an example embodiment of the present disclosure.

[0069] The rocket propellant tank 15 may include a housing operationally coupled to the rocket engine 20. The housing includes the plurality of compartments 36, 37,38, 39, 40, 36B, 37B, 38B, 39B, and 40B arranged longitudinally along the housing, as discussed earlier. Each of the plurality of compartments 36, 37, 38, 39,40, 36B, 37B, 38B, 39B, and 40B may be adapted to store the propellant. Further, the housing includes the plurality of separator walls 33, 34, 35, 41, 42 disposed adjacent to each of the plurality of compartments 36, 37, 38, 39, 40, 36B, 37B,38B, 39B, and 40B. Each of the plurality of separator walls 41, 42 may be adapted to define each of the plurality of compartments 36, 37, 38, 39, 40, 36B, 37B, 38B, 39B, and 40B. Further, the housing includes the set of control valves 45 coupled to each of the plurality of compartments 36, 37, 38, 39, 40, 36B, 37B, 38B, 39B, and 40B. The set of control valves 45 may be adapted to facilitate a controlled transfer of the propellant from each of the plurality of compartments 36, 37, 38, 39, 40, 36B, 37B, 38B, 39B, and 40B to the rocket engine 20.

[0070] In an embodiment, each of the plurality of separator walls 41, 42 adapted to balance pressure of the plurality of compartments 36, 37, 38, 39, 40, 36B, 37B, 38B, 39B, and 40B. Further, each of the plurality of separator walls 41, 42 may be adapted to serve as a thermal conductor between adjacent compartments of the plurality of compartments 36, 37, 38, 39, 40, 36B, 37B, 38B, 39B, and 40B to reduce temperature differentials.

[0071] In an embodiment, each of the plurality of compartments 36, 37, 38, 39, 40, 36B, 37B, 38B, 39B, and 40B may be integrated with a baffle structure formed by each of the plurality of separator walls 41, 42.

[0072] In an embodiment, the set of control valves 45 may be pressure-regulated valves and reinforced feeder lines 18, 19 adapted to operate under varying pressure ranges from 1 bar to 20 bar.

[0073] In an embodiment, the plurality of compartments 36, 37, 38, 39, 40, 36B, 37B, 38B, 39B, and 40B may include one or more oxidizer compartments to store liquid oxidizer and one or more fuel compartments to store liquid .

[0074] In an embodiment, an outer wall 31, 43 of each of the plurality of compartments 36, 37, 38, 39, 40, 36B, 37B, 38B, 39B, and 40B may be cylindrical or conical frustum. An inner wall the plurality of compartments 36, 37, 38, 39, 40, 36B, 37B, 38B, 39B, and 40B may be cylindrical, conical frustum, or any combination thereof. In an embodiment, the diameter of the rocket propellant tank 15 may be in a range of 3 to 20 meters.

[0075] In an exemplary embodiment, a modular rocket propulsion unit is disclosed. The modular rocket propulsion unit may be constructed from a plurality of tanks detachably coupled to each other via a detachable mechanism. Each of the plurality of tanks includes the housing operationally coupled to the rocket engine 20, as described earlier. In an embodiment, each tank may be referred to as the rocket propellant tank 15.

[0076] Figure 9 illustrates a schematic view of a rocket launch vehicle 70, in accordance with an example embodiment of the present disclosure.

[0077] The vehicle 70 may include a payload 61 with payload fairing 62 and a plurality of tanks 15 detachably coupled to the payload 61 via a payload separator ejection mechanism 60. In an embodiment, each of the plurality of tanks 15 may include the housing operationally coupled to the rocket engine 20, as described earlier.

[0078] Further, the vehicle 70 includes at least one combustion chamber 16, 46, coupled to each of the plurality of tanks 15 for receiving propellants from each of the plurality of compartments 59, 68, 72, 58, 69, 57, 79 of stage S2 of the vehicle and the plurality of compartments 64, 67, 63, 65, 66, 51, 49, 50, of stage SI. Further, the vehicle 70 includes the at least one nozzle 23 coupled to the at least one combustion chamber 16, 46, for ejecting exhaust gases from the at least one combustion chamber 16, 46.

