A blended wing body aircraft with a fuel cell and method of use

Abstract

A system for a blended wing body aircraft with a combustion engine is illustrated. The aircraft comprises a blended wing body, at least a fuel store located within the blended wing body and configured to store a fuel, wherein the fuel includes liquid hydrogen, and at least a propulsor 5 configured to propel the blended wing body aircraft. The at least a propulsor comprises a combustion engine configured to burn the fuel from the fuel store and produce mechanical work to power the at least a propulsor.

Description

A BLENDED WTNG BODY AIRCRAFT WITH A FUEL CELL AND METHOD OF USECROSS-REFERENCE TO RELATED APPLICATIONSThis application claims the benefit of priority of U.S. Nonprovisional Application Serial No. 17 / 478,724, filed on September 17, 2021 and entitled “A BLENDED WING BODY AIRCRAFT WITH A FUEL CELL AND METHOD OF USE,” U.S. Nonprovisional Application Serial No. 17 / 731,622, filed on April 28, 2022, and entitled “BLENDED WING BODY AIRCRAFT WITH A COMBUSTION ENGINE AND METHOD OF USE,” U.S. Nonprovisional Application Serial No. 17 / 841,153, filed on June 15, 2022 and entitled “A METHOD FOR FUELING AN AIRCRAFT,” U.S. Nonprovisional Application Serial No. 17 / 731,655, filed on April 28, 2022 and entitled “SYSTEMS AND METHODS FOR A BLENDED WING BODY AIRCRAFT WITH PERMANENT TANKS,” U.S. Nonprovisional Application Serial No. 17 / 731,728, filed on April 28, 2022 and entitled “FAN AIRCRAFT WITH FUEL TANKS STORED AFT OF A CABIN IN A MAIN BODY AND A METHOD FOR MANUFACTURING,” U.S. Nonprovisional Application Serial No. No. 17 / 867,851, filed on July 19, 2022, and entitled “METHODS AND SYSTEMS FOR VENTING A FUEL TANK ON A BLENDED WING BODY AIRCRAFT,” U.S. Nonprovisional Application Serial No. 17 / 731,685, filed on April 28, 2022, and entitled “AN AIRCRAFT FUELING APPARATUS AND A METHOD FOR ITS USE,” U.S. Nonprovisional Application Serial No. 17 / 731,772, filed on April 28, 2022, and entitled “AN AIRCRAFT WITH A MULTI-WALLED FUEL TANK AND A METHOD OF MANUFACTURING,” which are incorporated by reference herein in its entirety.FIELD OF THE INVENTIONThe present invention generally relates to the field of aircraft. In particular, the present invention is directed to a blended wing body aircraft with a fuel cell and method of use. BACKGROUNDHuman flight is a large contributor of greenhouse gases, the effects of which are compounded by their release high in the atmosphere. However, non-greenhouse gas generating energy storage methods are less energy dense, according to one or both of volumetric energy density and weight energy density. Presently, current aircraft designs are tightly constrained in both storage volume and weight. Hydrogen fuel has characteristics favorable as an aviation fuel. Storage of hydrogen fuel within an aircraft is limited by the location of the fuel reserves.SUMMARY OF THE DISCLOSUREIn an aspect a blended wing body aircraft with a fuel cell includes a blended wing body, at least a first fuel store located within a transitional portion of the blended wing body and configured to store a first fuel, at least a fuel cell configured to combine the first fuel with oxygen to produce electricity, at least a second fuel store located within a wing portion of the blended wing body and configured to store a second fuel, and at least a propulsor mechanically affixed to the aircraft and configured to propel the blended wing body aircraft.In another aspect a method of use of a blended wing body aircraft with a fuel cell includes storing a first fuel, using at least a first fuel store located within a transitional portion of a blended wing body of the blended wing body aircraft, combining the first fuel with oxygen to produce electricity, using at least a fuel cell, storing a second fuel, using at least a second fuel store located within a wing portion of the blended wing body, and propelling the aircraft, using at least a propulsor mechanically affixed to the blended wing body aircraft.In an aspect, a system for a blended wing body aircraft with a combustion engine is illustrated. The aircraft comprises a blended wing body, at least a fuel source located within the blended wing body and configured to store a fuel, wherein the fuel includes liquid hydrogen, at least a propulsor configured to propel the blended wing body aircraft and comprising a combustion engine, and at least an auxiliary power unit powered by the fuel and mechanically affixed to the aircraft. The combustion engine is configured to bum the fuel from the fuel source and produce mechanical work to use to power the at least a propulsor.In another aspect, a method of use of a system for a blended wing body aircraft with a combustion engine is shown. The method comprises storing a fuel using at least a fuel source located within a blended wing body of the blended wing body aircraft, propelling the aircraft, using at least a propulsor mechanically affixed to the blended wing body aircraft and comprising a combustion engine, burning, at the combustion engine, the fuel from the fuel source and producing, at the combustion engine, mechanical work to use to power the at least a propulsor.In another aspect, a method for fueling an aircraft, the method including storing liquified gas fuel using a fuel tank, wherein the fuel tank is configured to store liquified gas fuel, fueling an aircraft using a fuel line. Fueling the aircraft may additionally include the fuel tank with liquified gas fuel to a desired level, wherein a desired level comprises fuel for a plurality offlights plus reserves. Filling may also include removing the fuel line as a function of the desired level. The method may additionally include venting the fuel tank using a vent line in fluid connection to the fuel tank. The fuel tank may then be prepared for flight as a function of a desired level. Finally, the method includes flying a plurality of flights using the aircraft.In an aspect, a system for a blended wing body aircraft with permanent tanks is shown. The aircraft comprises a blended wing body and at least a tank permanently attached the blended wing body. The at least a tank is configured to store liquified gas fuel and further comprises at least a vent configured to vent gaseous fuel from the at least a tank and an insulation to reduce thermal transfer to the liquified gas fuel inside of the at least a tank.In another aspect, a method of use for a blended wing body aircraft with permanent tanks is illustrated. The method of use comprises permanently attaching at least a tank into a blended wing body of the blended wing body aircraft, storing liquified gas fuel inside the at least a tank, burning the liquified gas fuel to power the blended wing body aircraft, venting gaseous fuel from the at least a tank using at least a vent, and reducing thermal transfer to the liquified gas fuel inside of the at least a tank using an insulation.In an aspect, an aircraft with fuel tanks stored aft of a cabin in a main body including a blended wing body, comprising: a main body, and a cabin at least partially located within the main body, and a plurality of fuel tanks located at least partially aft of the cabin within the main body and configured to store liquified gas fuel.In another aspect, a method of manufacturing fuel tanks stored aft of the main body of an aircraft including receiving a blended wing body, comprising: a main body; and a cabin at least partially located within the main body, receiving a plurality of fuel tanks, locating the plurality of fuel tanks at least partially aft of the cabin within the main body, and storing, using the plurality of fuel tanks, liquified gas fuel.In an aspect, a liquified gas fuel tank for an aircraft includes a first compartment wherein the first compartment has a first cross-section describing a first continuously convex differentiable curve, the first compartment includes an inner volume configured to contain fuel, and the first compartment is configured to be pressurized, a second compartment wherein the second compartment has a second cross-section describing a second continuously convex differentiable curve which intersects the first continuously convex differentiable curve at anintersection, and the second compartment includes an additional inner volume fluidly connected to the inner volume of the first compartment, and a junction configured to structurally support each of the first compartment and the second compartment at the intersection.In another aspect, a method of manufacturing a liquified gas fuel tank for an aircraft includes constructing a first compartment wherein the first compartment has a first cross-section describing a first continuously convex differentiable curve, the first compartment includes an inner volume configured to contain fuel, and the first compartment is configured to be pressurized, constructing a second compartment wherein the second compartment has a second cross-section describing a second continuously convex differentiable curve which intersects the first continuously convex differentiable curve at an intersection, and the second compartment includes an additional inner volume fluidly connected to the inner volume of the first compartment, and connecting the first compartment and the second compartment with a junction configured to structurally support each of the first compartment and the second compartment at the intersection.In an aspect a method for venting a fuel tank includes connecting a vent line to a fuel tank in a grounded blended wing body (BWB) aircraft, wherein the fuel tank contains liquified gas fuel, heating the vent line, and collecting the gaseous fuel in an external fuel tank.In another aspect a system for venting a fuel tank includes a grounded blended wing body aircraft, a vent line connected to a fuel tank in the grounded BWB aircraft, and an external fuel tank configured to collect the gaseous fuel boil-off from the fuel tank in the grounded BWB.In an aspect, an aircraft fueling apparatus is disclosed. The apparatus includes at least a container comprising a fuel tank configured to store liquified gas fuel. The apparatus may also include a translocation device configured to carry the at least a container. An orientation guidance track may also be included in the apparatus. The orientation guidance track may be configured to direct a movement of the translocation device to a first positionIn another aspect, method of use for an aircraft fueling apparatus is shown. The method may include storing, using at least a container comprising a fuel tank configured to store liquified gas fuel. The method may also include using a translocation device configured to carry the at least a container. Using an orientation guidance track, the method may direct a movement of the translocation device to a first positionIn an aspect an aircraft with at least a multi-walled fuel tank includes a blended wing body, and at least a fuel tank attached to the blended wing body and configured to store liquified gas fuel, wherein the at least a fuel tank further comprises: an inner wall, an outer wall, an interstitial volume between the inner wall and the outer wall comprising of at least a reflective film layer and at least a structural insulation layer.In another aspect a method of manufacturing at least a multi-walled fuel tank for an aircraft includes receiving a blended wing body, receiving an inner wall, receiving an outer wall, inserting an interstitial volume comprising at least a reflective film layer and at least a structural insulation layer between the inner wall and the outer wall.The details of one or more variations of the subject matter described herein are set forth in the accompanying drawings and the description below. Other features and advantages of the subject matter described herein will be apparent from the description and drawings, and from the claims.DESCRIPTION OF DRAWINGSFor the purpose of illustrating the invention, the drawings show aspects of one or more embodiments of the invention. However, it should be understood that the present invention is not limited to the precise arrangements and instrumentalities shown in the drawings, wherein: FIG. l is a schematic illustration of an exemplary blended wing body aircraft with a fuel cell; FIG. 2 is a diagram of an exemplary blended wing body aircraft with a combustion engine; FIG. 3 A is an exemplary embodiment of at least a tank with mathematical equations to calculate the force and stress of the at least a tank;FIG. 3B is an isometric view of an exemplary embodiment of a conical tank;FIG. 3C is an isometric view of an exemplary embodiment of a curved axisymmetric tank;FIG. 3D is an isometric view of an exemplary embodiment of a double-curved tank;FIG. 3E is a front quarter view of an exemplary embodiment of a dual tank with different diameters;FIG. 3F is a quarter side view of an exemplary embodiment of a cambered, tapered tank;FIG. 3G illustrates an isometric and side quarter views of an exemplary embodiment of a dual cambered tank;FIG. 3H shows a quarter front view of an exemplary embodiment of a double tank;FIG. 31 is a front quarter view of an exemplary embodiment of a multi-bubble tank without trimming;FIG. 3J is a front quarter view of an exemplary embodiment of a multi-bubble tank trimmed with septa;FIG. 3K shows an exemplary embodiment of an overview of a vertical axis “air mattress” tank;FIG. 3L shows an exemplary embodiment of an overview of longitudinal arrangement of a vertical axis “air mattress” tank;FIG. 3M shows an exemplary embodiment of a tank having multiple compartments.FIG. 4 is a block diagram of an exemplary embodiment of a method for fueling an aircraft;FIG. 5 illustrates an isometric view of an exemplary embodiment of a tank with a plurality of tank support links;FIG. 6 illustrates a cross-section of an exemplary multi -chambered tank;FIG. 7 is a block diagram for the exemplary embodiment of a liquified gas fuel tank incorporated in an aircraft;FIG. 8 is a block diagram of a system for venting fuel tanks;FIG. 9 is a schematic of a fuel tank;FIG. 10A and B are schematics of a fuel line;FIG. 11 is a block diagram of an exemplary embodiment of an aircraft fueling apparatus;FIG. 12 is an exemplary depiction of a cross section of a fuel tank;FIG. 13 is a block diagram of a sensing system;FIG. 14 is a flow diagram of a method of use for a blended wing body aircraft with a fuel cell;FIG. 15 is a flow diagram of a method of use of a system for a blended wing body aircraft with a combustion engine;FIG. 16 is a flow diagram of a method of use for fueling an aircraft;FIG. 17 is a flow diagram of a method of use for a blended wing body aircraft with permanent tanks;FIG. 18 is a flow diagram of a method of manufacturing fuel tanks stored aft of the cabin in the main body of an aircraft;FIG. 19 is a flow diagram of a method of manufacturing includes locating the plurality of fuel tanks aft of the cabin within the main body;FIG. 20 is a block diagram for a method of use for an aircraft fueling apparatus;FIG. 21 is a flow diagram of a method for manufacturing a multi -walled fuel tank;FIG. 22 is a flow diagram of a method of manufacturing 2200 for a multi-walled fuel tank for an aircraft;FIG. 23 is a block diagram of a computing system that can be used to implement any one or more of the methodologies disclosed herein and any one or more portions thereof.The drawings are not necessarily to scale and may be illustrated by phantom lines, diagrammatic representations, and fragmentary views. In certain instances, details that are not necessary for an understanding of the embodiments or that render other details difficult to perceive may have been omitted. Like reference symbols in the various drawings indicate like elements.DETAILED DESCRIPTIONAt a high level, aspects of the present disclosure are directed to systems and methods for using a fuel cell within a blended wing body aircraft (i.e., blended wing aircraft). In an embodiment, a blended wing aircraft allows for an increase in volumetric storage space allowing of use of liquid hydrogen fuel, which has less energy per unit volume than conventional aircraft fuel.At a high level, aspects of the present disclosure are directed to systems and methods for a blended wing body aircraft with a combustion engine. Aspects of the present disclosure may include a blended wing body aircraft. Aspects of the present disclosure may include at least a fuel source located within the blended wing body and configured to store a fuel. Aspects of the present disclosure may also include at least a propulsor configured to propel the blended wing body aircraft. Propulsor may comprise a combustion engine, which differs from a fuel cell, as fuel cells do not convert the fuel into heat first. Aspects of the present disclosure may include at least an auxiliary power unit powered by the fuel and mechanically affixed to the aircraft.At a high level, aspects of the present disclosure are directed to systems and methods for fueling an aircraft. The method includes storing liquified gas fuel using a fuel tank, wherein the fuel tank is configured to store liquified gas fuel, fueling an aircraft using a fuel line. Fueling the aircraft may additionally include the fuel tank with liquified gas fuel to a desired level, wherein a desired level comprises fuel for a plurality of flights plus reserves. Filling may also includeremoving the fuel line as a function of the desired level. The method may additionally include venting the fuel tank using a vent line in fluid connection to the fuel tank. The fuel tank may then be prepared for flight as a function of a desired level. Finally, the method includes flying a plurality of flights using the aircraft.At a high level, aspects of the present disclosure include a system for a blended wing body aircraft with permanent tanks. Aspects of the present disclosure include a blended wing body. Aspects of the present disclosure include at least a tank permanently attached the blended wing body. Aspects of the present disclosure include storing liquid hydrogen fuel inside the at least a tank. Aspects of the present disclosure further comprise at least a vent configured to vent gaseous hydrogen from the at least a tank and an insulation to reduce thermal transfer to the liquid hydrogen fuel inside of the at least a tank.At a high level, aspects of the present disclosure include a liquified gas fuel tank. Aspects of the present disclosure include a first compartment. Aspects of the present disclosure include a first compartment having a first cross section describing a continuously convex differentiable curve. Aspects of the present disclosure include a first compartment having a first inner wall defining a first internal cavity of the first compartment and a first outer wall exterior to the first inner wall, and a first separation of the first inner wall and the first outer wall defines a first space. Aspects of the present disclosure include a first compartment having an inner volume containing liquified gas fuel and is pressurized. Aspects of the present disclosure further include a second compartment has a second cross-section describing a continuously convex differentiable curve. Aspects of the present disclosure includes a second compartment having a second inner wall defining a second internal cavity of the second compartment and a second outer wall exterior to the second inner wall, and a second separation of the second inner wall and the second outer wall defines a second space. Aspects of the present disclosure further include the first compartment and the second compartment are at least partially separated from one another. Aspects of the present disclosure include a second compartment having an inner volume containing liquified gas fuel and is pressurized. Aspects of the present disclosure include second compartment includes an additional inner volume containing liquified gas fuel. Moreover, aspects of the present disclosure include a junction, wherein the junction fluidly connects thefirst internal cavity and the second internal cavity a junction fluidly connecting a first internal cavity and a second internal cavity.At a high level, aspects of the present disclosure are directed to systems and methods for venting a fuel tank. In an embodiment, fuel tank is located on a blended wing body aircraft. Fuel tank is pressurized and contains liquified gas fuel. Over time, liquified gas fuel may warm and produce boil-off, located at the top of the fuel tank. A vent located at the top of the fuel tank may be connected to a vent line, configured to vent out the gaseous fuel from the fuel tank to an external fuel tank off of the aircraft.At a high level, aspects of the present disclosure are directed to an aircraft fueling apparatus and a method for its use. The apparatus includes at least a container. The apparatus may also include a translocation device configured to carry the at least a container. An orientation guidance track may also be included in the apparatus. The orientation guidance track may be configured to direct a movement of the translocation device to a first position.Aspects of the present disclosure can be used to power aircraft propulsors using a fuel cell. Aspects of the present disclosure can also be used to power an auxiliary power system using a fuel cell.Aspects of the present disclosure can be used to power aircraft propulsors using a combustion engine. Aspects of the present disclosure can also be used to power an auxiliary power unit using a fuel cell.Aspects of the present disclosure allow for use of non-greenhouse gas emitting fuels to power human flight. Exemplary embodiments illustrating aspects of the present disclosure are described below in the context of several specific examples.Aspects of the present disclosure include a method of use for a blended wing body aircraft with permanent tanks. Aspects of the present disclosure include permanently attaching at least a tank into a blended wing body of the blended wing body aircraft. Aspects of the present disclosure include storing liquid hydrogen fuel inside the at least a tank. Aspects of the present disclosure include burning the liquid hydrogen fuel to power the blended wing body aircraft. Aspects of the present disclosure include venting gaseous hydrogen from the at least a tank using at least a vent. Aspects of the present disclosure include reducing thermal transfer to the liquid hydrogen fuel inside of the at least a tank using an insulation.Aspects of the present disclosure include method of manufacturing a liquified gas fuel tank. Aspects of the present disclosure include constructing a first compartment and a second compartment of the liquified gas fuel tank. Aspects of the present disclosure include forming a first space between a first inner wall and a first outer wall by constructing the first inner wall having a first continuously convex differentiable curve cross section and constructing the first outer wall with the first continuously convex differentiable curve cross section. Aspects of the present disclosure include forming a second space between a second inner wall and a second outer wall by constructing the second inner wall having a second continuously convex differentiable curve cross section and constructing the second outer wall with the second continuously convex differentiable curve cross section. Aspects of the present disclosure include constructing the liquified gas fuel tank comprises connecting the first compartment and the second compartment via a junction, and the junction connects the first space to the second space.Aspects of the present disclosure are only used on a grounded aircraft. A grounded aircraft is an aircraft that is not ready to fly. Grounded aircrafts often occur during maintenance and during periods where the aircraft is not in use.In the following description, for the purposes of explanation, numerous specific details are set forth in order to provide a thorough understanding of the present invention. For purposes of description herein, relating terms, including “upper”, “lower”, “left”, “rear”, “right”, “front”, “vertical”, “horizontal”, and derivatives thereof relate to embodiments oriented as shown for exemplary purposes in FIG. 1. Furthermore, there is no intention to be bound by any expressed or implied theory presented in this disclosure.Referring now to FIG. 1, an exemplary embodiment of a blended wing body aircraft with a fuel cell is illustrated. System includes a computing device, computing device may include any computing device as described in this disclosure, including without limitation a microcontroller, microprocessor, digital signal processor (DSP) and / or system on a chip (SoC) as described in this disclosure. Computing device may include, be included in, and / or communicate with a mobile device such as a mobile telephone or smartphone. Computing device may include a single computing device operating independently, or may include two or more computing device operating in concert, in parallel, sequentially or the like; two or more computing devices may be included together in a single computing device or in two or more computing devices. Computingdevice may interface or communicate with one or more additional devices as described below in further detail via a network interface device. Network interface device may be utilized for connecting computing device to one or more of a variety of networks, and one or more devices. Examples of a network interface device include, but are not limited to, a network interface card (e.g., a mobile network interface card, a LAN card), a modem, and any combination thereof. Examples of a network include, but are not limited to, a wide area network (e.g., the Internet, an enterprise network), a local area network (e.g., a network associated with an office, a building, a campus or other relatively small geographic space), a telephone network, a data network associated with a telephone / voice provider (e.g., a mobile communications provider data and / or voice network), a direct connection between two computing devices, and any combinations thereof. A network may employ a wired and / or a wireless mode of communication. In general, any network topology may be used. Information (e.g., data, software etc.) may be communicated to and / or from a computer and / or a computing device, computing device may include but is not limited to, for example, a computing device or cluster of computing devices in a first location and a second computing device or cluster of computing devices in a second location, computing device may include one or more computing devices dedicated to data storage, security, distribution of traffic for load balancing, and the like, computing device may distribute one or more computing tasks as described below across a plurality of computing devices of computing device, which may operate in parallel, in series, redundantly, or in any other manner used for distribution of tasks or memory between computing devices, computing device may be implemented using a “shared nothing” architecture in which data is cached at the worker, in an embodiment, this may enable scalability of system 100 and / or computing device.With continued reference to FIG. 1, computing device may be designed and / or configured to perform any method, method step, or sequence of method steps in any embodiment described in this disclosure, in any order and with any degree of repetition. For instance, computing device may be configured to perform a single step or sequence repeatedly until a desired or commanded outcome is achieved; repetition of a step or a sequence of steps may be performed iteratively and / or recursively using outputs of previous repetitions as inputs to subsequent repetitions, aggregating inputs and / or outputs of repetitions to produce an aggregate result, reduction or decrement of one or more variables such as global variables, and / or divisionof a larger processing task into a set of iteratively addressed smaller processing tasks, computing device may perform any step or sequence of steps as described in this disclosure in parallel, such as simultaneously and / or substantially simultaneously performing a step two or more times using two or more parallel threads, processor cores, or the like; division of tasks between parallel threads and / or processes may be performed according to any protocol suitable for division of tasks between iterations. Persons skilled in the art, upon reviewing the entirety of this disclosure, will be aware of various ways in which steps, sequences of steps, processing tasks, and / or data may be subdivided, shared, or otherwise dealt with using iteration, recursion, and / or parallel processing.With continued reference to FIG. 1, an exemplary aircraft 100 is illustrated. Aircraft 100 may include a blended wing body 104. For the purposes of this disclosure, a “blended wing body aircraft” is an aircraft having a blended wing body. As used in this disclosure, A “blended wing body” (BWB), also known as a “blended body” or a “hybrid wing body” (HWB), is a fixed-wing aircraft body having no clear demarcation between wings and a main body of the aircraft at a leading edge of the wings. For example, a BWB 104 aircraft may have distinct wing and body structures, which are smoothly blended together with no clear dividing line or boundary feature between wing and fuselage. This contrasts with a flying wing, which has no distinct fuselage, and a lifting body, which has no distinct wings. A BWB 104 design may or may not be tailless. One potential advantage of a BWB 104 may be to reduce wetted area and any accompanying drag associated with a conventional wing-body junction. In some cases, a BWB 104 may also have a wide airfoil-shaped body, allowing entire aircraft to generate lift and thereby facilitate reduction in size and / or drag of wings. In some cases, a BWB 104 may be understood as a hybrid shape that resembles a flying wing, but also incorporates features from conventional aircraft. In some cases, this combination may offer several advantages over conventional tube-and-wing airframes. In some cases, a BWB airframe 104 may help to increase fuel economy and create larger payload (cargo or passenger) volumes within the BWB. BWB 104 may allow for advantageous interior designs. For instance, cargo can be loaded and / or passengers can board from the front or rear of the aircraft. A cargo or passenger area may be distributed across a relatively wide (when compared to conventional tube-wing aircraft) fuselage, providing a largeusable volume. Tn some embodiments, passengers seated within an interior of aircraft, real-time video at every seat can take place of window seats.With continued reference to FIG. 1, BWB 104 of aircraft 100 may include a nose portion. A “nose portion,” for the purposes of this disclosure, refers to any portion of aircraft 100 forward of the aircraft’s fuselage 116. Nose portion may comprise a cockpit (for manned aircraft), canopy, aerodynamic fairings, windshield, and / or any structural elements required to support mechanical loads. Nose portion may also include pilot seats, control interfaces, gages, displays, inceptor sticks, throttle controls, collective pitch controls, and / or communication equipment, to name a few. Nose portion may comprise a swing nose configuration. A swing nose may be characterized by an ability of the nose to move, manually or automatedly, into a differing orientation than its flight orientation to provide an opening for loading a payload into aircraft fuselage from the front of the aircraft. Nose portion may be configured to open in a plurality of orientations and directions.With continued reference to FIG. 1, BWB 104 may include at least a structural component of aircraft 100. Structural components may provide physical stability during an entirety of an aircraft’s 100 flight envelope, while on ground, and during normal operation Structural components may comprise struts, beams, formers, stringers, longerons, interstitials, ribs, structural skin, doublers, straps, spars, or panels, to name a few. Structural components may also comprise pillars. In some cases, for the purpose of aircraft cockpits comprising windows / windshields, pillars may include vertical or near vertical supports around a window configured to provide extra stability around weak points in a vehicle’s structure, such as an opening where a window is installed. Where multiple pillars are disposed in an aircraft’s 100 structure, they may be so named A, B, C, and so on named from nose to tail. Pillars, like any structural element, may be disposed a distance away from each other, along an exterior of aircraft 100 and BWB 104. Depending on manufacturing method of BWB 104, pillars may be integral to frame and skin, comprised entirely of internal framing, or alternatively, may be only integral to structural skin elements. Structural skin will be discussed in greater detail below.With continued reference to FIG. 1, BWB 104 may include a plurality of materials, alone or in combination, in its construction. At least a BWB 104, in an illustrative embodiment may include a welded steel tube frame further configured to form a general shape of a nosecorresponding to an arrangement of steel tubes. Steel may include any of a plurality of alloyed metals, including but not limited to, a varying amount of manganese, nickel, copper, molybdenum, silicon, and / or aluminum, to name a few. Welded steel tubes may be covered in any of a plurality of materials suitable for aircraft skin. Some of these may include carbon fiber, fiberglass panels, cloth-like materials, aluminum sheeting, or the like. BWB 104 may comprise aluminum tubing mechanically coupled in various and orientations. Mechanical fastening of aluminum members (whether pure aluminum or alloys) may comprise temporary or permanent mechanical fasteners appreciable by one of ordinary skill in the art including, but not limited to, screws, nuts and bolts, anchors, clips, welding, brazing, crimping, nails, blind rivets, pull- through rivets, pins, dowels, snap-fits, clamps, and the like. BWB 104 may additionally or alternatively use wood or another suitably strong yet light material for an internal structure.

