阅读说明：本技术 用于冷却高速交通工具的前缘的系统和方法 (System and method for cooling leading edge of high speed vehicle ) 是由 尼古拉斯·威廉·拉塞 格利高里·亚历山大·纳特苏伊 布莱恩·马根·拉什 于 2020-11-13 设计创作，主要内容包括：高超音速飞行器包括一个或多个前缘组件,一个或多个前缘组件被设计成管理在高速或高超音速操作期间在前缘处经受的热负荷。具体地,前缘组件可以包括渐缩至前缘或停滞点的外壁。外壁可以限定汽化物室和在汽化物室内的毛细管结构,用于使工作流体以液体或汽化物形式循环以冷却前缘。另外,定位在汽化物室内的热能存储储存器容纳用于吸收热能的相变材料。(Hypersonic aircraft include one or more leading edge assemblies designed to manage the thermal loads experienced at the leading edge during high speed or hypersonic operation. In particular, the leading edge assembly may include an outer wall that tapers to a leading edge or stagnation point. The outer wall may define a vapor chamber and a capillary structure within the vapor chamber for circulating a working fluid in liquid or vapor form to cool the leading edge. In addition, a thermal energy storage reservoir positioned within the vapor chamber contains a phase change material for absorbing thermal energy.)

1. A leading edge assembly for a hypersonic vehicle, the leading edge assembly comprising:

an outer wall tapering to a leading edge, wherein the outer wall at least partially defines a vapor chamber;

a capillary structure positioned on an inner surface of the outer wall within the vapor chamber; and

a thermal energy storage assembly positioned in thermal communication with the vapor chamber.

2. The leading edge assembly of claim 1, wherein the thermal energy storage assembly comprises:

a reservoir wall; and

a phase change material positioned within the reservoir wall.

3. The leading edge assembly of claim 2, wherein the phase change material comprises silicon or beryllium.

4. The leading edge assembly of claim 2, wherein the phase change material has a melting temperature greater than 1000 degrees celsius.

5. The leading edge assembly of claim 2, wherein the phase change material has a latent heat of fusion greater than 300 kJ/kg.

6. The leading edge assembly of claim 1, wherein the thermal energy storage assembly comprises a first chamber containing a first phase change material having a first melting temperature and a second chamber containing a second phase change material having a second melting temperature.

7. The leading edge assembly of claim 2, wherein the capillary structure is a wick, porous structure, or screen lining between the inner surface of the outer wall and an outer surface of the reservoir wall.

8. The leading edge assembly of claim 7, wherein the capillary structure comprises:

at least one liquid bridge extending away from the reservoir wall towards the outer wall for providing a shorter path to the leading edge relative to the capillary structure.

9. The leading edge assembly of claim 2, wherein the outer wall, the reservoir wall, and the capillary structure are additively manufactured as a single unitary component.

10. The leading edge assembly of claim 2, wherein the reservoir wall is a compliant containment structure that is capable of expanding or contracting depending on a state of the phase change material.

Technical Field

The present subject matter relates generally to leading edge technology for use in high speed vehicles, such as hypersonic aircraft.

Background

High speed vehicles often encounter thermal management problems due to the high thermal loads experienced during high speed operation, especially at the leading edge where the free air stream impinges the vehicle. For example, in applications involving hypersonic aircraft, the leading edge may include the nose, hood, and leading edges of wings and stabilizers. Especially when such traffic is concernedWhen the tool is operated in the hypersonic velocity range (e.g., Mach 5 or above), the leading edge may be subjected to high thermal loads (e.g., 500-²). Non-moderate exposure to such thermal loads can result in component degradation and/or failure.

Improvements in materials and manufacturing techniques have enabled hypersonic aircraft to operate at higher speeds and temperatures. For example, materials have been developed to increase the temperature that a component can withstand while maintaining its structural integrity. In this regard, for example, a nickel-base superalloy may be used for 800 ℃, a single crystal material may be used for 1200 ℃, and a refractory metal may be required for higher temperatures. In addition, various cooling techniques have been developed to provide cooling for the leading edge of hypersonic vehicles. However, the corresponding advances in vehicle speed and high speed flight duration have created a need to further improve the cooling capability and high temperature durability of the leading edge of high speed vehicles.

Therefore, improvements to hypersonic aircraft and propulsion technology would be useful. More specifically, improvements in techniques and methods for cooling leading edges or leading edges of hypersonic vehicles would be particularly beneficial.