[0079] In an embodiment, the vehicle 70 may include a supporting structure designed to attach additional propulsion modules, payload fairing 62 to create a multi-stage rocket. In an embodiment, the plurality of compartments 59, 68, 72, 58, 69, 57, 79 includes one or more oxidizer compartments to store liquid oxidizer and one or more fuel compartments to store liquid .

[0080] The rocket launch vehicle 70 a non-reusable rocket comprising two stages. The upper portion of the rocket is equipped with the payload fairing 62, which houses the payload 61, representing a spacecraft intended for orbital deploymentaround Earth. The payload separator ejection mechanism 60 may be activated to detach the payload 61 from the second stage S2.

[0081] The second stage S2 may include seven compartmentalized tanks as 59, 68, 58, and 72, and inner compartmentalized tanks 69, 57, and 79. The stage S2 may be equipped with a vacuum-optimized rocket engine featuring a combustion chamber 55, a nozzle 53, and feeder lines or turbopumps 56. Additionally, the second stage includes a stage separator system 54, to facilitate its separation from the first stage S 1.

[0082] The first stage SI of the vehicle 70 may include eight compartmentalized tanks 64, 67, 63, 65, 66, 51, 49, 50. In one example, six compartments are outer tanks 64, 67, 63, 66, 51, 49, and two compartments are inner tanks 50, 65. The engines may be mounted on a thrust framework 48, and consist of three engines supported by propellant feeder lines 47.

[0083] In an embodiment, the first stage SI may include the propellant tank layout 14, which may adopt any of the configurations described earlier. The propellant tank layout 14 may be equipped with at least three engines and each includes the at least one combustion chamber 16 and the at least one nozzle 23. The first stage SI includes multiple turbopumps and pump feed lines 18 and 19, respectively. The internal pressure within the fuel tank of the first-stage SI the propellant tank layout 14 may range from 1 to 20 bar. The use of turbopumps allows for moderate fuel storage pressures, thereby reducing the structural mass of the tank and enhancing overall efficiency.

[0084] The second-stage S2 may include a second-stage engine similar to that of the first stage SI. The engine may be encased within an interstage adapter 54, to improve the aerodynamic performance of the rocket system. The internal pressure within the fuel tank of the second-stage S2 typically ranges from 5 to 30 bar.

[0085] The operation of the vehicle 70 may include several phases of flight. The mission begins with the activation of the first-stage rocket engines, providingthrust to power the rocket through the initial ascent phase. During this phase, the rocket utilizes fuel and oxidizer stored in the compartments 64, 67, 63, 65, 66, 51, 49, 50, to generate sufficient thrust to overcome Earth’s gravity and propel the vehicle through the lower atmosphere.

[0086] A second step is first stage separation. Upon depletion of the first-stage fuel and oxidizer, the first-stage SI is shut down, and the first stage is ejected through an ejection mechanism 52. The separation process is initiated by explosive bolts or mechanical latches, enabling the first stage to detach and either fall away or, in the case of a reusable rocket stage, land safely for recovery.

[0087] A third step is ignition of the second stage S2. Following the separation of the first stage SI, the second-stage S2, featuring a single engine, is ignited. The second stage S2 may rely on fuel and oxidizer stored in the compartments 59, 68, 72, 58, 69, 57, 79s, to continue propelling the vehicle 70 through the upper atmosphere and into space.

[0088] A fourth step is separation of the payload fairing 62. The deployable aerodynamic fairing 62, is jettisoned at this step. The fairing serves to shield the payload 61 from aerodynamic forces, heating, and pressure during ascent.

[0089] A fifth step is jettisoning. Upon completion of the second-stage mission, the propulsion of the vehicle 70 is shut down. At this point, the second stage S2 is discarded in a manner similar to the first stage SI, leaving only the payload 61 in orbit.

[0090] Figure 10 illustrates a flowchart depicting a method 1000 for storing and carrying propellants in a rocket, in accordance with an example embodiment of the present disclosure

[0091] At step 1002, the method 1000 includes storing the liquid oxidizer in one or more oxidizer compartments of the plurality of compartments 36, 37, 38, 39, 40, 36B, 37B, 38B, 39B, and 40B.