[0001] With continued reference to FIG. 1, aircraft 100 may include monocoque or semi- monocoque construction. BWB 104 may include carbon fiber. Carbon fiber may include carbon fiber reinforced polymer, carbon fiber reinforced plastic, or carbon fiber reinforced thermoplastic (e.g., CFRP, CRP, CFRTP, carbon composite, or just carbon, depending on industry). “Carbon fiber,” as used in this disclosure, is a composite material including a polymer reinforced with carbon. In general, carbon fiber composites consist of two parts, a matrix and a reinforcement. In carbon fiber reinforced plastic, the carbon fiber constitutes the reinforcement, which provides strength. The matrix can include a polymer resin, such as epoxy, to bind reinforcements together. Such reinforcement achieves an increase in CFRP’s strength and rigidity, measured by stress and elastic modulus, respectively. In embodiments, carbon fibers themselves can each comprise a diameter between 5-10 micrometers and include a high percentage (i. e. , above 85%) of carbon atoms. A person of ordinary skill in the art will appreciate that the advantages of carbon fibers include high stiffness, high tensile strength, low weight, high chemical resistance, high temperature tolerance, and low thermal expansion. According to embodiments, carbon fibers may be combined with other materials to form a composite, when permeated with plastic resin and baked, carbon fiber reinforced polymer becomes extremely rigid. Rigidity may be considered analogous to stiffness which may be measured using Young’s Modulus. Rigidity may be defined in terms of stress and strain, wherein “stress” is force per area and “strain” is an elongation or deviation generally represented as a proportion or fraction of a length or angle. Forexample, the strain of a one-hundred-inch-long rod that is stretched to 101 inches is the one inch of stretch divided by the one-hundred-inch length, or a 1% strain. In the case of shear, strain may be measured by an angular deformation. Rigidity may be considered analogous to stiffness and, for linear displacements, may be quantified as Young’s modulus. Young’s modulus may be defined as stress divided by strain. Rigidity may be defined as a force necessary to bend and / or flex a material and / or structure to a given degree. For example, ceramics have high rigidity, which can be visualized by shattering before bending. In embodiments, carbon fibers may additionally, or alternatively, be composited with other materials like graphite to form reinforced carbon-carbon composites, which include high heat tolerances over 2000°C. A person of skill in the art will further appreciate that aerospace applications may require high-strength, low-weight, high heat resistance materials in a plurality of roles, such as without limitation fuselages, fairings, control surfaces, and structures, among others. Additionally, in some embodiments, composite construction may include three elements: carbon fiber, epoxy resin, and a second fiber. In some cases, second fiber may be a thread that is woven through a ply thickness to prevent delamination. Second fiber may be made of a high tensile strength synthetic material that may provide extra elongation before failure, such as Kevlar and / or an aramid fiber.With continued reference to FIG. 1, BWB 104 may include at least a fuselage. A “fuselage,” for the purposes of this disclosure, refers to a main body of an aircraft 100, or in other words, an entirety of the aircraft 100 except for nose, wings, empennage, nacelles, and control surfaces. In some cases, fuselage may contain an aircraft’s payload. At least a fuselage may comprise structural components that physically support a shape and structure of an aircraft 100. Structural components may take a plurality of forms, alone or in combination with other types. Structural components vary depending on construction type of aircraft 100 and specifically, fuselage. A fuselage 112 may include a truss structure. A truss structure may be used with a lightweight aircraft. A truss structure may include welded steel tube trusses. A “truss,” as used in this disclosure, is an assembly of beams that create a rigid structure, for example without limitation including combinations of triangles to create three-dimensional shapes. A truss structure may include wood construction in place of steel tubes, or a combination thereof. In some embodiments, structural components can comprise steel tubes and / or wood beams. An aircraft skin may be layered over a body shape constructed by trusses. Aircraft skinmay comprise a plurality of materials such as plywood sheets, aluminum, fiberglass, and / or carbon fiber.With continued reference to FIG. 1, in embodiments, at least a fuselage may comprise geodesic construction. Geodesic structural elements may include stringers wound about formers (which may be alternatively called station frames) in opposing spiral directions. A “stringer,” for the purposes of this disclosure is a general structural element that includes a long, thin, and rigid strip of metal or wood that is mechanically coupled to and spans the distance from, station frame to station frame to create an internal skeleton on which to mechanically couple aircraft skin. A former (or station frame) can include a rigid structural element that is disposed along a length of an interior of a fuselage orthogonal to a longitudinal (nose to tail) axis of aircraft 100. In some cases, a former form a general shape of at least a fuselage. A former may include differing cross- sectional shapes at differing locations along a fuselage, as the former is a structural component that informs an overall shape of the fuselage. In embodiments, aircraft skin can be anchored to formers and strings such that an outer mold line of volume encapsulated by the formers and stringers comprises a same shape as aircraft 100 when installed. In other words, former(s) may form a fuselage’s ribs, and stringers may form interstitials between the ribs. A spiral orientation of stringers about formers may provide uniform robustness at any point on an aircraft fuselage such that if a portion sustains damage, another portion may remain largely unaffected. Aircraft skin may be mechanically coupled to underlying stringers and formers and may interact with a fluid, such as air, to generate lift and perform maneuvers.With continued reference to FIG. 1, according to some embodiments, a fuselage can comprise monocoque construction. Monocoque construction can include a primary structure that forms a shell (or skin in an aircraft’s case) and supports physical loads. Monocoque fuselages are fuselages in which the aircraft skin or shell may also include a primary structure. In monocoque construction aircraft skin would support tensile and compressive loads within itself and true monocoque aircraft can be further characterized by an absence of internal structural elements. Aircraft skin in this construction method may be rigid and can sustain its shape with substantially no structural assistance form underlying skeleton-like elements. Monocoque fuselage may include aircraft skin made from plywood layered in varying grain directions, epoxy-impregnated fiberglass, carbon fiber, or any combination thereof.With continued reference to FIG. 1, according to some embodiments, a fuselage may include a semi-monocoque construction. Semi-monocoque construction, as used in this disclosure, is used interchangeably with partially monocoque construction, discussed above. In semi-monocoque construction, a fuselage may derive some structural support from stressed aircraft skin and some structural support from underlying frame structure made of structural components. Formers or station frames can be seen running transverse to a long axis of fuselage with circular cutouts which may be used in real-world manufacturing for weight savings and for routing of electrical harnesses and other modern on-board systems. In a semi-monocoque construction, stringers may be thin, long strips of material that run parallel to a fuselage’s long axis. Stringers can be mechanically coupled to formers permanently, such as with rivets. Aircraft skin can be mechanically coupled to stringers and formers permanently, such as by rivets as well. A person of ordinary skill in the art will appreciate that there are numerous methods for mechanical fastening of the aforementioned components like screws, nails, dowels, pins, anchors, adhesives like glue or epoxy, or bolts and nuts, to name a few. According to some embodiments, a subset of semi-monocoque construction may be unibody construction. Unibody, which is short for “unitized body” or alternatively “unitary construction”, vehicles are characterized by a construction in which body, floor plan, and chassis form a single structure, for example an automobile. In the aircraft world, a unibody may include internal structural elements, like formers and stringers, constructed in one piece, integral to an aircraft skin. In some cases, stringers and formers may account for a bulk of any aircraft structure (excluding monocoque construction). Stringers and formers can be arranged in a plurality of orientations depending on aircraft operation and materials. Stringers may be arranged to carry axial (tensile or compressive), shear, bending or torsion forces throughout their overall structure. Due to their coupling to aircraft skin, aerodynamic forces exerted on aircraft skin may be transferred to stringers. Location of said stringers greatly informs type of forces and loads applied to each and every stringer, all of which may be accounted for through design processes including, material selection, cross-sectional area, and mechanical coupling methods of each member. Similar methods may be performed for former assessment and design. In general, formers may be significantly larger in cross-sectional area and thickness, depending on location, than stringers.Both stringers and formers may comprise aluminum, aluminum alloys, graphite epoxy composite, steel alloys, titanium, or an undisclosed material alone or in combination.With continued reference to FIG. 1, in some cases, a primary purpose for a substructure of a semi-monocoque structure is to stabilize a skin. Typically, aircraft structure is required to have a very light weight and as a result, in some cases, aircraft skin may be very thin. In some cases, unless supported, this thin skin structure may tend to buckle and / or cripple under compressive and / or shear loads. In some cases, underlying structure may be primarily configured to stabilize skins. For example, in an exemplary conventional airliner, wing structure is an airfoil-shaped box with truncated nose and aft triangle; without stabilizing substructure, in some cases, this box would buckle upper skin of the wing and the upper skin would also collapse into the lower skin under bending loads. In some cases, deformations are prevented with ribs that support stringers which stabilize the skin. Fuselages are similar with bulkheads or frames, and stringers.With continued reference to FIG. 1, in some embodiments, another common structural form is sandwich structure. As used in this disclosure, “sandwich structure” includes a skin structure having an inner and outer skin separated and stabilized by a core material. In some cases, sandwich structure may additionally include some number of ribs or frames. In some cases, sandwich structure may include metal, polymer, and / or composite. In some cases, core material may include honeycomb, foam plastic, and / or end-grain balsa wood. In some cases, sandwich structure can be popular on composite light airplanes, such as gliders and powered light planes. In some cases, sandwich structure may not use stringers, and sandwich structure may allow number of ribs or frames to be reduced, for instance in comparison with a semi- monocoque structure. In some cases, sandwich structure may be suitable for smaller, possibly unmanned, unpressurized blended wing body aircraft.With continued reference to FIG. 1, stressed skin, when used in semi-monocoque construction, may bear partial, yet significant, load. In other words, an internal structure, whether it be a frame of welded tubes, formers and stringers, or some combination, is not sufficiently strong enough by design to bear all loads. The concept of stressed skin is applied in monocoque and semi-monocoque construction methods of at least a fuselage and / or BWB 104. In some cases, monocoque may be considered to include substantially only structural skin, and in thatsense, aircraft skin undergoes stress by applied aerodynamic fluids imparted by fluid. Stress as used in continuum mechanics can be described in pound-force per square inch (lbf / in2) or Pascals (Pa). In semi-monocoque construction stressed skin bears part of aerodynamic loads and additionally imparts force on an underlying structure of stringers and formers.With continued reference to FIG. 1, a fuselage may include an interior cavity. An interior cavity may include a volumetric space configurable to house passenger seats and / or cargo. An interior cavity may be configured to include receptacles for fuel tanks, batteries, fuel cells, or other energy sources as described herein. In some cases, a post may be supporting a floor (i.e., deck), or in other words a surface on which a passenger, operator, passenger, payload, or other object would rest on due to gravity when within an aircraft 100 is in its level flight orientation or sitting on ground. A post may act similarly to stringer in that it is configured to support axial loads in compression due to a load being applied parallel to its axis due to, for example, a heavy object being placed on a floor of aircraft 100. A beam may be disposed in or on any portion a fuselage that requires additional bracing, specifically when disposed transverse to another structural element, like a post, that would benefit from support in that direction, opposing applied force. A beam may be disposed in a plurality of locations and orientations within a fuselage as necessitated by operational and constructional requirements.With continued reference to FIG. 1, aircraft 100 may include at least a flight component 108. A flight component 108 may be consistent with any description of a flight component described in this disclosure, such as without limitation propulsors, control surfaces, rotors, paddle wheels, engines, propellers, wings, winglets, or the like. For the purposes of this disclosure, at least a “flight component” is at least one element of an aircraft 100 configured to manipulate a fluid medium such as air to propel, control, or maneuver an aircraft. In nonlimiting examples, at least a flight component may include a rotor mechanically connected to a rotor shaft of an electric motor further mechanically affixed to at least a portion of aircraft 100. In some embodiments, at least a flight component 108 may include a propulsor, for example a rotor attached to an electric motor configured to produce shaft torque and in turn, create thrust. As used in this disclosure, an “electric motor” is an electrical machine that converts electric energy into mechanical work.With continued reference to FIG. 1, for the purposes of this disclosure, “torque”, is a twisting force that tends to cause rotation. Torque may be considered an effort and a rotational analogue to linear force. A magnitude of torque of a rigid body may depend on three quantities: a force applied, a lever arm vector connecting a point about which the torque is being measured to a point of force application, and an angle between the force and the lever arm vector. A force applied perpendicularly to a lever multiplied by its distance from the lever's fulcrum (the length of the lever arm) is its torque. A force of three newtons applied two meters from the fulcrum, for example, exerts the same torque as a force of one newton applied six meters from the fulcrum. In some cases, direction of a torque can be determined by using a right-hand grip rule which states: if fingers of right hand are curled from a direction of lever arm to direction of force, then thumb points in a direction of the torque. One of ordinary skill in the art would appreciate that torque may be represented as a vector, consistent with this disclosure, and therefore may include a magnitude and a direction. “Torque” and “moment” are used interchangeably within this disclosure. Any torque command or signal within this disclosure may include at least the steady state torque to achieve the torque output to at least a propul sor.With continued reference to FIG. 1, at least a flight component may be one or more devices configured to affect aircraft’s 100 attitude. “Attitude”, for the purposes of this disclosure, is the relative orientation of a body, in this case aircraft 100, as compared to earth’s surface or any other reference point and / or coordinate system. In some cases, attitude may be displayed to pilots, personnel, remote users, or one or more computing devices in an attitude indicator, such as without limitation a visual representation of a horizon and its relative orientation to aircraft 100. A plurality of attitude datums may indicate one or more measurements relative to an aircraft’s pitch, roll, yaw, or throttle compared to a relative starting point. One or more sensors may measure or detect an aircraft’s 100 attitude and establish one or more attitude datums. An “attitude datum”, for the purposes of this disclosure, refers to at least an element of data identifying an attitude of an aircraft 100.With continued reference to FIG. 1, in some cases, aircraft 100 may include one or more of an angle of attack sensor and a yaw sensor. In some embodiments, one or more of an angle of attack sensor and a yaw sensor may include a vane (e.g., wind vane). In some cases, vane may include a protrusion on a pivot with an aft tail. The protrusion may be configured to rotate aboutpivot to maintain zero tail angle of attack. In some cases, pivot may turn an electronic device that reports one or more of angle of attack and / or yaw, depending on, for example, orientation of the pivot and tail. Alternatively, or additionally, in some cases, one or more of angle of attack sensor and / or yaw sensor may include a plurality of pressure ports located in selected locations, with pressure sensors located at each pressure port. In some cases, differential pressure between pressure ports can be used to estimate angle of attack and / or yaw.With continued reference to FIG. 1, in some cases, aircraft 100 may include at least a pilot control. As used in this disclosure, a “pilot control,” is an interface device that allows an operator, human or machine, to control a flight component of an aircraft. Pilot control may be communicatively connected to any other component presented in aircraft 100, the communicative connection may include redundant connections configured to safeguard against single-point failure. In some cases, a plurality of attitude datums may indicate a pilot’s instruction to change heading and / or trim of an aircraft 100. Pilot input may indicate a pilot’s instruction to change an aircraft’s pitch, roll, yaw, throttle, and / or any combination thereof. Aircraft trajectory may be manipulated by one or more control surfaces and propulsors working alone or in tandem consistent with the entirety of this disclosure. “Pitch”, for the purposes of this disclosure refers to an aircraft’s angle of attack, that is a difference between a plane including at least a portion of both wings of the aircraft running nose to tail and a horizontal flight trajectory. For example, an aircraft may pitch “up” when its nose is angled upward compared to horizontal flight, as in a climb maneuver. In another example, an aircraft may pitch “down”, when its nose is angled downward compared to horizontal flight, like in a dive maneuver. In some cases, angle of attack may not be used as an input, for instance pilot input, to any system disclosed herein; in such circumstances proxies may be used such as pilot controls, remote controls, or sensor levels, such as true airspeed sensors, pitot tubes, pneumatic / hydraulic sensors, and the like. “Roll” for the purposes of this disclosure, refers to an aircraft’s position about its longitudinal axis, that is to say that when an aircraft rotates about its axis from its tail to its nose, and one side rolls upward, as in a banking maneuver. “Yaw”, for the purposes of this disclosure, refers to an aircraft’s turn angle, when an aircraft rotates about an imaginary vertical axis intersecting center of earth and aircraft 100. “Throttle”, for the purposes of this disclosure, refers to an aircraft outputting an amount of thrust from a propulsor. In context of a pilot input, throttle may refer to a pilot’s inputto increase or decrease thrust produced by at least a propulsor. Flight components 108 may receive and / or transmit signals, for example an aircraft command signal. Aircraft command signal may include any signal described in this disclosure, such as without limitation electrical signal, optical signal, pneumatic signal, hydraulic signal, and / or mechanical signal. In some cases, an aircraft command may be a function of a signal from a pilot control. In some cases, an aircraft command may include or be determined as a function of a pilot command. For example, aircraft commands may be determined as a function of a mechanical movement of a throttle. Signals may include analog signals, digital signals, periodic or aperiodic signal, step signals, unit impulse signal, unit ramp signal, unit parabolic signal, signum function, exponential signal, rectangular signal, triangular signal, sinusoidal signal, sine function, or pulse width modulated signal. Pilot control may include circuitry, computing devices, electronic components or a combination thereof that translates pilot input into a signal configured to be transmitted to another electronic component. In some cases, a plurality of attitude commands may be determined as a function of an input to a pilot control. A plurality of attitude commands may include a total attitude command datum, such as a combination of attitude adjustments represented by one or a certain number of combinatorial datums. A plurality of attitude commands may include individual attitude datums representing total or relative change in attitude measurements relative to pitch, roll, yaw, and throttle.With continued reference to FIG. 1, in some embodiments, pilot control may include at least a sensor. As used in this disclosure, a “sensor” is a device that detects a phenomenon. In some cases, a sensor may detect a phenomenon and transmit a signal that is representative of the phenomenon. At least a sensor may include, torque sensor, gyroscope, accelerometer, magnetometer, inertial measurement unit (IMU), pressure sensor, force sensor, proximity sensor, displacement sensor, vibration sensor, among others. At least a sensor may include a sensor suite which may include a plurality of sensors that may detect similar or unique phenomena. For example, in a non-limiting embodiment, sensor suite may include a plurality of accelerometers, a mixture of accelerometers and gyroscopes, or a mixture of an accelerometer, gyroscope, and torque sensor. For the purposes of the disclosure, a “torque datum” is one or more elements of data representing one or more parameters detailing power output by one or more propulsors, flight components, or other elements of an electric aircraft. A torque datum may indicate thetorque output of at least a flight component 108. At least a flight component 108 may include any propulsor as described herein. In embodiment, at least a flight component 108 may include an electric motor, a propeller, a jet engine, a paddle wheel, a rotor, turbine, or any other mechanism configured to manipulate a fluid medium to propel an aircraft as described herein, an embodiment of at least a sensor may include or be included in, a sensor suite. The herein disclosed system and method may comprise a plurality of sensors in the form of individual sensors or a sensor suite working in tandem or individually. A sensor suite may include a plurality of independent sensors, as described herein, where any number of the described sensors may be used to detect any number of physical or electrical quantities associated with an aircraft power system or an electrical energy storage system. Independent sensors may include separate sensors measuring physical or electrical quantities that may be powered by and / or in communication with circuits independently, where each may signal sensor output to a control circuit such as a user graphical interface. In a non-limiting example, there may be four independent sensors housed in and / or on battery pack measuring temperature, electrical characteristic such as voltage, amperage, resistance, or impedance, or any other parameters and / or quantities as described in this disclosure. In an embodiment, use of a plurality of independent sensors may result in redundancy configured to employ more than one sensor that measures the same phenomenon, those sensors being of the same type, a combination of, or another type of sensor not disclosed, so that in the event one sensor fails, the ability of a battery management system and / or user to detect phenomenon is maintained and in a non-limiting example, a user alter aircraft usage pursuant to sensor readings.With continued reference to FIG. 1, at least a sensor may include a moisture sensor. “Moisture”, as used in this disclosure, is the presence of water, this may include vaporized water in air, condensation on the surfaces of objects, or concentrations of liquid water. Moisture may include humidity. “Humidity”, as used in this disclosure, is the property of a gaseous medium (almost always air) to hold water in the form of vapor. An amount of water vapor contained within a parcel of air can vary significantly. Water vapor is generally invisible to the human eye and may be damaging to electrical components. There are three primary measurements of humidity, absolute, relative, specific humidity. “Absolute humidity,” for the purposes of this disclosure, describes the water content of air and is expressed in either grams per cubic meters orgrams per kilogram. “Relative humidity”, for the purposes of this disclosure, is expressed as a percentage, indicating a present stat of absolute humidity relative to a maximum humidity given the same temperature. “Specific humidity”, for the purposes of this disclosure, is the ratio of water vapor mass to total moist air parcel mass, where parcel is a given portion of a gaseous medium. A moisture sensor may be psychrometer. A moisture sensor may be a hygrometer. A moisture sensor may be configured to act as or include a humidistat. A “humidistat”, for the purposes of this disclosure, is a humidity-triggered switch, often used to control another electronic device. A moisture sensor may use capacitance to measure relative humidity and include in itself, or as an external component, include a device to convert relative humidity measurements to absolute humidity measurements. “Capacitance”, for the purposes of this disclosure, is the ability of a system to store an electric charge, in this case the system is a parcel of air which may be near, adjacent to, or above a battery cell.With continued reference to FIG. 1, at least a sensor may include electrical sensors. An electrical sensor may be configured to measure voltage across a component, electrical current through a component, and resistance of a component. Electrical sensors may include separate sensors to measure each of the previously disclosed electrical characteristics such as voltmeter, ammeter, and ohmmeter, respectively. One or more sensors may be communicatively coupled to at least a pilot control, the manipulation of which, may constitute at least an aircraft command. Signals may include electrical, electromagnetic, visual, audio, radio waves, or another undisclosed signal type alone or in combination. At least a sensor communicatively connected to at least a pilot control may include a sensor disposed on, near, around or within at least pilot control. At least a sensor may include a motion sensor. “Motion sensor”, for the purposes of this disclosure refers to a device or component configured to detect physical movement of an object or grouping of objects. One of ordinary skill in the art would appreciate, after reviewing the entirety of this disclosure, that motion may include a plurality of types including but not limited to: spinning, rotating, oscillating, gyrating, jumping, sliding, reciprocating, or the like. At least a sensor may include, torque sensor, gyroscope, accelerometer, torque sensor, magnetometer, inertial measurement unit (IMU), pressure sensor, force sensor, proximity sensor, displacement sensor, vibration sensor, among others. At least a sensor may include a sensor suite which may include a plurality of sensors that may detect similar or unique phenomena. For example, in anon-limiting embodiment, sensor suite may include a plurality of accelerometers, a mixture of accelerometers and gyroscopes, or a mixture of an accelerometer, gyroscope, and torque sensor. The herein disclosed system and method may comprise a plurality of sensors in the form of individual sensors or a sensor suite working in tandem or individually. A sensor suite may include a plurality of independent sensors, as described herein, where any number of the described sensors may be used to detect any number of physical or electrical quantities associated with an aircraft power system or an electrical energy storage system. Independent sensors may include separate sensors measuring physical or electrical quantities that may be powered by and / or in communication with circuits independently, where each may signal sensor output to a control circuit such as a user graphical interface. In an embodiment, use of a plurality of independent sensors may result in redundancy configured to employ more than one sensor that measures the same phenomenon, those sensors being of the same type, a combination of, or another type of sensor not disclosed, so that in the event one sensor fails, the ability to detect phenomenon is maintained and in a non-limiting example, a user alter aircraft usage pursuant to sensor readings.With continued reference to FIG. 1, at least a flight component 108 may include wings, empennages, nacelles, control surfaces, fuselages, and landing gear, among others, to name a few. In embodiments, an empennage may be disposed at the aftmost point of an aircraft body 104. Empennage may comprise a tail of aircraft 100, further comprising rudders, vertical stabilizers, horizontal stabilizers, stabilators, elevators, trim tabs, among others. At least a portion of empennage may be manipulated directly or indirectly by pilot commands to impart control forces on a fluid in which the aircraft 100 is flying. Manipulation of these empennage control surfaces may, in part, change an aircraft’s heading in pitch, roll, and yaw. Wings comprise may include structures which include airfoils configured to create a pressure differential resulting in lift. Wings are generally disposed on a left and right side of aircraft 100 symmetrically, at a point between nose and empennage. Wings may comprise a plurality of geometries in planform view, swept swing, tapered, variable wing, triangular, oblong, elliptical, square, among others. Wings may be blended into the body of the aircraft such as in a BWB 104 aircraft 100 where no strong delineation of body and wing exists. A wing’s cross section geometry may comprise an airfoil. An “airfoil” as used in this disclosure, is a shape specifically designed such that a fluid flowingon opposing sides of it exert differing levels of pressure against the airfoil. In embodiments, a bottom surface of an aircraft can be configured to generate a greater pressure than does a top surface, resulting in lift. A wing may comprise differing and / or similar cross-sectional geometries over its cord length, e.g., length from wing tip to where wing meets the aircraft’s body. One or more wings may be symmetrical about an aircraft’s longitudinal plane, which comprises a longitudinal or roll axis reaching down a center of the aircraft through the nose and empennage, and the aircraft’s yaw axis. In some cases, wings may comprise controls surfaces configured to be commanded by a pilot and / or autopilot to change a wing’s geometry and therefore its interaction with a fluid medium. Flight component 108 may include control surfaces. Control surfaces may include without limitation flaps, ailerons, tabs, spoilers, and slats, among others. In some cases, control surfaces may be disposed on wings in a plurality of locations and arrangements. In some cases, control surfaces may be disposed at leading and / or trailing edges of wings, and may be configured to deflect up, down, forward, aft, or any combination thereof.In some cases, flight component 108 may include a winglet. For the purposes of this disclosure, a “winglet” is a flight component configured to manipulate a fluid medium and is mechanically attached to a wing or aircraft and may alternatively called a “wingtip device.” Wingtip devices may be used to improve efficiency of fixed-wing aircraft by reducing drag. Although there are several types of wingtip devices which function in different manners, their intended effect may be to reduce an aircraft's drag by partial recovery of tip vortex energy. Wingtip devices can also improve aircraft handling characteristics and enhance safety for aircraft 100. Such devices increase an effective aspect ratio of a wing without greatly increasing wingspan. Extending wingspan may lower lift-induced drag but would increase parasitic drag and would require boosting the strength and weight of the wing. As a result, according to some aeronautic design equations, a maximum wingspan made be determined above which no net benefit exits from further increased span. There may also be operational considerations that limit the allowable wingspan (e.g., available width at airport gates).Wingtip devices, in some cases, may increase lift generated at wingtip (by smoothing airflow across an upper wing near the wingtip) and reduce lift-induced drag caused by wingtip vortices, thereby improving a lift-to-drag ratio. This increases fuel efficiency in powered aircraftand increases cross-country speed in gliders, in both cases increasing range. U.S. Air Force studies indicate that a given improvement in fuel efficiency correlates directly and causally with increase in an aircraft's lift-to-drag ratio. The term "winglet" has previously been used to describe an additional lifting surface on an aircraft, like a short section between wheels on fixed undercarriage. An upward angle (i.e., can’t) of a winglet, its inward or outward angle (i.e., toe), as well as its size and shape are selectable design parameters which may be chosen for correct performance in a given application. A wingtip vortex, which rotates around from below a wing, strikes a cambered surface of a winglet, generating a force that angles inward and slightly forward. A winglet’s relation to a wingtip vortex may be considered analogous to sailboat sails when sailing to windward (i.e., close-hauled). Similar to the close-hauled sailboat’s sails, winglets may convert some of what would otherwise-be wasted energy in a wingtip vortex to an apparent thrust. This small contribution can be worthwhile over the aircraft's lifetime. Another potential benefit of winglets is that they may reduce an intensity of wake vortices. Wake vortices may trail behind an aircraft 100 and pose a hazard to other aircraft. Minimum spacing requirements between aircraft at airports are largely dictated by hazards, like those from wake vortices. Aircraft are classified by weight (e g., "Light," "Heavy," and the like) often base upon vortex strength, which grows with an aircraft’s lift coefficient. Thus, associated turbulence is greatest at low speed and high weight, which may be produced at high angle of attack near airports. Winglets and wingtip fences may also increase efficiency by reducing vortex interference with laminar airflow near wingtips, by moving a confluence of low-pressure air (over wing) and high-pressure air (under wing) away from a surface of the wing. Wingtip vortices create turbulence, which may originate at a leading edge of a wingtip and propagate backwards and inboard. This turbulence may delaminate airflow over a small triangular section of an outboard wing, thereby frustrating lift in that area. A fence / winglet drives an area where a vortex forms upward away from a wing surface, as the resulting vortex is repositioned to a top tip of the winglet.Referring now to FIG. 2, an exemplary top-down diagram of an exemplary blended wing aircraft 200 is with a combustion engine is illustrated. Aircraft 200 may include a blended wing body 204, first fuel store 212a-b, second fuel store 224a-b, propulsor 208, combustion engine 228, fuel cell 216, and auxiliary power system 220. Aircraft 200 may be the same aircraft asaircraft 100, or at least has all the same components and characteristics as described above. As described above, a blended wing body (BWB) is a fixed-wing aircraft body having no clear demarcation between wings and a main body of the aircraft. For example, a BWB 204 aircraft may have distinct wing and body structures, which are smoothly blended together with no clear dividing line or boundary feature between wing and fuselage. As used in this disclosure, a “transitional” portion of blended wing body 204 is the portion of the blended wing body 204 that includes the aircraft body between wing and a main body.With continued reference to FIG. 2, in further nonlimiting embodiments, an energy source as explained above with reference to FIG. 1 may include a fuel store. As used in this disclosure, a “fuel store” is an aircraft component configured to store a fuel. In some cases, a fuel store 204 may include a fuel tank. Fuel tank may include a fuel, which may include liquified gas fuel, liquid hydrogen, or natural gas. As used in this disclosure, a “liquified gas fuel” is a fuel that at standard atmospheric conditions or when utilized (e.g., combusted) is gas and is stored as a fuel. Liquified gas fuels include without limitation liquid hydrogen, propane, and liquified natural gas. As used in this disclosure, a “fuel” may include any substance that stores energy. Exemplary non-limiting fuels include hydrocarbon fuels, petroleum-based fuels., synthetic fuels, chemical fuels, Jet fuels (e.g., Jet-A fuel, Jet-B fuel, and the like), kerosene-based fuel, gasolinebased fuel, an electrochemical-based fuel (e.g., lithium-ion battery), a hydrogen-based fuel, natural gas-based fuel, and the like. As described in greater detail below fuel store may be located substantially within blended wing body 104 of aircraft 200, for example without limitation within a wing portion of blended wing body 108. Aviation fuels may include petroleum-based fuels, or petroleum and synthetic fuel blends, used to power aircraft 200. In some cases, aviation fuels may have more stringent requirements than fuels used for ground use, such as heating and road transport. Aviation fuels may contain additives to enhance or maintain properties important to fuel performance or handling. Fuel may include one or more of liquid hydrogen and liquid natural gas. In some cases, specific energy may be considered an important criterion in selecting fuel for an aircraft 200. Liquid fuel may include Jet-A. Presently Jet-A powers modern commercial airliners and is a mix of extremely refined kerosene and burns at temperatures at or above 49 °C (1120 °F). In this embodiment, fuel has a greater energy density per weight and a lesser energy density per volume than conventional kerosene-based fuel.Kerosene-based fuel has a much higher flash point than gasoline-based fuel, meaning that it requires significantly higher temperature to ignite.With continued reference to FIG. 2, aircraft 200 may include at least a first fuel store 212a-b. At least a first fuel store 212a-b may be configured to store a first fuel. First fuel may include any fuel taught in this disclosure, for example without limitation liquid hydrogen, liquid natural gas, gasoline-based fuels, kerosene-based fuels and the like. In some embodiments, first fuel store 212a-b may be at least partially located within a transitional portion of blended wing body 204. According to some embodiments, first fuel store may be configured to store one or more of liquid hydrogen and natural gas. For example, although weight energy density of liquid hydrogen is high, volume energy density of liquid hydrogen is lower than conventional aviation fuels. For this reason, in some cases, fuel store 212a-b may be located within a transitional portion of blended wing body 204 as greater volume for storage is available here, for example when compared to a wing portion. In some cases, liquid nitrogen may need to be stored at extremely cold temperatures, for instance without limitation at a temperature below -252°C. As liquid hydrogen warms it boils off and is lost. As a result, boil off rate is considered when employing liquid hydrogen as a fuel. In some cases, first fuel store 212, or any fuel store containing liquid hydrogen, may be heavily insulated. For example, in some cases, fuel store may include an inner wall and an outer wall with a vacuum chamber disposed between the inner wall and the outer wall. Vacuum within vacuum chamber prevents convective and conductive heat loss between inner and outer wall, so that substantially only radiative heat transfer may be possible between the two walls dramatically slowing heat transfer (and heating). Alternatively, or additionally, in some cases, an insulation may be located between inner wall and outer wall of fuel store. Exemplary non-limiting insulations include high loft materials, silica aerogel, polyurethane, polystyrene, fiberglass, and the like. In some cases, a reflective material may be used within a wall of fuel store to slow radiative heat transfer, for example without limitation metallic materials with high polish like foil.Continuing to refer to FIG. 2, aircraft 200 may further include a tank as a fuel store. In this disclosure, a “tank” is a container of fluids, for example for flammable fluids, such as fuel. In an embodiment, a tank stores fuel to power aircraft 200. Tank may be permanently attached to aircraft 200. As used in this disclosure, a tank may be "permanently attached" when it isconfigured to not be removed during ordinary use. For example, a tank permanently attached to aircraft may be removed during maintenance or overhaul but is otherwise a permanent flight component of the aircraft. Tank may include one or more compartments to store fuel in. Tank may be a part of fuel delivery system for an engine, in which the fuel may be stored inside the at least a tank and then propelled or released into an engine, such as without limitation a combustion engine. Tank may be a pressure vessel. A “pressure vessel” is a container configured to hold fluids at a pressure that may differ from an ambient pressure. Pressure vessel may be configured to be pressurized in order to allow flow of gaseous hydrogen from tank, for example without a need to pump. In an embodiment, but without limitation, tank may act as a pressure vessel to store the fuel at a high pressure above 5 psig, 15 psig, 50psig, or the like. Tank may be made of any material able to withstand such high pressure, such as but without limitation, aluminum, carbon fiber, composite materials, or the like. Furthermore, tank may further include an inner wall and an outer wall. Tank is further discussed herein with reference to FIGS. 3A-J.In some cases, a voluminous fuel store 212a-b, for instance located within a transitional portion of blended wing body 204, may be advantageous for liquid hydrogen (or liquid natural gas) storage as it slows a rate of temperature rise of fuel. For instance, heat transfer is a function of surface area of fuel store and may be understood according to Newton’s Law of Cooling. Whereas thermal compliance is a function of mass (volume multiplied by density). As a fuel store increases in size, its volume increases more than surface area. This phenomenon may be understood as square-cube law, stated thus when an object undergoes a proportional increase in size, its new surface area is proportional to the square of the multiplier and its new volume is proportional to the cube of the multiplier. For example, imagine a cubic fuel store increases from a first length, h, to a second length, h. An area of fuel store may increase thus:and a volume of fuel store increases thuswhere Ai is first surface area, A2 is second surface area, Vi is first volume, and V2 is second volume. For example, a cube with a side length of 1 meter has a surface area of 6 m2and avolume of 1 m3. If dimensions of cube were multiplied by 2, its surface area would be multiplied by the square of 2 and become 24 m2. Its volume would be multiplied by cube of 2 and become 8 m3. The original cube (Im sides) has a surface area to volume ratio of 6:1. The larger (2m sides) cube has a surface area to volume ratio of (24 / 8) 3: 1. As dimensions increase, volume will continue to grow faster than surface area. Square-cube principle applies to all solids, not just cubes.With continued reference to FIG. 2, modular aircraft 100200 may include an energy source which may include a fuel cell 216. As used in this disclosure, a “fuel cell” is an electrochemical device that combines a fuel and an oxidizing agent to create electricity. In some cases, fuel cells 216 may be different from most batteries in requiring a continuous source of fuel and oxygen (usually from air) to sustain the chemical reaction, whereas in a battery the chemical energy comes from metals and their ions or oxides that are commonly already present in the battery, except in flow batteries. Fuel cell 216 may produce electricity continuously for as long as fuel and oxygen are supplied. In some cases, oxygen may be provided by way of ambient or atmospheric air. Alternatively, or additionally, in some cases, aircraft 100 may additionally include at least an oxygen tank configured to store oxygen for use with fuel cell. In some cases, oxygen may be stored within at least an oxygen tank in a gaseous state. In some cases, at least a fuel cell 216 may be configured to combine first fuel with an oxidizing agent, such as oxygen to produce electricity. At least a fuel cell 216 may include any fuel cell described in this disclosure, including without limitation with reference to FIG. 1 above.With continued reference to FIG. 2, in some embodiments, fuel cell 216 may consist of different types. Commonly a fuel cell 216 consists of an anode, a cathode, and an electrolyte that allows ions, often positively charged hydrogen ions (protons), to move between two sides of the fuel cell 216. At anode, a catalyst causes fuel to undergo oxidation reactions that generate ions (often positively charged hydrogen ions) and electrons. Ions move from anode to cathode through electrolyte. Concurrently, electrons may flow from anode to cathode through an external circuit, producing direct current electricity. At cathode, another catalyst causes ions, electrons, and oxygen to react, forming water and possibly other products. Fuel cell 216 may be classified by type of electrolyte used and by difference in startup time ranging from 1 second for protonexchange membrane fuel cells (PEM fuel cells, or PEMFC) to 10 minutes for solid oxide fuelcells (SOFC). In some cases, energy source may include a related technology, such as flow batteries. Within a flow battery fuel can be regenerated by recharging. Individual fuel cells produce relatively small electrical potentials, about 0.7 volts. Therefore, in some cases, fuel cell 216 may be "stacked", or placed in a series, to create sufficient voltage to meet an application's requirements. In addition to electricity, fuel cell 216 may produce water, heat and, depending on the fuel source, very small amounts of nitrogen dioxide and other emissions. Energy efficiency of a fuel cell 216 is generally between 40 and 90%.Fuel cell 216 may include an electrolyte. In some cases, electrolyte may define a type of fuel cell 216. Electrolyte may include any number of substances like potassium hydroxide, salt carbonates, and phosphoric acid. Commonly a fuel cell 216 is fueled by hydrogen. Fuel cell 216 may feature an anode catalyst, like fine platinum powder, which breaks down fuel into electrons and ions. Fuel cell 216 may feature a cathode catalyst, often nickel, which converts ions into waste chemicals, with water being the most common type of waste. A fuel cell 216 may include gas diffusion layers that are designed to resist oxidization.With continued reference to FIG. 12, aircraft 200 may include an energy source which may include a cell such as a battery cell, or a plurality of battery cells making a battery module. An energy source may be a plurality of energy sources. The module may include batteries connected in parallel or in series or a plurality of modules connected either in series or in parallel designed to deliver both the power and energy requirements of the application. Connecting batteries in series may increase the voltage of an energy source which may provide more power on demand. High voltage batteries may require cell matching when high peak load is needed. As more cells are connected in strings, there may exist the possibility of one cell failing which may increase resistance in the module and reduce the overall power output as the voltage of the module may decrease as a result of that failing cell. Connecting batteries in parallel may increase total current capacity by decreasing total resistance, and it also may increase overall amp-hour capacity. The overall energy and power outputs of an energy source may be based on the individual battery cell performance, or an extrapolation based on the measurement of at least an electrical parameter. In an embodiment where an energy source includes a plurality of battery cells, the overall power output capacity may be dependent on the electrical parameters of each individual cell. If one cell experiences high self-discharge during demand, power drawn from anenergy source may be decreased to avoid damage to the weakest cell. An energy source may further include, without limitation, wiring, conduit, housing, cooling system and battery management system. Persons skilled in the art will be aware, after reviewing the entirety of this disclosure, of many different components of an energy source.With continued reference to FIG. 2, aircraft 200 may include multiple flight component 108 sub-systems, each of which may have a separate energy source. For instance, and without limitation, one or more flight components 108 may have a dedicated energy source. Alternatively, or additionally, a plurality of energy sources may each provide power to two or more flight components 108, such as, without limitation, a “fore” energy source providing power to flight components located toward a front of an aircraft 200, while an “aft” energy source provides power to flight components located toward a rear of the aircraft 200. As a further nonlimiting example, a flight component of group of flight components may be powered by a plurality of energy sources. For example, and without limitation, two or more energy sources may power one or more flight components; two energy sources may include, without limitation, at least a first energy source having high specific energy density and at least a second energy source having high specific power density, which may be selectively deployed as required for higher-power and lower-power needs. Alternatively, or additionally, a plurality of energy sources may be placed in parallel to provide power to the same single propulsor or plurality of propulsors, as explained below. Alternatively, or additionally, two or more separate propulsion subsystems may be joined using intertie switches (not shown) causing the two or more separate propulsion subsystems to be treatable as a single propulsion subsystem or system, for which potential under load of combined energy sources may be used as the electric potential. Persons skilled in the art, upon reviewing the entirety of this disclosure, will be aware of various combinations of energy sources that may each provide power to single or multiple propulsors in various configurations.With continued reference to FIG. 2, aircraft 200 may include a flight component 108 that includes at least a nacelle. For the purposes of this disclosure, a “nacelle” is a streamlined body housing, which is sized according to that which is houses, such as without limitation an engine, a fuel store, or a flight component. When attached by a pylon entirely outside an BWB airframe 104 a nacelle may sometimes be referred to as a pod, in which case an engine within the nacellemay be referred to as a podded engine. Tn some cases, an aircraft cockpit may also be housed in a nacelle, rather than in a conventional fuselage. At least a nacelle may substantially encapsulate a propulsor, which may include a motor or an engine. At least a nacelle may be mechanically connected to at least a portion of aircraft 200 partially or wholly enveloped by an outer mold line of the aircraft 200. At least a nacelle may be designed to be streamlined. At least a nacelle may be asymmetrical about a plane comprising the longitudinal axis of the engine and the yaw axis of modular aircraft 200.With continued reference to FIG. 2, a flight component 108 may include at least a propulsor 208. A “propulsor,” as used herein, is a component or device used to propel a craft by exerting force on a fluid medium, which may include a gaseous medium such as air or a liquid medium such as water. For the purposes of this disclosure, “substantially encapsulate” is the state of a first body (e g., housing) surrounding all or most of a second body. A motor may include without limitation, any electric motor, where an electric motor is a device that converts electrical energy into mechanical work for instance by causing a shaft to rotate. A motor may be driven by direct current (DC) electric power; for instance, a motor may include a brushed DC motor or the like. A motor may be driven by electric power having varied or reversing voltage levels, such as alternating current (AC) power as produced by an alternating current generator and / or inverter, or otherwise varying power, such as produced by a switching power source. A motor may include, without limitation, a brushless DC electric motor, a permanent magnet synchronous motor, a switched reluctance motor, and / or an induction motor; persons skilled in the art, upon reviewing the entirety of this disclosure, will be aware of various alternative or additional forms and / or configurations that a motor may take or exemplify as consistent with this disclosure. In addition to inverter and / or switching power source, a circuit driving motor may include electronic speed controllers or other components for regulating motor speed, rotation direction, torque, and / or dynamic braking. Motor may include or be connected to one or more sensors detecting one or more conditions of motor; one or more conditions may include, without limitation, voltage levels, electromotive force, current levels, temperature, current speed of rotation, position sensors, and the like. For instance, and without limitation, one or more sensors may be used to detect back-EMF, or to detect parameters used to determine back-EMF, as described in further detail below. One or more sensors may include a plurality of current sensors,voltage sensors, and speed or position feedback sensors. One or more sensors may communicate a current status of motor to a flight controller and / or a computing device; computing device may include any computing device as described in this disclosure, including without limitation, a flight controller.With continued reference to FIG. 2, aircraft 200 may include at least a propulsor 208 mechanically affixed to the aircraft 200. In some cases, at least a propulsor 208 may be configured to propel aircraft 200. Propulsor may include any propulsor described in this disclosure, for example with reference to FIG. 1. In some embodiments, at least a propulsor 208 may include at least a combustion engine that burns first fuel and produces mechanical work. Resulting mechanical work may be used to power at least a propulsor 208. In some embodiments, at least a propulsor 208 may include at least an electric motor operatively connected with fuel cell 216. Propulsor 208 may be operatively connected to fuel cell 216 by way of electrical communication, for example through one or more conductors. In some cases, at least a fuel cell 216 may be configured to power at least an electric motor of propulsor 208. In some embodiments, at least a propulsor 208 may include both a combustion engine and an electric motor.With continued reference to FIG. 2, a motor may be connected to a thrust element. Thrust element may include any device or component that converts mechanical work, for example of a motor or engine, into thrust in a fluid medium. Thrust element may include, without limitation, a device using moving or rotating foils, including without limitation one or more rotors, an airscrew or propeller, a set of airscrews or propellers such as contra-rotating propellers or corotating propellers, a moving or flapping wing, or the like. Thrust element may include without limitation a marine propeller or screw, an impeller, a turbine, a pump-jet, a paddle or paddlebased device, or the like. Thrust element may include a rotor. Persons skilled in the art, upon reviewing the entirety of this disclosure, will be aware of various devices that may be used as thrust element. A thrust element may include any device or component that converts mechanical energy (i.e., work) of a motor, for instance in form of rotational motion of a shaft, into thrust within a fluid medium. As another non-limiting example, a thrust element may include an eight- bladed pusher propeller, such as an eight-bladed propeller mounted behind the engine to ensure the drive shaft is in compression.With continued reference to FIG. 12, in nonlimiting embodiments, at least a flight component 108 may include an airbreathing engine such as a jet engine, turbojet engine, turboshaft engine, ramjet engine, scramjet engine, hybrid propulsion system, turbofan engine, or the like. At least a flight component 108 may be fueled by any fuel described in this disclosure, for instance without limitation Jet-A, Jet-B, diesel fuel, gasoline, or the like. In nonlimiting embodiments, a jet engine is a type of reaction engine discharging a fast-moving jet that generates thrust by jet propulsion. While this broad definition can include rocket, waterjet, and hybrid propulsion, the term jet engine, in some cases, refers to an internal combustion airbreathing jet engine such as a turbojet, turbofan, ramjet, or pulsejet. In general, jet engines are internal combustion engines. As used in this disclosure, a “combustion engine” is a mechanical device that is configured to convert mechanical work from heat produced by combustion of a fuel. In some cases, a combustion engine may operate according to an approximation of a thermodynamic cycle, such as without limitation a Carnot cycle, a Cheng cycle, a Combined cycle, a Brayton cycle, an Otto cycle, an Allam power cycle, a Kalina cycle, a Rankine cycle, and / or the like. In some cases, a combustion engine may include an internal combustion engine. An internal combustion engine may include heat engine in which combustion of fuel occurs with an oxidizer (usually air) in a combustion chamber that comprises a part of a working fluid flow circuit. Exemplary internal combustion engines may without limitation a reciprocating engine (e g., 4-stroke engine), a combustion turbine engine (e.g., jet engines, gas turbines, Brayton cycle engines, and the like), a rotary engine (e g., Wankel engines), and the like. In this embodiment, engine of aircraft 200 may be a combustion engine 228. Combustion engine 228 is further explained below. In nonlimiting embodiments, airbreathing jet engines feature a rotating air compressor powered by a turbine, with leftover power providing thrust through a propelling nozzle — this process may be known as a Brayton thermodynamic cycle. Jet aircraft may use such engines for long-distance travel. Early jet aircraft used turbojet engines that were relatively inefficient for subsonic flight. Most modern subsonic jet aircraft use more complex high-bypass turbofan engines. In some cases, they give higher speed and greater fuel efficiency than piston and propeller aeroengines over long distances. A few air-breathing engines made for highspeed applications (ramjets and scramjets) may use a ram effect of aircraft's speed instead of a mechanical compressor. An airbreathing jet engine (or ducted jet engine) may emit a jet of hotexhaust gases formed from air that is forced into the engine by several stages of centrifugal, axial or ram compression, which is then heated and expanded through a nozzle. In some cases, a majority of mass flow through an airbreathing jet engine may be provided by air taken from outside of the engine and heated internally, using energy stored in the form of fuel. In some cases, a jet engine may include are turbofans. Alternatively, and / or additionally, jet engine may include a turbojet. In some cases, a turbofan may use a gas turbine engine core with high overall pressure ratio (e.g., 40: 1) and high turbine entry temperature (e.g., about 1800 K) and provide thrust with a turbine-powered fan stage. In some cases, thrust may also be at least partially provided by way of pure exhaust thrust (as in a turbojet engine). In some cases, a turbofan may have a high efficiency, relative to a turbojet. In some cases, a jet engine may use simple ram effect (e.g., ramjet) or pulse combustion (e.g., pulsejet) to give compression. Persons skilled in the art, upon reviewing the entirety of this disclosure, will be aware of various devices that may be used as a thrust element.Still referring to FIG. 2, a propulsor 208 of aircraft 200 may comprise a combustion engine. As used in this disclosure, a “combustion engine” is a mechanical device that is configured to convert mechanical work from heat produced by combustion of a fuel. In some cases, a combustion engine 228 may operate according to an approximation of a thermodynamic cycle, such as without limitation a Carnot cycle, a Cheng cycle, a Combined cycle, a Brayton cycle, an Otto cycle, an Allam power cycle, a Kalina cycle, a Rankine cycle, and / or the like. In some cases, a combustion engine 228 may include an internal combustion engine 228. An internal combustion engine 228 may include heat engine in which combustion of fuel occurs with an oxidizer (usually air) in a combustion chamber that comprises a part of a working fluid flow circuit. Exemplary internal combustion engine 228 may without limitation a reciprocating engine (e.g., 4-stroke engine), a combustion turbine engine (e.g., jet engines, gas turbines, Brayton cycle engines, and the like), a rotary engine (e.g., Wankel engines), and the like. Combustion engine is configured to burn the fuel from the fuel source to produce mechanical work. Resulting mechanical work may be used to power the propulsor. Additionally, at least an electric motor of the propulsor may be operatively connected with a fuel cell by way of electrical communication, for example through one or more conductors.Still referring to FIG. 2, in some embodiments aircraft 200 may additionally include an auxiliary power system 220 operatively connected with at least a fuel cell 216. As used in this disclosure, an “auxiliary power system” is a power system, such as without limitation an electrical circuit or mechanical power source, that provides electrical energy to non-propulsor flight components of an aircraft. Exemplary non-limiting non-propulsor flight component include an avionic system, a flight control system, an environmental control system, and anti-ice system, a lighting system, a fuel system, a braking system, and / or a landing gear system. Auxiliary power system 220 may be operatively connected to fuel cell 216 by way of electrical communication, for example through one or more conductors. In some cases, at least a fuel cell 216 may be configured to power auxiliary power system 220. In some cases, auxiliary power system 220 may include a motor configured to convert electric energy to mechanical work. In some cases, motor may be used to operate a compressor, for instance of air conditioning or refrigeration system. In some cases, auxiliary power system 220 may include a motor that is configured to start a combustion engine of at least a propulsor 208.Still referring to FIG. 2, in some embodiments, aircraft 200 may additionally include a second fuel store 224a-b. In some cases, second fuel store 224a-b may be configured to store a second fuel. In some cases, second fuel may be different than a first fuel. Second fuel may include any fuel described in this disclosure, including without limitation kerosene-based fuels. In some cases, at least a propulsor 208 may include at least a combustion engine configured to burn second fuel thereby producing mechanical work, which is used to power the at least a propulsor 208.Still referring to FIG. 2, in some embodiments, one or more of first fuel store 212a-b and second fuel store 224a-b may include at least a fuel environment control mitigation. As used in this disclosure, a “fuel environment mitigation” is any design parameter selected to control an environmental factor associated with fuel within a fuel store. In some cases, fuel environment control mitigation may include a design parameter that affects one or more of fuel pressure, fuel temperature, fuel phase, and the like. For example, in some cases, a fuel environment control mitigation may include insulation to control fuel temperature. Additionally, or alternatively, in some cases, fuel environment control mitigation may include a pressure vessel within which fuel pressure may be controlled.Now referring to FIG. 3A, shown is an exemplary embodiment of a fuel tank 408. A key component that may heavily impact shape of fuel tank 408 is pressure. Fuel tank 408 may be far lighter if pressure is resisted in tension as compared to bending. In tension, thin-walled tanks may be used. Thin-walled tanks may be lighter than thick-walled tanks as less material is used. In some embodiments, tension in a tank may be generally achieved by shapes that provide a circular cross section, including spheres, cylinders, and cones. In an embodiment, a pressurized tank of a given volume may be made as a sphere to place the tank in tension. Alternatively, a tank made as a cube may require tank walls to operate in bending; thus, the cube tank would likely be vastly heavier than a sphere tank of similar volume. Cube tanks may be heavier as thick walls may be necessary to resist bending. Accordingly, in some embodiments, any tank geometry may provide tank walls acting in tension. Tank geometry is further described in FIGS. 3B-J. Force and stress of tank 408 may be considered in a cylindrical portion 304 and an end of the cylindrical portion. Stress within cylindrical portion may be calculated according to: nr2P Pr Stress = - = —2nrt 2t where, r is radius of the cylinder, P is pressure, and t is thickness of tank wall. Force at end of the cylinder may be found according to:Force = nr2PA key component that may heavily impact shape of the tank may be pressure. Tank may be far lighter if pressure is resisted in pure tension as compared to bending. In some embodiments, pure tension in a tank may be generally achieved by shapes that provide a circular cross section, including spheres, cylinders, and cones. In an embodiment, a pressurized tank of a given volume may be made as a sphere to place the tank in pure tension. Alternatively, a tank made as a cube may require tank walls to operate in bending; thus, the cube tank would likely be vastly heavier than a sphere tank of similar volume. Accordingly, in some embodiments, any tank geometry may provide tank walls acting in tension.