Disclosure of Invention

Aspects and advantages of the invention will be set forth in part in the following description, or may be obvious from the description, or may be learned through practice of the invention.

In one exemplary embodiment of the present disclosure, a leading edge assembly for a hypersonic vehicle is provided. The leading edge assembly includes: an outer wall tapering to a leading edge, wherein the outer wall at least partially defines a vapor chamber; a capillary structure positioned on an inner surface of the outer wall within the vapor chamber; and a thermal energy storage assembly positioned in thermal communication with the vapor chamber.

According to another exemplary embodiment, a leading edge assembly for a hypersonic vehicle is provided. The leading edge assembly includes: an outer wall tapering to a leading edge, wherein the outer wall at least partially defines a vapor chamber: a capillary structure positioned on an inner surface of the outer wall within the vapor chamber; and a thermal energy storage assembly positioned within the vapor chamber, the thermal energy storage assembly including a reservoir wall containing a phase change material.

These and other features, aspects, and advantages of the present invention will become better understood with reference to the following description and appended claims. The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate embodiments of the invention and together with the description, serve to explain the principles of the invention.

Drawings

A full and enabling disclosure of the present invention, including the best mode thereof, directed to one of ordinary skill in the art, is set forth in the specification, which makes reference to the appended figures.

FIG. 1 is a close-up cross-sectional schematic view of a leading edge of a hypersonic vehicle according to an exemplary embodiment of the present disclosure.

FIG. 2 is a close-up cross-sectional schematic view of a leading edge of a hypersonic vehicle according to another exemplary embodiment of the present disclosure.

FIG. 3 is a close-up cross-sectional schematic view of a leading edge of a hypersonic vehicle according to another exemplary embodiment of the present disclosure.

FIG. 4 is a close-up cross-sectional schematic view of a leading edge of a hypersonic vehicle according to another exemplary embodiment of the present disclosure.

FIG. 5 is a close-up cross-sectional schematic view of a leading edge of a hypersonic vehicle according to another exemplary embodiment of the present disclosure.

FIG. 6 is a close-up cross-sectional schematic view of a leading edge of a hypersonic vehicle according to another exemplary embodiment of the present disclosure.

Repeat use of reference characters in the present specification and drawings is intended to represent same or analogous features or elements of the invention.

Detailed Description

Reference now will be made in detail to embodiments of the invention, one or more examples of which are illustrated in the drawings. The detailed description uses numerical and letter designations to refer to features in the drawings. The same or similar reference numbers have been used in the drawings and the description to refer to the same or similar parts of the invention.

The word "exemplary" is used herein to mean "serving as an example, instance, or illustration. Any embodiment described herein as "exemplary" is not necessarily to be construed as preferred or advantageous over other embodiments. Furthermore, each example is provided by way of illustration of the invention, and not limitation of the invention. Indeed, it will be apparent to those skilled in the art that various modifications and variations can be made in the present invention without departing from the scope thereof. For instance, features illustrated or described as part of one embodiment, can be used with another embodiment to yield a still further embodiment. Thus, it is intended that the present invention cover the modifications and variations of this invention provided they come within the scope of the appended claims and their equivalents.

As used herein, the terms "first," "second," and "third" may be used interchangeably to distinguish one element from another, and are not intended to denote the position or importance of a single element. The singular forms "a," "an," and "the" include plural references unless the context clearly dictates otherwise. Unless otherwise specified herein, the terms "coupled," "secured," "attached," and the like, refer to both being directly coupled, secured, or attached, as well as indirectly coupled, secured, or attached through one or more intermediate components or features.

The terms "forward" and "aft" refer to relative positions within a gas turbine engine or vehicle, and to the normal operating attitude of the gas turbine engine or vehicle. For example, for a gas turbine engine, forward refers to a position closer to the engine inlet, and aft refers to a position closer to the engine nozzle or exhaust. The terms "upstream" and "downstream" refer to relative directions with respect to fluid flow in a fluid path. For example, "upstream" refers to the direction from which the fluid flows, and "downstream" refers to the direction to which the fluid flows.

Approximating language, as used herein throughout the specification and claims, is applied to modify any quantitative representation that could permissibly vary without resulting in a change in the basic function to which it is related. Accordingly, a value modified by a term or terms, such as "about", "approximately" and "substantially", are not to be limited to the precise value specified. In at least some instances, the approximating language may correspond to the precision of an instrument for measuring the value, or the precision of a method or machine for constructing or manufacturing the component and/or system. For example, approximate language may refer to being in the range of 10%.

