Aerospace vehicle thermal protection systems

Abstract

Thermal protection systems for aerospace vehicles, such as hypersonic vehicles or space vehicles, are disclosed herein. In some embodiments, the disclosed thermal protection systems can be used to protect propellant tanks (e.g., composite propellant tanks) and vehicle primary structures (e.g., composite primary structures) from exceeding upper design temperature limits during vehicle ascent, reentry into Earth's atmosphere, and / or high velocity flight through an atmosphere. In some embodiments, a thermal protection system can include a framework of elongated composite members forming a plurality of cells. One or more layers of insulation material such as aerogel can be positioned in each of the cells. The cells can be covered by an external skin of the vehicle. In some embodiments, the framework is attached to a propellant tank and the insulation material is in contact with the external surfaced of the propellant tank.

Description

AEROSPACE VEHICLE THERMAL PROTECTION SYSTEMSCROSS-REFERENCE TO RELATED APPLICATION(S)

[0001] The present application claims priority to U.S. Provisional Patent Application No. 63 / 553,746, filed February 15, 2024, which is incorporated herein by reference in its entirety.TECHNICAL FIELD

[0002] The present disclosure is generally related to space and hypersonic vehicle thermal protection systems and, more particularly, to thermal protection systems for insulating propellant tanks and / or other portions of space and / or hypersonic vehicles, e.g., during ascent, reentry into Earth’s atmosphere, and / or high velocity flight through an atmosphere.BACKGROUND

[0003] “ Thermal protection system” (“TPS”) traditionally refers to insulative and / or ablative materials on the exterior of an aerospace vehicle that are designed to maintain the relatively low operating temperatures of internal vehicle structures. Conventional TPSs only provide thermal protection and do not perform a structural function, so the weight of the TPS is detrimental to the overall vehicle system mass. Systems that support structural functions are typically considered “hot structure” and — although they may provide a modest thermal barrier — they lack significant insulative properties and often rely on active cooling methods that also tend to increase mass.

[0004] Some space vehicles have used ceramic tiles or blankets adhesively bonded to external surfaces of the vehicle to provide thermal protection during reentry into Earth’s atmosphere. Ceramic tiles, however, tend to be fragile, brittle, and easily damaged. Other space vehicles with cryogenic propellant tanks have used foam insulation that is sprayed onto or bonded to the tank wall to limit condensation, air liquefaction and cryopumping at or near the tank surface. Further examples of existing space vehicle thermal protection systems include Toughened Uni-piece Fibrous Reinforced Oxidation-resistant Composite (TUFROC) tiles, Alumina Enhanced Thermal Barrier (AETB) tiles (e.g., AETB-8 / AETB-12 / TUFI tiles), Ceramic Matrix Composite (CMC) shingles, metallic armor, and metallic honeycomb panels. Exterior surfaces of some hypersonic vehicles may include nickel-based superalloy metals and / or CMC components towithstand aerothermal heating. Some such metals include, for example, HASTELLOY from Haynes International, INCONEL from Special Metals Corporation, WASPALOY, or RENE.

[0005] If a vehicle’ s external surface is metal with a relatively high upper service temperature limit, the TPS does not have to provide as much thermal insulation as would be required for vehicles with composite materials forming the external surfaces. For composite vehicles, a ceramic tile system would require a thicker, heavier, and less durable TPS to stay within the upper service temperature limit of the composite materials. Metallic systems have been explored in standoff tile structures and armor tiles that could also serve a dual purpose in carrying some of the aerodynamic flight loads, but these systems often encapsulated insulative material in foils, resulting in the combined weight of the metal structure and insulation exceeding the vehicle’s operability.

[0006] Various hypersonic and / or space vehicle thermal protection systems are described in the following publications, each of which is incorporated herein by reference in its entirety: “Ceramic Matrix Composite (CMC) Thermal Protection Systems (TPS) and Hot Structures for Hypersonic Vehicles,” D. E. Glass, 15thAIAA Space Planes and Hypersonic Systems and Technologies Conference, April 28-May 1, 2008, Dayton, OH; “Advanced Metallic Thermal Protection System Development - AIAA 2002-0504,” M.L. Blosser et al., 40thAerospace Sciences Meeting & Exhibit, January 14-17, 2002, Reno, NV; “Reusable Metallic Thermal Protection Systems Development,” M. L. Blosser et al., March 25, 1998, NASA Langley Research Center, Hampton, VA; “From IXV to Space Rider: CMC Thermal Protection System Evolutions,” J. Valverde et al., HT-CMC / 10th, September 22-26, 2019, Bordeaux, FR; “From IXV to Space Rider: CMC Thermal Protection System Evolutions,” J. Valverde et al., DOI: 10.13009 / EUCASS2019-991, 8thEuropean Conference for Aeronautics and Space Sciences (EUCASS), 2019; and “A Competitive Thermal Protection System For Hypersonic Vehicles,” S. lanelli et al., IAC-13-C2.4.1 xl9563, 64thIntT Astronautical Congress, September 23-27, 2013, Beijing China.

[0007] Existing materials and / or TPSs may not provide sufficient insulation of underlying structure to maintain structural integrity of an aerospace vehicle, and / or they may be too fragile, too heavy, too difficult to maintain, and / or have other disadvantages that limit their utility in certain applications, such as in the context of reusable vehicles.BRIEF DESCRIPTION OF THE DRAWINGS

[0008] Figures 1A and IB are partially schematic, top front isometric and bottom front isometric views, respectively, of an aerospace vehicle having a thermal protection system configured in accordance with embodiments of the present technology.

[0009] Figure 2A is a partially schematic top view of the aerospace vehicle of Figures 1A and IB, and Figure 2B is an enlarged cross-sectional view taken substantially along line 2B-2B in Figure 2A to illustrate a portion of a propellant tank and thermal protection system configured in accordance with embodiments of the present technology.

[0010] Figures 3A-3C illustrate a series of partially exploded isometric views of the portion of the propellant tank and thermal protection system shown in Figure 2B, and Figure 3D illustrates a corresponding assembled isometric view of the portion of the propellant tank and thermal protection system shown in Figure 2B, configured in accordance with embodiments of the present technology.

[0011] Figures 4A-4C are partially exploded isometric, assembled isometric, and partially- schematic cross-sectional views, respectively, of a portion of the thermal protection system shown in Figure 3D, configured in accordance with embodiments of the present technology.

[0012] Figure 5 is an isometric view of a portion of an aerospace vehicle propellant tank having a thermal protection system configured in accordance with embodiments of the present technology.

[0013] Figures 6A-6C are isometric, top, and cross-sectional views, respectively, of a portion of the thermal protection system of Figure 5.

[0014] Figures 6D and 6E are enlarged cross-sectional views showing details of portions of the thermal protection system shown in Figure 5, configured in accordance with embodiments of the present technology.DETAILED DESCRIPTION

[0015] The following disclosure describes various embodiments of thermal protection systems for aerospace vehicles, such as hypersonic vehicles and / or space vehicles. In some embodiments, the thermal protection systems described herein can protect internal vehicle structures and / or components (e.g., composite propellant tanks) from exceeding their upper design temperature limits during ascent, reentry into Earth’s atmosphere, and / or transit throughan atmosphere. For example, as described in greater detail below, a composite propellant tank configured in accordance with embodiments of the present disclosure can include a plurality of structural, load-carrying members arranged in an external framework on an outer surface of the tank. In some embodiments, the tank can also include an internal structure (e.g., an internal truss structure) that is structurally attached to the external framework. The thermal protection systems described herein can include one or more layers of thermal insulation material that are positioned on the outer surface of the tank within the external framework, and exterior panels or outer skin that covers the insulation material and is structurally attached to the framework. This exterior panel can provide both thermal protection and structural support for the tank. Accordingly, a thermal protection system configured in accordance with embodiments of the present technology (hereinafter occasionally abbreviated as “TPS”) can also improve the mass ratio of an aerospace vehicle.

[0016] Unlike conventional TPSs that are generally non- structural reusable ceramics or limited use ablatives, embodiments of the TPSs described herein can include one or more layers of insulation material covered by a reusable - and structural — outer skin that can withstand the heat of reentry and insulate composite structures (e.g., composite tanks for holding cryogenic propellants). In some embodiments, this structural outer skin (which can also be referred to as an outer “panel,” “layer,” etc.) can be made from Ceramic Matrix Composites (CMCs) such as carbon fiber-reinforced silicon carbide (C / SiC), carbon fiber reinforced carbon (CFRC), carboncarbon (C / C), and / or reinforced carbon-carbon (RCC). Additional embodiments can include Silicon Carbide / Silicon Carbide (SiC / SiC) and / or Aluminum oxide / Aluminum oxide (AI2O3 / AI2O3). In some embodiments, a C / SiC external skin is strong enough to contribute to the structural performance of the vehicle and durable enough to resist damage. Although the C / SiC is based in a ceramic matrix (e.g., a SiC matrix), the inclusion of the carbon fiber to create the composite system allows the damage to be absorbed by the fibers rather than propagating cracks through the ceramic matrix as a purely ceramic tile would. The durability of this layer can allow re-flight of the TPS with reduced (e.g., minimal) requirements for inspection and maintenance.

[0017] In some embodiments, the one or more insulation layers directly below the structural outer skin can include light weight insulators (e.g., Opacified Fibrous Insulation (OFI) and / or inorganic aerogels, etc.) designed to withstand high external temperatures (e.g., temperatures up to 2,800-3,000 degrees F (1550-1650 degrees C)) while significantly reducing the temperature that the underlaying primary structure and propellant tank wall are exposed to during reentry. Insome embodiments, the layer of insulation closest to the tank wall can be made of a polymeric material (e.g., a polymeric aerogel) that can provide a flexible, hydrophobic layer that prevents or at least inhibits moisture, condensation, icing and / or cryopumping at or near the external surface of the tank, which are common adverse effects typically associated with conventional cryogenic tanks. In some embodiments, the one or more insulation layers can include a hybrid of inorganic and organic aerogel, which can provide sufficient insulation in fewer layers and / or in fewer quantities of types of layers in the overall TPS.