[0002] Still referring to FIG. 3A, the walls of tank may operate at or below a limit stress. A “limit stress” is a threshold stress below which tank may operate at to avoid failure or damage. Stress may be defined in hoop direction or longitudinal direction. As shown in FIG. 3A, stress of a thin-wall cylindrical tank may be calculated in the hoop direction by multiplying the pressure(e g., in lb . / in2) by the radius (e g., in inches), and then dividing that value by the tank wall thickness (e.g., in inches). Also as shown in FIG. 3A, the stress of a thin-wall cylindrical tank may be calculated in the longitudinal direction by dividing the hoop stress in half. The mathematical equation shown for stress may be found by first calculating a force by multiplying pi by the radius squared and the pressure. The force is then divided by 2*7i*r*t to calculate the stress. Thus, the equation simplifies to half the hoop stress, or P*r / 2*t. The maximum stress in a cylindrical tank is the vector sum of the hoop and longitudinal stresses. A thin-wall hemisphere provides equal stress everywhere of P*r / 2*t. A cylindrical tank may be fabricated with hemispherical end caps. As used in this disclosure, an “end cap” is an end of a tank and / or a compartment of a tank. End caps, like tanks themselves, may have curvature. Generally, end caps may have a continuously convex differentiable curvature, for example spherical. In some cases, end caps may be nearly spherical, elliptical, cylindrical, or the like. In some cases, an end cap may be spherical and have a radius similar to that of the cylinder which it is capping. In this case, the end cap may be considered ideally selected for resisting stresses, for example with a hemispherical end cap. However more volume within the aircraft cabin is likely wasted from the bulging end cap. Instead in some cases, end cap may be selected to compromise resistance to stress in exchange for improved cabin volume utilization. For example, where end caps are spherical and have twice the radius of the cylinder they are capping, stresses in all parts of the tank may be similar or below that of those stresses within the cylindrical portion of the tank (assuming tank skin thickness is constant) and the cylindrical portion may be lengthened to increase tank volume. If the end caps have twice the radius of the cylinder, and the skin thickness is everywhere the same, then the stresses in all parts of the tank may be similar. In an embodiment, the fuel tanks of the plurality of fuel tanks each include domed end caps. Domed end caps may have twice the radius of the cylindrical portion of the fuel tank 408, therefore allowing the fuel tanks to have equal stress and pressure throughout the whole tank.Now referring to just FIG. 3B, an isometric view of an exemplary embodiment of a conical tank is illustrated. Tank may be a conical tank. A “conical” tank is a type of tapered tank that has a cone-shaped tank geometry. Conical pure-tension tank shapes may include spherical and cylindrical shapes, possibly with spherical end caps. For example, a tank could be conicalwith a spherical end cap. This might resemble an ice cream cone. This cone may be truncated, with another spherical end cap on the opposing end.Now referring to FIG. 3C, an isometric view of an exemplary embodiment of a curved axisymmetric tank is exhibited. Tank may be a curved axisymmetric tank. Another type of tapered tank, a “curved axisymmetric tank” has a circular cross-section with compound curvature on the sides or ends. Curved axisymmetric tank may be capped with hemispheres. A curved axisymmetric tank may more efficiently fill a volume with variable depth.Now referring to FIG. 3D, an isometric view of an exemplary embodiment of a doublecurved tank is presented. Tank may be a double-curved tank. A “double-curved tank” occurs when a tapered tank is merged with a similar or mirror-image tank with a central septum. The two tanks comprising a double-curved tank may be intersected along their length and a septum may be placed at the tank junction to address the resulting tension. In this disclosure, a “central septum” is a partition centrally located in a system separating two compartments. In some cases, central septum may not be parallel to the tank axis; for example, it may be favorable to fill a volume of constant width with a curved, variable-height ceiling. Tank axis may then be adjusted so that the tank wall on the outer side of the tank may be a selected distance from the compartment wall.Now referring to FIG. 3E, a double-curved tank need not have identical compartments; a front quarter view of an exemplary embodiment of a dual tank with different diameters is illustrated. Tank may be a dual tank with different diameters. In some compartments in this embodiment, height on one side of the compartment may be lower than on the other. Two or more compartments in tank may have different diameters if the ceiling height is different across the compartment. The two or more tanks may be joined with one or more septa that may form a curved surface as seen in top view, see cambered tanks below. In an embodiment, the shapes of the compartments may differ. In this case, the two merged tanks (dual tank) may have different diameters to maximize their height along their length.Now referring to FIG. 3F, some tapered tanks may be sheared so that instead of following a straight centerline, tank follows a curved camber line. Shown in FIG. 3F is a quarter side view of an exemplary embodiment of a cambered, tapered tank. The camber line enables the tank to more efficiently fit a volume with variable depth but one flat side, for example a floor, whileconforming more closely to, for example, a curved ceiling. This results in a centerline that may be curved as seen in the figure. An example may be shown in wireframe and surfaced views. This tank geometry may have a flat bottom and a curved top.Now referring to FIG. 3G, an isometric and side quarter views of an exemplary embodiment of a dual-cambered tank is presented. Tank may be a dual-cambered tank. A “dualcambered” tank is the same as normal cambered tank, but the camber line may be curved from the top view as well as the side view. Two cambered lines can place the outer surface of the tank at a selected distance from the compartment wall. The bottom and right edge of tank may be straight while the top view and side view of camber line may be curved. Circular cross sections of the tank may be sheared so that they remain circles in the lateral -vertical plane, for example. Alternatively, circular cross sections may be orthogonal to the camber line. From a stress standpoint, wherever the camber line curvature of the tank is modest, the stress difference is probably very small.Now referring to FIG. 3H, a quarter front view of an exemplary embodiment of a double tank is shown. The double tank, as explained above, may have a curved septum separating the two compartments of tank.Now referring to FIGS. 31 and 3J, a front quarter view of an exemplary embodiment of a multi-bubble tank is illustrated in both figures. Given a rectangular compartment cross-section with a longitudinally oriented tank, this cross-section may be occupied by a single circular crosssection tank. Or, as noted above, a double tank may be used to provide greater cross-section area within the rectangular compartment. Additional bubbles may be added to fill in the four corners. A “multi-bubble tank” is a tank that has more than two compartments attached together. In some cases, two more bubbles may be added to fill in the valleys between the two main tanks. Multibubble tanks may have any number of compartments, but there may be a diminishing return on increasing complexity; either engineering judgment or actual engineering may be applied. In an embodiment, a muti-bubble tank may have four lobes added to fill in the corners of a notional envelope indicated by the lines in the figure. This provides a more valuable tank volume for a given compartment volume. Each compartment may have a circular cross section as shown, which may be trimmed to the large, main lobes. The main lobes may then be trimmed to small lobes. Each junction may be then faced with a septum. This “trimming” can be seen in FIG. 3 J,wherein the tank is trimmed with septa. On the other hand, the multi-bubble tank in FIG. 31 does not have trimming. Tank may or may not have trimming.Now referring to FIG. 3K, an exemplary embodiment of an overview of a vertical axis “air mattress” tank is exhibited. An “air mattress” tank refers to a multi-bubble tank, where each of a lobes’ upper and lower surfaces have a common center point. When the lobes of at least a tank 408 do not form a near-circular shape, the surface of at least a tank 408 may deform in order to achieve pure tension. As shown in the figure, the “air mattress” may extend laterally. Additionally, for example, “air mattress” may extend longitudinally as shown in FIG. 3L. In an embodiment and as described above, at least a tank 408 may be a vertically oriented multibubble tank wherein its pressurized walls and its septa may extend from a lower outer mold line to an upper outer mold line of a blended wing body. A top and bottom of at least a tank 408 may be closed out by end caps inset from the outer mold line to provide room for the outer mold line skin’s supporting structure. This may assist BWB 104 in resisting pressurization and simultaneously carrying the shear stress that is otherwise carried by structural members of BWB 104. Also, as shown, at least a tank 408 may extend across the full width of the cabin and outboard cargo bays.With continued reference to FIG. 3K, in some embodiments, at least a tank 408 may connect the at least a tank to at least an upper skin and at least a lower skin of the blended wing body. Furthermore, the upper and lower skins of BWB 104 may be structurally supported by one or more septa of at least a tank 408. In this disclosure, the “upper and lower skins” of the BWB 104 refer to the upper and lower surfaces of an outer mold line of BWB 104. In some cases, center body of BWB 104 may be pressurized, relative ambient pressure which is typically below atmospheric (sea level) pressure at elevation. Therefore, in some cases, an upper skin and a lower skin may be forced away from one another by a resultant pressure differential. In some embodiments, additional structural support between walls of at least a tank 408 (e.g., septa and / or tank walls) and upper and lower skins of BWB 104 may resist pressure. In some cases, use of at least a tank 408 as a structural element of BWB 104 may reduce weight. In FIG. 3K, walls of at least a tank 408 may be shown to extend to upper and lower skins, while end caps of at least tank 408 may be shown inset to provide clearance from the structure that supports the skins.Now referring to FIG. 3L, an exemplary embodiment of an overview of longitudinal arrangement of a vertical axis “air mattress” tank is shown. As said above, the “air mattress” extends laterally but could also, for example, extend longitudinally. FIG. 3L shows at least a tank 408 within a center bay. There are many ways to arrange vertical “air mattress” type tanks and this longitudinal arrangement may provide mutual support between BWB 104 and at least a tank 408. In some embodiments, a length of at least tank 408 into BWB 104 may be varied by either varying a diameter of at least a tank 408 or by doubling (or more) a number of rows of at least a tank 408. In a doubled tank, there may be a center lateral septum that replaces what is shown as the forward curved tank walls; this septum may be planar. In some cases, doubling (or more) the number of rows of at least a tank 408 may result in a smaller tank wall radius and a thinner and lighter tank wall (when compared to a single row of equal total longitudinal dimension).