Here and throughout the specification and claims, range limitations are combined and interchanged, such ranges are identified and include all the sub-ranges contained therein unless context or language indicates otherwise. For example, all ranges disclosed herein are inclusive of the endpoints, and the endpoints are independently combinable with each other.

In general, aspects of the present subject matter are directed to leading edge assemblies for high speed aircraft or vehicles, such as hypersonic aircraft. As used herein, the term "hypersonic" generally refers to air velocities of about mach 4 up to about mach 10 or more, e.g., above mach 5. However, it should be understood that aspects of the present subject matter are not limited to hypersonic flight, but may alternatively be applicable to applications involving other high speed vehicles, projectiles, objects, and the like. The description herein of a leading edge assembly for use on a hypersonic aircraft is merely an example intended to facilitate explanation of various aspects of the present subject matter. The present subject matter is not limited to such exemplary embodiments and applications.

Notably, as noted above, high speed vehicles, such as hypersonic aircraft, are typically subjected to extremely high temperatures and thermal gradients during high speed or hypersonic operation. Temperature gradients caused by high heat fluxes are often a more serious problem than the temperature itself. For example, the thermal conductivity of a structural material in combination with heat flux sets a temperature gradient within the material, and under high thermal loads, the gradient can cause mechanical stress, resulting in plastic deformation or cracking of the material. The thermal load of the structural material should be reduced to maintain the structural integrity of the component.

As mentioned above, the leading edge of such high speed vehicles is often subjected to the highest thermal loads. For example, a hypersonic vehicle may include a plurality of leading edge assemblies (e.g., generally identified herein by reference numeral 300) that are subject to high thermal loads during hypersonic flight. In this regard, the leading edge assembly 300 may be disposed on the forward end of an aircraft wing, nose cone, vertical stabilizer, a hood for a propulsive engine, or other leading edge or surface of a hypersonic aircraft. In accordance with exemplary embodiments of the present subject matter, the leading edge assembly 300 includes features for mitigating the effects of such thermal loading, for example, by carrying heat out of the area.

It is noted that it is generally desirable to make the leading edge assembly 300 as sharp or pointed as possible, for example, to reduce drag on hypersonic vehicles. However, referring now to FIG. 1, when leading edge assembly 300 is formed with a small radius of curvature, extremely high temperature and thermal gradients are experienced at the forward or leading edge of leading edge assembly 300, also referred to herein as stagnation lines, stagnation points 302, or similar terms. In this regard, when a hypersonic vehicle travels in air at a hypersonic speed, a free stream of air (e.g., identified herein by reference numeral 304) passes over and around the leading edge assembly 300, thereby generating a large thermal load. Aspects of the present subject matter are directed to thermal management techniques and features for cooling the leading edge assembly 300, particularly with emphasis on the region near the stagnation point 302 where the most severe thermal management issues typically arise.

It should be appreciated that the leading edge assembly 300 illustrated herein is a simplified cross-sectional view of an exemplary leading edge. The size, configuration, geometry, and application of such leading edge techniques may vary while remaining within the scope of the present subject matter. For example, the leading edge assembly 300 described herein defines a radius of between about 1mm and 3 mm. However, according to alternative embodiments, the leading edge assembly may have any other suitable radius.

Cooling techniques and thermal management features are described herein for cooling portions of one or more parts of a hypersonic aircraft subject to high temperatures and thermal gradients, such as the leading edge of a wing, nose, propulsion engine, or other part of the hypersonic aircraft. However, it should be understood that aspects of the present subject matter may be used to manage thermal loads, such as high temperatures and thermal gradients, within any component and in any suitable application. In this regard, for example, aspects of the present subject matter may be applied to any other hypersonic vehicle or any other technology or system having components exposed to high temperatures and/or large temperature gradients.

Additionally, although various techniques, component configurations, and systems for cooling a leading edge assembly 300 of a hypersonic vehicle are described herein, it should be understood that variations and modifications may be made to these techniques without departing from the scope of the present subject matter. Additionally, one or more of such techniques may be used in combination with one another to achieve improved cooling and thermal management. In this regard, although each cooling technique is described separately for the purpose of clearly describing how each technique operates, the described embodiments are merely examples intended for purposes of illustration and explanation and are not intended to limit the scope of the present subject matter in any way.