[0018] Certain details are set forth in the following description and in Figures 1A-6E to provide a thorough understanding of various embodiments of the present technology. In other instances, well-known structures, materials, operations and / or systems often associated with aerospace vehicle structures, propulsion systems, propellant systems, insulation systems, control systems, flight sequences, etc. are not shown or described in detail in the following disclosure to avoid unnecessarily obscuring the description of the various embodiments of the technology. Those of ordinary skill in the art will recognize, however, that the present technology can be practiced without one or more of the details set forth herein, or with other structures, methods, components, and so forth. The terminology used below is to be interpreted in its broadest reasonable manner, even though it is being used in conjunction with a detailed description of certain examples of embodiments of the technology. Indeed, certain terms may even be emphasized below; however, any terminology intended to be interpreted in any restricted manner will be overtly and specifically defined as such in this Detailed Description.

[0019] The accompanying Figures depict embodiments of the present technology and are not intended to be limiting of its scope. The sizes of various depicted elements are not necessarily drawn to scale, and these various elements may be arbitrarily enlarged to improve legibility. Component details may be abstracted in the Figures to exclude details such as position of components and certain precise connections between such components when such details are unnecessary for a complete understanding of how to make and use the present technology. Many of the details, dimensions, angles and / or other features shown in the Figures are merely illustrative of particular embodiments of the present technology. Accordingly, other embodiments can have other details, dimensions, angles and features without departing from the spirit or scope of the present disclosure. In addition, those of ordinary skill in the art will appreciate that further embodiments of the present technology can be practiced without several of the details described below. In general, identical reference numbers in the Figures identify identical, or at least generally similar, elements.A. Overview of Structures and Implementations

[0020] Figures 1 A and IB are partially schematic, top front and bottom front isometric views, respectively, of a vehicle 100 (e.g., an aerospace vehicle) having a thermal protection system configured in accordance with embodiments of the present technology. Referring to Figures 1 A and IB together, the vehicle 100 can be a Horizontal Takeoff, Horizontal Landing (HTHL) single-stage-to-orbit (SSTO) spaceplane having a pair of wings 104 (identified individually as a left wing 104a and a right wing 104b), which can be swept wings, and which can extend outwardly from a fuselage 102 to provide lift during flight in Earth’s atmosphere. The vehicle 100 can include any suitable aerodynamic control surfaces. For example, the trailing edge portion of each wing 104a, 104b can include an elevon for vehicle pitch and roll control. In some embodiments, the vehicle 100 can include one or more (e.g., a pair of) vertical stabilizers (identified individually as a left vertical stabilizer 110a and a right vertical stabilizer 110b) having corresponding rudders on trailing edge portions thereof for providing the vehicle 100 with yaw control. The fuselage 102 can include a crew cabin 112 in a forward portion thereof.

[0021] With reference to Figure 1A, in some embodiments, the fuselage 102 carries a propulsion system 108 having one or more rocket engines 114 to create thrust for the vehicle 100 (three rocket engines 114 are shown in Figure 1 A, but other embodiments can include more or fewer rocket engines). The rocket engines 114 can be configured to burn various types of propellants (e.g., liquid propellants) carried in one or more tanks attached to or carried within the vehicle 100. Suitable propellants can include, but are not limited to: fuels such as methane, hydrogen, or jet fuel (e.g., kerosene-types such as Jet-A); oxidizers such as oxygen, dinitrogen tetroxide, or hydrogen peroxide; hypergolic propellants; and / or other types of propellants that can be cryogenic (e.g., subcooled cryogenic propellants), petroleum-based, and / or otherwise suitable for rocket propulsion. In some embodiments, the engines 114 can be at least generally similar in structure and function to engines described in U.S. Patent No. 11,643,994, filed on December 23, 2020, and titled "ROCKET PROPULSION SYSTEMS AND ASSOCIATED METHODS," which is incorporated herein by reference in its entirety. The technology disclosed herein is not limited to any particular number or arrangement of engines or any particular type of engine or propellants. Accordingly, it will be understood that vehicles configured in accordance with the present technology can include more or fewer engines using other types of propellants.

[0022] In addition to the foregoing structures and systems, the vehicle 100 can include other structures and systems necessary or helpful for performing various missions, such as missions that include, for example, horizontal take-off from, and landing on, Earth, and deliveries of people and / or cargo to Low Earth Orbit (LEO). By way of example, such systems can include: a landing gear system; orbital maneuvering system (OMS) engines; thrusters positioned at various locations on the exterior of the vehicle 100 to provide attitude control while on orbit; a controller having one or more processors that can control various operations and functions of the vehicle 100 in accordance with computer-readable instructions stored on system memory or external systems; a communications system for, e.g., wireless communications (including crew communications, digital communications between processing devices, etc.) between the vehicle 100 and, e.g., ground control, ground stations, orbiting stations, etc.; an electrical power and distribution system; a navigation system; one or more flight control systems for affecting actuation of the vehicle control surfaces, engine throttles, landing gear, etc.; avionics; a hydraulic system for, e.g., control surface and landing gear actuation; an environmental control system for maintaining, e.g., air conditioning, etc. for human occupancy; etc.

[0023] In some embodiments, the vehicle 100 can take off from Earth with the assistance of a launch sled and, as a result, the landing gear can be retractable and stowable into associated gear bays during or after takeoff. When returning to Earth, the vehicle 100 can reenter Earth’s atmosphere in a gliding attitude with the landing gear deployed for landing in a manner that is at least generally similar to conventional commercial aircraft. For example, in some embodiments the vehicle 100 can be at least generally similar in structure and function to the vehicle described in U.S. Patent No. 11,059,608, filed on January 16, 2020, and titled “EARTH TO ORBIT TRANSPORTATION SYSTEM,” which is incorporated herein by reference in its entirety. Although several embodiments of the present technology are disclosed herein as including space vehicles, such embodiments are for example only, and embodiments of the present technology can include, and / or can be implemented in or on, other vehicles for which thermal protection is helpful or necessary, such as hypersonic vehicles for high speed travel within an atmosphere.

[0024] In some embodiments, the fuselage 102 and / or the wings 104a, 104b can include an external skin 106 (also identified as an upper surface external skin 106a and a lower surface external skin 106b) that is exposed to the environment outside of the vehicle during flight and other operations. In some embodiments, at least a portion of the external skin 106 can be composed of ceramic matrix composite (CMC) materials, such as C / SiC. For example, asdescribed in greater detail below, in some embodiments, at least a portion of the external skin 106a that forms the upper surface of the wings 104a, 104b, at least a portion of the external skin 106b that forms the lower surface of the wings 104a, 104b, and / or the lower surface of the fuselage 102, can include C / SiC material that can form an outer layer of the TPS of the vehicle 100. In some embodiments, other portions of the external skin 106 (e.g., portions that are exposed to relatively high temperatures during vehicle reentry) can also include C / SiC materials, while still other portions of the external skin 106 (e.g., portions that see relatively low temperatures during vehicle reentry) can include other composite materials, such as lightweight and durable carbon fiber-reinforced polymer materials (which can also be referred to as graphite- reinforced polymers, such as graphite in an epoxy matrix). In some embodiments, all or portions of the vehicle primary structures (including, for example, the internal primary structures of the wings 104a, 104b, the fuselage 102, the vertical stabilizers 110a, 110b, etc.) and / or other external surfaces, etc., can also include composite materials (e.g., carbon fiber-reinforced polymer materials). It will be understood, however, that in other embodiments, all or portions of the primary structures and / or external surfaces of the vehicle 100 can be at least partially or entirely constructed with other materials, including other composite materials (e.g., sandwich composite structures) and metals, such as aluminum, titanium, stainless steel, etc.

[0025] The foregoing structures and systems of the vehicle 100 are provided by way of example only and, unless otherwise specified herein, they are non-exclusive. Accordingly, it will be understood that some embodiments of the vehicle 100 and other vehicles configured in accordance with the present disclosure can include other structures and systems, while other embodiments may not include one or more of these structures and systems. Moreover, although Figures 1A and IB illustrate an example of an aerospace vehicle configured in accordance embodiments of the present technology, the present technology is not limited to use with a particular space vehicle configuration. Accordingly, in other embodiments, aerospace vehicles configured in accordance with the present technology and / or incorporating portions of the technology disclosed herein can have other configurations and / or can be constructed from other materials. Although the vehicle 100 is depicted by way of example as an HTHL SSTO spaceplane, the present technology is not limited to use with such vehicles and it is expected that the technology can be used on virtually any type of space vehicle without departing from the present disclosure. For example, vehicles configured in accordance with the present technology can include other shapes or configurations such as a stacked rocket without wings, and / or with fins instead of wings.

[0026] Figure 2A is a partially schematic top view of the vehicle 100. Figure 2B is a cross- sectional view taken substantially along line 2B-2B in Figure 2A to illustrate a portion of one of the wings (e.g., the wing 104a) having a propellant tank and a TPS configured in accordance with embodiments of the present technology. Referring first to Figure 2A, in some embodiments, the vehicle 100 can include a first propellant tank 210 positioned within a portion of the fuselage 102 and / or within the wings 104a, 104b. The vehicle 100 can further include a second propellant tank 212, which can be positioned generally within the fuselage 102 (e.g., above the first propellant tank 210). For example, some embodiments, the first propellant tank 210 can have a planform shape that at least generally conforms to at least a portion of the planform of the wings 104a, 104b and the fuselage 102, and the second propellant tank 212 can have a cylindrical shape with hemispherical end caps.