[0003] Referring now to FIG. 3M, a general case for a tank 700 having multiple compartments is illustrated. FIG 3N illustrates a cross section of a tank 700 having three compartments, a first compartment A, a second compartment B, and a third compartment C. Each compartment is approximated by a continuously convex differentiable curve, e.g., a circular section. The cross-section has three intersections between walls for each compartment, AB, AC, and BC. A junction 720, i.e., septum, runs from each intersection. FIG. 3M illustrates a general case, wherein each junction runs from an actual point of intersection toward an imaginary intersection between walls of intersecting compartments. For example, junction between compartment A and compartment B starts at intersection AB runs toward, a second imaginary intersection between curves AB (as represented by broken lines). In some embodiments, arranging the septa in this manner ensures that stress at each junction is substantially in tension. As can be seen in FIG. 3M, each junction runs (toward the imaginary intersection) until reaching a support (another junction or compartment wall). In some cases, this arrangement ensures that each junction is substantially in tension at each intersection of junctions.Referring now to FIG. 4, an exemplary embodiment of a method 400 for fueling an aircraft is illustrated. As used in the current disclosure, a “fuel tank” is a container specifically designed to hold fuel. A fuel tank may also be referred to as a “tank.” A fuel tank 408 of a plurality of fuel tanks may include one or more compartments to store fuel in. A fuel tank 408 of a plurality of fuel tanks may be a part of fuel delivery system for an engine, in which the fuelmay be stored inside a fuel tank 408 and then propelled or released into an engine, such as without limitation a combustion engine. A fuel tanks may be configured to be removably attached to an aircraft 412. A fuel tank 408 may be configured to contain enough fuel for multiple flights or trips. A fuel tank 408 may also be configured to hold an amount of fuel that is predetermined by the flight plan of the aircraft. In some cases, fueling may be performed at a slow rate over a long period of time. For example, in some cases, fueling may take one hour, two hours, three hours, five hours, or eight hours. In some cases, aircraft may be fueled overnight, or when substantially no flights are scheduled for a long period of time, i.e., not between changeovers between flights.With continued reference to FIG. 4, method 400 may include a fuel tank 408. In an embodiment, a fuel tank 408 stores fuel to power aircraft 412. Fuel tank 408 may be permanently attached to aircraft 412. Fuel tank 408 may be a pressure vessel. Fuel tank 408 may further have a tank geometry. “Tank geometry” refers to an overall shape and arrangement of fuel tank 408. Fuel tank 408 may have a multi-lobe geometry, such that the multi-lobe geometry includes one or more curvatures. As used herein, “multi-lobe geometry” refers to a shape include multiple curvatures or “lobes.” A “lobe” as used herein, is a curved section of the multi-lobe geometry. Multi-lobe geometry may have multiple lobes, wherein the start of one lobe is marked by a discontinuation in the curvature. A discontinuation in the curvature may occur when the curvature of a tank switches between two non-adjacent values instantaneously or substantially instantaneously. The curvature of the tank may switch substantially instantaneously when there is a weld or other fastener or a division such as a septum. In other embodiments, the curvature may switch substantially instantaneously, when, for example, two lobes of a tank meet, but the intersection has been rounded or otherwise altered in order to, among other things, reduce stress concentrations at the intersection point. For example, the fuel tank 408 may have one or more surfaces with spherical and / or cylindrical shapes. Each lobe of the multi-lobe tank geometry may be configured to have a different radius. Fuel tank 408 may include variable diameter and length. A lobe may be defined as a single sphere or cylinder of the plurality of spheres and / or cylinders. According to some embodiments, tank geometry for a blended wing body aircraft may be driven by at least five objectives: (1) to provide as much fuel volume as possible while using little payload floor space, (2) to provide a fuel tank shape that resists pressure and that is lightweight,(3) to provide a fuel tank shape that can be insulated between outer wall and inner wall, (4) to provide a fuel center of gravity that is not widely misaligned with the aircraft’s center of gravity, and (5) to provide fuel tanks that are compatible with a passenger cabin. In this disclosure, a “fuel center of gravity” is the center of gravity of the fuel inside fuel tank 408. In some cases, fuel center of gravity may affect overall aircraft center of gravity. Fuel center of gravity may be associated with a volume within airplane that has sufficient volume to store a practical quantity of fuel in discrete tanks. Fuel tank 408 may be arranged inside blended wing body as a function of its fuel center of gravity in relation to center of gravity of aircraft 412. Tank geometry is further discussed herein with reference to FIGS. 3A-J.With continued reference to FIG. 4, aircraft 412 includes a fuel tank 408, which may further include a liquified gas fuel, and may include at least a vent, and an insulation. Fuel tank 408 may be a permanent tank. A “permanent tank” is a tank that is permanently attached to aircraft 412. A permanent tank may be added to an existing airplane or attached during the airplane’s construction. In either case, the intention of a permanent tank may be to leave fuel tank 408 in place for an extended period. A permanent tank may integrate fuel tank 408 and BWB 104 structure so that the BWB 104 supports fuel tank 408, and / or fuel tank 408 supports BWB 104. As a non-limiting example, Fuel tank 408 may support BWB 104 by assisting the airframe of the BWB 104 in resisting pressurization and carrying shear that would otherwise be carried by the ribs of the BWB 104. In some embodiments, curved surfaces of fuel tank 408 may not be well suited to assisting structural surfaces of BWB 104. However, in some embodiments, a septa of multi-lobe tanks may be well-suited to support the structure of aircraft 412. In some embodiments, septa of multi-bubble tanks may be well-suited to carry structural loads, for example loads intrinsic to the tank as well as loads extrinsic to the tank (e.g., loads resulting from pressurization of aircraft). As used herein, “septa” are partitions between objects, such as two fuel tanks. The plurality of fuel tanks may be divided by septa such that there is a septum between each fuel tank. In an embodiment, the septa may extend from a lower outer mold line to an upper outer mold line of a fuel tank 408. A top and bottom of fuel tank 408 may be closed out by spherical end caps inset from the outer mold line to provide room for the outer mold line skin’s supporting structure. This may assist aircraft 412 in resisting pressurization andsimultaneously carrying the shear stress that is otherwise carried by the ribs of the airframe of aircraft 412.Still referring to FIG. 4, fuel tank 408 is filled with a liquified gas fuel. Liquified gas fuel may include a liquid hydrogen fuel. Liquified gas fuels include without limitation liquid hydrogen, propane, and liquified natural gas. Liquified gas fuel may be stored as a compressed gas or in liquid form. Liquid hydrogen fuel may be used for airliner-scale systems, liquid hydrogen fuel may a higher density than gaseous hydrogen; even so, about four times as much liquified gas fuel may be needed compared to the volume of Jet-A fuel needed. Additionally, liquified gas fuel may also allow for reduced tank pressure and tank weight. Liquified gas fuel may be kept at extremely low, cold temperatures, for example below its critical point of 33 Kelvin. Gas turbine engines, such as combustion engines, may operate on gaseous hydrogen and may transition liquified gas fuel into a gas before consumption; gaseous hydrogen may be supplied to the engine at high pressure, but typically no fuel pump is used. In some embodiments, the gas may squirt into a combustion chamber due to compressor stages. Hydrogen vapor from tank ullage may be combined with the hydrogen vapor from an output of a fuel heater. In some cases, this combined channel may then lead to engine for combustion. When boiled, a rate of boil-off for liquified gas fuel may be determined as a function of heat energy applied to liquified gas fuel. Boil-off for liquified gas fuel can be selected, but fuel tank 408 may be insulated with an insulation to control application of heat to tank contents. Fuel tank 408 may also be refrigerated to remove heat from liquified gas fuel. The degree of insulation may be selected to provide a desired rate of boil-off. Insulation is further described below. The selected rate of boil-off for liquified gas fuel may generally be less than the rate at which liquified gas fuel is consumed by the engines and possibly other aircraft systems, such as without limitation an APU. Excess boil-off may be dumped overboard or outside of aircraft through a vent.Continuing to refer to FIG. 4, fuel tank 408 may include at least a vent. As used in this disclosure, a “vent” is an opening and / or aperture configured to allow one or more fluids to pass. In an embodiment, at least a vent may be configured to vent gasses from fuel tank 408. In some embodiments, these gasses may include gaseous hydrogen. Gaseous hydrogen may result from boil-off of liquified gas fuel as the fuel warms. In an embodiment, and without limitation, at least a vent may be configured to vent boil-off from fuel tank 408. In some cases, at least a vent mayinclude a check valve. As used in this disclosure, a “check valve” is a valve that permits flow of a fluid only in certain (e.g., one) directions. In some cases, check valve may be configured to allow flow of fluids substantially only away from fuel tank 408 while preventing back flow of vented fluid to fuel tank 408. At least a vent may also include a pressure regulator. A “pressure regulator” is a type of valve that controls the pressure of a fluid. Venting gasses from a fuel tank 408 prevents over-pressurizing or other events that may cause catastrophic damage or harm. It may also desirable, when aircraft 412 is grounded, to connect a system of lines and tanks to at least a vent to collect the boiled-off hydrogen. In some cases, the collected gaseous hydrogen or fuel gasses can be compressed by a pump into storage tanks and then cooled to liquid temperatures for reuse as aircraft fuel. In embodiments a vent may be in fluid connection to both a fuel tank 408 and an interface port. In other embodiments, a vent may be attached to an interface port and a vent line 424.Still referring to FIG. 4, fuel tank 408 may include insulation. In this disclosure, “insulation” is a component configured to reduce heat transfer. Insulation may be used to reduce thermal transfer to liquified gas fuel inside of a fuel tank 408. Heat may be transferred to fuel tank 408 by at least two means: conduction and radiation. To reduce conduction, insulation may include a vacuum to separate an interior volume of fuel tank 408 from an exterior with an evacuated vessel. Another means to reduce conduction in fuel tank 408 may be an insulating material that inhibits conduction. Insulating materials include fiberglass wool, plastic or ceramic foam, aerogel, and other materials. Heat flow through an insulator can be inversely proportional to its thickness. For example, an insulator that is twice as thick may conduct heat at half the rate. Heat transfer by radiation may be reduced by reflective coatings., Reflective coatings may be located proximal surfaces of fuel tank 408 and / or on intermediate materials within insulation. For example, a dewar may be coated with a mirror-like material that reflects heat radiation; many thermos bottles are silvered for this reason. A “dewar” is double-walled flask of metal or silvered glass with a vacuum between the walls and configured to hold a liquid. Another means to reduce radiation may be to sandwich multiple layers of thin reflective foil within an insulative material such as foam. Additionally, a modest amount of insulation may be needed to limit boil-off to a rate below that needed to provide fuel vapor to the engines in cruise. On the ground, a system to capture boil-off liquified gas fuel may be provided. Another characteristic requirement forinsulation includes surface area of a fuel tank 408. A very large tank may provide a large volume of liquified gas fuel per unit surface area. Boil-off rate may be determined by heat transferred into fuel tank 408. Heat transferred into the fuel tank 408 may be a function of tank surface area and / or insulation effectiveness. For example, heat transferred into fuel tank 408 may be proportional to tank surface area and / or inversely proportional to insulation effectiveness. For a given insulation and storage volume, fuel tank 408 having a larger surface area may result in more boil-off. Or, for a given boil-off rate and storage volume, fuel tank 408 having a larger surface area will need more insulation. As explained above, degree of insulation may be selected to provide a desired rate of boil-off. Insulation may also include a chamber located between the inner wall and the outer wall foams, aerogels, reflective materials, and the like of fuel tank 408. Chamber may contain gas such as air, nitrogen, argon, or the like. In some cases, gas may be actively pumped into the chamber to ensure that the gas within the chamber is clean and dry and thereby not conducive to condensation, freezing or contamination.Continuing to refer to FIG. 4, fuel tank 408 of the plurality of fuel tanks are located aft of the cabin in the main body. As used herein, “cabin” is the portion of the aircraft that holds the crew, passengers, and cargo. Fuel tanks may be stored at least partially aft of the cabin, near the propulsors. For the purposes of this disclosure, the fuel tanks are stored at least partially aft of the cabin if at least a portion of the fuel tanks extends behind or “aft” of the rearmost portion of cabin. In an embodiment, because of the low density of liquified gas fuel, storing fuel tanks behind the cabin of the aircraft 412 may not substantially affect the longitudinal center of gravity of the aircraft 412. FIG. 3D shows a depiction of a multi-lobe tank configuration. In a multi-lobe tank configuration, in some embodiments, each fuel tank 408 does not form a completely circular shape, such that the surface of fuel tank 408 may deform to achieve tension. Multi-lobe tank configuration may be beneficial as it provides more tank volume compared to a singular spherical tank. Multi-lobe tank configuration may be derived by adding tanks in the junctions between tanks. Junctions are discussed in further detail in FIG. 5. Fuel tanks, as discussed above, may provide structural support to the aircraft by acting as load bearing columns between the floor and ceiling of the aircraft 412. In an embodiment, fuel tanks may be mounted within the airframe. As used herein, an “airframe” provides structure to an aircraft. Airframe may be a part of the structural components of the aircraft 412. The plurality of fuel tanks may span across thefull width of the aircraft 412. In embodiments, as a non-limiting example where the ceiling of the aircraft is downward sloping, each fuel tank 408 of the plurality of fuel tanks may vary in diameter and length. Fuel tanks may be stored vertically or horizontally. In an embodiment and as described above, fuel tank 408 may be a vertically oriented multi-lobe tank wherein its pressurized walls and its septa may extend from a lower outer mold line to an upper outer mold line of a fuel tank 408. A top and bottom of fuel tank 408 may be closed out by spherical end caps inset from the outer mold line to provide room for the outer mold line skin’s supporting structure. This may assist aircraft 412 in resisting pressurization and simultaneously carrying the shear stress that is otherwise carried by the ribs of the airframe of aircraft 412. Fuel tanks may be stored in rows aft of the cabin. In an embodiment, there may be two rows of fuel tanks. In doubling the rows, each fuel tank 408 had a smaller tank wall radius, therefore each fuel tank wall is thinner and lighter.With continued reference to FIG. 4, a Computing device 404 may engage a flight preparation 416 as a function of a fdl datum or a desired level. As used in the current disclosure, a “flight preparation” is a process of preparing the fuel tank for flight. The flight preparation 416 may include detaching fuel line 420 and vent line 424 as a function of fdl datum of desired level. Fuel line 420 and vent line 424 may be attached when fdl datum indicates fuel tank 408 while actively engaged in the fueling process. Fuel line 420 and vent line 424 may be detached when fdl datum indicates fueling process has been completed or the plane is in flight or flight preparations. The fueling process is generally completed when the fuel tank 408 is full or has enough fuel to complete one or a plurality of flight plans. The flight preparation 416 may include adjusting the vent for flight conditions, so that the vent may safely vent fuel gasses during flight. The flight preparation 416 may include capping the fuel tank. In some cases, flight preparation 416 may include preparing actively conditioning a cavity of the fuel tank. For example, the fuel tank may include a chamber between an inner wall and an outer wall of the tank. In some cases, clean dry gas (e.g., air, nitrogen, argon, etc.) may be pumped into the chamber. The clean dry gas may prevent condensation and freezing of humidity against the inner wall, thereby aiding in insulating the liquified gas fuel. The clean dry air may also be vented from the chamber and / or recirculated.Continuing to refer to FIG. 4, a fuel line 420 may be used to transfer liquid nitrogen fuel into fuel tank 408. As used in the current disclosure, a “fuel line” is a hose configured to transport fuel from a fuel reservoir to a fuel tank. In embodiments, a fuel line may be comprised of a hose, tube, or conduit. A fuel line may be made of rubber, plastic, steel, copper, or any combination thereof. Fuel lines may be configured to be able to contain pressurized liquified gas fuel. In some embodiments, liquified gas fuel may include liquid hydrogen fuel. Fuel line 420 may be in fluid connection with a liquified gas reservoir. Liquified gas fuel may be pumped from a liquified gas reservoir through fuel line 420 into fuel tank 408. The fuel line 420 may be connected to the fuel tank 408 through an airtight or a watertight seal. A seal may be made of a rubber compound. In embodiments, Fuel line 420 may be removably attached to a Fuel tank 408. For the purposes of this disclosure, “removably attached” means attached to an object such that it may be removed without damaging the object. The attachment of the fuel line 420 may be determined as a function fill datum. Computing device 404 may be detach a fuel line 420 unlocking the latching mechanism between fuel line 420 and fuel tank 408. The latching mechanism may only be unlocked when there is no fuel passing through the fuel line 420. Once the latching mechanism is unlocked the fuel line 420 may be configured to automatically be stored on a fuel line storage device. This device may store fuel line 420 on a reel or spool as a function of fill datum.Continuing to refer to FIG. 4, a vent line 424 in fluid connection with a fuel tank 408. As used in this disclosure, elements are in “fluid connection” when a fluid may move from one element to another, notwithstanding any temporary obstructions such as valves. In embodiments, a vent line 424 may be configured to be attached to a vent of a fuel tank 408. A vent line may be configured to capture gas from fuel tank 408 as a result of a vent boil off. A vent line 424 may be made of rubber, plastic, steel, copper, or any combination thereof. Vent lines 424 may be configured to be able to contain pressurized liquified gas fuel. Vent line 424 may be in fluid connection with a reservoir of liquified gas. In embodiments, Vent line 424 may be removably attached to a fuel tank 408. The attachment of the fuel line 420 may be determined as a function fill datum.Continuing to reference FIG. 4, fuel tank 408 is connected to a vent line 424. As used herein, a “vent line” is a tether or a bundle of tethers, e.g., hose, tubing, cables, wires, and thelike, which is configured to removably attach with a mating component of a vent of a fuel tank 408. Vent line 424 may be made of a rigid or a flexible material. For example, vent line 424 may be composed of polypropylene, polycarbonate, acrylonitrile butadiene styrene, polyethylene, nylon, polystyrene, polyether ether ketone, and the like. Vent line 424 may also be composed of metals such as carbon fiber, aluminum, titanium, copper, or the like. Vent line 424 may be attached to a vent located at the top of the tank 408 such that vent line 424 directs gaseous fuel from the vent to an external fuel tank. External fuel tank may be consistent with any fuel tank as disclosed herein. Vent line 424 may be configured to direct gaseous fuel overboard, wherein overboard means off of the aircraft. While aircraft is grounded, aircraft engine and other systems are not using hydrogen vapor. Because of this, boil-off from the liquid fuel is collected through a vent line 424 to an external fuel tank.Continuing to reference FIG. 4, vent line 424 may be insulated to prevent the accumulation of frozen gases, such as frozen water vapor. The boiling point of hydrogen is - 423.2 F, which may be the freezing point of other gases. For example, water / water vapor freezes at 32 F. In an embodiment, insulation may include vermiculite, fiberglass, X-Aerogel, crosslinked Aerogel, firebrick, or any other insulation with a low thermal conductivity coefficient. A low thermal conductivity may be any thermal conductivity less than 0.10, 0.01, or 0.001 W / m-k. Insulation may limit heat transfer by conduction. Insulation may surround the vent line. In some embodiments, insulation may surround the circumference of vent line. In some embodiments, insulation may partially surround the circumference of the vent line. Alternatively, or additionally, vent line 424 may include an inner wall and an outer wall, wherein a vacuum resides in between. As used herein, an “inner wall” is the inner barrier of a vent line that would be in contact with the gaseous fuel. As used herein, an “outer wall” is the outer barrier of a vent line that is in contact with an environment outside the vent line. In some embodiments, there may be insulation between the outer and inner wall. In some embodiments, there may be a vacuum as an insulation between the outer and inner wall.With continued reference to FIG. 4, additionally or alternatively, a heat exchanger may be disposed proximal vent line 424. As used in this disclosure, a “heat exchanger” is a device that is configured to transfer heat from a first fluid or medium to a second fluid or medium. In some cases, a heat exchanger may include a hot fluid channel, which may be in thermalcommunication with the vent line 424. A “channel,” as used herein, is a component that is substantially impermeable to gases and contains and / or directs a flow of hot fluid. A Channel may act as a heat exchanger between the hot fluid and the cold gaseous fuel. Hot fluids may include oils, air, water, propylene glycol, ethylene glycol, or the like. Additionally, or alternatively, heat may be provided by a powered heat source such as a radiator, electric resistance heater, combustion heater, or the like. Powered heat source may be placed in thermal communication with the vent line 424, for instance to heat the cold gaseous fuel. Powered heat source may be placed in thermal communication with hot fluid channel to heat the hot fluid that may, for example, heat the cold gaseous fuel. In some embodiments, powered heat source may be in thermal communication with a length of vent line, for example to heat the cold gaseous fuel. In some embodiments, powered heat source may be in thermal communication with a length of hot fluid channel, for example, to heat the hot fluid. In some embodiments, heat elements may be in thermal communication with the vent line to prevent an accumulation of frozen air and / or water vapor.With continued reference to FIG. 4, fuel tank 408 may be configured to be filled to a desired level 432. As used in the current disclosure, a “desired level” is when the tank is filled with enough fuel for a plurality of flights 436 plus additional reserve fuel. A sensor 428 may be used to determine when a tank 408 is at a desired level 432. A desired level 432 may include enough fuel to several flights in one day. In an embodiment, a desired level 432 may include filling the tank to capacity or filling the tank to a predetermined fraction of the capacity of the fuel tank 408. Desired level 432 may be determined as a function of weight or refueling time. Reserve fuel may include approximately 10%-20% more fuel than what is required to complete a plurality of flights. A desired level 432 may additionally comprise filling the fuel tank to capacity. As used in the current disclosure, “filling the fuel tank” comprises transferring fuel from a fuel reservoir into a fuel tank. Fill datum may also be considered when determining if a tank is at a desired level 432.With continued reference to FIG. 4, method 400 includes a sensor 428. Sensor 428 may be connected to fuel line 420, vent line 424, fuel tank 408, or components thereof. As used in this disclosure, a “sensor” is a device that is configured to detect a physical characteristic and / or a phenomenon and convert the detection into a signal, such as an output sensor signal. In one ormore embodiments, sensor 428 may be configured to transmit information, such as fill datum, related to a detection. As used in the current disclosure, a “fill datum” is a datum related to the fueling process of fuel tank 408. For example, and without limitation, fill datum may include information related to failure of the fuel tank 408, fuel line 420, and vent line 424. In some embodiments, fill datum may be related to any abnormal phenomena, such as and without limitation, temperature, voltage, current, pressure, fuel levels and the like as it relates to fuel tank 408. For instance, and without limitation, fill datum may include a high / low temperature of liquified gas fuel or fuel tank 408 directly. In another instance, and without limitation, sensor 428 may detect various characteristics of fuel line 420, vent line 424, fuel tank 408, and / or their surrounding environment that allows for fill datum to be indirectly determined, such as by computing device 404. For example, and without limitation, sensor 428 may determine a voltage, current, and thermal conditions of a surrounding environment of Fuel tank 408 to determine temperature of liquified gas fuel. Fill datum may additionally include the amount of fuel currently in fuel tank 408 in relation to the fuel tank’s 408 capacity. Fuel datum may also be the rate or volume of fuel that is passing through fuel line 420. This may be done using a flow sensor placed within fuel line 420. Or alternatively, the amount of fuel to complete a flight plan or a plurality of flight plans. As another non-limiting example, fill datum may include a pressure of fuel tank 408 that that is generated as a function of liquid nitrogen fuel. Fill datum may include an indication that a fuel tank 408 may need to be prepared for flight. Fill datum may be related to the fuel tank 408 needing maintenance. Fill datum may also include the weight of the fuel. For example, and without limitation, sensor 428 may include a humidistat, hygrometer, voltage sensor, current sensor, multimeter, voltmeter, ammeter, electrical current sensor, resistance sensor, impedance sensor, capacitance sensor, a Wheatstone bridge, displacements sensor, vibration sensor, Daly detector, electroscope, electron multiplier, Faraday cup, galvanometer, Hall effect sensor, Hall probe, magnetic sensor, optical sensor, magnetometer, magnetoresistance sensor, MEMS magnetic field sensor, metal detector, planar Hall sensor, thermal sensor, thermocouple, resistance thermometer, semiconductor-based temperature sensors, thermistor, fuel flow sensor, fuel level sensor, and the like, among others. Sensor(s) 728 may efficaciously include, without limitation, any of the sensors disclosed in the entirety of the present disclosure.With continued reference to FIG. 4, in one or more embodiments, sensor 428 may be communicatively connected to computing device 404, fuel line 420, vent line 424, and fuel tank 408. As used in this disclosure, “communicatively connected” means connected by way of a connection, attachment, or linkage between two or more relata which allows for reception and / or transmittance of information therebetween. For example, and without limitation, this connection may be wired or wireless, direct, or indirect, and between two or more components, circuits, devices, systems, and the like, which allows for reception and / or transmittance of data and / or signal(s) therebetween. Data and / or signals therebetween may include, without limitation, electrical, electromagnetic, magnetic, video, audio, radio, and microwave data and / or signals, combinations thereof, and the like, among others. A communicative connection may be achieved, for example and without limitation, through wired or wireless electronic, digital, or analog, communication, either directly or by way of one or more intervening devices or components. Further, communicative connection may include electrically coupling or connecting at least an output of one device, component, or circuit to at least an input of another device, component, or circuit. For example, and without limitation, via a bus or other facility for intercommunication between elements of a computing device. Communicative connecting may also include indirect connections via, for example and without limitation, wireless connection, radio communication, low power wide area network, optical communication, magnetic, capacitive, or optical coupling, and the like. In some instances, the terminology “communicatively coupled” may be used in place of communicatively connected in this disclosure.With continued reference to FIG. 4, in one or more embodiments, sensor 428 may include one or more sensors. In some embodiments, sensor 428 may be a sensor suite. For example, and without limitation, sensor 428 may include a plurality of sensors, such as for redundancy purposes or confidence level reassurance of each detection by a sensor of the plurality of sensors. Sensor 428 may be a contact or a non-contact sensor. For example, and without limitation, sensor 428 may be physically attached to fuel line 420, vent line 424, and / or fuel tank 408. In other embodiments, sensor 428 may be remote to fuel line 420, vent line 424, and fuel tank 408.With continued reference to FIG. 4, in one or more embodiments, sensor 428 may include a temperature sensor or probe. A temperature sensor may include a thermocouple, thermometer, pyrometer, resistance temperature detector (RTD), platinum resistance temperature detector (PRTD), thermistor, negative temperature coefficient (NTC) thermistor, semiconductor based integrated circuit (IC), microbolometers, local temperature sensor, remote digital temperature sensor, infrared sensor, infrared or visible spectrum imaging device, thermophile infrared sensor, any combination thereof, and the like. Temperature, for the purposes of this disclosure, and as would be appreciated by someone of ordinary skill in the art, is a measure of the heat energy of a system. Temperature, as measured by any number or combinations of sensors present within sensor 428, may be measured in Fahrenheit (°F), Celsius (°C), Kelvin (°K), or another scale alone or in combination. The temperature measured by sensors may comprise electrical signals, such as output sensor signal, which are transmitted to their appropriate destination, such as Computing device 404, using a wireless and / or wired connection.With continued reference to FIG. 1, in other embodiments, sensor 428 may also include other sensors to detect the amount of fuel in fuel tank 408. This may occur my measuring the amount of fuel that pass through a fuel line 420. This may also take into account the amount of fuel that has vented through vent line 424. In embodiments, Sensor 428 may include a hydrogen sensor. A hydrogen sensor may be configured to detect both gaseous and liquid hydrogen. Sensor 428 may be configured to determine the capacity of fuel tank 408 for liquid hydrogen. Sensor 428 may a be able to determine the amount of liquid hydrogen in the fuel tank 408. In one or more embodiments, sensor 428 may include a plurality of independent sensors, where any number of the described sensors may be used to detect any number of physical quantities associated with liquid / gaseous hydrogen fuel, fuel line 420, vent line 424, fuel tank 408, and / or components thereof. Independent sensors may include separate sensors measuring physical quantities that may be powered by and / or in communication with circuits independently, where each may signal sensor output to a control circuit such as a user graphical interface. In an embodiment, use of a plurality of independent sensors may result in redundancy configured to employ more than one sensor that measures the same phenomenon, those sensors being of the same type, a combination of, or another type of sensor not disclosed, so that in the event onesensor fails, the ability of sensor 428 to detect phenomenon, such as fill datum of fuel tank 408, may be maintained.With continued reference to FIG. 4, in some embodiments, sensor 428 may include a pressure sensor. “Pressure,” for the purposes of this disclosure, and as would be appreciated by someone of ordinary skill in the art, is a measure of force required to stop a fluid from expanding and is usually stated in terms of force per unit area. The pressure sensor that may be included in sensor 428 may be configured to measure an atmospheric pressure and / or a change of atmospheric pressure. In some embodiments, the pressure sensor may include an absolute pressure sensor, a gauge pressure sensor, a vacuum pressure sensor, a differential pressure sensor, a sealed pressure sensor, and / or other unknown pressure sensors or