Additionally, according to exemplary embodiments of the present subject matter, some or all of the components described herein may be formed using an additive manufacturing process, such as a 3-D printing process. The use of such a process may allow certain components of the hypersonic vehicle, such as the leading edge assembly 300, to be integrally formed as a single unitary component, or as any suitable number of sub-components. As used herein, the term "additive manufacturing" or "additive manufacturing technique or process" generally refers to a manufacturing process in which successive layers of material are provided on top of each other to "build up" a three-dimensional part layer by layer. The continuous layers are typically fused together to form a unitary component that may have various integrated sub-components.

Although additive manufacturing techniques are described herein as enabling the manufacture of complex objects by building the object, typically point-by-point, layer-by-layer in a vertical direction, other manufacturing methods are possible and within the scope of the present subject matter. For example, although the discussion herein refers to adding material to form a continuous layer, one skilled in the art will appreciate that the methods and structures disclosed herein may be practiced with any additive manufacturing technique or fabrication technique. For example, embodiments of the invention may use additive layer processing, subtractive layer processing, or hybrid processing.

Suitable additive manufacturing techniques according to the present disclosure include, for example, Fused Deposition Modeling (FDM), Selective Laser Sintering (SLS), 3D printing such as by inkjet, laser jet, and binder jet, Stereolithography (SLA), Direct Selective Laser Sintering (DSLS), Electron Beam Sintering (EBS), Electron Beam Melting (EBM), laser engineered web formation (LENS), laser mesh fabrication (LNSM), Direct Metal Deposition (DMD), Digital Light Processing (DLP), Direct Selective Laser Melting (DSLM), Selective Laser Melting (SLM), Direct Metal Laser Melting (DMLM), and other known processes.

The additive manufacturing processes described herein may be used to form components using any suitable material. For example, the material may be metal, concrete, ceramic, epoxy, or any other suitable material in solid, liquid, powder, sheet material, wire, or any other suitable form or combination thereof. More specifically, in accordance with exemplary embodiments of the present subject matter, the additive manufactured components described herein may be formed partially, entirely, or in some combination of materials including, but not limited to, pure Metals, nickel alloys, chromium alloys, titanium alloys, magnesium alloys, aluminum alloys, and nickel or cobalt-based superalloys (e.g., available from Special Metals Corporation under the designation nickel or cobalt-based superalloy)The product of (1). These materials are examples of materials suitable for use in the additive manufacturing processes described herein, and may be generally referred to as "additive materials.

Additionally, the additive manufacturing processes disclosed herein allow a single component to be formed from multiple materials. Accordingly, the components described herein may be formed from any suitable mixture of the above materials. For example, a component may include multiple layers, segments, or parts that are formed using different materials, processes, and/or on different additive manufacturing machines. In this manner, components having different materials and material properties may be constructed for the needs of any particular application. Additionally, although the components described herein are constructed entirely from additive manufacturing processes, it should be understood that in alternative embodiments all or a portion of these components may be formed via casting, machining, and/or any other suitable manufacturing process. Indeed, any suitable combination of materials and manufacturing methods may be used to form these components.

Still referring to FIG. 1, a leading edge assembly 300 in accordance with exemplary embodiments of the present subject matter will be described in more detail. Specifically, fig. 1 provides a cross-sectional view of a leading edge assembly 300, which leading edge assembly 300 may be positioned at a leading edge (e.g., a forward end, an upstream end, etc.) of any component of a hypersonic aircraft. For example, the leading edge assembly 300 may be, for example, the leading edge of an air intake duct of a hypersonic propulsion engine, the leading edge of a turbine engine, the leading edge of a wing of an aircraft, a nose of an aircraft, the forward end of a vertical stabilizer, etc.

As explained herein, the leading edge assembly 300 may be subjected to large thermal loads during hypersonic flight operations. As used herein, the term "thermal load" or the like is generally intended to refer to the high temperatures, temperature gradients, or heat fluxes experienced within components of hypersonic or high speed vehicles. In accordance with exemplary embodiments of the present subject matter, the leading edge assembly 300 forms or is provided with thermal conditioning features or techniques for managing these thermal loads.