[0092] At step 1004, the method 1000 includes storing a liquid fuel in one or more fuel compartments of the plurality of compartments 36, 37, 38, 39, 40, 36B, 37B, 38B, 39B, and 40B. The one or more oxidizer compartments and the one or more fuel compartments are positioned adjacent to each other separated by the plurality of separator walls 41, 42.

[0093] The present invention offers several advantages, including enhanced efficiency and performance through the use of compartmentalized propellant tank systems, which allow for optimized fuel storage and reduced structural mass.

[0094] Space Utilization Efficiency: The present disclosure incorporates a compartmentalized configuration wherein multiple propellants are stored in adjacent, separate tanks within the same rocket stage. This arrangement optimally utilizes the available space, facilitating the design and implementation of large- diameter rocket configurations, which are challenging to achieve with conventional prior art designs.

[0095] Enhanced Structural Integrity: The compartmentalized tank design employs shared common walls between adjacent tanks, resulting in improved structural rigidity. The inner tank serves as a structural support for the outer tank, thereby reducing the overall weight and enhancing the structural efficiency of the rocket stage.

[0096] Thermal Isolation Features: The present disclosure provides exemplary embodiments enabling both insulated and non-insulated configurations for the shared walls between tanks. This flexibility is tailored to accommodate the specific boiling and freezing points of the respective propellant and oxidizer, ensuring optimal thermal management.

[0097] Thermally Independent Fluid System Design: The present disclosure is not required to function as a single fluid system, particularly in the context of large- diameter rockets, where such integration offers minimal advantages. Instead, theinvention separates fluids into distinct tanks within a compartmentalized layout, yielding significant practical benefits, especially for thermal management.

[0098] In one embodiment, the coldest fluid is stored in the outer tanks, with the tank walls exposed to the atmosphere insulated to minimize heat transfer. This arrangement reduces the need for active thermal conditioning of the central tanks, as the surrounding cooler fluid serves as a natural thermal buffer, shielding the central tanks from external atmospheric temperatures.

[0099] In an alternative configuration, the coldest propellant is stored in the central compartmentalized tank, with the warmer propellant occupying the outer tanks. This design leverages the higher temperature of the surrounding warmer fluid to insulate the central tank, reducing the thermal gradient experienced by the coldest fluid. Consequently, heat transfer is minimized, and the desired temperature of the coldest fluid is maintained more efficiently.

[0100] The compartmentalized tank design offers significant advantages over traditional stacked tank configurations, where each tank is directly exposed to atmospheric temperature and thus more vulnerable to external thermal fluctuations.

[0101] In contrast, the compartmentalized layout incorporates a deliberate thermal management strategy. By strategically positioning fluids, either in the inner or outer tanks, engineers may optimize insulation, regulate heat transfer, and maintain the required thermal conditions for each propellant. This design flexibility enhances thermal control and overall system performance, particularly in large-diameter rocket applications where the scale and complexity of thermal management present heightened challenges.

[0102] Configuration Flexibility and Thermal Management: The selection of whether the coldest fluid is stored in the inner or outer tanks is dictated by mission-specific requirements, the thermal properties of the propellants, and the overarching design objectives of the rocket. Irrespective of this choice, the compartmentalized tank layout offers a substantial enhancement in thermalmanagement compared to conventional designs, thereby increasing the reliability and efficiency of the rocket's propulsion system.

[0103] Mass Reduction: The present disclosure incorporates outer compartmentalized tanks surrounding inner compartmentalized tanks, separated by a shared wall. This shared wall not only segregates the fuel and oxidizer but also provides structural reinforcement, facilitating a lightweight construction of the propellant tank system. Rockets designed with this innovation significantly improve payload capacity by enhancing overall efficiency and minimizing the dead weight of the launch system. By reducing the need for extensive structural reinforcements, as seen in conventional vertically stacked designs, this configuration achieves considerable mass savings.

[0104] Enhanced Stability Through Reduced Fuel Sloshing: The internal compartments in the tank layout effectively mitigate liquid propellant sloshing during flight, contributing to improved rocket stability and control. This design is particularly advantageous during critical flight phases, such as Max-Q during ascent or Earth-return manoeuvres in reusable rockets, ensuring precise flight dynamics and operational reliability.