alone or in a combination thereof. The pressure sensor may include a barometer. In some embodiments, the pressure sensor may be used to indirectly measure fuel flow, Fuel tank level, liquified gas fuel level, and altitude. In some embodiments, the pressure sensor may be configured to transform a pressure into an analogue electrical signal. In some embodiments, the pressure sensor may be configured to transform a pressure into a digital signal.With continued reference to FIG. 4, in one or more embodiments, sensor 428 may include a moisture sensor. In one or more embodiments, sensor 428 may include a sensor suite which may include an array of sensors that may detect similar or unique phenomena. For example, in a non-limiting embodiment, sensor suite may include a plurality of temperature sensors, such as, for example, thermometers or a mixture of thermistors and thermometers. Method 400 may include a plurality of sensors in the form of individual sensors or a sensor suite working in tandem or individually. A sensor suite may include a plurality of independent sensors, as described in this disclosure, where any number of the described sensors may be used to detect any number of physical or electrical quantities associated with a fuel tank 408 of an aircraft. Independent sensors may include separate sensors measuring physical or electrical quantities that may be powered by and / or in communication with circuits independently, where each may signal sensor output to a control circuit such as Computing device 404. In an embodiment, use of a plurality of independent sensors may result in redundancy configured to employ more than one sensor that measures the same phenomenon, those sensors being of thesame type, a combination of, or another type of sensor not disclosed, so that in the event one sensor fails, the ability to detect a phenomenon is maintained.With continued reference to FIG. 4, fuel tank 408 may be configured hold enough fuel to successfully complete a plurality of flights 436. As used in the current disclosure, a “plurality of flights” may include conducting two or more flights in succession using the aircraft. In some embodiments, aircraft may receive substantially no fuel in between flights of the plurality of flights. In embodiments, a fuel tank 408 may be configured to refuel prior to the start of the plurality of flights 436 (and after the end of the plurality of flights). In other embodiments, the fuel tank 408 may be configured to refuel then make a plurality of flights 436 without refueling in between flights. A benefit of this embodiment may include shorter layover times between the plurality of flights 436, because the aircraft does not need to refuel. Once the plurality of flights 436 has been completed the aircraft may then return to the original airport / refueling location for refueling or another airport that offers liquified gas fuel. The plurality of flights 436 may begin and end at the same location or airport. In an embodiment, the plurality of flights 436 may begin at a first refueling station and terminate at a second at a refueling station. As used in the current disclosure, a “refueling station” is an airport that is equipped to refuel aircraft that use liquified gas fuel. The tank 408 may only be filled at locations with a refueling station. In an embodiment, a first refueling station may be located in California and a second refueling station may be located in Texas. A multitude of flights may occur in between the aircrafts departure from a first refueling station to its arrival at a second refueling station. A first and second refueling station may be located at the same or different airports. Refueling commence at any time after the arrival of the aircraft at the second refueling station. Refueling may occur at a continuous rate over an extended period. This may include refueling the fuel tank 408 overnight or refueling over several hours ranging from .5-10 hours. Refueling may also take place in 30 minutes or more. Once the tanks have been refueled, the aircraft then may be prepared for flight. In some embodiments, methods presently described provide an improvement of flight fueling regimes, as changeover times may be reduced and more flights operating on liquified gas fuel may be offered (even between airports which lack those refueling capabilities). In some cases, a computing device may be configured to determine desired level 432, for example as a function of plurality of flights, fuel consumption, and fuel reserves.Now referring to FIG. 5, an isometric view of an exemplary embodiment of a tank 500 with a plurality of tank support links is shown. Fuel tank 408 needs to be mounted into aircraft 100, usually inside the body, in order to be a permanently attached. As explained above, fuel tank 408 may be a permanent tank, meaning it may be mounted to the aircraft for an extended period of time. Fuel tank 408 and aircraft 100 may be constructed separately, and then fuel tank 408 may be mounted within. In an embodiment, fuel tank 408 and BWB 104 may be independent because the structural load paths of Method 400 may not pass-through fuel tank 408, and the structural load paths of fuel tank 408 may not pass through BWB 104, except insofar as necessary to restrain fuel tank 408 within BWB 104. Fuel tank 408 may alternatively, or additionally, provide additional structural support to the airframe. Given this constraint, there may be at least two methods for tank support links to mount fuel tank 408 to BWB 104: rigid and linked mounts.Still referring to FIG. 5, fuel tank 408 may be mounted to BWB 104 using a rigid mount. Fuel tank 408 may be a standalone structure and may be connected to the BWB 104 with one or more rigid connections. A “rigid mount” or “rigid connection” is a type of link that does not allow for free movement in any direction; the link is rigid and / or non-movable. For example, if fuel tank 408 is a longitudinally mounted cylindrical tank, it may have a series of feet on either side, which may be connected to the structure of fuel tank 408. These feet may also be connected, for example, to a compartment floor structure. Additionally, the rigid mount may impose loads on BWB 104 and fuel tank 408. In an embodiment, during a flight maneuver, Method 400 may stretch, compress, or deform slightly, which may ultimately and slightly alter the location of the example feet mounting points. Overall, this ensues some deformation of the structure of fuel tank 408 and may impose additional loads on the tank that result in an unfavorably heavier design. Additional feet may be provided to distribute the load of fuel tank 408 more widely into BWB 104. For example, feet may also connect the tank to compartment walls, the compartment ceiling structure, and the compartment aft pressure bulkhead. This arrangement may not be intended to reinforce BWB 104 by its connection to fuel tank 408. Also, it may not be intended to reinforce fuel tank 408 by its connection to BWB 104.Still referring to FIG. 5, fuel tank 408 may be mounted to BWB 104 using a linked mount. A standalone tank structure may be connected to the airframe in such a way that airframedeformation may not result in tank deformation, and vice-versa. This can be achieved through two ways to mount the tank: through its feet as described above, and through a series of links that may have hinges or ball-joints. A “linked mount” is a way to attach an object to another object that allows movement in one or more directions. With reference to the feet of fuel tank 408 as described above, to avoid structure deformation, the feet may be mounted to the airframe with rubber fittings that provide compliance. Exemplary feet include mounts manufactured by LORD Corp., of Williston, Vermont. A series of links that may have hinges or ball-joints may also mount fuel tank 408 to Method 400 without structural deformation occurring. For example, and as shown in the figure, fuel tank 408 may be mounted at three points forming a triangle: a first point may be a rigid connection that provides location in three axes, a second point may be a link that provides substantially two (vertical and lateral) location, and a third point may be a link that provides substantially one (vertical) location. These three points are seen in the figure as links 504, 508, and 512, respectively. Link 504 may support vertically, laterally, and longitudinally. Link 508 may support vertically and laterally because it may be pivoted about a lateral axis. Link 512 may provide only vertical support because it has ball joints at each end. A combination of links 504 and 508 may resist yaw and pitch motion. A combination of links 508 and 512 may resist roll. Altogether, motion of fuel tank 408 may be restrained against motion and rotation through the use of tank support links. A relative change in length between fuel tank 408 and BWB 104 may be accommodated by Links 508 and 512 pivoting fore-aft. A relative change in width may be accommodated by Link 512 pivoting laterally. A relative change in height may be unconstrained. Torsion applied to fuel tank 408 by Links 508 and three may be accommodated by a spherical or cylindrical connection at link 504. One skilled in the art can provide alternate ways to achieve these objectives.Now referring to FIG. 6, an exemplary embodiment of a cross-sectional view of multilobe tanks that may be placed aft of the main body of aircraft 100. The multi-lobe geometry of the multi-lobe tanks 600 may provide tension for each fuel tank 408 of the plurality of multi-lobe tanks Tension may be achieved by equal pressure in each lobe of the geometry. Multi-lobe geometry may include convex junctions joining each lobe. Convex junction 616 can be seen in FIG. 6.Now referring to FIG. 7, a block diagram of an exemplary embodiment of at least a tank for liquified gas fuel is shown with continued reference to FIG. 1, at least a tank 700, further includes a first compartment 708, a second compartment 716, liquified gas fuel 712, and a junction 720. A permanent tank may integrate at least a tank 700 and aircraft 100 structure so that the aircraft 100 supports at least a tank 700, and / or at least a tank 700 supports aircraft 100. In some embodiments, curved surfaces of at least a tank 700 may not be well suited to assisting structural surfaces of aircraft 100 and / or maximizing efficient use of volume within an outer mold line (OML) of the aircraft.Still referring to FIG. 7, at least a tank 700 is filled with and / or configured to be filled with a fuel; fuel is liquified gas fuel 712. Liquified gas fuel may include liquid hydrogen, propane, natural gas, or the like. Liquified gas may be used for airliner-scale systems. Liquified gas fuel 712 may have a higher density than gaseous fuel; even so, liquified gas fuel 712 may require about four times the volume of Jet-A fuel. Additionally, liquified gas fuel 712 may also allow for reduced tank pressure and tank weight as compared to gas fuel. Liquified gas fuel 712 may be kept at extremely low, cold temperatures, for example below its critical point of 33 Kelvin. Gas turbine engines, such as combustion engines, may operate on gaseous fuel and may transition liquified gas fuel 712 into a gas before consumption; gaseous fuel may be supplied to the engine at high pressure; in some embodiments no fuel pump may be used to provide fuel to the engine. In some embodiments, gas may squirt into a combustion chamber due to compressor stages. Vapor from tank ullage may be combined with the vapor from an output of a fuel heater. In some cases, this combined channel may then lead to an engine for combustion. When boiled, a rate of boil-off for liquified gas fuel 712 may be determined as a function of heat energy applied to liquified gas fuel 712. Boil-off rate for liquified gas fuel 712 may be selected based upon anticipated fuel consumption. At least a tank 700 may be insulated with an insulation to control application of heat to tank contents, and thereby achieve a selected boil-off rate. At least a tank 700 may also be refrigerated to remove heat from liquified gas fuel 712. The degree of insulation may be selected to provide a desired rate of boil-off. A selected rate of boil-off for liquified gas fuel 712 may generally be less than a rate at which liquified gas fuel 712 is consumed by the engines and possibly other aircraft systems, such as without limitation an Auxiliary Power Unit (APU). Excess boil-off may be dumped overboard or outside of aircraft through a vent.Continuing to refer to FIG. 7, at least a tank may include a first compartment 708. In this disclosure, a “compartment” is a separate section of a structure that can contain a substance. First compartment 708 may be used to store liquified gas fuel 712. First compartment 708 may be constructed such that the first compartment 708 has a first cross-section describing a continuously convex differentiable curve. As used in this disclosure, a “continuously convex differentiable curve” is a continuous convex curve which is differentiable at substantially all points along the curve; where this phrase is used in this application to refer to a physical component (e.g., tank 700), intention is to convey an overall or approximate shape of the physical component, for instance with ordinary manufacturing tolerances or practical limits. Exemplary non-limiting continuously convex differentiable curves include sections of circles, ellipses, conic curves, parabolas, and the like. In some cases, a continuously convex differentiable curve may be a circular or elliptical section having an eccentricity inclusive of 0.5 and 1.0 and an ellipticity inclusive of 0.5 and 1.0. First compartment 708 may be constructed as such to reduce a number of vertices on any surface defining the first compartment 708. That is, a surface of first compartment 708 may have a smooth surface with no comers. Additionally, first compartment 708 may include a first inner wall that defines a first internal cavity of first compartment 708. First inner wall may be made of any suitable material as described herein. For example, first inner wall may be a material with certain heating properties that allow heat to be transferred and / or insulated from a heat source to liquified gas fuel 712 stored in an inner volume in first internal cavity. Further, the first compartment may include a first outer wall exterior to first inner wall. First outer wall may be made of the same material as first inner wall or a different material. First inner wall and first outer wall may be separated by a first distance. Thus, a first space is defined therein. In other words, “first space” is a volume of a wall of first compartment 708 and may be a hollow space such that the first space may capture any fluid leaks from an internal cavity of first compartment 708.Still referring to FIG. 7, at least a tank 700 may include a second compartment 716. Second compartment 716 may include similar features discussed herein regarding first compartment 708. That is, second compartment 716 may be constructed in a similar or same manner and with the same considerations as first compartment 708 as discussed above. It should be noted that discussion herein regarding first compartment 708 and second compartment 716may also be applied to an nth compartment. For example, at least a tank 700 may include first compartment 708, second compartment 716, a third compartment, and so forth. In some instances, the number of compartments may depend on the shape and structure of aircraft 100. For example, to optimize volume of aircraft 100 filled with at least a tank 700, multiple compartments with cross-sections describing a continuously convex differentiable curve may be utilized rather than a 3-D shape to conform to a desired volume.Still referring to FIG. 7, at least a tank 700 may include a junction 720. A “junction,” as used in this disclosure, is an intersection where two or more objects are joined together. Junction 720 may be configured to structurally support at least a tank 700. In some cases, junction 720 may intersect with a first wall of first compartment 708 and a second wall of second compartment 716; and each of the first wall, the second wall, and the junction may be configured to be substantially loaded in tension when one or both of the first compartment 708 and the second compartment 716 are pressurized. In some instances, a junction 720 may be disposed at a section of at least a tank 700 where first compartment 708 and second compartment 716 about each other. That is, junction 720 may serve as a physical boundary between first compartment 708 and second compartment 716. Junction 720 may include a structural member that runs from a first intersection between walls of a first and second compartment to a second intersection between walls of a first and second compartment. In some instances, junction 720 may include one or more apertures and / or vents. One or more apertures may enable fluid communication between first compartment 708 and second compartment 716. To ensure that no additional stress is added to walls of at least a tank 700, each compartment of the at least a tank 700 may have an equalized pressure. That is, each compartment of at least a tank 700 may have similar or substantially equal pressure, thereby preventing walls of the at least a tank 700 from deforming or experiencing uneven loading. In some embodiments, a junction may be referred to as a septum. In some cases, a septum may be configured to carry structural loads. For example, loads intrinsic to the tank as well as loads extrinsic to the tank (e.g., loads resulting from pressurization of aircraft). As used in this disclosure, a “septum” is a structural element within a tank between chambers (e.g., a first compartment 708 and a second compartment 716). Further disclosure related to junctions 1120 (i.e., septa) may be found with reference to FIGS. 3C-D, 3G I, and 3K. In some cases, a septum may be perforated or otherwise allow for fluidic communicationbetween two or more compartments on opposing sides of the septum. Relationship between at least a tank 700 and aircraft 100 is further described herein with reference to FIG. 3L. In some embodiments, junction 720 joins a first internal cavity and a second internal cavity of first compartment 708 and second compartment 716, respectively. “A cavity” is a section of at least a tank 700 that is vacant or that stores substances and / or that is filled with ambient air or gas. Junction 720 may enable fluid communication between first internal cavity and second internal cavity.Still referring to FIG. 7, comprises one or more vents 724. In an embodiment, one or more vents 724 may be configured to vent gaseous fuel from at least a tank 700. Gaseous fuel may result from boil-off of liquified gas fuel 712 as the fuel warms. In an embodiment, and without limitation, one or more vents 724 may be configured to vent boil-off from at least a tank 700. In some cases, one or more vents 724 may include a check valve. In some cases, check valve may be configured to allow flow of fluids substantially only away from at least a tank 700 while preventing back flow of vented fluid to at least a tank 700. One or more vents 724 may also include a pressure regulator. Venting gaseous fuel from at least a tank 700 prevents overpressurizing or other events that may cause catastrophic damage or harm. It may also desirable, when aircraft 100 is grounded, to connect a system of lines and tanks to one or more vents 724 to collect the boiled-off fuel. In some cases, the collected gaseous fuel can be compressed by a pump into storage tanks and then cooled to liquid temperatures for reuse as aircraft fuel. In some instances, junction 720 may include one or more actuatable vents 724 that may be opened or closed to facilitate fluid communication between first internal cavity and second internal cavity. For example, there may be a pressure difference between first internal cavity and second internal cavity. To equalize pressure in both cavities, one or more actuatable vents 724 formed in junction 720 may be switched to an open position to facilitate fluid transfer from one cavity to another. Alternatively, first compartment and second compartment may be maintained in fluidic communication, thereby allowing for substantially equal pressure at substantially all times. In some cases, a pressure gradient may exist such that no additional outside intervention may be needed to transfer fluid from first compartment 708 to second compartment 716 or vice-versa.Still referring to FIG. 7, at least a tank 700 may include an insulation. Insulation may be used to reduce thermal transfer to liquified gas fuel inside of at least a tank 700. Heat may betransferred to at least a tank 700 by at least three means: convection, conduction, and radiation. To reduce conduction, insulation may include a vacuum to separate an interior volume of at least a tank 700 from an exterior with an evacuated vessel. Another means to reduce conduction and / or convection in at least a tank 700 may be an insulating material that inhibits conduction and / or convection. Insulating materials include fiberglass wool, plastic or ceramic foam, aerogel, and other materials. Heat flow through an insulator may be inversely proportional to its thickness. For example, an insulator that is twice as thick may conduct heat at half the rate. Heat transfer by radiation may be reduced by reflective, scattering, and / or absorbing coatings. Reflective coatings may be located proximal surfaces of at least a tank 700 and / or on intermediate materials within insulation.In some embodiments, and continuing to refer to FIG. 7, a tank 700 may have a 3-D shape to conform to a desired volume, which may allow the tank 700 to fit within an entire volume of aircraft 100. In some cases, tank 700 may include compartments with cross-sections describing a continuously convex differentiable curve such that a surface of the compartments is a smooth surface. By having a smooth surface (i.e., a continuously differentiable surface), stress on a wall of a compartment may be reduced. In addition, reducing stress on a wall of a compartment may allow for easier construction of at least a tank 700. For instance, at least a tank 700 may be constructed to have thin walls. Indeed, at least a tank 700 being constructed with a thin wall may allow for thin wall approximations when calculating hoop stress as discussed herein.Still referring to FIG. 7, in some embodiments, constructing the at least tank 700 with thinner walls of may lessen the cost of construction by using thin and / or less materials. Also, a total weight of at least a tank 700 may be lessened by using less materials and / or lighter materials. Additionally, or alternatively, a total weight of at least a tank 700 may be lessened relative to an original design because of a thin wall. That is, effects of constructing at least a tank 700 with a heavier material may be mitigated by constructing a thin wall because the construction would inherently use less material and thus have a lower total weight.Further referring to FIG. 7, at least a tank 700 may include a sensing component 728. Sensing component 728 may include a controller 732, one or more blowers or pumps 736, one or more sensors 740, one or more valves 744, or any combination thereof. Controller 732 may beany computing device described in this disclosure. Controller 732 may be communicatively coupled to at least tank 700. In addition, controller 732 may be communicatively coupled to other components of sensing component 728 in order to control fluid communication of at least a tank 700. For example, the sensing component 728 may receive a first indication, which may indicate that a first gas concentration exceeds a threshold gas concentration, from one or more sensors 740. In response to receiving the first indication, sensing component 32 may send a signal to one or more valves 744, via controller 732, to switch from a closed position to an open position. Moreover, in response to one or more valves 744 switching from a closed to open position, controller 732 may send an activation signal to one or more pumps 736 to activate the one or more pumps 736 to direct boiled-off gas from a first space and a second space through one or more valves 744. For instance, in some cases, an inner wall of tank 700 may slowly leak gaseous fuel, into a space between the inner wall and outer wall. In some cases, controller 732 may be configured to detect, using sensors 740, gas levels within this space an evacuated the gas as needed using one or more of pumps 736 and valves 744.Generally, sensing component 728 may send signals to various components within aircraft 100. In some instances, sensing component 728 may require communication with one or more sensors 740 to determine what actions to signals to send out. It should be noted that one or more sensors 740 may be gas sensors, pressure sensors, velocity sensors, or any combination thereof. One or more sensors 740 may be disposed within various parts of aircraft 100 to provide more an accurate state of operation metric.Still referring to FIG. 7, at least a tank 700 may include conduits that may fluidly couple one or more valves 744 to a first space and a second space and / or a first compartment 708 and a second compartment 716, as described herein. Conduits may connect one or more valves 744 to an internal space to facilitate a transport of boil-off gas to a fuel cell. In addition, conduits may facilitate a transport of any gas fuel that may leak into a first space into a first internal space, a secluded reservoir, air exterior to at least a tank 700 and / or a vehicle containing at least a tank 700, or both.Referring now to FIG. 8, an exemplary embodiment of a system 800 for venting a fuel tank is illustrated. System may include a computing device such as a controller. Controller 804 may include any computing device as described in this disclosure. System 800 includes a fueltank 808. Fuel tank 808 may include a vent 812. Fuel tank 808 is connected to a vent line 816. Vent line 816 may be attached to a vent 812 located at the top of the tank 208 such that vent line 816 directs gaseous fuel from the vent 812 to an external fuel tank 820. External fuel tank 820 may be consistent with any fuel tank as disclosed herein. External fuel tank 820 (also referred to as “external tank”) is discussed in further detail below.Continuing to reference FIG. 8, vent line 816 may include a pump 814 to move gaseous fuel from the fuel tank 808 to the external tank 820. Pump may include a substantially constant pressure pump (e.g., centrifugal pump) or a substantially constant flow pump (e.g., positive displacement pump, gear pump, and the like). Pump can be hydrostatic or hydrodynamic. As used in this disclosure, a “pump” is a mechanical source of power that converts mechanical power into fluidic energy. A pump may generate flow with enough power to overcome pressure induced by a load at a pump outlet. A pump may generate a vacuum at a pump inlet, thereby forcing fluid from a reservoir into the pump inlet to the pump and by mechanical action delivering this fluid to a pump outlet. Hydrostatic pumps are positive displacement pumps. Hydrodynamic pumps can be fixed displacement pumps, in which displacement may not be adjusted, or variable displacement pumps, in which the displacement may be adjusted. Exemplary non-limiting pumps include gear pumps, rotary vane pumps, screw pumps, bent axis pumps, inline axial piston pumps, radial piston pumps, and the like. Pump may be powered by any rotational mechanical work source, for example without limitation and electric motor or a power take off from an engine. Pump may be in fluidic communication with at least a reservoir. In some cases, reservoir may be unpressurized and / or vented. Alternatively, reservoir may be pressurized and / or sealed.Continuing to reference FIG. 8, system 800 may include a sensor 824. In some embodiments, sensor 824 may be configured to detect the pressure within tank 208. In some embodiments, sensor 824 may be a pressure sensor. Sensor 824 may detect when pressure within the tank 208 increases above a threshold pressure. In an embodiment, a Controller 804 may be communicatively connected to sensor 824. Controller 804 may be onboarding the aircraft or remote from the aircraft. Sensor 824 may send pressure data to Controller 804 and Controller 804 may determine when to open vent 812 based on the pressure. A threshold may be set by a user to determine when vent 812 should open. Threshold pressure may be a 0.01%, 0.1%, or 1%increase in pressure. In some embodiments, threshold pressure may be 3% or 5% greater than the operating pressure inside tank 208.With continued reference to FIG. 8, additionally, sensor 824 may detect when a vent line 816 is connected to vent 812. Vent 812 may be a check valve, as discussed above, that may only open when a vent line 816 is connected to the vent 812. Additionally, vent line 816 may only be connected to vent 812 when aircraft is grounded. “Grounded” as used herein, is when the aircraft is prevented from flying. An aircraft may be grounded during maintenance. An aircraft may be grounded when the aircraft is not in operation to carry cargo or passengers. Vent line 816 may be physically connected to vent 812 by a user, such as a maintenance worker. Sensor 824 may be consistent with any sensor as discussed above.Continuing to reference FIG. 8, sensor 824 may be a motion sensor. Motion sensor may detect when a vent line is connected to a vent. A motion sensor refers to a device or component configured to detect physical movement of an object or grouping of objects. One of ordinary skill in the art would appreciate, after reviewing the entirety of this disclosure, that motion may include a plurality of types including but not limited to: spinning, rotating, oscillating, gyrating, jumping, sliding, reciprocating, or the like. A motion sensor may include, torque sensor, gyroscope, accelerometer, position, sensor, magnetometer, inertial measurement unit (IMU), pressure sensor, force sensor, proximity sensor, displacement sensor, vibration sensor, or the like. In one or more embodiments, sensor 824 may include various other types of sensors configured to detect a pressure of a tank 208 and / or a vent line connection. For instance, sensor 824 may include photoelectric sensors, radiation sensors, infrared sensors, and the like. Sensor 824 may include contact sensors, non-contact sensors, or a combination thereof. In one or more embodiments, sensor 824 may include digital sensors, analog sensors, or a combination thereof. Sensor 824 may include digital-to-analog converters (DAC), analog-to-digital converters (ADC, A / D, A-to-D), a combination thereof, or other signal conditioning components used in transmission of measurement data to a destination, such as computing device 148, over a wireless and / or wired connection.Still referring to FIG. 8, Vent line 812 connects the fuel tank 808 to the external tank 820. External tank 820 collects the gaseous fuel boil-off from the fuel tank 808 in the grounded aircraft 100. External tank 820 may collect the gaseous fuel boil-off and re-liquefy it back intoliquid fuel to be reused in the aircraft. The process of re-liquefying the gaseous fuel may include using a refrigeration cycle to lower the temperature of the gaseous fuel back into a liquid state. Re-liquefying gaseous fuel may include cooling down the gas below the boiling point of 2 IK and pressurizing it to around 13 bars. Refrigeration cycle may include condensers, throttle valves, compressors, and heat exchangers. As used herein, a “refrigeration cycle” is a thermodynamic cycle wherein heat is transferred from a body with lower temperature to a body with higher temperature. In an embodiment, refrigeration cycle may include using a compressor to compress the gaseous fuel from a low-pressure state to a high-pressure state. A condenser may use a refrigerant fluid to remove heat from the high-pressure vapor gas. Refrigerant may act as a heat exchanger wherein the gas cools down and the refrigerant increases in heat. A throttle valve may be used to create a drop in pressure in the refrigerant. In some embodiments, a pump, as disclosed above, may be used to move the refrigerant fluid. A pump, as disclosed above, may be used to compress gaseous fuel. External tank 820 may store the re-liquefied gaseous fuel, and / or the liquid hydrogen may be transferred into additional storage tanks. Additional storage tanks may be consistent with any fuel tank as discussed above. In some embodiments, air in vent line and / or downstream system may pose a problem. For example, if air reaches a cold portion of the airplane’s internal vent line, the air may freeze and block the vent. Also, in some cases, moisture (e.g., water vapor) in air may freeze at an even higher temperature (e.g., 32° F) and block vent line. Accordingly, in some embodiments, system is designed to prevent air from getting into the airplane’s vent system. In some cases, a valve near a fitting that couples to an external vent line may be used to vent air and / or moisture. In some cases, an ongoing (and slow) flow of gaseous hydrogen from tank may help to prevent ingress of air and / or moisture, as the line is self-purging and pressurized, thereby preventing unpressurized air and / or moisture from entering through a valve.Now referring to FIG. 9, an exemplary embodiment of a tank 900. Tank 900 may be consistent with any tank as discussed herein. Tank 900 may be an embodiment of tank 408, tank 808, or external tank 820. A key component that may heavily impact shape of tank 900 is pressure. Tank 900 may be far lighter if pressure is resisted in pure or predominantly tension as compared to bending. In pure tension, thin-walled tanks may be used. Thin-walled tanks may be lighter than thick-walled tanks as less material is used. In some embodiments, pure tension in atank may be generally achieved by shapes that provide a circular cross section, including spheres, cylinders, and cones. In an embodiment, a pressurized tank of a given volume may be made as a sphere to place the tank in pure tension. Alternatively, a tank made as a cube may require tank walls to operate in bending; thus, the cube tank would likely be vastly heavier than a sphere tank of similar volume. Cube tanks may be heavier as thick walls may be necessary to resist bending. Accordingly, in some embodiments, any tank geometry may provide tank walls acting in tension. Tank 900 may include a vent 904. Vent 904 may be consistent with any vent as discussed herein. Vent 904 may be located on top of the tank 900 or at the bottom of tank 900. Vent 904a, located at the top of the tank 300, may provide