For example, as described in more detail below with reference to fig. 1, the leading edge assembly 300 may include features for providing or distributing a cooling fluid or material within the outer wall 320 of the leading edge assembly 300 to move thermal energy from a relatively hot location, e.g., proximate the stagnation point 302, to a relatively cold region downstream of the stagnation point. In this manner, the temperature gradient experienced within the leading edge assembly 300 may be reduced. Fig. 2-4 provide cooling techniques for the leading edge assembly 300 according to alternative embodiments. It should be appreciated that the thermal conditioning features and techniques described herein for each exemplary leading edge assembly 300 may be used alone, or in combination with any other leading edge technique described herein, to condition thermal loads on one or more leading edge assemblies 300 of a hypersonic vehicle or any other surface of any other component subject to high thermal loads.

As described above, the outer wall 320 and other components of the leading edge assembly 300 may be formed from any suitable material. According to an exemplary embodiment, such materials are selected to withstand the high thermal loads experienced by the leading edge of a hypersonic aircraft. For example, the outer wall 320 may be constructed from at least one of aluminum, titanium aluminide, tungsten alloys, nickel superalloys, refractory metals, single crystal metals, ceramics, Ceramic Matrix Composites (CMC), ultra high temperature ceramics (UHTC, including high melting point diborides, nitrides, etc.), or carbon-carbon composites. Additionally or alternatively, the outer wall 320 may include composites such as silicon carbide (SiC), SiC composites, carbon fiber reinforced SiC matrices, and other carbide matrix composites, composites with and without surface coatings, and/or high entropy alloys, including refractory materials, platinum group metals, hafnium alloys, and the like. However, it may still be desirable in certain applications to provide additional cooling capability for thermal management of the high thermal loads experienced by the leading edge assembly 300. Further, as described above, additive manufacturing techniques may be used to print the leading edge assembly 300 (e.g., including the outer wall 320) as a single, unitary component, and may facilitate improved cooling techniques and leading edge features.

As shown in the depicted embodiment, the outer wall 320 is generally formed by a first wall section 322 and a second wall section 324 that intersect or join at the stagnation point 302. More specifically, first wall section 322 and second wall section 324 each include an outer surface that together form an outer surface 326 and an inner surface that together form an inner surface 328. Additionally, the first wall section 322 and the second wall section 324 may be angled relative to one another such that the leading edge assembly 300 tapers from a rearward end 330 of the leading edge assembly 300 to a forward end 332 of the leading edge assembly 300 (e.g., corresponding to the stagnation point 302). In other words, the leading edge assembly 300 is wider or taller near the rearward end 330 of the leading edge assembly 300 and narrows as it approaches the stagnation point 302. Notably, the taper angle may vary depending on aerodynamic and other considerations while remaining within the scope of the present subject matter. For example, according to an exemplary embodiment, the leading edge assembly 300 may be asymmetric, e.g., defining a sharper angle on one side.

As described above, for the illustrated embodiment, the outer surfaces 326 of the first and second wall sections 322, 324 intersect at the stagnation point 302 and generally form the leading edge portion 340 of the outer wall 320 and define at least a portion of the outer and inner surfaces 326, 328. For the illustrated embodiment, the leading edge assembly 300 may include a rearward partition 350, the rearward partition 350 being positioned at the rearward end 330 of the outer wall 320 and extending substantially perpendicular to the longitudinal direction L. Specifically, as shown, rearward bulkhead 350 extends between and connects first wall section 322 and second wall section 324 at rearward end 330 of leading edge assembly 300 between first wall section 322 and second wall section 324. Further, aft bulkhead 350 may be connected to the rest of the hypersonic vehicle, may be a replacement component, or the like.

In this manner, first wall section 322, second wall section 324, and rearward partition 350 may generally define a cavity or vapor chamber 352 surrounded and defined by inner surface 328. Thus, according to an exemplary embodiment, vapor chamber 352 may be a closed constant volume chamber or reservoir. According to an exemplary embodiment, the vapor chamber 352 may be filled or loaded with a working fluid 354, the working fluid 354 being used to transfer thermal energy within the leading edge assembly 300. Additionally, the outer wall 320 and the rearward partition 350 may be sealed or impermeable walls such that the vapor chamber 352 may be a sealed chamber for containing the working fluid 354. When the leading edge assembly 300 is configured as described herein, the leading edge assembly 300 may generally function as a heat pipe, or may be used as a heat exchanger that transfers thermal energy through evaporation and condensation of a working fluid, such as the working fluid 354, as will be described in more detail below.