[0027] By way of example only, in some embodiments, both the first propellant tank 210 and the second propellant tank 212 can be cryogenic propellant tanks. For example, in some embodiments, the first propellant tank 210 can be configured to contain cryogenic oxidizer, such as liquid oxygen (LOX), and the second propellant tank 212 can be configured to contain cryogenic fuel, such as liquid methane. In some embodiments, one or both cryogenic propellants can be subcooled. For example, the LOX can be subcooled to a temperature of from -192 degrees C to -198 degrees C, to a temperature of about -196 degrees C, or to a temperature of -195.79 degrees C. Subcooling the LOX can reduce the vapor pressure and increase the density of the LOX. In some embodiments, reducing the vapor pressure in this manner enables the tank 210 (or other propellant tanks on the vehicle 100) to have a non-symmetrical, non-spherical or irregular shape that conforms to, or at least generally conforms to, an internal volume of the vehicle 100 (e.g., an internal volume of the wings 104a, 104b, the fuselage 102, etc.) without requiring that the tank 210 have a rounded, cylindrical, or other similar shape (e.g., a pressure- stabilized shape) as typically used for propellant tanks associated with conventional space vehicles. By way of example only, in some embodiments, the first propellant tank 210 can be configured to withstand a maximum internal design pressure of 20 psia, and the second propellant tank 212 can be configured to withstand a maximum internal design pressure of 65 psia. It will be appreciated that these design pressures, as well as the particular tank arrangement and / or shapes shown in Figure 2 A, are provided by way of example only, and in other embodiments, space vehicles configured in accordance with the present technology can include one or more tanks having other shapes, positions, arrangements, design pressures, etc., and can be configured to contain other types of rocket propellants including, for example, other cryogenicand / or non-cryogenic propellants. For example, in some embodiments, the first propellant tank 210 can be a fuel tank, and the second propellant tank 212 can be an oxidizer tank, and / or one or both of the tanks 210, 212 can carry other cryogenic and / or noncryogenic propellants, including Jet-A, refined petroleum (e.g., RP-1), and / or liquid hydrogen.

[0028] Figure 2B is an enlarged cross-sectional view taken substantially along line 2B-2B in Figure 2A and showing a portion of the propellant tank 210 contained within the wing 104a. For purposes of illustration, the view in Figure 2B shows a portion of the propellant tank 210 aft of a cut plane A-A (which may align with a plane transversing the vehicle 100 normal to an axis of the vehicle 100 extending forward and aftward, for example), but the basic features of the propellant tank 210 and the associated insulation system described herein can also apply to other portions of the propellant tank 210. As also noted above, subcooled propellants can be stored with reduced pressure in the tank, which enables some embodiments of the propellant tank 210 to have a shape that follows or conforms to interior regions of the vehicle 100 and can have, for example, relatively flat tank walls or skins) rather than a more traditional pressure-vessel design (e.g., the propellant tank 210 does not require cylindrical or spherical elements).

[0029] In some embodiments, the propellant tank 210 can include a tank skin 214 defining the internal volume of the propellant tank 210, and an optional internal truss structure 218 supporting the tank skin 214. The tank skin 214 can be formed with materials suitable for providing structural strength and withstanding the applied temperatures. For example, the tank skin 214 can be manufactured by laying-up and co-curing multiple plies of composite laminate material. Suitable composite materials can include carbon fiber reinforced polymers, such as a Bismaleimide resin with carbon fiber (e.g., an M65 resin combined with IM7 carbon fiber). The tank skin 214 can have a thickness between 0.02 inch and 0.18 inch, between 0.04 inch and 0.12 inch, between 0.06 inch and 0.10 inch, or about 0.086 inch, or other thicknesses. In some embodiments, the tank skin 214 can be designed for a maximum operating temperature of 300 degrees F. In other embodiments, the tank skin 214 can be designed for other maximum operating temperatures and can be formed from other composite materials and, in other embodiments, the propellant tanks 210, 212 can be formed with metallic materials including, for example, aluminum, stainless steel, etc.

[0030] In some embodiments, the truss structure 218 can include a framework of elongated truss members 218a that are bonded and / or fastened together to provide an internal structure for supporting the tank skin 214. For example, in some embodiments, the truss structure 218 canbe formed with composite materials, such as carbon fiber-reinforced silicon carbide (C / SiC). By way of example only, in some embodiments, components of the truss structure 218 can be manufactured via chemical vapor infiltration (CVI) densification of the C / SiC material using a hand layup and autoclave compaction and curing process. In other embodiments, components of the truss structure 218 can be manufactured using other suitable processes and / or other suitable materials. As generally shown in Figure 2B, the truss structure 218 can be mechanically attached to the tank skin 214, for example, via a plurality of attachment fittings 226 (e.g., clevis fittings) on the truss structure 218 that can be bolted, pinned, or otherwise mechanically fastened to corresponding attachment fittings 228 (e.g., lug fittings) that extend through openings in the skin 214 as described in greater detail below. In addition to the foregoing components, in some embodiments, the propellant tank 210 can further include one or more internal tank spars 224 that extend between an upper wall 214a and a lower wall 214b of the tank skin 214 and can include, in some embodiments, generally flat or planar structural members (e.g., with composite materials). In some embodiments, the aft portion of the propellant tank 210 (i.e. the end of the propellant tank 210 aft of the cut plane A- A) can have a height H. By way of example only, in some embodiments the height H can be between 24 inches and 120 inches, between 36 inches and 96 inches, between 60 inches and 84 inches, or about 71 inches.

[0031] The tank 210 can include a plurality of external structural members 216 that are affixed (e.g., adhesively bonded) to the outer surface of the tank skin 214 to provide additional structural support for the tank 210. In some embodiments, the structural members 216 are elongated members and extend spanwise (i.e., transverse to the longitudinal axis X of the vehicle 100, see Figure 2A) and may have an I-beam cross-sectional shape. In other embodiments, the structural members 216 can have various other cross-sectional shapes (e.g., C-section, L-section, hat section, etc.) and can be referred to as frames, ribs, beams, spars, etc. As described in greater detail below, in some embodiments, the tank 210 can further include additional structural members oriented longitudinally (i.e., parallel to the longitudinal axis X of the vehicle 100) that together with the structural members 216 form an external framework 220 of the tank 210, which can also be referred to as an exoskeleton 220 of the tank 210. Although a propellant tank is described herein as an example of a vehicle component protected by the TPS according to embodiments of the present technology, in further embodiments, TPSs disclosed herein can be positioned and / or configured to protect other vehicle components.

[0032] In some embodiments, the tank 210 can be thermally insulated by an arrangement of insulation materials 230 positioned between the structural members 216 and covered by theadjacent external skins 106a, 106b. The upper surface external skin 106a and the lower surface external skin 106b, can be structurally attached to the corresponding adjacent structural members 216. By covering the insulation materials 230 with the external skins 106a, 106b (which can be formed with CMC materials such as panels of C / Sic material, as discussed above) and structurally attaching the external skins 106a, 106b to the structural members 216, the external skins 106a, 106b provide thermal protection for the tank 210 and additional structural support for the tank 210. Further details regarding the arrangement of structural members 216, the insulation materials 230, and the external skins 106a, 106b are described below with reference to Figures 3 A-3D.

[0033] Figures 3A-3C are a series of partially exploded isometric views of the portion of the wing 104a and the propellant tank 210 shown in Figure 2B. Figure 3D is an assembled isometric view of the portion of the wing 104a and the propellant tank 210 shown in Figure 2B. Referring first to Figure 3 A, in some embodiments, the structural members 216 can be oriented spanwise. In some embodiments, in addition to the spanwise structural members 216 configured to be attached to the upper wall 214a and the lower wall 214b of the tank skin 214, the exoskeleton 220 can further include a plurality of longitudinally-oriented structural members 317. In some embodiments, the longitudinally-oriented structural members 317 can be attached to at least the upper wall 214a of the tank skin 214. The longitudinally-oriented structural members 317 can be at least generally similar in structure and function to the spanwise structural members 216, and can be structurally attached to the spanwise structural members 216 (e.g., at end portions thereof) using mechanical fasteners and associated fittings, adhesive bonding, etc. In some embodiments, some or all of the structural members 216, 317 can be made from composite materials, such as carbon fiber-reinforced polymer (e.g., epoxy) materials, or CVI C / SiC materials in, for example, an I-beam configuration. In other embodiments, some or all of the structural members 216, 317 can be formed from other types of composite materials and / or suitable metallic materials. In some embodiments, the longitudinally-oriented structural members 317 can alternatively, or additionally, be implemented on the lower end of the structure, e.g., attached to the lower wall 214b of the tank skin 214. In some embodiments, the framework 220 comprises a grid of the structural members 216, 317.

[0034] In some embodiments, one or more (e.g., each) of the spanwise structural members 216 includes a plurality of the spaced apart fittings 228 (e.g., lug fittings) that extend inwardly through corresponding openings 322 in the tank skin 214. As noted above, the fittings 228 enable the spanwise structural members 216 to be mechanically fastened to the correspondingfittings 226 on the internal truss structure 218. For example, in some embodiments, the spanwise structural members 216 can be bonded to the tank skin 214 with the corresponding fittings 228 extending through the openings 322, and then the truss structure 218 can be attached to structural members (e.g., the spanwise structural members 216) via the fittings 226 as further illustrated in Figure 3B.

[0035] In Figure 3B, the exoskeleton 220 has been bonded to the exterior surfaces of the tank skin 214 and attached to the internal truss structure 218 via connections between the attachment fittings 228, 226 (see also, Figure 2B). In some embodiments, an aft wall 214c of the tank 210 can be positioned against and / or attached to (e.g., bonded to) a rear spar 320, which may be a composite wing spar.