gaseous fuel to an engine of aircraft 100. Vent 904b, located at the bottom of the tank 900 provides liquid fuel that may be heated to provide gaseous fuel to an engine of the aircraft 100. Vent line 812 may connect to vent 904a, as vent 904a may be near the vapor ullage. Because liquid fuel and gaseous fuel are both very cold, the heating of liquid fuel may be provided by waste heat from another system of the aircraft 100. “Waste heat”, as used herein, is excess heat from a system. In an embodiment, heat generated from friction in the propulsors may be used to provide heating to the liquified gas fuel, and the liquified gas fuel may provide cooling to the propulsors.Now referring to FIG. 10A, a depiction of a vent line 1004 connected to two tanks is shown. Vent line 1000 may be connected to fuel tank 808 and external tank 820. Vent line may be consistent with any vent line as discussed herein. Vent line may include a mating component to mate with a vent 1008 on fuel tank 808 and external tank 820. As used in this disclosure, a “mating component” is a component that is configured to mate with at least another component, for example in a certain (i.e., mated) configuration. Mating component may include a mechanical or electromechanical mechanism. For example, without limitation mating may include an electromechanical device used to join electrical conductors and create an electrical circuit. In some cases, mating component may include gendered mating components. Gendered mating components may include a male component, such as a plug, which is inserted within a female component, such as a socket. In some cases, mating between mating components may be removable. In some cases, mating between mating components may be permanent. In some cases, mating may be removable, but require a specialized tool or key for removal. Mating may be achieved by way of one or more of plug and socket mates, pogo pin contact, crown springmates, and the like. In some cases, mating may be keyed to ensure proper alignment of vent line 1004. In some cases, mating may be lockable. Vent line 1004 may be placed on vent 408 when aircraft 100 is grounded. Vent line 1004 may be removed once aircraft 100 is ready for flight.Now referring to FIG. 10B, a cross sectional view 1012 of a vent line 1004. Vent line 404 may include an inner wall 1016 in contact with the fuel, an outer wall 1020 in contact with the external environment, and insulation 1024 in between the inner and outer wall. Insulation may be a vacuum. Insulation may also be fiberglass, aerogel, or the like. There may be a heating element 1028 surrounding the outer wall to prevent a buildup of frozen water vapor and other gases. Heating element may be a hot liquid run through a channel surrounding the vent line 1004. As used herein, a “heating element” is a component that emits heat to its surroundings.Referring now to FIG. 11, an exemplary embodiment of an apparatus 1100 for aircraft fueling. Apparatus 100 may include a container 1104. Container 1104 As used in the current disclosure, a “container” is an enclosure for holding matter. In embodiments, a container 1104 may include one or a plurality of fuel tanks. In embodiments, a plurality of fuel tanks (also referred to as a “tank”). Container 1104 may be a pressure vessel. Pressure vessel may be configured to be pressurized in order to allow flow of gaseous hydrogen from a container 1104, for example without a need to pump. Furthermore, a container 1104 may further include an inner wall and an outer wall. There may be insulation between the outer and inner wall. A container 1104 may also include safety valves, closures, vessel threads, or any other features that can be found on fuel tanks. Container 1104 may further have a tank geometry.Continuing to refer to FIG. 11, container 1104 may be comprised of a multi-lobe tank configuration. In a multi -lobe tank configuration, each container 1104 does not form a completely circular shape, such that the surface of container 1104 may deform to achieve pure tension. Multi-lobe tank configuration may be beneficial as it provides more tank volume compared to a singular spherical tank. Multi-lobe tank configuration may be derived by adding tanks in the junctions between tanks. Junctions are discussed in further detail in FIG. 3B. Container 1104 or Fuel tanks, as discussed in more detail below, may provide structural support to the aircraft by acting as load bearing columns between the floor and ceiling of the aircraft 100. In an embodiment, fuel tanks may be mounted within the airframe. As used herein, an “airframe” provides structure to an aircraft. Airframe may be a part of the structural components of theaircraft 100. The plurality of fuel tanks may span across the full width of the aircraft 100. Because the ceiling of the aircraft is downward sloping, each container 1104 of the plurality of fuel tanks may vary in diameter and length. Fuel tanks may be stored vertically or horizontally. In an embodiment and as described above, container 1104 may be a vertically oriented multilobe tank wherein its pressurized walls and its septa may extend from a lower outer mold line to an upper outer mold line of a container 1104. A top and bottom of container 1104 may be closed out by spherical end caps inset from the outer mold line to provide room for the outer mold line skin’s supporting structure. Fuel tanks may be stored in rows aft of the cabin. In an embodiment, there may be two rows of fuel tanks. In doubling the rows, each container 1104 had a smaller tank wall radius, therefore each fuel tank wall is thinner and lighter.Continuing to refer to FIG. 11, container 1104 may be configured to be transported using a translocation device 1108. As used in the current disclosure, a “translocation device” is a device configured to carry container 1104. A translocation device 1108 may include a platform mechanically attached to a plurality of wheels. Container 1104 may be removably attached to the platform of translocation device 1108. Additionally, translocation device 1108 and / or the plurality of wheels may be configured to be removably attached to an orientation guidance track 1116. In embodiments, a translocation device 1108 may include a cart or basket attached to a platform.Continuing to refer to FIG. 11, translocation device 1108 may be removably attached to an orientation guidance track 1116. As used in the current disclosure, “orientation guidance track” is a track or conveyor system on which a translocation device 1108 travels. In some embodiments, orientation guidance track 1116 is an apparatus for transporting translocation device 1108 from a first position 1120 to a second position 1124. In embodiments, orientation guidance track 1116 may also transport a translocation device 1108 from a second position 1124 to a first position 1120. An orientation guidance track 1116 may include a conveyor system, a system of tracks guided by rollers, railroad style tracks, and the like. Orientation guidance track 1116 may be configured be stowed or removed during flight of aircraft 100. In some embodiments, Orientation guidance track 1116 may be configured to transport an object in a straight path. In other embodiments, Orientation guidance track 1116 may be configured to transport an object along a curved path. In some embodiments, Orientation guidance track 1116may be configured to transport an object along a nonsymmetrical path. In some embodiments, Orientation guidance track 1116 may be configured to transport an object along a symmetrical path. Orientation guidance track 1116 may be configured to be in communication with other components of aircraft 100. An orientation guidance track 1116 may be configured to be in a circular formation, serpentine formation, rectangular formation, L-shaped formation, inverted L- shaped formation, and the like. In other embodiments, an orientation guidance track 1116 may be oriented longitudinally within the aircraft. When orientation guidance track 1116 is oriented longitudinally it may run parallel with the fuselage of the aircraft. When orientation guidance track 1116 is oriented longitudinally it may run parallel to the longitudinal axis of the fuselage of the aircraft.Still referring to FIG. 11, orientation guidance track 1116 may include a conveyor system. Conveyor system may be configured to transport one or more objects from one location to another location. In some embodiments, conveyor system may be configured to transport one or more translocation device 1108 to one or more locations. Conveyor system may include, but is not limited to, a roller bed conveyor, belt conveyor, curved bel conveyor, incline conveyor, decline conveyor, specialty conveyor belt and the like. In some embodiments, the conveyor may include, but is not limited to, a pneumatic, vibrating, flexible, spiral, or vertical conveyor. In some embodiments, conveyor system may be configured to transport polymer sheets 104 from a first location to a second location. In some embodiments, conveyor system may be configured to transport translocation device 1108 to a plurality of locations.Continuing to refer to FIG. 11, an orientation guidance track 1116 may be configured to restrain the lateral or vertical movement of both translocation device 1108 and container 1104. A locking mechanism may be used to restrain the movement of both translocation device 1108 and container 1104. As used in the current disclosure, a “locking mechanism” is a mechanism that can lock an object onto a track. Locking mechanism may removably attach translocation device 1108 to an orientation guidance track 1116. Locking mechanism may have an engaged configuration and a disengaged configuration. In the engaged configuration, locking mechanism may lock translocation device 1108 to orientation guidance track 1116 such that it cannot be removed. In the disengaged configuration, locking mechanism may not prevent translocation device 1108 from being removed from orientation guidance track. A locking mechanism mayinclude mating a portion of the translocation device 1108 to an orientation guidance track 1 116 using a male / female connection. In other embodiments, an orientation guidance track 1116 may use a plurality of straps. Straps may be constructed of any flexible material and / or set of materials, including without limitation membranes or sheets of polymer material, natural materials such as leather, and / or natural or artificial textiles. Strap may be effectively fire- resistant, wherein the strap is effectively melt-resistant, have high tensile strength, and / or have high electrical resistivity, as defined above. Strap may be constructed using elastic and / or inelastic materials, which may be combined in various ways; for instance and without limitation, strap may be constructed loosely woven mesh or webbing of effectively inelastic fiber such as meta-aramid fibers, aramid fibers, KEVLAR, NOMEX, or the like that may permit stretching within a certain range owing to slackness of fibers when embedded elastic strands are elastically neutral; thus, elasticity of strap may be equivalent to elastic component within such a range.Continuing to refer to FIG. 11, container 1104 may be configured to be located in a first position 1120. As used in the current disclosure “first position” is when container 1104 is located aft of the cabin in the main body. As used herein, “cabin” is the portion of the aircraft that holds the crew, passengers, and cargo. First position 1120 may be at least partially aft of the cabin, near the propulsors. In an embodiment, because of the low density of liquid hydrogen fuel, storing container 1104 in first position 1120 may not substantially affect the longitudinal center of gravity of the aircraft 100 when it is in first position 1120. First position 1120 may include a position located inside of aircraft 100. Container 1104 may be configured to be. Container 1104 may be configured to be located in a second position 1124. As used in the current disclosure, “second position” is when container 1104 is located to the exterior to the body of the aircraft.Continuing to refer to FIG 11, orientation guidance track 1116 may be configured to create a path between a first position 1120 and a second position 1124. For example, orientation guidance track 1116 may comprise a pair of tracks that run from first position 1120 to second position 1124, or vice versa. For example, orientation guidance track may comprise a single track that runs from first position 1120 to second position 1124. In some embodiments, orientation guidance track may comprise a conveyor belt that runs from first position 1120 to second position 1124. In embodiments, orientation guidance track 1116 may be configured to begin aft of the main body of the aircraft. In some embodiments, orientation guidance track1116may linkup with or combine with an exterior orientation guidance track that is outside of the aircraft. Orientation guidance track 1116 may be aligned with the location of container 1104 in such manner wherein the container is easily transferred from a first position 1120 inside of the aircraft on to the orientation guidance track 1116. For example, container 1104 may rest on orientation guidance track. As another example, orientation track may be located directly adjacent to container 1104. Large portions of the orientation guidance track 1116 may be configured to be on the exterior of the aircraft. Once a container 1104 is positioned on an orientation guidance track 1116 may be transported to the second position 1124. The second position 1124 may be located exterior of the aircraft to either side of the aircraft. The orientation guidance track 1116 may use an L-shaped track or conveyer to reach the second position 1124. Orientation guidance track 1116 may include curves, bends, turns, and the like to reach second position. In embodiments, a second position 1124 may be an aircraft fueling station. An orientation guidance track 1116 may be used to transport container 1104 into a position to make refueling more time and energy efficient. For example, orientation guidance track 1116 may be used to transport container 1104 from its location inside of the aircraft to the aircraft fueling station. An orientation guidance track 1116 may be configured to have any one of a plurality of geometries / configurations, as mentioned herein above, in order to have the most efficient path to transport container 1104 from a first position 1120 to a second position 1124. A second position 1124 may also be a container storage / maintenance facility.Now referencing FIG. 12, a cross sectional view 1202 of fuel tank 1200. Tank 1200 includes an inner wall 1204. Inner wall 1204 may be in direct contact with the fuel in the fuel tank 1200. Inner wall 1204 may be configured to hold the fuel. Inner wall 1204 may also be direct or indirect contact with an interstitial volume 1208 found between the inner wall 1204 and the outer wall 1212. As used herein, an “interstitial volume” is an intervening space between two barriers. Barriers may include the outer and inner wall. Inner wall 1204 may be made from carbon polymer composite (carbon epoxy), aluminum, composite, or the like, or a combination thereof. In an embodiment, inner wall 1204 may include an aluminum liner in combination with the carbon epoxy material. Carbon epoxy may be used over aluminum as it is lighter and has a lower thermal expansion coefficient as compared to aluminum. However, carbon epoxy may leak more than aluminum. Thermal expansion coefficient may be an important factor in the inner wallmaterial as exposure to low-temperature (e.g., -252.87°C at 1 .013 bar of pressure) liquified gas fuel in the fuel tank 1200 causes the inner wall 1204 to contract. Aluminum may substantially vary in size when exposed to the variations in temperature of a typical liquified gas fuel tank. The coefficient of thermal expansion (a) of aluminum is approximately 23 x 106° C-1. Change in length (L) for a given piece of aluminum subjected to a temperature increase is determined by the following equation, wherein t is the temperature increase.AL = aLtThe contraction of the inner wall 1204 may cause pressure, such as without limitation ununiform pressures, on the outer wall 1212 or the interstitial volume 1208.Continuing to reference FIG. 12, tank 1200 may include a sensing system connected to a controller to monitor and control leakage from the tank. Sensing system is described with respect to FIG. 13. As discussed above, inner wall 1204 may leak gaseous fuel (e.g., gaseous hydrogen). Outer wall 1212 may contain the leakage such that the gaseous fuel does not escape the tank 1200. However, leaked gaseous fuel may need to be purged from the interstitial volume 1208 to maintain a gaseous fuel concertation (e.g., hydrogen concentration) below that which supports combustion. In some cases, tank 1200 may be configured to prevent hydrogen volumetric concentration from exceeding 4% within interstitial volume. Above this threshold value, in some cases, the hydrogen may become combustible when mixed with oxygen in the air. In some embodiments, a threshold for purging hydrogen may be set lower than 4%, for instance 1%, 0.1%, or 0.01% to account for a margin of safety. In an embodiment, sensing system may purge interstitial volume to keep hydrogen gas at or below 1%, 0.1%, or 0.01% hydrogen volumetric concentration of the interstitial volume 1208. Purged gaseous fuel may be released overboard. Sensing system may include a gas sensor, as discussed in FIG. 1, that may detect gaseous fuel in the interstitial volume. Gas sensor may measure gaseous fuel concentration (e.g., hydrogen concentration) in the interstitial volume. A gas sensor may include optical fiber surface plasmon resonance (SPR) sensors. As non-limiting examples, SPR sensors may include fiber brig gratings coated with a palladium layer, a micromirror, or a tapered wire coated with palladium. Gas sensor may also include electrochemical hydrogen sensors, microelectromechanical system (MEMS) hydrogen sensors, thin film sensors, thick film sensors, chemochromic hydrogen sensors, diode based Schottky sensors, or the like. Alternatively, or additionally, gas sensor maybe configured to monitor purge flow rate. “Purge flow rate”, as used in this disclosure, is the rate that of gas being vented out. Purge flow rate may be measured in mass flow rate, pressure (e.g., pressure difference), volumetric flow rate, gas velocity, or the like. Purge flow rate may be monitored using a velocity sensor physically or communicatively connected to the gas sensor. Purge flow rate may be measured using one or more pressure sensors. Purge flow rate may be measured using a pitot tube. Purge flow rate may be monitored at a vent. Purge flow rate may be used to identify excessive gas leakage, which may help avoid sudden failure. Sensing system may include a vent 1216 attached to tank 1200. As used in this disclosure, a “vent” is an opening and / or aperture configured to allow one or more fluids to pass. Gas sensor may be communicatively connected to a controller that may control a vent 1216 to ventilate the interstitial volume 1208 of gaseous discharge. In an embodiment, controller may signal the vent to switch from a closed position to an open position when a threshold gas concentration value is detected. Threshold gas concentration value may be 1%, 0.1%, or 0.01% hydrogen volumetric concentration of the interstitial volume 1208. In some embodiments, threshold gas concentration may be 3%, or 5% and greater. In some cases, vent 1216 may include a check valve. A check valve may be used to prevent backflow of gases. As used in this disclosure, a “check valve” is a valve that permits flow of a fluid only in certain (e.g., one) directions. In some cases, check valve may be configured to allow flow of fluids substantially only away from tank 1200 while preventing back flow of vented fluid to tank 1200. Vent may also include a pressure regulator. A “pressure regulator” is a type of valve that controls the pressure of a fluid. Venting gaseous fuel from tank 1200 prevents over-pressurizing or other events that may cause catastrophic damage or harm. It may also desirable, when aircraft 100 is grounded, to connect a system of lines and tanks to the vent to collect the boiled-off fuel. In some cases, the collected gaseous fuel can be compressed by a pump into storage tanks and then cooled to liquid temperatures for reuse as aircraft fuel. In some cases, vent 1216 may be configured to prevent condensation resulting from venting of gas. For instance, and without limitation, vent 1216 may include insulation, cascading pressure vessels, gradual throttling valves, clean dry air mixing, and the like to prevent one or more of gas expansion, temperature drop, and / or condensation.Continuing to reference FIG. 12, tank 1200 may include a gap 1220. Gap 1220 may be in the interstitial volume 1208. As used herein, “gap” refers to the portion of the interstitial volumethat is not taken up by a structure, such as insulation. In some cases, gap 1220 may be configured to resist conductive and / or convective heat transfer, for instance between inner and outer walls of tank 1200. The gap may be ventilated. In an embodiment, gap may be connected to the vent 1216 such that the vent may ventilate fuel gas out of the gap 1220. The gap may allow the gaseous fuel that leaks from the inner wall 1204 to be purged from the tank 1200. Gap may be located between the outside face of the inner wall 1204 and insulation. In some cases, vented gas from the gap 1220 may be vented overboard outside of aircraft. Alternatively, or additionally, in some cases, vented gases from the gap 1220 may be vented into cabin (and then vented using cabin ventilation systems). In some cases, purge gas may be pumped into gap 1220. Purge gas may include air, nitrogen, recirculated gas, or the like. In some cases, purge gas may be filtered and / or dried prior to pumping into gap 1220, for example to prevent condensation and / or contamination of an inner or outer tank wall which could lead to decreased insulation.Continuing to reference FIG. 12, tank 1200 includes at least a reflective film layer and at least a structural insulation layer. At least a reflective film layer and at least a structural insulation layer may form insulation for the tank 1200. As used herein, “reflective film” is a thin layer of reflective material that lowers heat transfer. Reflective film layer may be a coating, for instance on a surface of inner and / or outer wall of a tank. Reflective film layer may include, without limitation, gold, nickel, silver, or the like. Reflective film layer may be a coating on the inner or outer wall. Reflective film layer may include a high polish surface on one or more of inner or outer walls. A high polish surface may have a Ra value less than 10 microns, less than 1 micron, or less than 0.5 microns. In some cases, reflective film layer may be configured to limit radiative heat transfer, for example from inner wall to outer wall. Interstitial volume 1208 may include several layers of Reflective film 1228 and structural insulation 232. Reflective film 1228 may be sandwiched in between layers of structural insulation 1232 as shown in FIG. 2. In an embodiment, there may be 5 layers of insulation, wherein a “layer of insulation” is defined as a combination of one layer of reflective film and one layer of structural insulation. The number of layers of insulation may be determined by the thermal resistance requirement of the interstitial volume 1208. Thermal resistance is a property of the material’s thermal conductivity, thickness, and area.Continuing to reference FIG. 12, at least a reflective film layer may be used in the insulation. Reflective film 1228 may include a metal sheet, for example a sheet of an aluminum alloy, silver, gold, titanium, stainless steel, or the like. Reflective film 1228 may include a metalized plastic film. Plastic film may include Mylar, Kapton, or Tedlar. Reflective film 1228 may use nickel, aluminum, gold, silver, a combination thereof, or other metals. Reflective film 1228 may use polyimide or polyester, or the like. Reflective film 1228 may reduce heat transfer from radiation due to the reflective nature. Reflective film 1228 may be a multi-layer insulation (MLI). Persons skilled in the art, upon reviewing the entirety of this disclosure, will be aware of various materials used in reflective films. Reflective film 1228 may be used in combination with a structural insulation 1232 to decrease heat transfer between the inner wall 1204 and the outer wall 1208. Insulation may be used to prevent air or water from freezing on the outer surface of the inner wall 1204.Still referencing FIG. 12, insulation may also include at least a structural insulation layer. There may be a plurality of structural insulation layers, wherein the Reflective film 1228 is in between each layer of structural insulation 1232. “Structural insulation,” as used herein, is a form of self-supporting insulation within the interstitial volume. In some cases, structural insulation may support other insulation layers (e.g., reflective film layer) within the interstitial volume. Structural insulation 1232 may, in some cases, support the outer wall 1212 against the vacuum in the interstitial volume 1208, for example by providing compressive forces between inner and outer walls. Vacuum may be used to evacuate volume between the inner wall 1204 and the outer wall 1212. Vacuum may also be result of a pressure differential (i.e., higher pressure outside of outer wall than in interstitial volume). Additionally, structural insulation 1232 may support the inner wall 1204 against the high pressure from the liquified gas fuel. Structural insulation 1232 may be divided into blocks such that the structural insulation 1232 is discontinuous. Gaps between the structural insulation 1232 may provide gaps to allow evacuation and leak detection. Gaps between structural insulation 1232 may provide fluidic pathways for gaseous discharge. Structural insulation 1232 may be composed from a porous or non-porous insulation. Structural insulation 1232 may be composed of, as non-limiting examples, Aerogel, vermiculite, fiberglass, X- Aerogel, crosslinked Aerogel, firebrick, or any other insulation with a low thermalconductivity coefficient. A low thermal conductivity may be any thermal conductivity less than 0.10, 0.01, or 0.001 W / m-k. Structural insulation 1232 may limit heat transfer by conduction.Still referencing FIG. 12, tank 1200 includes an outer wall 1212. Outer wall 1212 may be in between an outside environment (e.g., aircraft cabin environment) and interstitial volume 1208. Outer wall 1212 may be offset from the inner wall 1204 such that the interstitial volume 1208 is in between. Outer wall 1212 may provide damage protection from external elements such as temperature, transportation, or the like that may occur within the aircraft. Tank 1200 may be stored within the aircraft 100. In some embodiments, tank 1200 may be stored within the outer mold line (skin) of the aircraft body (such as blended wing body 104). Alternatively, tank 1200 may be mounted outside the aircraft within a nacelle. Outer wall 1212 may prevent permeation of gases originating from within inner wall 1204 and / or interstitial volume 1208. Outer wall 1212 may be composed of one or more of steel and aluminum to prevent the permeation of gases from the inner wall 1204 of the tank 1200. Outer wall 1212 may also be composed of other material, such as composites, carbon polymer composites, titanium, and the like. Outer wall 1212 may be composed of materials that are able to withstand the compression loads from a pressure differential between in the interstitial volume 1208 and an environment outside the outer wall. In an embodiment, outer wall 1212 may be subjected to compression loads because the outer wall 1212 may shrink due to the vacuum in the interstitial volume. In some embodiments, outer wall 1212 may be corrugated so that the outer wall 1212 may accept changes in its diameter with flexure. In an embodiment, corrugated outer wall 1212 may minimize thickness of the outer wall 1212 because a corrugated outer wall may be stronger against compression loads than a non-corrugated outer wall, therefore less material is needed for a corrugated outer wall. In some embodiments, outer wall 1212 may include a corrugated piece of material sandwiched between two non-corrugated pieces of material. The non-corrugated pieces of material may provide additional strength and insulation to the outer wall 1212 and increase the durability of it.Now referring to FIG. 13, a block diagram of a sensing system 1300 for aircraft 100. Sensing system may include a controller 1304. Controller 1304 may include any computing device as described in this disclosure. Controller 1304 may be connected to a gas sensor 1308. Gas sensor 1308 and controller 1304 may be communicatively connected. Gas sensor 1308 maybe consistent with any gas sensor disclosed as part of this disclosure. Controller may be consistent with any controller disclosed as part of this disclosure. Gas sensor 1308 may detect gas concentrations such as gaseous fuel concentrations. Gas sensor may monitor purge flow rate by monitoring the velocity of gas leaving vent 1312. Velocity, mass flow rate, or the like may be measured using a pressure sensor or velocity sensor communicatively connected or physically connected to the gas sensor 308. Vent may be consistent with any vent as discussed above, Gas sensor 1308 may be consistent with any gas sensor as described above. Gas sensor 1308 may cause the controller 1304 to issue signals to a vent 1312, such as a purge valve or pump. For example, the signals may cause the vent 1312 to open if the gaseous fuel concentration exceeds a threshold value, such as the threshold values discussed above. In another example, gas sensor 1308 may send a signal that causes the vent 1312 to open if the gas sensor 1308 detects excessive gas leakage, which may help avoid sudden failure of the tank.Continuing with reference to FIG. 13, in some embodiments controller 1304 may control gaseous fuel (i.e., fuel vapor) concentration. For example, in some cases, gaseous fuel concentration within an interstitial volume (i.e., gap 1220) between an inner wall and an outer wall may be controlled. In some embodiments, gaseous fuel concentration may be adjusted by controlled admission of air or other gases into gap 1220. In some cases, admission of air or other gases may be controlled by vent 1308. Vent 1308 may be configured to allow of admission of air or other gases (e.g., nitrogen) in to gap 1220. In some cases, air may come from within aircraft cabin and / or from outside of aircraft. In some cases, other gases may help prevent combustion of gaseous fuel. For instance, if interstitial volume 1208 is filled with a mixture of nitrogen and fuel vapor, combustion is not possible due to absence of oxygen. In some cases, using administrating an inert gas into interstitial volume 1208 may be favorable because greater concentrations of fuel vapor may be permitted. In some cases, nitrogen may be supplied by an onboard inert gas generator system. In some cases, an onboard inert gas generator system may be in place to reduce flammability of fuel vapor, generally. In some embodiments, gas admission rate may be controlled in concert with interstitial volume pressure to result in a combination of desired pressure and fuel vapor concentration. For instance, in some embodiments, controller 1304 may control pressure within the interstitial volume 1208. Generally lower gas pressure withininterstitial volume 1208 will increase thermal insulation. In some cases, interstitial volume pressure may be controller by way of pump rate (e.g., vacuum pump rate).Referring now to FIG. 14, a method 1400 of use for a blended wing body aircraft with a fuel cell is illustrated by way of a flow diagram. At step 1405, method 1400 may include storing a first fuel, using at least a first fuel store. First fuel may include any fuel described in this disclosure, for example with reference to FIGS. 1 - 2. First fuel store may include any fuel store described in this disclosure, for example with reference to FIGS. 1 - 2. In some embodiments, first fuel may include one or more of liquid hydrogen and natural gas. In some embodiments, first fuel store may be at least partially located within a transitional portion of blended wing body. In some embodiments, first fuel store may include at least a fuel environment control mitigation.With continued reference to FIG. 14, at step 1410, method 1400 may include combining first fuel with an oxidizing agent to produce electricity, using at least a fuel cell. Fuel cell may include any fuel cell described in this disclosure, for example with reference to FIGS. 1 - 2. Oxidizing agent may include any oxidizing agent, such as without limitation oxygen, described in this disclosure, for example with reference to FIGS. 1 - 2.With continued reference to FIG. 14, at step 1415, method 1400 may include storing a second fuel store, using at least a second fuel store located within a wing portion of the blended wing body. At least a second fuel store may include any fuel store described in this disclosure, including with reference to FIGS. 1 - 2. In some embodiments, second fuel may include one or more of a gasoline-based fuel and a kerosene based fuel.With continued reference to FIG. 14, at step 1420, method 1400 may include propelling aircraft, using at least a propulsor mechanically affixed to the aircraft, wherein the aircraft has a blended wing body. Propulsor may include any propulsor described in this disclosure, for example with reference to FIGS. 1 - 2. Blended wing body may include any blended wing body described in this disclosure, for example with reference to FIGS. 1 - 2.Still referring to FIG. 14, in some embodiments, method 1400 may additionally include burning, using at least a combustion engine of at least a propulsor, second fuel, and producing, using the at least a combustion engine, mechanical work which is used to power the at least apropulsor. Combustion engine may include any combustion engine described in this disclosure, for example with reference to FIGS. 1 - 2.Still referring to FIG. 14, in some embodiments, method 1400 may additionally include powering, using at least a fuel cell, at least an electric motor of at least a propulsor, operatively connected with the at least a fuel cell. Electric motor may include any electric motor described in this disclosure, for example with reference to FIGS. 1 - 2.Still referring to FIG. 14, in some embodiments, method 1400 may additionally include burning, using at least a combustion engine