The working fluid 354 is generally any fluid or gas that may be circulated within the vapor chamber 352 to transfer thermal energy from a relatively hotter region of the leading edge assembly 300 (i.e., near the stagnation point 302) to a relatively cooler region of the leading edge assembly 300 (e.g., a region downstream of the stagnation point 302). The working fluid should generally be selected such that it is compatible with the leading edge assembly 300 and suitable for the desired operating range. For example, according to an exemplary embodiment, the working fluid 354 may include at least one of water, steam, acetone, methanol, ethanol, toluene, and the like. According to still other embodiments, the working fluid 354 may include one or more of lithium, sodium, silver, and the like. As described in more detail below, the working fluid 354 may be configured to evaporate from a liquid state to a gaseous state to absorb thermal energy, and to condense from the gaseous state back to the liquid state to discharge the thermal energy into a cooler region or surface of the leading edge assembly 300.

According to the illustrated embodiment, vapor chamber 352 generally extends between a condenser section or condenser area 360 at one end of vapor chamber 352 and an evaporator section or evaporator area 362 at an opposite end of vapor chamber 352. Specifically, as shown, the evaporator region 362 is positioned near the forward end 332 of the leading edge assembly 300, e.g., near the stagnation point 302, where temperature and heat flux are generally highest. Conversely, the condenser zone 360 may be generally positioned near the aft end 330 of the leading edge assembly 300, where the temperature is relatively low compared to the stagnation point 302.

Although the condenser and evaporator regions 360, 362 are illustrated as being located at the aft and forward ends, respectively, of the leading edge assembly 300, it should be understood that the regions serving as evaporator or condenser surfaces may vary depending on operating conditions, for example. For example, depending on certain operating conditions, the condenser region 360 may extend along the entire outer wall 320 except for, for example, the leading edge region at the forward end 332 immediately adjacent the stagnation point 302.

During operation, the working fluid 354 contained in the vapor chamber 352 of the leading edge assembly 300 absorbs thermal energy at the evaporator region 362, e.g., stagnation point 302. The working fluid 354 may evaporate and travel in a gaseous state from the evaporator region 362 to the condenser region 360. In the condenser zone 360, the gaseous working fluid 354 condenses to a liquid state, thereby releasing thermal energy. The working fluid 354 can then flow back to the evaporator region 362 in liquid form, for example, via capillary flow as described below. In this manner, vapor chamber 352 and working fluid 354 generally act as a heat pipe, transferring thermal energy from the portion of leading edge assembly 300 that is subject to the highest thermal load toward the region of leading edge assembly 300 that is subject to a relatively low thermal load. After the heat is transferred to the rearward facing surface (e.g., proximate the condenser region 360), the heat may be expelled from the leading edge assembly 300 in the form of thermal radiation.

It should be understood that the terms "liquid" and "vapor" generally refer herein to the phase or state of the working fluid 354 as it passes within the vapor chamber 352 between the condenser region 360 and the evaporator region 362. However, it should be understood that the present subject matter does not require that all of the working fluid 354 in the condenser zone 360 be liquid, and vice versa, that all of the working fluid 354 in the evaporator zone 362 be vapor. Depending on the current operating conditions of the leading edge assembly 300, the working fluid 354 may be in any suitable state without departing from the scope of the present subject matter.

As best shown in the enlarged portion of fig. 1, the leading edge assembly 300 may further include a capillary structure 364, the capillary structure 364 being positioned within the vapor chamber 352 for circulating the working fluid 354. Specifically, as shown, capillary structure 364 is positioned on inner surface 328 of outer wall 320 within vapor chamber 352. In this regard, the capillary structures 364 may line or cover all or a portion of the perimeter of the inner surface 328 for transporting the condensed working fluid 354 toward the stagnation point 302 of the leading edge 300.

The capillary structure 364 may generally be any component, feature, material, or structure configured to transport the liquid working fluid 354 from the condenser region 360 to the evaporator region 362 via capillary flow or force. For example, the capillary structure 364 can be a porous or mesh membrane 366 (shown in FIG. 1). Alternatively, the capillary structure 364 may be an array of capillaries, offset walls, a porous structure, a wick, a screen, a honeycomb structure, or any other structure configured to facilitate the flow of the liquid working fluid 354 toward the evaporator region 362.