[0036] Turning to Figure 3C, after the exoskeleton 220 has been attached to the external surfaces of the tank skin 214, the insulation materials 230 described above with reference to Figure 2B can be installed, for example, within the exoskeleton 220 between the adjacent structural members 216, 317. Next, with additional reference to Figure 3D, one or more panels forming the upper surface external skin 106a and the lower surface external skin 106b can be attached to the exoskeleton 220. In some embodiments, the external skin 106a, 106b can be mechanically attached to the structural members 216, 317 of the exoskeleton 220 via suitable fasteners (e.g., flush head / countersunk fasteners such as flush mounted countersunk C / SiC screws, rivets, screws, bolts, etc.). Additionally or alternatively, in some embodiments, the external skin 106a, 106b can be bonded to the structural members 216 and 317 with one or more adhesive materials. As described in additional detail below, external skin 106a, 106b (which can include, for example, C / SiC panels) in combination with the insulation materials 230 can form at least part of a TPS 340 that provides both thermal protection and structural support for the tank 210 and / or other internal structures of the vehicle 100.B. Further Details of Thermal Protection Systems

[0037] Figure 4A is an exploded isometric view of the TPS 340 configured in accordance with embodiments of the present technology. Figure 4B is an assembled isometric view of the TPS 340 shown in Figure 4A. Figure 4C is a cross-sectional view of the TPS 340 shown in Figure 4A. In some embodiments, the insulation materials 230 can be seated within cells 402 of the exoskeleton 220. Although Figures 4A-4C show an arrangement of the insulation materials 230 in a single cell 402 of the exoskeleton 220 for purposes of illustration, it will be understood that (as shown in, e.g., Figures 2B and 3C) in most embodiments all, or at least a portion, of theadjacent cells 402 of the exoskeleton 220 can also contain the same arrangement, or at least a generally similar arrangement, of insulation materials 230. In other embodiments, adjacent cells 402 can contain different arrangements of insulation materials 230.

[0038] With reference to Figures 4A and 4B, in some embodiments, the insulation materials 230 can include one or more layers of insulation material configured to achieve insulative properties corresponding to design requirements. For example, in some embodiments, a first layer 432 of insulation material can be positioned directly against the exterior surface of the tank skin 214 between the adjacent structural members 216, 317 of the exoskeleton 220. By way of non-limiting example, the first layer 432 of insulation can include a polymeric aerogel that provides a flexible hydrophobic layer that inhibits (e.g., prevents) moisture, condensation, icing and / or cryopumping from occurring at or near the exterior surface of the tank skin 214. In some embodiments, the first layer 432 of insulation can include a polymeric aerogel developed by NASA Glenn Research Center (GRC) and / or provided by Aerogel Technologies, LLC, of Boston, MA, 02136 USA. In other embodiments, other types of insulation materials can be used for the first layer 432, or this layer can be omitted. In some embodiments, the insulation (e.g., the first layer 432) may be spaced apart from the exterior surface of the tank skin 214.

[0039] In some embodiments, the insulation materials 230 can further include a second layer 434 of insulation material. In some embodiments, the second layer 434 of insulation material can include an inorganic aerogel such as an aluminosilicate aerogel configured to withstand high temperatures. Such a material may be developed by NASA GRC and / or commercially available from Aerogel Technologies, LLC, or by HeetShield, Inc. of Flagstaff, AZ 86001 USA. In some embodiments, the layers 432 and 434 and be combined into a single layer including a hybrid inorganic and organic aerogel material, which can simplify construction and reduce dusting as compared to an inorganic layer alone.

[0040] In some embodiments, the insulation materials 230 can further include a third layer 436 of insulation material that is positioned above the second layer 434. In some embodiments, the third layer 436 can include high temperature opacified fibrous insulation (OFI), which may be commercially available from HeetShield, Inc., or other suitable or similarly configured materials. In some embodiments, the second layer 434 and the third layer 436, which can be positioned directly below the external skin 106a (or above the external skin 106b for the lower portion of the vehicle) can include lightweight insulation materials configured to withstand high temperatures and insulate the underlying / adjacent materials. The insulative properties of thevarious layers enable composite materials such as carbon fiber reinforced polymers to be used for the tank skin 214 and structural systems of the vehicle. By way of comparison only, in some embodiments, one or more of the layers 432, 434, and 436 of insulation material can have lower thermal conductivities than traditional ceramic tiles (e.g., thermal conductivities that are about 30-50% less than ceramic tiles), so the overall insulative structure can be thinner than that of other TPSs, such as insulative structures formed with traditional ceramic tiles.

[0041] Once the insulation materials 230 have been positioned between the adjacent structural members 216, 317 of the exoskeleton 220, the external skin 106 can be attached (e.g., mechanically fastened and / or adhesively bonded) to the structural members 216, 317 by, for example, a plurality of fasteners 442 (e.g., flush head / countersunk rivets, screws, etc.). For example, in some embodiments in which the structural members 216, 317 have I-beam cross sectional shapes, the fasteners 442 can extend through the external skin 106 and into corresponding caps or flanges 438a, 438b of the structural members 216 and 317, respectively.

[0042] As noted above, in some embodiments, the external skin 106 can include a CMC material, such as one or more panels or layers of C / SiC material. By way of example, a C / SiC panel can be formed by hand layup and CVI matrix densification of fabric material (e.g., woven carbon fiber fabric, such as T300 woven carbon fiber) that undergoes an autoclave compaction and curing process. In some embodiments, the C / SiC panel can be formed by a plurality of fabric plies (e.g., 24-plies) laid up in, e.g., 0° + / - 60° ply orientations, and the panel can have a thickness of from 0.04 inch to 0.50 inch, from 0.06 inch to 0.40 inch, from 0.08 inch to 0.30 inch, from 0.10 inch to 0.25 inch, from 0.10 to 0.20 inch, or 0.15 inch. In some embodiments, the C / SiC materials forming the external skin 106 can be sourced from General Electric Edison Works of Cincinnati, OH, U.S.A. In some embodiments, the external skin 106 can be configured to withstand a maximum temperature of from 2800-3000 degrees F (1550-1650 degrees C). In other embodiments, the external skin 106 can be formed from other types of materials and / or using other processes, including other CMC materials, other composite materials, metallic materials, etc., and / or the external skin 106 can be configured to withstand other maximum temperatures, including temperatures above or below 3000 degrees F.

[0043] In some embodiments, the external skin 106 can be formed with multiple adjacent skin panels 440, an example of one being shown in Figure 4A (other panels would be repeated adjacent to the illustrated panel 440). For purposes of illustration, the panel 440 of external skin 106 is depicted in Figures 4A and 4B as being approximately the size of a single cell 402 of theexoskeleton 220. However, the external skin 106 can be provided in larger panels that cover greater areas of the exoskeleton 220 (and hence the vehicle 100). The size of such panels can be dictated by various factors including manufacturability, cost, etc. In some embodiments, the upper surface external skin 106a can include the same material, composition, and / or construction as the lower surface external skin 106. In other embodiments, the upper surface external skin 106a can include different materials, composition, and / or construction as the lower surface external skin 106.

[0044] Figure 4C is a partially-schematic cross-sectional view of the TPS 340 configured in accordance with embodiments of the present technology. In some embodiments, the structural members 216,317 can have a height Hl (e.g., from a first flange to the opposite flange) between 1 inch and 10 inches, between 2 inches and 6 inches, between 3 inches and 5 inches, or about 4 inches. The first layer 432 of insulation can have a thickness T1 of from 0.10 inch to 2 inches, from 0.20 inch to 1 inch, or about 0.50 inch. The second layer 434 of insulation can have a thickness T2 of from 0.50 inch to 6 inches, from 1 inch to 4 inches, or about 2 inches. The third layer 436 of insulation can have a thickness T3 of from 0.50 inch to 5 inches, from 1 inch to 3 inches, or about 1.5 inches. The foregoing dimensions are for example only. Other embodiments of the present technology can include insulative materials and / or layers of insulative materials having different dimensions or compositions. Moreover, in other embodiments, one or more of the layers of insulation material can be omitted and / or replaced with a different type of material. Accordingly, the insulation materials 230 can include more or fewer layers of insulation and / or other varieties, arrangements, and / or compositions than those shown in the Figures or described herein. In some embodiments, one or more of the layers of insulation may be bonded together or otherwise attached. In yet other embodiments, one or more of the layers can be enclosed within a sheet material, such as, for example, a flexible mylar sheet, a metallic foil sheet, etc. In further embodiments, instead of multiple layers of insulated materials, the TPS 340 can include a single layer of insulation material with a nonhomogeneous composition through its thickness. In general, embodiments of the present technology can include various other arrangements of insulation materials between the outer skin 106 and the tank skin 214.

[0045] As described above, the vehicle 100 (see Figures 1A, IB) can be constructed largely of lightweight composite materials that react flight loads (e.g., propulsive and aerodynamic loads), hold cryogenic fluids (in the case of the propellant tanks), and withstand reentry temperatures. In general, however, some of these composite materials (e.g., the composite materials forming the tank skin 214) typically have a relatively low upper service temperaturelimit (e.g., an upper service temperature limit of 300 degrees F or less) as compared to metals. In some embodiments, the TPS 340 described herein can provide a highly efficient, passive, low- maintenance and reusable thermal protection system that can ensure that these composite materials do not exceed their upper service temperature limit from aerodynamic heating during reentry of the vehicle 100 into Earth’s atmosphere and / or during hypersonic flight through an atmosphere. Additionally, the TPS 340 can maintain an external temperature of the propellant tanks low enough to prevent or at least reduce the buildup of condensate and ice, air liquefaction, and cryopumping. Other advantages of some embodiments of the TPS 340 can include: the durability of the CMC outer layer (e.g., the external skin 106) can enable reflight of the vehicle with minimal or at least reduced inspection and maintenance; the relatively low maintenance and reusability of the design can reduce the time required for vehicle refurbishment between flights; the design can provide flexibility for the use of lightweight low service temperature composite structures for flight and reentry systems, especially those intended for reuse; the design may be lighter (e.g., 15-30% lighter) than conventional thermal protections system (e.g.., ceramic tiles and metallic systems); and the design provides both thermal protection and structural strength, unlike traditional thermal protection systems which are generally non- structural reusable ceramics or limited use ablatives.