of at least a propulsor, second fuel; producing, using the at least a combustion engine, mechanical work which is used to power the at least a propulsor; and powering, using at least a fuel cell, at least an electric motor of the at least a propulsor, operatively connected with the at least a fuel cell.Still referring to FIG. 14, in some embodiments, method 1400 may additionally include powering, using at least a fuel cell, an auxiliary power system operatively connected with the at least a fuel cell. Auxiliary power system may include any auxiliary power system described in this disclosure, for example with reference to FIGS. 1 - 2. In some cases, method 1400 may further include powering, using auxiliary power system, one or more of an avionic system, a flight control system, an environmental control system, and anti-ice system, a lighting system, a fuel system, a braking system, and a landing gear system.Still referring to FIG. 14, in some embodiments, method 1400 may additionally include storing, using a second fuel store, a second fuel, burning, using at least a combustion engine of at least a propulsor, the second fuel, and producing, using the at least a combustion engine, mechanical work which is used to power the at least a propulsor. Second fuel may include any fuel described in this disclosure, for example with reference to FIGS. 1 - 2.Referring now to FIG. 15, an exemplary method 1500 of use of a system for a blended wing body aircraft with a combustion engine. Blended wing body aircraft may be any of the blended wing body aircrafts described herein with reference to FIGS. 1 and 2. Combustion engine may be any of the combustion engines described herein with reference to FIG. 2.Still referring to FIG. 15, at step 1505, method 1500 includes storing a fuel using at least a fuel store located within a blended wing body of the blended wing body aircraft. At least a fuel store may comprise at least a fuel environment control mitigation. At least a fuel store may bepressurized. Fuel may be any of the fuels described herein with reference to FIG. 1 . At least a fuel source may be any of the fuel sources described herein with reference to FIG. 1. Blended wing body aircraft may be any of the blended wing body aircrafts described herein with reference to FIGS. 1 and 2.Still referring to FIG. 15, at step 1510, method 1500 includes propelling the aircraft, using at least a propulsor mechanically affixed to the blended wing body aircraft and comprising a combustion engine. Blended wing body aircraft may be any of the blended wing body aircrafts described herein with reference to FIGS. 1 and 2. At least a propulsor may be any of the propulsors described herein with reference to FIGS. 1 and 2. Combustion engine may be any of the combustion engines described herein with reference to FIG. 2.Still referring to FIG. 15, at step 1515, method 1500 includes burning, at the combustion engine, the fuel from the fuel store. Combustion engine may be any of the combustion engines described herein with reference to FIG. 2. Fuel may be any of the fuels described herein with reference to FIG. 1. At least a fuel source may be any of the fuel sources described herein with reference to FIG. 1.Still referring to FIG. 15, at step 1520, method 1500 includes producing, at the combustion engine, mechanical work to use to power the at least a propulsor. Combustion engine may be any of the combustion engines described herein with reference to FIG. 2. At least a propulsor may be any of the propulsors described herein with reference to FIGS. 1 and 2.Still referring to FIG. 15, method 1500 may include the use of a fuel cell. At least a fuel cell may be powered by the fuel. At least a fuel cell may be configured to power at least an electric motor. At least a fuel cell may comprise an oxygen source. Still referring to FIG. 15, method 1500 may include at least an auxiliary power unit powered by the fuel and mechanically affixed to the aircraft. An auxiliary power unit may be operatively connected with the at least a fuel cell or may be the actual at least a fuel cell. At least a fuel cell is configured to provide electrical power to aircraft systems. At least a fuel cell starts the airplane and as an emergency backup in the event of an engine failure. At least a fuel cell 216 may also power the combustion engine.Now referring to FIG. 16, a flow diagram for a method 1600 of method of use for fueling an aircraft. At step 1605, method 1600 includes storing liquified gas fuel using a fuel tank. Fueltank may include any fuel tank, as described above in reference to FIGS. 1-13. Liquified gas fuel may include any liquified gas fuel, as described above in reference to FIGS. 1-13.At step 1610, method 1600 includes fueling an aircraft using a fuel line. Fuel line may include any fuel line, as described above in reference to FIGS. 1-13. Fueling an aircraft may also include filling fuel tank with liquified gas fuel to a desired level, wherein a desired level comprises fuel for a plurality of flights plus reserves and removing the fuel line as function of the fuel level in the aircraft. Filling the fuel tank may occur as described above in reference to FIGS. 1-13.At step 1615 method 1600 includes venting the fuel tank, using a vent line in fluid connection to the fuel tank, as described above in reference to FIGS. 1-13. Vent line may include any vent line, as described above in reference to FIGS. 1-13.At step 1620 method 1600 includes preparing the fuel tank for flight as a function of a desired level, as described above in reference to FIGS. 1-13.At step 1625 method 1600 includes flying, using the aircraft, a plurality of flights using the liquid gas fuel, as described above in reference to FIGS. 1-13. Plurality of flights may include any plurality of flights, as described above in reference to FIGS. 1-13. Plurality of flights may include flights to and / or from airports different from that which fuels the aircraft. In some cases, flights may include commercial airline flights. In some cases, plurality of flights may originate and terminate at same airport, where aircraft is fueled.With continued reference to method 1600, the aircraft may include a blended wing body aircraft. The fuel tank may be configured to have a multi-lobe geometry. The fuel tank may include a septum. Additionally, the multi-lobe geometry may provide tension for each container of the multi-lobe geometry. The fill data may include the amount of fuel necessary for the flight. A flight preparation may be engaged as a function of a flight status. A vent line may be configured to capture gaseous hydrogen from the fuel tank. The vent line may be configured to release pressure from the fuel tank. The aircraft may be configured to engage in flight using liquified gas fuel.Now referring to FIG. 17, a block diagram for a method 1700 of use for a blended wing body aircraft with permanent tanks is illustrated. Blended wing body aircraft may be any of theblended wing body aircraft described herein with reference to FIGS. 1-13. Permanent tank may be any of the tanks described herein with reference to FIGS. 1-13.Still referring to FIG. 17, at step 1705, method 1700 of use comprises permanently attaching at least a tank into a blended wing body 1017 of blended wing body aircraft. At least a tank is a pressure vessel. At least a tank further comprises a shape having a plurality of curved surfaces. At least a tank further comprises at least a septum. At least a tank 116 further comprises an inner wall and an outer wall. Blended wing body aircraft may be any of the blended wing body aircraft described herein with reference to FIGS. 1-13. At least a tank 116 may be any of the tanks described herein with reference to FIGS. 1-13.Still referring to FIG. 17, at step 1710, method 1700 of use comprises storing liquid hydrogen fuel inside the at least a tank. Liquid hydrogen fuel may be any of the liquid hydrogen fuels as described herein with reference to FIG. IB. At least a tank may be any of the tanks described herein with reference to FIGS. 1-13.Still referring to FIG. 17, at step 1715, method 1700 of use comprises burning the liquid hydrogen fuel to power blended wing body aircraft. Liquid hydrogen fuel may be any of the liquid hydrogen fuels as described herein with reference to FIGS. 1-13. Blended wing body aircraft may be any of the blended wing body aircraft described herein with reference to FIGS. 1- 13.Still referring to FIG. 17, at step 1720, method 1700 of use comprises venting gaseous hydrogen from the at least a tank using at least a vent. Venting gaseous hydrogen from the at least tank prevents over-pressurizing. At least a tank may be any of the tanks described herein with reference to FIGS. 1-13. At least a vent may be any of the vents described herein with reference to FIGS. 1-13.Still referring to FIG. 17, at step 1725, method 1700 of use comprises reducing thermal transfer to the liquid hydrogen fuel inside of the at least a tank using an insulation. Insulation of the at least a tank includes a chamber located between the inner wall and the outer wall of the at least a tank. Chamber contains gas that is actively pumped, for example into or out of the chamber. In some cases, actively pumping gas from chamber may reduce density of the gas in chamber, thereby reducing thermal heat transfer through the gas by way of convection and / or convection and improving insulative properties. In some cases, flowing gas from chamber, forexample through active pumping, may prevent the gas from exceeding a threshold concentration of hydrogen gas. In some cases, threshold concentration of hydrogen gas may be selected in order to prevent combustion and / or explosion, for instance within the chamber. Liquid hydrogen fuel may be any of the liquid hydrogen fuels as described herein with reference to FIGS. 1-13. At least a tank may be any of the tanks described herein with reference to FIGS. 1-13. Insulation may be any of the insulation described herein with reference to FIGS. 1-13.Now referring to FIG. 18, a block diagram for a method 1800 of method of manufacturing fuel tanks stored aft of the cabin in the main body of an aircraft is illustrated. Step 1805 of method 1800 includes receiving a blending wing body. Blended wing body aircraft may be any of the blended wing body aircraft described herein with reference to FIGS. 1-13.Still referring to FIG. 18, at step 1810, method 1800 of manufacturing includes receiving a plurality of fuel tanks. Plurality of fuel tanks may include a variety of diameters and lengths. A tank of the plurality of fuel tanks may have an inner wall and an outer wall. Plurality of fuel tanks may include multi-lobe geometry consistent with any multi-lobe geometric as discussed in this disclosure. Plurality of fuel tanks may provide structural support to the aircraft. Fuel tank may be any of the tanks described herein with reference to FIGS. 1-13.Still referring to FIG. 18, at step 1815, method 1800 of manufacturing includes locating the plurality of fuel tanks aft of the cabin within the main body. Plurality of fuel tanks may be permanently attached to the blended wing body. Tank and blended wing body may be constructed separately, and then tank may be independently mounted within. “Independently mounted,” as used herein, means mounted separately from the structure of the aircraft or other elements within the aircraft. Plurality of fuel tanks may be integrated within a structure of the blended wing body. For example, fuel tanks may be mounted such that the fuel tanks may support the airframe of the aircraft. This may entail mechanically connecting the fuel tanks to the airframe. This integration of the fuel tanks makes the fuel tanks “permanently attached” as removing the fuel tanks may damage the airframe of the aircraft. Integration within the structure may make the fuel tanks and the aircraft mutually dependent. Fuel tanks may be mounted with rigid mounts, or linked mounts, or the like. Mounting fuel tanks may be consistent with any mounting as discussed in FIGS. 1-13.Still referring to FIG. 18, at step 1820, method 1800 of manufacturing includes storing, using the plurality of fuel tanks, liquified gas fuel. Liquified gas fuel may be any of the liquified gas fuels as described herein with reference to FIGS. 1-13. Fuel tank may be any of the tanks described herein with reference to FIGS. 1-13.Now referring to FIG. 19, a block diagram for a method 1900 of method of manufacturing fuel tanks stored aft of the cabin in the main body of an aircraft is illustrated. Step 1905 of method 1900 includes receiving a blending wing body. Blended wing body aircraft may be any of the blended wing body aircraft described herein with reference to FIGS. 1-13.Still referring to FIG. 19, at step 1910, method 1900 of manufacturing includes receiving a plurality of fuel tanks. Plurality of fuel tanks may include a variety of diameters and lengths. A tank of the plurality of fuel tanks may have an inner wall and an outer wall. Plurality of fuel tanks may include multi-lobe geometry consistent with any multi-lobe geometric as discussed in this disclosure. Plurality of fuel tanks may provide structural support to the aircraft. Fuel tank may be any of the tanks described herein with reference to FIGS. 1-13.Still referring to FIG. 19, at step 1915, method 1900 of manufacturing includes locating the plurality of fuel tanks aft of the cabin within the main body. Plurality of fuel tanks may be permanently attached to the blended wing body. Tank and blended wing body may be constructed separately, and then tank may be independently mounted within. “Independently mounted,” as used herein, means mounted separately from the structure of the aircraft or other elements within the aircraft. Plurality of fuel tanks may be integrated within a structure of the blended wing body. For example, fuel tanks may be mounted such that the fuel tanks may support the airframe of the aircraft. This may entail mechanically connecting the fuel tanks to the airframe. This integration of the fuel tanks makes the fuel tanks “permanently attached” as removing the fuel tanks may damage the airframe of the aircraft. Integration within the structure may make the fuel tanks and the aircraft mutually dependent. Fuel tanks may be mounted with rigid mounts, or linked mounts, or the like. Mounting fuel tanks may be consistent with any mounting as discussed in FIGS. 1-13.Still referring to FIG. 19, at step 1920, method 1900 of manufacturing includes storing, using the plurality of fuel tanks, liquified gas fuel. Liquified gas fuel may be any of the liquifiedgas fuels as described herein with reference to FIGS. 1 -13. Fuel tank may be any of the tanks described herein with reference to FIGS. 1-13.Now referring to FIG. 20, a method 2000 for venting a fuel tank. Step 20020 of method 2000 includes connecting a vent line to a fuel tank in a grounded blended wing body aircraft, wherein the fuel tank contains liquified gas fuel, and the vent line contains gaseous fuel. Fuel tank may be pressurized. The addition of gaseous fuel from boil-off may increase the pressure in the fuel tank. Tank may include a vent configured to release pressure buildup from the fuel tank. Vent may be connected to a vent line. Vent line may direct gaseous fuel overboard, such as into an external tank. Vent line may be insulated to prevent the accumulation of frozen gases, such as frozen water vapor. Vent line may include an inner wall and an outer wall, wherein a vacuum resides in between. Vacuum may act as an insulation to prevent an accumulation of frozen gases on the outer wall. Vent line may be consistent with any vent line in FIG. 1-13.Step 2010 of method 2000 includes heating the vent line. Heating the vent line may warm the extremely cold vented gas. A hot fluid channel may surround the vent line to heat the vent line. Heat may be provided by a powered heat source such as a radiator, electric resistance heater, combustion heater, or the like. Hot fluid channel may be consistent with any hot fluid channel as discussed in FIGS. 1-13.Step 20120 of method 2000 includes collecting the gaseous fuel in an external fuel tank. A distal end of the vent line may be connected to an external tank, located overboard the aircraft. External fuel tank may be configured to re-liquefy the gaseous fuel to be used in the aircraft. Reliquefying the hydrogen gas includes using a refrigeration cycle to change the gas back into liquid fuel. Refrigeration cycle includes condensers, throttle valves, compressors, and heat exchangers.Now referring to FIG. 21, a block diagram for a method 2100 of method of use for an aircraft fueling apparatus. At Step 2105 of method 2100 includes receiving at least a container, as described above in reference to FIGS. 1-13.Still referring to FIG. 21, at step 2110, method 2100 includes using a translocation device configured to carry the at least a container, to a first position, as described above in reference to FIGS. 1-13.Still referring to FIG. 21, at step 2121, method 2100 includes directing the translocation device using an orientation guidance track, wherein the orientation guidance track is configured to direct the movement of the translocation device to a first position, as described above with reference to FIGS . 1-13.Still referring to FIG. 21, the at least a container may include a fuel tank. The at least a container may also include a septum. The at least a container may be configured to have a multilobe geometry. The multi-lobe geometry may provide a pure tension for each fuel tank of the multi-lobe geometry. The orientation guidance track may be oriented longitudinally within the aircraft. The orientation guidance track may also be configured to restrain movement of the translocation device. The orientation guidance may be configured to be in a L-shape. The aircraft may include a blended wing body aircraft.Now referring to FIG. 22, a method of manufacturing 2200 for a multi-walled fuel tank for an aircraft is shown. Step 2205 of method 2200 includes receiving a blended wing body. Blended wing body may be consistent with any blended wing body as discussed herein and in reference to FIGS. 1-13. Step 2210 of method 2200 includes manufacturing at least a fuel tank configured to store liquified gas fuel. Fuel tank may contain liquified gas fuel. Manufacturing a fuel tank includes receiving an inner wall. Inner wall may include an aluminum liner to help prevent the permeation of gases. Inner wall may be consistent with any inner wall as discussed herein and with reference to FIGS. 1-13. Manufacturing a fuel tank also includes receiving an outer wall. Outer wall may be offset from the inner wall. The offset space may include an interstitial volume. Outer wall may be corrugated to increase the strength of the wall and decrease the thickness of the wall. Outer wall may be configured to change sizes due to the evacuation of the interstitial volume. The vacuum may put compression loads on the outer wall causing the outer wall to shrink. Outer wall may be consistent with any outer wall as discussed herein, and with reference to FIGS. 1-13.Continuing to reference FIG. 22, manufacturing a fuel tank also includes forming an interstitial volume including at least a reflective film layer and at least a structural insulation layer between the inner wall and the outer wall. Interstitial volume may include a gap to allow leaked gases to be purged. Gap may be found between the blocks of structural insulation as shown in FIGS. 1-13. Insulation may include layers of reflective film sandwiched in betweenstructural insulation that is divided into blocks. Reflective film layer may include a metalized plastic film. In an embodiment, the metal in the reflective film layer may include aluminum, copper, or the like. Structural insulation may include aerogel. Interstitial volume may be evacuated to form a vacuum. Interstitial volume may be evacuated to an absolute pressure of 0 to 5 PSI. In an embodiment, interstitial volume may be evacuated through the use of a vacuum pump. Vacuum pump may be connected to the interstitial volume through a vent such as a valve. Structural insulation, gap, metalized plastic film, and interstitial volume may be consistent with any structural insulation, gap, metalized plastic film, and interstitial volume, respectively, as disclosed herein.Step 2215 of method 2200 includes attaching the at least a fuel tank to the blended wing body. Fuel tank and blended wing body may be constructed separately, then fuel tank may be independently mounted within. “Independently mounted,” as used herein, means mounted separately from the structure of the aircraft or other elements within the aircraft. Plurality of fuel tanks may be integrated within a structure of the blended wing body. For example, fuel tanks may be mounted such that the fuel tanks may support the airframe of the aircraft. This may entail mechanically connecting the fuel tanks to the airframe. This integration of the fuel tanks makes the fuel tanks permanently attached as removing the fuel tanks may damage the airframe of the aircraft. Integration within the structure may make the fuel tanks and the aircraft mutually dependent. Fuel tanks may be mounted with rigid mounts, or linked mounts, or the like.It is to be noted that any one or more of the aspects and embodiments described herein may be conveniently implemented using digital electronic circuitry, integrated circuitry, specially designed application specific integrated circuits (ASICs), field programmable gate arrays (FPGAs) computer hardware, firmware, software, and / or combinations thereof, as realized and / or implemented in one or more machines (e.g., one or more computing devices that are utilized as a user computing device for an electronic document, one or more server devices, such as a document server, etc.) programmed according to the teachings of the present specification, as will be apparent to those of ordinary skill in the computer art. These various aspects or features may include implementation in one or more computer programs and / or software that are executable and / or interpretable on a programmable system including at least one programmable processor, which may be special or general purpose, coupled to receive data and instructionsfrom, and to transmit data and instructions to, a storage system, at least one input device, and at least one output device. Appropriate software coding may readily be prepared by skilled programmers based on the teachings of the present disclosure, as will be apparent to those of ordinary skill in the software art. Aspects and implementations discussed above employing software and / or software modules may also include appropriate hardware for assisting in the implementation of the machine executable instructions of the software and / or software module.Such software may be a computer program product that employs a machine-readable storage medium. A machine-readable storage medium may be any medium that is capable of storing and / or encoding a sequence of instructions for execution by a machine (e.g., a computing device) and that causes the machine to perform any one of the methodologies and / or embodiments described herein. Examples of a machine-readable storage medium include, but are not limited to, a magnetic disk, an optical disc (e.g., CD, CD-R, DVD, DVD-R, etc ), a magnetooptical disk, a read-only memory “ROM” device, a random access memory “RAM” device, a magnetic card, an optical card, a solid-state memory device, an EPROM, an EEPROM, Programmable Logic Devices (PLDs), and / or any combinations thereof. A machine-readable medium, as used herein, is intended to include a single medium as well as a collection of physically separate media, such as, for example, a collection of compact discs or one or more hard disk drives in combination with a computer memory. As used herein, a machine-readable storage medium does not include transitory forms of signal transmission.Such software may also include information (e.g., data) carried as a data signal on a data carrier, such as a carrier wave. For example, machine-executable information may be included as a data-carrying signal embodied in a data carrier in which the signal encodes a sequence of instruction, or portion thereof, for execution by a machine (e.g., a computing device) and any related information (e.g., data structures and data) that causes the machine to perform any one of the methodologies and / or embodiments described herein.Examples of a computing device include, but are not limited to, an electronic book reading device, a computer workstation, a terminal computer, a server computer, a handheld device (e.g., a tablet computer, a smartphone, etc.), a web appliance, a network router, a network switch, a network bridge, any machine capable of executing a sequence of instructions thatspecify an action to be taken by that machine, and any combinations thereof. In one example, a computing device may include and / or be included in a kiosk.FIG. 23 shows a diagrammatic representation of one embodiment of a computing device in the exemplary form of a computer system 2300 within which a set of instructions for causing a control system to perform any one or more of the aspects and / or methodologies of the present disclosure may be executed. It is also contemplated that multiple computing devices may be utilized to implement a specially configured set of instructions for causing one or more of the devices to perform any one or more of the aspects and / or methodologies of the present disclosure. Computer system 2300 includes a processor 2304 and a memory 2308 that communicate with each other, and with other components, via a bus 2312. Bus 2312 may include any of several types of bus structures including, but not limited to, a memory bus, a memory controller, a peripheral bus, a local bus, and any combinations thereof, using any of a variety of bus architectures.Memory 2308 may include various components (c. ., machine-readable media) including, but not limited to, a random-access memory component, a read only component, and any combinations thereof. In one example, a basic input / output system 2316 (BIOS), including basic routines that help to transfer information between elements within computer system 2300, such as during start-up, may be stored in memory 2308. Memory 2308 may also include (e.g., stored on one or more machine-readable media) instructions (e.g., software) 2320 embodying any one or more of the aspects and / or methodologies of the present disclosure. In another example, memory 2308 may further include any number of program modules including, but not limited to, an operating system, one or more application programs, other program modules, program data, and any combinations thereof.Computer system 2300 may also include a storage device 2324. Examples of a storage device e.g., storage device 2324) include, but are not limited to, a hard disk drive, a magnetic disk drive, an optical disc drive in combination with an optical medium, a solid-state memory device, and any combinations thereof. Storage device 2324 may be connected to bus 2312 by an appropriate interface (not shown). Example interfaces include, but are not limited to, SCSI, advanced technology attachment (ATA), serial ATA, universal serial bus (USB), IEEE 1394 (FIREWIRE), and any combinations thereof. In one example, storage device 2324 (or one ormore components thereof) may be removably interfaced with computer system 2300 (e.g., via an external port connector (not shown)). Particularly, storage device 2324 and an associated machine-readable medium 2328 may provide nonvolatile and / or volatile storage of machine- readable instructions, data structures, program modules, and / or other data for computer system 2300. In one example, software 2320 may reside, completely or partially, within machine- readable medium 2328. In another example, software 2320 may reside, completely or partially, within processor 2304.Computer system 2300 may also include an input device 2332. In one example, a user of computer system 2300 may enter commands and / or other information into computer system 2300 via input device 2332. Examples of an input device 2332 include, but are not limited to, an alpha-numeric input device (e.g, a keyboard), a pointing device, a joystick, a gamepad, an audio input device (e.g., a microphone, a voice response system, etc.), a cursor control device (e.g, a mouse), a touchpad, an optical scanner, a video capture device (e.g, a still camera, a video camera), a touchscreen, and any combinations thereof. Input device 2332 may be interfaced to bus 2312 via any of a variety of interfaces (not shown) including, but not limited to, a serial interface, a parallel interface, a game port, a USB interface, a FIREWIRE interface, a direct interface to bus 2312, and any combinations thereof. Input device 2332 may include a touch screen interface that may be a part of or separate from display 2336, discussed further below. Input device 2332 may be utilized as a user selection device for selecting one or more graphical representations in a graphical interface as described above.A user may also input commands and / or other information to computer system 2300 via storage device 2324 (e.g, a removable disk drive, a flash drive, etc.) and / or network interface device 2340. A network interface device, such as network interface device 2340, may be utilized for connecting computer system 2300 to one or more of a variety of networks, such as network 2344, and one or more remote devices 2348 connected thereto. Examples of a network interface device include, but are not limited to, a network interface card (e.g., a mobile network interface card, a LAN card), a modem, and any combination thereof. Examples of a network include, but are not limited to, a wide area network (e.g., the Internet, an enterprise network), a local area network (e.g., a network associated with an office, a building, a campus or other relatively small geographic space), a telephone network, a data network associated with a telephone / voiceprovider (e.g, a mobile communications provider data and / or voice network), a direct connection between two computing devices, and any combinations thereof. A network, such as network 2344, may employ a wired and / or a wireless mode of communication. In general, any network topology may be used. Information (e.g., data, software 2320, etc.) may be communicated to and / or from computer system 2300 via network interface device 2340.Computer system 2300 may further include a video display adapter 2352 for communicating a displayable image to a display device, such as display device 2336. Examples of a display device include, but are not limited to, a liquid crystal display (LCD), a cathode ray tube (CRT), a plasma display, a light emitting diode (LED) display, and any combinations thereof. Display adapter 2352 and display device 2336 may be utilized in combination with processor 2304 to provide graphical representations of aspects of the present disclosure. In addition to a display device, computer system 2300 may include one or more other peripheral output devices including, but not limited to, an audio speaker, a printer, and any combinations thereof. Such peripheral output devices may be connected to bus 2312 via a peripheral interface 2356. Examples of a peripheral interface include, but are not limited to, a serial port, a USB connection, a FIREWIRE connection, a parallel connection, and any combinations thereof.The foregoing has been a detailed description of illustrative embodiments of the invention. Various modifications and additions can be made without departing from the spirit and scope of this invention. Features of each of the various embodiments described above may be combined with features of other described embodiments as appropriate in order to provide a multiplicity of feature combinations in associated new embodiments. Furthermore, while the foregoing describes a number of separate embodiments, what has been described herein is merely illustrative of the application of the principles of the present invention. Additionally, although particular methods herein may be illustrated and / or described as being performed in a specific order, the ordering is highly variable within ordinary skill to achieve embodiments as disclosed herein. Accordingly, this description is meant to be taken only by way of example, and not to otherwise limit the scope of this invention.In the descriptions above and in the claims, phrases such as “at least one of’ or “one or more of’ may occur followed by a conjunctive list of elements or features. The term “and / or” may also occur in a list of two or more elements or features. Unless otherwise implicitly orexplicitly contradicted by the context in which it is used, such a phrase is intended to mean any of the listed elements or features individually or any of the recited elements or features in combination with any of the other recited elements or features. For example, the phrases “at least one of A and B;” “one or more of A and B;” and “A and / or B” are each intended to mean “A alone, B alone, or A and B together.” A similar interpretation is also intended for lists including three or more items. For example, the phrases “at least one of A, B, and C;” “one or more of A, B, and C;” and “A, B, and / or C” are each intended to mean “A alone, B alone, C alone, A and B together, A and C together, B and C together, or A and B and C together.” In addition, use of the term “based on,” above and in the claims is intended to mean, “based at least in part on,” such that an unrecited feature or element is also permissible.The subject matter described herein can be embodied in systems, apparatus, methods, and / or articles depending on the desired configuration. The implementations set forth in the foregoing description do not represent all implementations consistent with the subject matter described herein. Instead, they are merely some examples consistent with aspects related to the described subject matter. Although a few variations have been described in detail above, other modifications or additions are possible. In particular, further features and / or variations can be provided in addition to those set forth herein. For example, the implementations described above can be directed to various combinations and sub-combinations of the disclosed features and / or combinations and sub-combinations of several further features disclosed above. In addition, the logic flows depicted in the accompanying figures and / or described herein do not necessarily require the particular order shown, or sequential order, to achieve desirable results. Other implementations may be within the scope of the following claims.