[0105] These features collectively contribute to a highly reliable and efficient rocket launch.

[0106] The benefits and advantages which may be provided by the present invention have been described above with regard to specific embodiments. These benefits and advantages, and any elements or limitations that may cause them to occur or to become more pronounced are not to be construed as critical, required, or essential features of any or all of the embodiments.

[0107] While the present invention has been described with reference to particular embodiments, it should be understood that the embodiments are illustrative and that the scope of the invention is not limited to these embodiments. Many variations, modifications, additions, and improvements to the embodimentsdescribed above are possible. It is contemplated that these variations, modifications, additions, and improvements fall within the scope of the invention.

Claims

We Claim:

1. A rocket propellant tank and layout (15), comprising: a housing operationally coupled to a rocket engine (20), wherein the housing comprises: a plurality of compartments (17a, 17b, 17c, 17d, 17e, 17f) arranged longitudinally along the housing, wherein each of the plurality of compartments (17a, 17b, 17c, 17d, 17e, 17f) is adapted to store a propellant; a plurality of separator walls (24, 41, 42, 32, 33, 34) disposed adjacent to each of the plurality of compartments (17a, 17b, 17c, 17d, 17e, 17f), wherein each of the plurality of separator walls (24, 41, 42, 32, 33, 34) is adapted to define each of the plurality of compartments (17a, 17b, 17c, 17d, 17e, 17f); and a set of control valves (45) coupled to each of the plurality of compartments (17a, 17b, 17c, 17d, 17e, 17f), wherein the set of control valves (45) are adapted to facilitate a controlled transfer of the propellant from each of the plurality of compartments (17a, 17b, 17c, 17d, 17e, 17f) to the rocket engine (20).

2. The rocket propellant tank and layout (15) as claimed in claim 1, wherein each of the plurality of separator walls (24, 41, 42, 32, 33, 34) is adapted to balance pressure of the plurality of compartments (17a, 17b, 17c, 17d, 17e, 17f) and is adapted to serve as a thermal conductor between adjacent compartments of the plurality of compartments (17a, 17b, 17c, 17d, 17e, 17f) to reduce temperature differentials.

3. The rocket propellant tank and layout (15) as claimed in claim 1, wherein each of the plurality of compartments (17a, 17b, 17c, 17d, 17e, 17f) is integrated with a baffle structure formed by each of the plurality of separator walls (24, 41, 42, 32, 33, 34).

4. The rocket propellant tank and layout (15) as claimed in claim 1, wherein the set of control valves (45) comprise pressure-regulated valves and reinforced feeder lines (18, 19) adapted to operate under varying pressure ranges from 1 bar to 20 bar.

5. The rocket propellant tank and layout (15) as claimed in claim 1, wherein the plurality of compartments (17a, 17b, 17c, 17d, 17e, 17f) includes one or more oxidizer compartments to store liquid oxidizer and one or more fuel compartments to store liquid .

6. The rocket propellant tank and layout (15) as claimed in claim 1, wherein an outer wall (31, 43) of each of the plurality of compartments (17a, 17b, 17c, 17d, 17e, 17f) is cylindrical or conical frustum, and an inner wall the plurality of compartments (17a, 17b, 17c, 17d, 17e, 17f) is cylindrical, conical frustum, or any combination thereof.

7. The rocket propellant tank and layout (15) as claimed in claim 1, wherein the diameter of the rocket propellant tank (15) is in a range of 3 to 20 meters.

8. A modular rocket propulsion unit (14), comprising: a plurality of tanks detachably coupled to each other via a detachable mechanism, wherein each of the plurality of tanks comprise: a housing operationally coupled to a rocket engine (20), wherein the housing having: a plurality of compartments (17a, 17b, 17c, 17d, 17e, 17f) arranged longitudinally along the housing, wherein each of the plurality of compartments (17a, 17b, 17c, 17d, 17e, 17f) is adapted to store a propellant;a plurality of separator walls (24, 41, 42, 32, 33, 34) disposed adjacent to each of the plurality of compartments (17a, 17b, 17c, 17d, 17e, 17f), wherein each of the plurality of separator walls (24, 41, 42, 32, 33, 34) is adapted to define each of the plurality of compartments (17a, 17b, 17c, 17d, 17e, 17f); and a set of control valves (45) coupled to each of the plurality of compartments (17a, 17b, 17c, 17d, 17e, 17f), wherein the set of control valves (45) are adapted to facilitate a controlled transfer of the propellant from each of the plurality of compartments (17a, 17b, 17c, 17d, 17e, 17f) to the rocket engine (20).