Still referring to fig. 1, leading edge assembly 300 may further include a thermal energy storage assembly 370, thermal energy storage assembly 370 being positioned within vapor chamber 352 or in thermal communication with vapor chamber 352. As explained in more detail below, the thermal energy storage assembly 370 is generally configured to absorb thermal energy from the working fluid 354, particularly if extremely high temperatures are experienced at the leading edge assembly 300. In this regard, for example, during normal high heat operation of leading edge assembly 300, vapor chamber 352 and working fluid 354 may operate to reduce the temperature of stagnation point 302 and forward end 332 of outer wall 320 to a suitably low temperature for maintaining structural integrity. However, under extreme heating conditions, the working fluid 354 may not be able to transfer sufficient heat for maintaining structural integrity. However, as will be described in more detail below, the thermal energy storage assembly 370 may be configured to provide additional condenser surface area for absorbing thermal energy from the superheated working fluid 354 under these extreme heating conditions.

Still referring to the figures, thermal energy storage assembly 370 generally includes a reservoir wall 372, reservoir wall 372 defining an interior chamber 374 within vapor chamber 352. According to the illustrated embodiment, the thermal energy storage assembly 370, or more specifically, the interior chamber 374 is positioned adjacent the aft bulkhead 350 of the leading edge assembly 300. Although interior chamber 374 is described and illustrated herein as being positioned within vapor chamber 352, it should be understood that any suitable thermal communication may be used between interior chamber 374 and vapor chamber 352, according to alternative embodiments, while remaining within the scope of the present subject matter. . Thus, according to an alternative embodiment, interior chamber 374 may instead be positioned immediately adjacent to vapor chamber 352 and in thermal contact with vapor chamber 352.

According to an exemplary embodiment, the interior chamber 374 may contain a phase change material 376, the phase change material 376 generally configured to absorb thermal energy from the working fluid 354. In general, phase change material 376 may be any material or substance selected to change its state or phase (e.g., melt and/or solidify) at a desired temperature. When this phase change occurs, the phase change material 376 may absorb or release a large amount of thermal energy (generally referred to herein as latent heat). Specifically, for example, the phase change material 376 may be selected such that it melts when the leading edge assembly 300 is subjected to extremely high temperatures. When such a predetermined critical temperature is reached, phase change material 376 melts, thereby absorbing a large amount of heat from working fluid 354.

The phase change material 376 may generally be any material selected to change phase at a desired temperature for cooling the leading edge assembly 300. For example, according to an exemplary embodiment, phase change material 376 may include silicon or beryllium. According to certain example embodiments, the phase change material may have a melting temperature above a predetermined threshold or within a desired range. For example, according to an exemplary embodiment, the melting temperature of phase change material 376 may be greater than 500 ℃, greater than 1000 ℃, greater than 1200 ℃, greater than 1500 ℃, or higher. Additionally or alternatively, the phase change material 376 may have a melting temperature of less than 3000 ℃, less than 2500 ℃, less than 2000 ℃, less than 1500 ℃, less than 1000 ℃ or less. Other melting temperatures are also possible and within the scope of the present subject matter.

According to an exemplary embodiment, the phase change material 376 may be selected based at least in part on latent heat of fusion. For example, the phase change material 376 may be selected to have a latent heat of fusion between about 100 and 1000kJ/kg, between about 200 and 800kJ/kg, between about 250 and 500kJ/kg, or greater than about 300 kJ/kg. According to an exemplary embodiment, the phase change material 376 may be selected to have a latent heat of fusion greater than 800kJ/kg, greater than 1000kJ/kg, or higher. It should be understood that the phase change material may be selected based on melting temperature, latent heat of fusion, some combination of these two parameters, or based on any other suitable material property.

Additionally, although the exemplary embodiment illustrated herein includes a single interior chamber 374 for housing a single phase change material 376, it should be understood that, according to alternative embodiments, the thermal energy storage assembly 370 may include multiple chambers (not shown), each of which may house one or more different phase change materials 376. In this manner, for example, if the leading edge assembly 300 is designed to operate in two different extreme temperature regions, the phase change material 376 may be selected for absorbing thermal energy within each region. In particular, first phase change material 376 may be selected such that it has a first melting temperature and is designed to absorb latent heat when working fluid 354 reaches the first melting temperature. Additionally, the second phase change material 376 may be selected such that it has a second melting temperature and is designed to absorb latent heat when the working fluid 354 reaches the second melting temperature. Additional phase change materials, melting temperatures, and chamber configurations are possible and within the scope of the present subject matter.