Claims

WHAT TS CLAIMED IS:

1. A blended wing body aircraft with a fuel cell, the aircraft comprising: a blended wing body; at least a first fuel store located within the blended wing body and configured to store a first fuel; at least a fuel cell configured to combine the first fuel with oxygen to produce electricity; at least a second fuel store located within the blended wing body and configured to store a second fuel; and at least a propulsor mechanically affixed to the aircraft and configured to propel the blended wing body aircraft.

2. The aircraft of claim 1, wherein the first fuel comprises one or more of liquid hydrogen and natural gas.

3. The aircraft of claim 1, wherein the second fuel comprises one or more of a kerosene based fuel and a gasoline based fuel.

4. The aircraft of claim 1, wherein the at least a propulsor comprises at least a combustion engine that burns the second fuel and produces mechanical work which is used to power the at least a propulsor.

5. The aircraft of claim 1, wherein the at least a propulsor comprises at least an electric motor operatively connected with the at least a fuel cell; and wherein the at least a fuel cell is configured to power the at least an electric motor.

6. The aircraft of claim 1, wherein the at least a propulsor comprises: at least a combustion engine that burns the second fuel and produces mechanical work which is used to power the at least a propulsor; and at least an electric motor operatively connected with the at least a fuel cell; and wherein the at least a fuel cell is configured to power the at least an electric motor.

7. The aircraft of claim 1, further comprising: an auxiliary power system operatively connected with the at least a fuel cell; and wherein the at least a fuel cell is configured to power the auxiliary power system.

8. The aircraft of claim 7, wherein the auxiliary power system is configured to power one or more of an avionic system, a flight control system, an environmental control system, ananti-ice system, a lighting system, a fuel system, a braking system, and a landing gear system.

9. The aircraft of claim 7, wherein the at least a propulsor comprises a combustion engine and the auxiliary power system is further configured to start the internal combustion engine.

10. The aircraft of claim 1, wherein the at least a first fuel store is located within a transitional portion area of the blended wing body and the at least a second fuel store is located within a wing portion of the blended wing body.

11. A method of use of a blended wing body aircraft with a fuel cell, the method comprising: storing a first fuel, using at least a first fuel store located within a blended wing body of the blended wing body aircraft; combining the first fuel with oxygen to produce electricity, using at least a fuel cell; storing a second fuel, using at least a second fuel store located within the blended wing body; and propelling the aircraft, using at least a propulsor mechanically affixed to the blended wing body aircraft.

12. The method of claim 11, wherein the first fuel comprises one or more of liquid hydrogen and natural gas.

13. The method of claim 11, wherein the second fuel comprises one or more of a kerosenebased fuel and a gasoline based fuel.

14. The method of claim 11, further comprising: burning, using at least a combustion engine of the at least a propulsor, the second fuel; and producing, using the at least a combustion engine, mechanical work which is used to power the at least a propulsor.

15. The method of claim 11, further comprising: powering, using the at least a fuel cell, at least an electric motor of the at least a propulsor, operatively connected with the at least a fuel cell.

16. The method of claim 11, further comprising: burning, using at least a combustion engine of the at least a propulsor, the second fuel;producing, using the at least a combustion engine, mechanical work which is used to power the at least a propulsor; and powering, using the at least a fuel cell, at least an electric motor of the at least a propulsor, operatively connected with the at least a fuel cell.

17. The method of claim 11, further comprising: powering, using the at least a fuel cell, an auxiliary power system operatively connected with the at least a fuel cell.

18. The method of claim 17, further comprising: powering, using the auxiliary power system, one or more of an avionic system, a flight control system, an environmental control system, and anti-ice system, a lighting system, a fuel system, a braking system, and a landing gear system.

19. The method of claim 17, further comprising: starting, using the auxiliary power unit, a combustion engine of the at least a propulsor.

20. The method of claim 11, wherein the at least a first fuel store is located within a transitional portion area of the blended wing body and the at least a second fuel store is located within a wing portion of the blended wing body.

21. A system for a blended wing body aircraft with a combustion engine, the aircraft comprising: a blended wing body; at least a fuel store located within the blended wing body and configured to store a fuel, wherein the fuel includes liquid hydrogen; at least a propulsor configured to propel the blended wing body aircraft, the at least a propulsor comprising a combustion engine configured to: burn the fuel from the fuel store; and produce mechanical work to power the at least a propulsor.

22. The system of claim 21, wherein the at least a fuel store comprises at least a fuel environment control mitigation.

23. The system of claim 21, wherein the at least a fuel store is pressurized.

24. The system of claim 21, further comprising at least an auxiliary power system powered by the fuel and mechanically affixed to the aircraft.

25. The system of claim 21, further comprising at least a fuel cell powered by the fuel.

26. The system of claim 25, wherein the at least a fuel cell is configured to power at least an electric motor.

27. The system of claim 25, wherein the at least a fuel cell comprises an oxygen source.

28. The system of claim 25, wherein the at least a fuel cell is configured to provide electrical power to aircraft systems.

29. The system of claim 24, wherein the at least a fuel cell is configured to power one or more of an avionic system, a flight control system, an environmental control system, an anti-ice system, a lighting system, a fuel system, a braking system, and a landing gear system.

30. The system of claim 24, wherein the at least a fuel cell is configured to start the combustion engine.

31. A method of use of a system for a blended wing body aircraft with a combustion engine, the method comprising: storing a fuel using at least a fuel store located within a blended wing body of the blended wing body aircraft; propelling the aircraft, using at least a propulsor mechanically affixed to the blended wing body aircraft, the at least at propulsor comprising a combustion engine; burning, at the combustion engine, the fuel from the fuel store; and producing, at the combustion engine, mechanical work to power the at least a propulsor.

32. The method of claim 31, wherein the at least a fuel store comprises at least a fuel environment control mitigation.

33. The method of claim 31, wherein the at least a fuel store is pressurized.

34. The method of claim 31, further comprising at least an auxiliary power system powered by the fuel and mechanically affixed to the aircraft.

35. The method of claim 31, further comprising at least a fuel cell powered by the fuel.

36. The method of claim 35, wherein the at least a fuel cell is configured to power at least an electric motor.

37. The method of claim 35, wherein the at least a fuel cell comprises an oxygen source.

38. The method of claim 35, wherein the at least a fuel cell is configured to provide electrical power to aircraft systems.

39. The method of claim 31, wherein the at least a fuel cell is configured to power one or more of an avionic system, a flight control system, an environmental control system, an anti-ice system, a lighting system, a fuel system, a braking system, and a landing gear system.

40. The method of claim 34, wherein the at least a fuel cell is configured to start the combustion engine.

41. A method for fueling and using an aircraft, wherein the method comprises: storing, using a fuel tank, liquified gas fuel; fueling, using a fuel line, an aircraft, wherein fueling further comprises: filling the fuel tank with liquified gas fuel to a desired level, wherein the desired level comprises fuel for a plurality of flights plus reserves; and removing the fuel line as function of the desired level in the aircraft; and preparing the fuel tank for flight as a function of the desired level; and flying, using the aircraft, a plurality of flights using the liquid gas fuel.

42. The method of claim 41, wherein the aircraft comprises a blended wing body aircraft.

43. The method of claim 41, wherein the fuel tank is configured to have a multi-lobe geometry.

44. The method of claim 43, wherein the fuel tank comprises a septum.

45. The method of claim 43, wherein the multi-lobe geometry provides tension for each container of the multi-lobe geometry.

46. The method of claim 41, wherein the fuel tank is prepared as a function of a flight status.

47. The method of claim 41, further comprising venting, using a vent line, the fuel tank, wherein the vent line is in fluid connection to the fuel tank.

48. The method of claim 47, wherein the vent line is configured to capture gaseous hydrogen from the fuel tank.

49. The method of claim 47, wherein the vent line is configured to release pressure from the fuel tank.

50. The method of claim 47, wherein the vent line includes a pressure regulator.51 . The method of claim 47, wherein the vent line is insulated to prevent the accumulation of frozen gases.

52. The method of claim 47, wherein the vent line includes an inner wall and an outer wall wherein a void resides in between.

53. The method of claim 47, wherein the vent line further comprises a heat exchanger.

54. The method of claim 47, wherein the preparing the fuel tank for flight further includes detaching the vent line from the fuel tank as a function of the desired level.

55. The method of claim 41, wherein the preparing of the fuel tank for flight further includes detaching the fuel line from the fuel tank as a function of the desired level.

56. The method of claim 41, wherein the plurality of flights is further configured to: commence at a first refueling station; and terminate at a second refueling station.

57. The method of claim 41, wherein substantially no fueling occurs between the flights of the plurality of flights.

58. The method of claim 41, wherein fueling occurs over a period of time, exceeding 30 minutes.

59. The method of claim 41, wherein the fuel tank may comprise a chamber between an inner wall and an outer wall of the tank.

60. The method of claim 41, wherein the liquified gas fuel is comprised of liquid hydrogen.

61. A system for a blended wing body aircraft with built-in, non-removable fuel tanks, the aircraft comprising: a blended wing body; and at least a fuel tank permanently attached to the blended wing body and configured to store liquified gas fuel, wherein the at least a fuel tank further comprises: at least a vent configured to vent gaseous fuel from the at least a fuel tank; and an insulation to reduce thermal transfer to the liquified gas fuel inside of the at least a fuel tank.

62. The system of claim 61, wherein the at least a fuel tank is configured to carry structural loads to a skin of the blended wing body.

63. The system of claim 62, wherein the at least a fuel tank further comprises a shape having a plurality of curved surfaces.

64. The system of claim 63, wherein the at least a fuel tank further comprises at least a septum.

65. The system of claim 61, wherein venting gaseous fuel from the at least fuel tank prevents over-pressurizing.

66. The system of claim 61, wherein the insulation of the at least a fuel tank includes a chamber located between an inner wall and an outer wall of the at least a fuel tank.

67. The system of claim 66, wherein the chamber contains gas that is actively pumped.

68. The system of claim 61, further comprising at least a propulsor configured to propel the blended wing body aircraft.

69. The system of claim 68, further comprising a combustion engine configured to: burn the liquified gas fuel from the at least a fuel tank; and produce mechanical work to use to power the at least a propulsor.

70. The system of claim 64, wherein one or more of the plurality of curved surfaces and the at least a septum connect the at least a fuel tank to at least an upper skin and at least a lower skin of the blended wing body.

71. A method of use for a blended wing body aircraft with built-in, non-removable fuel tanks, the method of use comprising: permanently attaching at least a fuel tank into a blended wing body of the blended wing body aircraft; storing liquified gas fuel inside the at least a fuel tank; powering the blended wing body aircraft by burning the liquified gas fuel; venting gaseous fuel from the at least a fuel tank using at least a vent; and reducing thermal transfer to the liquified gas fuel inside of the at least a fuel tank using an insulation.

72. The method of claim 71, wherein the at least a fuel tank is configured to carry structural loads to a skin of the blended wing body.

73. The method of claim 72, wherein the at least a fuel tank further comprises a shape having a plurality of curved surfaces.

74. The method of claim 73, wherein the at least a fuel tank further comprises at least a septum.

75. The method of claim 71, wherein venting gaseous hydrogen from the at least fuel tank prevents over-pressurizing.

76. The method of claim 71, wherein the insulation of the at least a fuel tank includes a chamber located between an inner wall and an outer wall of the at least a fuel tank.

77. The method of claim 76, wherein the chamber contains gas that is actively pumped.

78. The method of claim 71, further comprising at least a propulsor configured to propel the blended wing body aircraft.

79. The method of claim 79, further comprising a combustion engine configured to: burning the liquified gas fuel from the at least a fuel tank; and producing mechanical work to use to power the at least a propulsor.

80. The method of claim 74, wherein the plurality of curved surfaces and the at least a septum connect the at least a fuel tank to at least an upper skin and at least a lower skin of the blended wing body.

81. An aircraft with fuel tanks stored aft of a cabin in a main body, the aircraft comprising: a blended wing body, comprising: a main body; and a cabin at least partially located within the main body; and a plurality of fuel tanks located at least partially aft of the cabin within the main body and configured to store liquified gas fuel.

82. The aircraft of claim 81, wherein the plurality of fuel tanks is permanently attached to the blended wing body.

83. The aircraft of claim 81, wherein a fuel tank of a plurality of fuel tanks has a multi-lobe geometry.

84. The aircraft of claim 83, wherein the multi-lobe geometry provides tension within a wall of the fuel tank of the plurality of fuel tanks.

85. The aircraft of claim 81, wherein a fuel tank of the plurality of fuel tanks is a tapered tank.

86. The aircraft of claim 81 , wherein a fuel tank of the plurality of fuel tanks further comprises a septum.

87. The aircraft of claim 81, wherein a fuel tank of the plurality of fuel tanks further comprises an inner wall and an outer wall.

88. The aircraft of claim 87, wherein the insulation of the fuel tank of the plurality of fuel tanks includes a chamber located between the inner wall and the outer wall of the fuel tank.

89. The aircraft of claim 88, wherein the chamber contains gas, wherein the gas is actively pumped.

90. The aircraft of claim 81, wherein a fuel tank of a plurality of fuel tanks comprises a variable diameter and length.

91. The aircraft of claim 81, wherein the plurality of fuel tanks provides structural support to the blended wing body.

92. The aircraft of claim 81, wherein the plurality of fuel tanks extend across a full width of the cabin.

93. The aircraft of claim 81, wherein the plurality of fuel tanks comprise a plurality of domed end caps.

94. A method of manufacturing fuel tanks stored aft of the main body of an aircraft, the method comprising: receiving a blended wing body, comprising: a main body; and a cabin at least partially located within the main body; receiving a plurality of fuel tanks; locating the plurality of fuel tanks at least partially aft of the cabin within the main body; and storing, using the plurality of fuel tanks, liquified gas fuel.

95. The method of claim 84, further comprising permanently attaching the plurality of fuel tanks to the blended wing body.

96. The method of claim 85, wherein the plurality of fuel tanks provides structural support to the aircraft.

97. The method of claim 84, wherein a fuel tank of the plurality of fuel tanks has a multi-lobe geometry.

98. The aircraft of claim 81, wherein a fuel tank of the plurality of fuel tanks further comprises an inner wall and an outer wall.

99. The method of claim 84, wherein the plurality of fuel tanks extends across a full width of the cabin.

100. The method of claim 84, wherein permanently attaching comprises: independently mounting the plurality of fuel tanks; and integrating the plurality of fuel tanks within a structure of the blended wing body.

101. A liquified gas fuel tank incorporated in an aircraft to fuel comprising: a first compartment wherein: the first compartment has a first cross-section describing a first continuously convex differentiable curve; the first compartment includes an inner volume configured to contain fuel; and the first compartment is configured to be pressurized; a second compartment wherein: the second compartment has a second cross-section describing a second continuously convex differentiable curve which intersects the first continuously convex differentiable curve at an intersection; and the second compartment includes an additional inner volume fluidly connected to the inner volume of the first compartment; and a junction configured to structurally support each of the first compartment and the second compartment at the intersection.

102. The liquified gas fuel tank of claim 101, wherein: the first compartment comprises a first inner wall defining a first internal cavity of the first compartment and a first outer wall exterior to the first inner wall, wherein a first separation of the first inner wall and the first outer wall defines a first space; and the second compartment comprises a second inner wall defining a second internal cavity of the second compartment and a second outer wall exterior to the second innerwall, wherein a second separation of the second inner wall and the second outer wall defines a second.

103. The liquified gas fuel tank of claim 102, wherein a thickness of the first space is less than 1 / 10 of a first radial distance of the first continuously convex differentiable curve; and wherein the second space is less than 1 / 10 of a second radial distance of the second continuously convex differentiable curve.

104. The liquified gas fuel tank of claim 101, wherein a first wall of the first compartment, a second wall of the second compartment, and the junction are all configured to be in tension when the first compartment is pressurized.

105. The liquified gas fuel tank of claim 101, wherein the first compartment and the second compartment are configured to have equal pressure.

106. The liquified gas fuel tank of claim 102, further comprising a sensing component configured to monitor and control leaks.

107. The liquified gas fuel tank of claim 106, wherein the sensing component comprises a gas evacuation element, one or more sensors, and a controller, wherein the gas evacuation element is communicatively connected to the controller, and wherein the gas evacuation element is configured to sense a gas concentration and purge gas from one or more of the first space and the second space.

108. The liquified gas fuel tank of claim 107, wherein the controller is configured to receive a second indication from the one or more sensors, wherein the second indication indicates that a second gas concentration is below the threshold gas concentration, wherein the controller is configured to deactivate the gas evacuation element based on the second indication.

109. The liquified gas fuel tank of claim 107, wherein a plurality of conduit lines is connected to the sensing component via the gas evacuation element, and wherein the plurality of conduit lines is configured to receive a boiled-off gas.

110. The liquified gas fuel tank of claim 102, wherein the first outer wall and the second outer wall prevent ambient air from entering the first internal cavity and the second internal cavity, respectively.

111. A method of manufacturing a liquified gas fuel tank for an aircraft, the method comprising: constructing a first compartment wherein: the first compartment has a first cross-section describing a first continuously convex differentiable curve; the first compartment includes an inner volume configured to contain fuel; and the first compartment is configured to be pressurized; constructing a second compartment wherein: the second compartment has a second cross-section describing a second continuously convex differentiable curve which intersects the first continuously convex differentiable curve at an intersection; and the second compartment includes an additional inner volume fluidly connected to the inner volume of the first compartment; and connecting the first compartment and the second compartment with a junction configured to structurally support each of the first compartment and the second compartment at the intersection.

112. The method of claim 111, wherein: constructing the first compartment comprises forming a first inner wall defining a first internal cavity of the first compartment and a first outer wall exterior to the first inner wall, wherein a first separation of the first inner wall and the first outer wall defines a first space; and constructing the second compartment comprises forming a second inner wall defining a second internal cavity of the second compartment and a second outer wall exterior to the second inner wall, wherein a second separation of the second inner wall and the second outer wall defines a second.

113. The method claim 112, wherein a thickness of the first space is less than 1 / 10 of a first radial distance of the first continuously convex differentiable curve; and wherein the second space is less than 1 / 10 of a second radial distance of the second continuously convex differentiable curve.

114. The method of claim 111, wherein a first wall of the first compartment, a second wall of the second compartment, and the junction are all configured to be in tension when the first compartment is pressurized.

115. The method of claim 111, wherein the first compartment and the second compartment are configured to have equal pressure.

116. The method of claim 112, further comprising installing a sensing component configured to monitor and control leaks.

117. The method of claim 116, wherein the sensing component comprises a gas evacuation element, one or more sensors, and a controller, wherein the gas evacuation element is communicatively connected to the controller, and wherein the gas evacuation element is configured to sense a gas concentration and purge gas from one or more of the first space and the second space.

118. The method of claim 117, wherein the controller is configured to receive a second indication from the one or more sensors, wherein the second indication indicates that a second gas concentration is below the threshold gas concentration, wherein the controller is configured to deactivate the gas evacuation element based on the second indication.

119. The method of claim 117, wherein a plurality of conduit lines is connected to the sensing component via the gas evacuation element, and wherein the plurality of conduit lines is configured to receive a boiled-off gas.

120. The method of claim 112, wherein the first outer wall and the second outer wall prevent ambient air from entering the first internal cavity and the second internal cavity, respectively.

121. A method for venting a fuel tank, the method comprising: connecting a vent line to a fuel tank in a grounded liquid gas aircraft, wherein the fuel tank contains liquid fuel; heating the vent line; and collecting the gaseous fuel, from the fuel tank, in an external fuel tank, wherein the external fuel tank is fluidly connected to the vent line.

122. The method of claim 121 , further comprising re-liquefying the gaseous fuel in the external fuel tank to be reused in the liquid gas aircraft, wherein re-liquefying the gaseous fuel in the external fuel tank comprises chilling the gaseous fuel.

123. The method of claim 121, wherein the fuel tank is pressurized.

124. The method of claim 121, wherein the fuel tank comprises a vent configured to release pressure buildup from the fuel tank, wherein the vent line is connected to the vent.

125. The method of claim 121, wherein the vent line includes a pressure regulator.

126. The method of claim 121, wherein heating the vent line comprises using a heat exchanger to heat the vent line.

127. The method of claim 121, wherein the vent line is insulated to prevent the accumulation of frozen gases.

128. The method of claim 121, wherein the vent line includes an inner wall and an outer wall wherein a vacuum resides in between.

129. The method of claim 122, wherein re-liquefying the hydrogen gas includes using a refrigeration cycle comprising condensers, throttle valves, compressors, and heat exchangers.

130. The method of claim 121, wherein heating the vent line includes using a powered heat source.

131. A system for venting a fuel tank, the system comprising: a grounded blended wing body aircraft; a vent line connected to a fuel tank in the grounded liquid gas aircraft, wherein the vent line includes a heating element; and an external fuel tank configured to collect the gaseous fuel boil-off from the fuel tank in the grounded liquid gas aircraft.

132. The system of claim 131, wherein the system is further configured to re-liquefy the gaseous fuel to be reused in the liquid gas aircraft.

133. The system of claim 131, wherein the fuel tank is pressurized.

134. The system of claim 131, wherein the fuel tank includes a vent configured to release pressure buildup from the fuel tank.

135. The system of claim 131, wherein the vent line includes a pressure regulator.

136. The system of claim 131, further comprising heat exchanger configured to heat the vent line.

137. The system of claim 131, wherein the vent line is insulated to prevent the accumulation of frozen gases.

138. The system of claim 131, wherein the vent line includes an inner wall and an outer wall wherein a vacuum resides in between.

139. The system of claim 132, wherein re-liquefying the hydrogen gas includes using a refrigeration cycle comprising condensers, throttle valves, compressors, and heat exchangers.

140. The system of claim 131, wherein the heating element comprises a powered heat source.

141. An aircraft fueling apparatus, wherein the apparatus comprises: at least a container comprising a fuel tank configured to store a liquified gas fuel; a translocation device configured to carry the at least a container; and an orientation guidance track, wherein the orientation guidance track is configured to direct a movement of the translocation device to a first position.

142. The apparatus of claim 141, wherein the apparatus is configured to transport the at least a container from a first position to a second position.

143. The apparatus of claim 141, wherein the at least a container is configured to have a multilobe geometry.

144. The apparatus of claim 143, wherein the at least a container comprises a septum.

145. The apparatus of claim 143, wherein the multi -lobe geometry provides tension in a wall of a container.

146. The apparatus of claim 141, wherein the fuel comprises liquid hydrogen fuel.

147. The apparatus of claim 141, wherein the orientation guidance track is located at least partially within the aircraft and the orientation guidance track is oriented longitudinally within the aircraft.

148. The apparatus of claim 141, wherein the orientation guidance track is configured to restrain movement of the translocation device.

149. The apparatus of claim 141, wherein the orientation guidance track is configured to be in a L-shape.

150. The apparatus of claim 141 , wherein the aircraft is a blended wing body aircraft.

151. A method of use for an aircraft fueling apparatus, wherein the method comprises: receiving at least a container comprising a fuel tank; transporting the container, using a translocation device configured to carry the at least a container, to a first position; and directing the translocation device using an orientation guidance track, wherein the orientation guidance track is configured to direct the movement of the translocation device to a first position.

152. The method of claim 151, wherein transporting the container comprises transporting the at least a container from a first position to a second position.

153. The method of claim 151, wherein the at least a container is configured to have a multilobe geometry.

154. The method of claim 153, wherein the at least a container comprises a septum.

155. The method of claim 153, wherein the multi-lobe geometry provides tension in a wall of a container.

156. The method of claim 151, wherein the fuel comprises liquid hydrogen fuel.

157. The method of claim 151, wherein the orientation guidance track is oriented longitudinally within the aircraft.

158. The method of claim 151, wherein the orientation guidance track is configured to restrain movement of the translocation device.

159. The method of claim 151, wherein the orientation guidance track is configured to be in a L-shape.

160. The method of claim 151, wherein the aircraft comprises a blended wing body aircraft.

161. An aircraft with at least a multi-walled fuel tank, the aircraft comprising: a blended wing body; and at least a fuel tank attached to the blended wing body and configured to store liquified gas fuel, wherein the at least a fuel tank further comprises: an inner wall; an outer wall;an interstitial volume between the inner wall and the outer wall comprising of at least a reflective film layer and at least a structural insulation layer.

162. The aircraft of claim 1, wherein the inner wall comprises carbon epoxy.

163. The aircraft of claim 1, wherein the outer wall comprises one or more of aluminum and steel.

164. The aircraft of claim 1, wherein the interstitial volume includes a gap.

165. The aircraft of claim 1, wherein the at least a fuel tank comprising a sensing system comprising at least a gas sensor, wherein the at least a gas sensor is located within a gap of the interstitial volume.

166. The aircraft of claim 5, further comprising a vent fluidly connected to the gap and configured to vent gas from the gap.

167. The aircraft of claim 5, wherein the at least a gas sensor is configured to monitor purge flow rate by measuring velocity of a gas.

168. The aircraft of claim 6, wherein the at least a gas sensor is configured to: detect fuel gas concentration; and control, using the vent, venting of the gas from the gap, when a threshold gas concentration value is detected.

169. The aircraft of claim 1, wherein the at least a structural insulation layer is divided into blocks.

170. The aircraft of claim 1, wherein the at least a structural insulation layer includes aerogel.

171. The aircraft of claim 1, wherein the at least a reflective film layer comprises a metalized plastic film.

172. A method of manufacturing a multi -walled fuel tank for an aircraft, the method comprising: receiving a blended wing body; manufacturing at least a fuel tank configured to store liquified gas fuel, wherein manufacturing the at least a fuel tank comprises: receiving an inner wall; receiving an outer wall; andforming an interstitial volume between the inner wall and the outer wall comprising at least a reflective film layer and at least a structural insulation layer between the inner wall and the outer wall; and attaching the at least a fuel tank to the blended wing body.

173. The method of claim 11, wherein the at least a structural insulation layer is divided into blocks.

174. The method of claim 11, wherein the interstitial volume is evacuated to an absolute pressure of about 0 to 5 PSI.

175. The method of claim 14, wherein the outer wall is corrugated.

176. The method of claim 15, wherein the outer wall is configured to change sizes due to the evacuation of the interstitial volume.

177. The method of claim 11, wherein the inner wall comprises an aluminum liner.

178. The method of claim 11, wherein the at least a structural insulation layer includes aerogel.

179. The method of claim 11, wherein the at least a reflective film layer comprises a metalized plastic film.

180. The method of claim 11, wherein the interstitial volume includes a gap.