[0046] In some embodiments, the TPS 340 can form the exterior surfaces of the vehicle 100 that surround and / or cover the propellant tanks 210 and 212, as well as other portions of the vehicle 100 that experience relatively high heating during vehicle ascent, reentry, and / or travel through an atmosphere, such as the windward surfaces of the vehicle 100 (e.g., the wing leading edges, the underside and nose of the fuselage, etc.). In some embodiments, the TPS 340 can cover all, or at least a portion, of the first propellant tank 210 and / or all, or at least a portion of, the second propellant tank 212. Additionally, the arrangement and / or thicknesses of the insulating materials (e.g., the materials forming the layers 432, 434, 436 and / or the upper and lower external skins 106a, 106b) may vary among different areas of the vehicle 100, depending on location, heating profile, and / or other factors. In some embodiments, some portions of the vehicle 100 (e.g., those that experience relatively lower temperatures) may not include all of the insulating materials shown in Figure 4A or disclosed elsewhere herein. Accordingly, the thicknesses, stack-up, and / or arrangement of insulation materials may vary among different areas of the vehicle 100. Moreover, depending on weight allocations and / or other considerations, some portions of the vehicle 100 may include alternate TPS solutions.c. Further Embodiments of Thermal Protection Systems

[0047] Figure 5 is an isometric view of a portion of a propellant tank 510 having a TPS 540 configured in accordance with embodiments of the present technology. In some embodiments, the propellant tank 510 can be the same as, or at least generally similar in structure and function to, the propellant tank 210 and can be used in place of the propellant tank 210 on or in the aerospace vehicle 100. For example, the propellant tank 510 can include a tank skin 514 that can be formed from composite materials, such as carbon fiber reinforced polymers (e.g., an M65 resin combined with IM7 carbon fiber) or other composite materials disclosed herein. In some embodiments, the tank 510 can include an internal truss structure that is at least generally similar in structure and function to the internal truss structure 218 described above with reference to, e.g., Figure 2B. Like the internal truss structure 218 described above, such an internal truss structure in the propellant tank 510 would be positioned within the propellant tank 510 to support the tank skin 514, and it can include a framework of truss members that are bonded and / or fastened together. In some embodiments, the internal truss structure can be formed from composite materials. However, as described below in further detail, instead of connecting the internal truss structure to external structure through openings in the tank skin 514, the external structure may be attached to the external surface of the tank skin 514 without structural features penetrating the tank skin 514.

[0048] In some embodiments, the propellant tank 510 further includes a plurality of first structural members 516 that can be affixed (e.g., adhesively bonded) to the outer surface of the tank skin 514 and extend in, e.g., a spanwise direction (i.e., transverse to the longitudinal axis X of the vehicle 100). The propellant tank 510 can further include a plurality of second structural members 517 that can be affixed (e.g., adhesively bonded) to the outer surface of the tank skin 514 and extend in, e.g., a longitudinal direction. In some embodiments, some or all of the first structural members 516 and / or the second structural members 517 can each include multiple component pieces. For example, as described in greater detail below, in some embodiments, multi -piece first structural members 516 may be continuous, or at least generally continuous members along a spanwise direction, and multi-piece second structural members 517 can be segmented to extend e.g., longitudinally, between the multi-piece first structural members 516 such that end portions 517a of segments of the second structural members 517 can be structurally attached to the first structural members 516 via brackets, fittings, fasteners, adhesives, and / or other suitable connections to form an external framework attached to the tank skin 514.

[0049] In some embodiments, each of the second structural members 517 can include an elongated first structural element 518 and a pair of elongated second structural elements 519. The first structural element 518 can have an inverted “T” cross-sectional shape and each of the second structural elements 519 can have an inverted “L” cross-sectional shape. The second structural elements 519 can be arranged back-to-back with their vertical flanges 519a attached to an upstanding vertical flange 518a of the first structural element 518, which is positioned between the vertical flanges 519a of the second structural elements 519. In other embodiments, the first structural element 518 can have other cross-sectional shapes, such as “L” cross-sectional shapes.

[0050] In some embodiments, each of the first structural members 516 can be formed by attaching a pair of the second structural elements 519 to one of the first structural elements 518 in a similar manner or the same manner as described above with reference to the second structural members 517. The first structural element 518 and / or the second structural element 519 can be made from composite materials, such as carbon fiber-reinforced polymer (e.g., epoxy) materials and / or CVI C / SiC materials. The first structural elements 518 can be bonded (e.g., via adhesive and / or co-curing) to the outer surface of the tank skin 514, and / or they can be attached to the tank skin 514 using other methods of structural attachments (e.g., fasteners). In other embodiments, the first structural elements 518 and / or the second structural elements 519 can be manufactured from other types of composite materials and / or suitable metallic materials.

[0051] One or more external skin panels 506 can be attached to horizontal flanges 519b of the second structural elements 519 (via e.g., a plurality of mechanical fasteners and / or adhesive bonding). In some embodiments, the skin panels 506 can be the same as, or at least generally similar in structure and function to, the panels forming the external skin 106 (e.g., a skin panel 440) described above, e.g., they can be made of ceramic matrix composite (CMC) materials, such as C / SiC. Accordingly, in some embodiments the skin panels 506 can form at least a portion of the external surface of the wings 104a, b and the fuselage 102 of the vehicle 100 (Figure 1). Although each of the skin panels 506 shown in Figure 5 is depicted as having the same general size as one cell of the external frame structure formed by the first structural members 516 and the second structural members 516, in some embodiments, individual skin panels 506 can span multiple cells and the sizes of the skin panels 506 can vary depending on, for example, the strength of the panel across the unsupported span, the amount of thermal growth and / or shrinkage during use, the ability to maintain a workable gap size between the edges of adjacent panels, manufacturing considerations, and / or other factors. In some embodiments, eachof the first structural members 516 and / or the second structural members 517 can be formed with a series of segments that are structurally attached together in an end-to-end arrangement to form a continuous member that extends around an entire external surface of the tank 510 or other underlying structure.

[0052] Many features of the structural members 516,517 can be at least generally similar in structure and function to corresponding features of the structural members 216,317 described above with reference to, e.g., Figures 2B-4C. For example, the structural members 516, 517 can be connected to each other in a gridwork to form a structural exoskeleton that constitutes at least part of the primary structure of the tank 510 to carry the internal and external pressure forces applied to the tank 510 during operation of the vehicle 100. In one aspect of this embodiment, however, the first structural elements 518 can be bonded to the tank skin 514 without any lugs, fittings, or other features that penetrate through the tank skin 514 to be mechanically attached to the internal truss structure within the tank 510. Nor do any fasteners and / or other structures penetrate the tank skin 514 to structurally attach the external structural elements (i.e., the first structural elements 518) to the internal truss structure. Omitting the use of fasteners or any penetrations through the tank skin 514 reduces the likelihood of propellant leak paths through the tank skin 514. Additionally, the multiple-piece configurations of the structural members 516, 517 can limit or prevent a “thermal short” from the hot side of the TPS 340 (i.e., the external skin panels 506) to the cold side of the TPS 340 (i.e., the external surface of the tank skin 514). For example, the structural members 516,517 can include thermal insulators between the “hot” second structural elements 519 and the “cold” first structural elements 518.

[0053] Figures 6A and 6B are isometric and top views, respectively, of a portion of the TPS 540 described above with regard to Figure 5. Figure 6C is a cross-sectional view taken along the line 6C-6C in Figure 6B. Figures 6A and 6B show an isolated portion of the TPS 540 to illustrate the arrangement of parts in a single cell defined by the structural members 516, 517, however, adjacent cells of the TPS 540 can be constructed in a similar manner to form a continuous thermal protection system that covers (or at least substantially covers) the tank 510, as demonstrated in Figure 6C by the presence of two adjacent skin panels 506a, 506b and corresponding other components.

[0054] Referring to Figures 6A, 6B, and 6C, in some embodiments, the first structural elements 518 are configured as “L” shaped members having a single flange (518b in Figure 6C) bonded to the tank skin 514 and an upstanding flange (518a in Figure 6C) that is mechanicallyattached to the second structural element(s) 519. As shown on the right side of Figure 6C, in some embodiments, two of the second structural elements 519 can be mounted to the upstanding flange 518a of the first structural element 518 in a back-to-back arrangement so that the horizontal flanges 519b of the second structural elements 519 project away from each other and provide structural attachment areas for the corresponding skin panels 506a, 506b.

[0055] In some embodiments, because the second structural elements 519 will experience a hotter environment during vehicle operation than the first structural elements 518, the second structural elements 519 can be formed from, e.g., a metallic material configured to experience relatively high temperatures while maintaining, or at least substantially maintaining, its mechanical properties. For example, in some embodiments, the second structural elements 519 can be formed from Inconel (and can be referred to as “hot” brackets) while the first structural elements 518 can be formed from composite materials as noted above (and can be referred to as “cold” brackets). Some embodiments can include a resilient (e.g., compressible) elongate seal 652 positioned between the back-to-back second structural elements 519 to resist (e.g., prevent) hot gases from passing through the gaps between the skin panels 506a, 506b and flowing toward the “cold side” of the TPS 540.