9. A rocket launch vehicle (70) comprising: a payload (61); a plurality of tanks detachably coupled to the payload (61) via a payload separator ejection mechanism (60), wherein each of the plurality of tanks comprise: a housing operationally coupled to a rocket engine (20), wherein the housing having: a plurality of compartments (17a, 17b, 17c, 17d, 17e, 17f) arranged longitudinally along the housing, wherein each of the plurality of compartments (17a, 17b, 17c, 17d, 17e, 17f) is adapted to store a propellant; a plurality of separator walls (24, 41, 42, 32, 33, 34) disposed adjacent to each of the plurality of compartments (17a, 17b, 17c, 17d, 17e, 17f), wherein each of the plurality of separator walls (24, 41, 42, 32, 33, 34) is adapted to define each of the plurality of compartments (17a, 17b, 17c, 17d, 17e, 17f); anda set of control valves (45) coupled to each of the plurality of compartments (17a, 17b, 17c, 17d, 17e, 17f), wherein the set of control valves (45) are adapted to facilitate a controlled transfer of the propellant from each of the plurality of compartments (17a, 17b, 17c, 17d, 17e, 17f) to the rocket engine (20); at least one combustion chamber (16), coupled to each of the plurality of tanks for receiving propellants from each of the plurality of compartments (17a, 17b, 17c, 17d, 17e, 17f); and at least one nozzle (23) coupled to the at least one combustion chamber (16), for ejecting exhaust gases from the at least one combustion chamber (16).

10. The rocket launch vehicle (70) as claimed in claim 9, comprising: a supporting structure designed to attach additional propulsion modules, payload fairing (62) to create a multi-stage rocket.

11. The rocket launch vehicle (70) as claimed in claim 9, wherein the set of control valves (45) comprise pressure-regulated valves and reinforced feeder lines (18, 19) adapted to operate under varying pressure ranges from 1 bar to 20 bar.

12. The rocket launch vehicle (70) as claimed in claim 9, wherein the plurality of compartments (17a, 17b, 17c, 17d, 17e, 17f) includes one or more oxidizer compartments to store liquid oxidizer and one or more fuel compartments to store liquid .

13. A method (1000) for storing and carrying propellants in a rocket, comprising:storing, (1002) a liquid oxidizer in one or more oxidizer compartments of a plurality of compartments (17a, ,17b, 17c, 17d, 17e, 17f); and storing, (1004) a liquid fuel in one or more fuel compartments of the plurality of compartments (17a, 17b, 17c, 17d, 17e, 17f), wherein the one or more oxidizer compartments and the one or more fuel compartments are positioned adjacent to each other separated by a plurality of separator walls (24, 41, 42, 32, 33, 34), wherein each compartment having a set of control valves (45) coupled to each of the plurality of compartments (17a, 17b, 17c, 17d, 17e, 17f), the set of control valves (45) are adapted to facilitate a controlled transfer of the propellant from each of the plurality of compartments (17a, 17b, 17c, 17d, 17e, 17f) to a rocket engine (20).

14. The method (1000) as claimed in claim 14, wherein each of the plurality of separator walls (24, 41, 42, 32, 33, 34) is adapted to balance pressure of the plurality of compartments (17a, 17b, 17c, 17d, 17e, 17f) and is adapted to serve as a thermal conductor between adjacent compartments of the plurality of compartments (17a, 17b, 17c, 17d, 17e, 17f) to reduce temperature differentials.

15. The method (1000) as claimed in claim 14, wherein each of the plurality of compartments (17a, 17b, 17c, 17d, 17e, 17f) is integrated with a baffle structure formed by each of the plurality of separator walls (24, 41, 42, 32, 33, 34).

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