[0056] With reference to Figure 6C, in some embodiments, the TPS 540 can further include one or more layers of insulation material positioned between the external skin panels 506a, 506b and the tank skin 514 to achieve insulation properties corresponding to design requirements. For example, the insulation materials can include a first layer 632 of insulation material positioned directly against the exterior surface of the tank skin 514 between the adjacent structural members 516, 517. For example, the first layer 632 of insulation material can be a polymeric aerogel (e.g., polymeric aerogel described above) that provides a flexible, hydrophobic layer that prevents, or at least inhibits moisture, condensation, icing and / or cryopumping from occurring at or near the exterior surface of the tank skin 514. In other embodiments, other types of insulation materials can be used for the first layer 632, or this layer can be omitted.

[0057] In some embodiments, the insulation materials can further include a second layer 634 of insulation material that can be an inorganic aerogel (e.g., inorganic aerogel described above). The insulation materials can further include a third layer 636 of insulation material that is positioned above the second layer 634. In some embodiments, the third layer 636 can include a high temperature OFI described above. In some embodiments, the second layer 634 and the third layer 636, which can be positioned directly beneath the external skin panels 506a, 506b (orabove external skin panels for the lower portion of the vehicle) can include lightweight insulation materials configured to withstand high temperatures and insulate the underlying / adjacent materials, such as composite materials used for the tank skin 514 and / or the primary structural systems of the vehicle 100. The insulation materials in the TPS 540 can be the same as, or at least generally similar in structure and function, to the insulation materials 230 described above with respect to the TPS 340.

[0058] Figures 6D and 6E are enlarged cross-sectional views showing details of portions of the TPS 540 shown in Figures 5-6C (see Figure 6C, showing the locations of the details shown in Figures 6D and 6E). In particular, Figure 6D illustrates an external skin panel 506a attached to the second structural element 519, and Figure 6E illustrates a first structural element 518 attached to a second structural element 519, in accordance with embodiments of the present technology. Referring first to Figure 6D, in some embodiments, the external skin panel 506a can be mechanically fastened to the second structural element 519 by a plurality of fasteners 642 (one such fastener is shown in Figure 6D). The head 642a of the fastener 642 can be seated against one or more (e.g., two) washers 656 which can in turn be seated against an annular spacer 654 positioned on the outer surface of the skin panel 506a. The shank 642b of the fastener 642 can extend through two additional spacers 654 positioned between the inner surface of the skin panel 506a and the horizontal flange 519b of the second structural element 519, and through a further spacer 654 and two additional washers 656 positioned on the “cold” side of the horizontal flange 519b of the second structural element 519. A nut 643 can be threadably engaged with the shank 642b of the fastener 642 and suitably tightened. In some embodiments, the washers 656 can be configured to allow for a limited amount of expansion (e.g., thermal expansion) and movement between the parts while maintaining a sufficient compression force to hold the corresponding parts together. For example, in some embodiments, the washers 656 can be Belleville washers or other suitable spring washers known in the art. In a further aspect of this embodiment, the spacers 654 can include low thermal conductivity materials, e.g., ceramic materials, that can at least partially insulate the second structural element 519 from the high temperatures experienced by the outer skin panel 506a during ascent, descent, and / or hypersonic travel of the vehicle 100. The spacers 654 can also be configured to allow for expansion and contraction of the tank 510 and / or underlying structures resulting from operational changes in internal / external pressure, temperature, etc.

[0059] In some embodiments, the fasteners 642 and the nuts 643 can include HI-LOK pins and collars available from Lisi Aerospace. However, other embodiments can include othersui table types of fasteners to structurally attach the outer skin panel 506a to the second structural element 519. For example, in some embodiments, the fasteners 642 can include a flush / countersunk head to reduce aerodynamic heating during vehicle operation and, in some embodiments, can include countersunk C / SiC screws or other types of rivets, screws, bolts, etc. In yet other embodiments, the external skin panels 506a can be structurally bonded or adhesively bonded to the second structural elements 519 in place of, or in addition to, mechanical fasteners.

[0060] With reference to Figure 6C, in some embodiments and / or in some portions of the TPS 540, two of the second structural elements 519 can be attached to the vertical flange 518a of the first structural element 518 in a back-to-back arrangement. With reference to Figure 6E, in some embodiments and / or portions of the TPS 540, only a single second structural element 519 is attached to the vertical flange 518a of the first structural element 518. In some embodiments, the vertical flange 519a of the second structural element 519 can be attached to the vertical flange 518a of the first structural element 518 via a plurality of fasteners 644 (one is shown in Figure 6E) that threadably engage corresponding nuts (or collars) 645. For each fastener, one or more annular spacers 654 can be positioned: (a) between the head 644a of each fastener 644 and the vertical flange 519a of the second structural element 519 (e.g., one spacer 654); (b) between the vertical flange 518a of the first structural element 518 and the vertical flange 519a of the second structural element 519 (e.g., two spacers 654); and / or (c) between the nut 645 and the vertical flange 518a of the first structural element 518 (e.g., one spacer). As described above, the spacer 654 can be made from low thermal conductivity materials (e.g., ceramics) that provide a thermal break between the “hot” second structural element 519 and the “cold” first structural element 518 and / or can allow for expansion and contraction of this joint as a result of changes in internal / external pressures, temperature gradients, etc. during vehicle operation. Although one second structural element 519 is shown in Figure 6E, an additional second structural element 519 may be included opposite the second structural element 519 already shown, and attached in a similar manner.

[0061] Embodiments of the TPS 540 can provide many advantages. For example, the first structural elements 518 can provide a structural exoskeleton that strengthens the tank 510 and helps carry the internal and external pressure loads during operation of the vehicle 100 while also functioning as part of the vehicle TPS. Although the first structural elements 518 are not mechanically attached to an internal truss structure of the tank 510 via penetrations and / or mechanical fasteners extending through the tank skin 514, the first structural elements 518 can nevertheless be bonded to the outer surface of the tank skin 514 directly adjacent to members ofthe internal truss structure, which themselves may be bonded to the inner surface of the tank skin 514 (e.g., in a similar manner as the bonding between the first structural elements 518 to the tank skin 514) so that the internal truss structure and the external first structural elements 518 together provide structural reinforcement for the tank skin 514 from opposing sides of the tank skin 514. A further advantage of some embodiments of the TPS 540 is that the spacers 654 can thermally insulate (or at least partially thermally insulate) the “hot” portions of the TPS 540 (e.g., the external skin panels 506 and the second structural elements 519) from the “cold” portions such as the first structural element 518, the tank skin 514, etc., which enables use of composite materials for the tank skin 514 and reduces the thermal and / or pressure stresses that can be induced between these components. These features, in combination with the insulation material (e.g., in the layers 632, 634, 636), can provide a robust and thermally efficient TPS for the vehicle 100.D. Additional Examples

[0062] Several aspects of the present technology are set forth in the following examples:1. An aerospace vehicle configured for hypersonic flight and / or reentry from space, the vehicle comprising: an external skin; a vehicle component positioned within an interior region defined by the external skin; a framework positioned between the vehicle component and the external skin, wherein the framework forms one or more cells; and one or more layers of insulation material positioned between the vehicle component and the external skin in the one or more cells.2. The aerospace vehicle of example 1 wherein the one or more layers of insulation material comprise a first layer of a first material, a second layer of a second material, and a third layer of a third material.3. The aerospace vehicle of example 2 wherein each of the first layer, the second layer, and the third layer comprises a different material than each of the other of the first layer, the second layer, and the third layer.4. The aerospace vehicle of example 2 wherein the first layer comprises polymeric aerogel, the second layer comprises inorganic aerogel, and the third layer comprises opacified fibrous insulation.5. The aerospace vehicle of example 4 wherein the inorganic aerogel includes aluminosilicate aerogel.6. The aerospace vehicle of any one of examples 2-5 wherein the first layer is positioned against the vehicle component.7. The aerospace vehicle of example 1 wherein the one or more layers comprises a first layer of a hybrid inorganic and organic aerogel, and a second layer of opacified fibrous aerogel.8. The aerospace vehicle of example 1 wherein the one or more layers comprises a single layer of a nonhomogeneous material comprising at least aerogel.9. The aerospace vehicle of any one of examples 1-8 wherein at least one layer of the one or more layers is enclosed within a metallic or polymer material.10. The aerospace vehicle of any one of examples 1-9 wherein the external skin is attached to the framework.11. The aerospace vehicle of any one of examples 1-10 wherein the external skin is further attached to the vehicle component via one or more fasteners and / or an adhesive material.12. The aerospace vehicle of any one of examples 1-11 wherein the vehicle component is a propellant tank.13. The aerospace vehicle of example 12 wherein the one or more layers of insulation material are positioned directly against an outer surface of the propellant tank.14. The aerospace vehicle of example 12 or example 13 wherein the propellant tank comprises a tank skin formed with composite material.15. The aerospace vehicle of any one of examples 12-14 wherein the propellant tank defines an internal volume, the propellant tank comprises an internal truss structure supporting the tank skin, and the internal truss structure comprises a plurality of elongated truss members.16. The aerospace vehicle of example 15 wherein the internal truss structure is attached to the tank skin.17. The aerospace vehicle of any one of examples 14-16 wherein the framework is attached to the tank skin via one or more adhesive materials.18. The aerospace vehicle of any one of examples 14-17 wherein the internal truss structure is attached to the framework via one or more attachment fittings, wherein at least one of the one or more attachment fittings extends through the tank skin.19. The aerospace vehicle of example 18 wherein the one or more attachment fittings comprise clevis fittings and lug fittings.20. The aerospace vehicle of any one of examples 14-17 wherein the framework and the internal truss structure are attached to the tank skin without structural connections extending through the tank skin to connect the internal truss structure to the framework.21. The aerospace vehicle of any one of examples 12-20 wherein the propellant tank is positioned in a wing of the aerospace vehicle, and wherein the wing comprises the external skin, the framework, and the one or more layers of insulation material.22. The aerospace vehicle of any one of examples 1-21 wherein the framework comprises a plurality of elongated structural members attached to the vehicle component.23. The aerospace vehicle of example 22 wherein the elongated structural members comprise one or more of the following: I-beam, C-beam, L-beam, hat-section beam.24. The aerospace vehicle of example 22 or example 23 wherein: the plurality of elongated structural members comprises (a) a plurality of first structural members oriented along a longitudinal direction relative to the aerospace vehicle, and (b) a plurality of second structural members oriented transversely to the first structural members; and each of the first structural members is attached to a second structural member of the plurality of second structural members.25. The aerospace vehicle of any one of examples 1-24 wherein the external skin comprises composite materials.26. The aerospace vehicle of any one of examples 1-25 wherein the external skin comprises a ceramic matrix composite material.27. The aerospace vehicle of example 26 wherein the ceramic matrix composite material comprises carbon fiber-reinforced silicon carbide.28. The aerospace vehicle of any one of examples 1-27 wherein the external skin comprises a plurality of panels of a composite material, and wherein each panel covers one or more cells of the framework.29. The aerospace vehicle of example 28 wherein the plurality of panels are positioned on an upper surface of a wing of the vehicle and on a lower surface of the wing of the vehicle.30. The aerospace vehicle of any one of examples 1-29 wherein the aerospace vehicle is a spaceplane configured to travel from Earth to an orbit around Earth via a single-stage launch from Earth.31. The aerospace vehicle of any one of examples 1-30 wherein the framework comprises: a plurality of first structural members affixed to an outer surface of the vehicle component; anda plurality of second structural members affixed to the outer surface of the vehicle component, wherein each second structural member extends extending transversely relative to each first structural member.32. The aerospace vehicle of example 31 wherein one or more end portions of the second structural members are structurally attached to one or more of the first structural members.33. The aerospace vehicle of example 31 or example 32 wherein, for each of first structural member of the plurality of first structural members: the first structural member comprises a first structural element and two second structural elements; the first structural element comprises a T-shaped cross-section or an L-shaped crosssection having a first flange extending transversely from a first base; each of the two second structural elements comprises an L-shaped cross-section having a second flange extending transversely from a second base; and the second flanges are attached to opposing sides of the first flange.34. The aerospace vehicle of any one of examples 31-33 wherein, for each second structural member of the plurality of second structural members: the second structural member comprises a first structural element and two second structural elements; the first structural element comprises a T-shaped cross-section or an L-shaped crosssection having a first flange extending transversely from a first base; each of the two second structural elements comprises an L-shaped cross-section having a second flange extending transversely from a second base; and the second flanges are attached to opposing sides of the first flange.35. The aerospace vehicle of example 33 or example 34, further comprising a resilient seal positioned between each of the second flanges and the first flange.36. The aerospace vehicle of example 31 or example 32 wherein: each of the first structural members and the second structural members comprises at least a first structural element and two second structural elements; the second structural elements are attached to the external skin; and the first structural elements are attached to the outer surface of the vehicle component.37. The aerospace vehicle of any one of examples 33-36 wherein: the first structural elements comprise composite materials; and the second structural elements comprise a metallic material.38. The aerospace vehicle of any one of examples 33-37, further comprising one or more spacers between the second structural elements and the external skin, wherein the one or more spacers comprise a ceramic material.39. The aerospace vehicle of any one of examples 31-38 wherein each of the first structural members and the second structural members comprises composite materials.40. The aerospace vehicle of example 39 wherein the first structural members and the second structural members comprise carbon fiber-reinforced polymer material or carbon fiber-reinforced silicon carbide.41. The aerospace vehicle of any one of examples 31-40 wherein at least part of the external skin is attached to the first structural members and the second structural members.42. An aerospace vehicle configured for hypersonic flight and / or reentry from space, the vehicle comprising: a propellant tank having a tank wall; an arrangement of elongated structural members attached to an outer surface of the tank wall, wherein the arrangement of elongated structural members forms a framework; a ceramic matrix composite (CMC) panel offset from the tank wall and structurally attached to the arrangement of elongated structural members; andone or more layers of insulation material positioned adjacent to the outer surface of the tank wall between the tank wall and the CMC panel.43. The aerospace vehicle of example 42 wherein the propellant tank is a composite propellant tank.44. The aerospace vehicle of example 42 or example 43, further comprising an internal arrangement of structural members positioned within the propellant tank.45. The aerospace vehicle of example 44 wherein the internal arrangement of structural members is structurally attached to the framework.46. The aerospace vehicle of example 44 wherein the internal arrangement of structural members and the framework are connected via the tank wall, without a connection between the internal arrangement of structural members and the framework extending through the tank wall.47. The aerospace vehicle of any one of examples 42-46, further comprising one or more rocket engines connected to the propellant tank and configured to use propellant from the propellant tank to create thrust for the aerospace vehicle.48. The aerospace vehicle of any one of examples 42-47, further comprising: a fuselage; a first wing extending outwardly from the fuselage; a second wing extending outwardly from the fuselage, opposite the first wing; and a plurality of aerodynamic control surfaces.49. The aerospace vehicle of any one of examples 42-48, further comprising two vertical stabilizers.50. The aerospace vehicle of any one of examples 42-49, further comprising an external skin, wherein the external skin comprises the CMC panel.51. A thermal protection system for an aerospace vehicle, wherein the aerospace vehicle is configured for hypersonic flight and / or reentry from space, the thermal protection system comprising: a framework comprising one or more cells; a plurality of layers of insulation material positioned in each cell of the one or more cells, wherein the layers of insulation material comprise at least one of the following: (a) a layer of polymeric aerogel, (b) a layer of inorganic aerogel, or (c) fibrous insulation; and a plurality of skin panels, wherein each skin panel is positioned over at least one cell of the one or more cells, and wherein each skin panel comprises ceramic matrix composite material.52. A method of making a thermal protection system for a hypersonic or space, the method comprising: forming a plurality of elongated structural members, wherein forming the structural members comprises forming one or more of the structural members using one or more composite materials; arranging the structural members to form a framework of first elongated structural members of the plurality of elongated structural members and second elongated structural members of the plurality of elongated structural members, wherein the first elongated structural members are oriented transversely relative to the second elongated structural members, and wherein the framework comprises a plurality of cells between the structural members; positioning one or more layers of insulation material in each cell of plurality of cells, wherein the one or more layers of insulation material comprise at least one of the following: (a) a layer of polymeric aerogel, (b) a layer of inorganic aerogel, or (c) fibrous insulation; positioning one or more composite panels at least one of the one or more cells, wherein the one or more composite panels are configured to form an exterior surface of the vehicle; and attaching the one or more composite panels to the framework.53. The method of example 52, further comprising attaching the framework to a propellant tank of the vehicle.54. An aerospace vehicle, comprising: a propellant tank having a tank wall; an external arrangement of structural elements attached to an outer surface of the tank wall; a ceramic matrix composite (CMC) panel offset from the tank wall and structurally attached to the external arrangement of structural elements; and one or more layers of insulation material positioned on the outer surface of the tank wall between the tank wall and the CMC panel.55. The aerospace vehicle of example 54 wherein the propellant tank is a composite propellant tank.56. The aerospace vehicle of example 54 or example 55, further comprising an internal arrangement of structural members positioned within the propellant tank and structurally attached to the external arrangement of structural elements.57. The aerospace vehicle or method of any one of examples 54-56, wherein the aerospace vehicle is a hypersonic vehicle for transiting an atmosphere or a space vehicle configured to ascend to space and / or descend from space.58. An aerospace vehicle having a thermal protection system configured according to any one or more of examples 1-57.59. A method for insulating an aerospace vehicle as described herein.E. Conclusion

[0063] As used herein, the term "and / or" when used in the phrase "A and / or B" means "A, or B, or both A and B." A similar manner of interpretation applies to the term "and / or" when used in a list of more than two terms. The word "or" in reference to a list of two or more items, covers all of the following interpretations of the word: any of the items in the list, all of the items in thelist, and any combination of the items in the list. As used herein, the terms “about,” “approximately,” and the like refer to values within 10% of the stated value. As used herein, the terms “connected,” “coupled,” “attached,” or any variant thereof means any connection, coupling, or attachment, either direct or indirect, between two or more elements. Additionally, the words “herein,” “above,” "below," and words of similar import, when used in this application, refer to this application as a whole and not to any particular portions of this application. Where the context permits, singular or plural terms may also include the plural or singular term, respectively. Numerical adjectives including “first” and “second,” or the like, as used in the present disclosure, do not convey hierarchy or specific features or functions. Rather, such numerical adjectives are intended to aid the reader in distinguishing between elements which may have similar nomenclature, but which may differ in position, orientation, or structure. Accordingly, such numerical adjectives may be used differently in the claims. To the extent any materials incorporated herein by reference conflict with the present disclosure, the present disclosure controls.

[0064] From the foregoing, it will be appreciated that some embodiments of the present technology have been described herein for purposes of illustration, but various modifications can be made without deviating from the disclosed technology. Methods according to embodiments of the present technology can be performed in other suitable orders different from those disclosed herein, and / or one or more steps of the methods can be performed simultaneously, or one or more steps can be omitted. Further, while advantages associated with some embodiments of the technology have been described in the context of those embodiments, some embodiments may also exhibit said advantages, and not all embodiments need necessarily exhibit such advantages to fall within the scope of the present technology. Accordingly, the present disclosure and associated technology can encompass other embodiments not expressly described or shown herein.

Claims

CLAIMSI / We claim:

1. An aerospace vehicle configured for hypersonic flight and / or reentry from space, the vehicle comprising: an external skin; a vehicle component positioned within an interior region defined by the external skin; a framework positioned between the vehicle component and the external skin, wherein the framework forms one or more cells; and one or more layers of insulation material positioned between the vehicle component and the external skin in the one or more cells.

2. The aerospace vehicle of claim 1 wherein the one or more layers of insulation material comprise a first layer of a first material, a second layer of a second material, and a third layer of a third material.

3. The aerospace vehicle of claim 2 wherein each of the first layer, the second layer, and the third layer comprises a different material than each of the other of the first layer, the second layer, and the third layer.

4. The aerospace vehicle of claim 2 wherein the first layer comprises polymeric aerogel, the second layer comprises inorganic aerogel, and the third layer comprises opacified fibrous insulation.

5. The aerospace vehicle of claim 4 wherein the inorganic aerogel includes aluminosilicate aerogel.

6. The aerospace vehicle of claim 2 wherein the first layer is positioned against the vehicle component.

7. The aerospace vehicle of claim 1 wherein the one or more layers comprises a first layer of a hybrid inorganic and organic aerogel, and a second layer of opacified fibrous aerogel.

8. The aerospace vehicle of claim 1 wherein the one or more layers comprises a single layer of a nonhomogeneous material comprising at least aerogel.

9. The aerospace vehicle of claim 1 wherein at least one layer of the one or more layers is enclosed within a metallic or polymer material.

10. The aerospace vehicle of claim 1 wherein the external skin is attached to the framework.

11. The aerospace vehicle of claim 10 wherein the external skin is further attached to the vehicle component via one or more fasteners and / or an adhesive material.

12. The aerospace vehicle of claim 1 wherein the vehicle component is a propellant tank.

13. The aerospace vehicle of claim 12 wherein the one or more layers of insulation material are positioned directly against an outer surface of the propellant tank.

14. The aerospace vehicle of claim 12 wherein the propellant tank comprises a tank skin formed with composite material.

15. The aerospace vehicle of claim 14 wherein the propellant tank defines an internal volume, the propellant tank comprises an internal truss structure supporting the tank skin, and the internal truss structure comprises a plurality of elongated truss members.

16. The aerospace vehicle of claim 15 wherein the internal truss structure is attached to the tank skin.

17. The aerospace vehicle of claim 16 wherein the framework is attached to the tank skin via one or more adhesive materials.

18. The aerospace vehicle of claim 15 wherein the internal truss structure is attached to the framework via one or more attachment fittings, wherein at least one of the one or more attachment fittings extends through the tank skin.

19. The aerospace vehicle of claim 18 wherein the one or more attachment fittings comprise clevis fittings and lug fittings.

20. The aerospace vehicle of claim 15 wherein the framework and the internal truss structure are attached to the tank skin without structural connections extending through the tank skin to connect the internal truss structure to the framework.

21. The aerospace vehicle of claim 12 wherein the propellant tank is positioned in a wing of the aerospace vehicle, and wherein the wing comprises the external skin, the framework, and the one or more layers of insulation material.

22. The aerospace vehicle of claim 1 wherein the framework comprises a plurality of elongated structural members attached to the vehicle component.

23. The aerospace vehicle of claim 22 wherein the elongated structural members comprise one or more of the following: I-beam, C-beam, L-beam, hat-section beam.

24. The aerospace vehicle of claim 22 wherein: the plurality of elongated structural members comprises (a) a plurality of first structural members oriented along a longitudinal direction relative to the aerospace vehicle, and (b) a plurality of second structural members oriented transversely to the first structural members; and each of the first structural members is attached to a second structural member of the plurality of second structural members.

25. The aerospace vehicle of claim 1 wherein the external skin comprises composite materials.

26. The aerospace vehicle of claim 1 wherein the external skin comprises a ceramic matrix composite material.

27. The aerospace vehicle of claim 26 wherein the ceramic matrix composite material comprises carbon fiber-reinforced silicon carbide.

28. The aerospace vehicle of claim 1 wherein the external skin comprises a plurality of panels of a composite material, and wherein each panel covers one or more cells of the framework.

29. The aerospace vehicle of claim 28 wherein the plurality of panels are positioned on an upper surface of a wing of the vehicle and on a lower surface of the wing of the vehicle.

30. The aerospace vehicle of claim 1 wherein the aerospace vehicle is a spaceplane configured to travel from Earth to an orbit around Earth via a single-stage launch from Earth.

31. The aerospace vehicle of claim 1 wherein the framework comprises: a plurality of first structural members affixed to an outer surface of the vehicle component; and a plurality of second structural members affixed to the outer surface of the vehicle component, wherein each second structural member extends extending transversely relative to each first structural member.

32. The aerospace vehicle of claim 31 wherein one or more end portions of the second structural members are structurally attached to one or more of the first structural members.

33. The aerospace vehicle of claim 31 wherein, for each of first structural member of the plurality of first structural members: the first structural member comprises a first structural element and two second structural elements; the first structural element comprises a T-shaped cross-section or an L-shaped crosssection having a first flange extending transversely from a first base;each of the two second structural elements comprises an L-shaped cross-section having a second flange extending transversely from a second base; and the second flanges are attached to opposing sides of the first flange.

34. The aerospace vehicle of claim 31 wherein, for each second structural member of the plurality of second structural members: the second structural member comprises a first structural element and two second structural elements; the first structural element comprises a T-shaped cross-section or an L-shaped crosssection having a first flange extending transversely from a first base; each of the two second structural elements comprises an L-shaped cross-section having a second flange extending transversely from a second base; and the second flanges are attached to opposing sides of the first flange.

35. The aerospace vehicle of claim 34, further comprising a resilient seal positioned between each of the second flanges and the first flange.

36. The aerospace vehicle of claim 31 wherein: each of the first structural members and the second structural members comprises at least a first structural element and two second structural elements; the second structural elements are attached to the external skin; and the first structural elements are attached to the outer surface of the vehicle component.

37. The aerospace vehicle of claim 36 wherein: the first structural elements comprise composite materials; and the second structural elements comprise a metallic material.

38. The aerospace vehicle of claim 36, further comprising one or more spacers between the second structural elements and the external skin, wherein the one or more spacers comprise a ceramic material.

39. The aerospace vehicle of claim 31 wherein each of the first structural members and the second structural members comprises composite materials.

40. The aerospace vehicle of claim 39 wherein the first structural members and the second structural members comprise carbon fiber-reinforced polymer material or carbon fiber- reinforced silicon carbide.

41. The aerospace vehicle of claim 31 wherein at least part of the external skin is attached to the first structural members and the second structural members.

42. An aerospace vehicle configured for hypersonic flight and / or reentry from space, the vehicle comprising: a propellant tank having a tank wall; an arrangement of elongated structural members attached to an outer surface of the tank wall, wherein the arrangement of elongated structural members forms a framework; a ceramic matrix composite (CMC) panel offset from the tank wall and structurally attached to the arrangement of elongated structural members; and one or more layers of insulation material positioned adjacent to the outer surface of the tank wall between the tank wall and the CMC panel.

43. The aerospace vehicle of claim 42 wherein the propellant tank is a composite propellant tank.

44. The aerospace vehicle of claim 43, further comprising an internal arrangement of structural members positioned within the propellant tank.

45. The aerospace vehicle of claim 44 wherein the internal arrangement of structural members is structurally attached to the framework.

46. The aerospace vehicle of claim 44 wherein the internal arrangement of structural members and the framework are connected via the tank wall, without a connection between the internal arrangement of structural members and the framework extending through the tank wall.

47. The aerospace vehicle of claim 42, further comprising one or more rocket engines connected to the propellant tank and configured to use propellant from the propellant tank to create thrust for the aerospace vehicle.

48. The aerospace vehicle of claim 42, further comprising: a fuselage; a first wing extending outwardly from the fuselage; a second wing extending outwardly from the fuselage, opposite the first wing; and a plurality of aerodynamic control surfaces.

49. The aerospace vehicle of claim 48, further comprising two vertical stabilizers.

50. The aerospace vehicle of claim 42, further comprising an external skin, wherein the external skin comprises the CMC panel.

51. A thermal protection system for an aerospace vehicle, wherein the aerospace vehicle is configured for hypersonic flight and / or reentry from space, the thermal protection system comprising: a framework comprising one or more cells; a plurality of layers of insulation material positioned in each cell of the one or more cells, wherein the layers of insulation material comprise at least one of the following: (a) a layer of polymeric aerogel, (b) a layer of inorganic aerogel, or (c) fibrous insulation; and a plurality of skin panels, wherein each skin panel is positioned over at least one cell of the one or more cells, and wherein each skin panel comprises ceramic matrix composite material.

52. A method of making a thermal protection system for a hypersonic or space, the method comprising: forming a plurality of elongated structural members, wherein forming the structural members comprises forming one or more of the structural members using one or more composite materials;arranging the structural members to form a framework of first elongated structural members of the plurality of elongated structural members and second elongated structural members of the plurality of elongated structural members, wherein the first elongated structural members are oriented transversely relative to the second elongated structural members, and wherein the framework comprises a plurality of cells between the structural members; positioning one or more layers of insulation material in each cell of plurality of cells, wherein the one or more layers of insulation material comprise at least one of the following: (a) a layer of polymeric aerogel, (b) a layer of inorganic aerogel, or (c) fibrous insulation; positioning one or more composite panels at least one of the one or more cells, wherein the one or more composite panels are configured to form an exterior surface of the vehicle; and attaching the one or more composite panels to the framework.

53. The method of claim 52, further comprising attaching the framework to a propellant tank of the vehicle.

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