According to the illustrated embodiment, reservoir wall 372 (which may also include wick or capillary structure 364) extends substantially parallel to outer wall 320. However, according to an exemplary embodiment, portions of leading edge assembly 300 may include features for improving thermal communication between working fluid 354, outer wall 320, reservoir wall 372, and/or phase change material 376. For example, to improve the thermal contact area between working fluid 354 and reservoir wall 372, reservoir wall 372 may have a wavy profile, such as shown in fig. 2. Alternatively, as shown in fig. 3, reservoir wall 372 may include a plurality of heat exchange fins 380, the plurality of heat exchange fins 380 being mounted on an outer surface 382 of reservoir wall 372 and extending into working fluid 354 to provide improved thermal communication between reservoir wall 372 and working fluid 354.

Although exemplary heat exchange features are illustrated herein, it should be understood that any other suitable heat exchange features may be used while remaining within the scope of the present subject matter. For example, fig. 2 shows a wavy wall and fig. 3 shows an exchange fin 380, the exchange fin 380 extending substantially perpendicular to the outer surface 382 of the reservoir wall 372, e.g., in a spanwise direction (e.g., into the page as shown in fig. 3). Rather, as shown in FIG. 4, the heat exchange fins 380 may alternatively extend in an airflow direction, e.g., parallel to the airflow 304. Additionally or alternatively, reservoir walls 372 and outer walls 320 may include any suitable number, type, geometry, and configuration of surface aberrations, protrusions, fins, or other suitable features for increasing heat transfer rates. Further, while such heat exchange features are illustrated on outer surface 382 of reservoir wall 372, it should be understood that such features may also be used on inner surface 328 of outer wall 320, inner surface 384 of reservoir wall 372, or any other suitable surface of leading edge assembly 300.

Additionally, it should be understood that the relative volume of interior chamber 374 with respect to vapor chamber 352 may vary while remaining within the scope of the present subject matter. For example, the volume of vapor chamber 352 (excluding the area filled by interior chamber 374) may be equal to the volume of interior chamber 374. According to still other embodiments, the volume of vapor chamber 352 may be about 1.5 times, 2 times, 3 times, 5 times, or more greater than the volume of interior chamber 374. Additionally or alternatively, the volume of vapor chamber 352 may be about 10 times, 5 times, 3 times, or less than the volume of interior chamber 374.

In addition, according to still other embodiments, the volume of the interior chamber 374 may be configured to expand or contract as desired, depending on the state of the phase change material 376. In this regard, for example, reservoir wall 372 may be a compliant containment structure made of an elastic material that is capable of expanding or contracting depending on the state of phase change material 376. Alternatively, reservoir wall 372 may include a plurality of wall segments joined by expansion joints, flexible regions, or other suitable joining mechanisms for allowing phase change material 376 to expand or contract.

The leading edge assembly 300 may further include features for improving circulation of the working fluid 354. For example, referring again to fig. 2, the capillary structure 364 may further include a plurality of liquid bridges 388, the plurality of liquid bridges 388 providing a faster path for the working fluid 354 to reach the evaporator region 362. In this regard, the liquid bridge 388 may comprise a porous microstructure, a solid bridge overlaid in a porous microstructure, or any other suitable capillary structure for providing a shorter path for the working fluid 354 to move toward the evaporator region 362. According to an exemplary embodiment, the liquid bridge 388 may extend away from the reservoir wall 372, e.g., toward the outer wall 320, for providing a shorter path to the leading edge 320 relative to the capillary structure 364 on the reservoir wall 372. It should be appreciated that the number, size, location, and configuration of the liquid bridges 388 may vary, while remaining within the scope of the present subject matter, according to an exemplary embodiment.

Additionally, the thermal energy storage assembly 370 may take any suitable shape and may cover any suitable area within the outer wall 320, for example, to ensure uniform melting and improved heat distribution of the phase change material 376. For example, referring to, for example, fig. 5, the reservoir wall 372 may define a plurality of span fins 390, the plurality of span fins 390 disrupting the phase change material 376 and improving the surface area to volume ratio of the thermal energy storage assembly 370. In this regard, as shown, the reservoir wall 372 may divide the volume of the interior chamber 374 into multiple regions having different numbers, sizes, shapes, and configurations of fins 390. Although span fins 390 are shown, it should be understood that the fins or projecting structures may extend in any other suitable direction and may have any other suitable size or shape according to alternative embodiments.

Referring now to fig. 6, the thermal energy storage assembly 370 may further include conductive paths 392 extending within the internal chamber 374 in any suitable pattern or configuration for distributing thermal energy throughout the internal chamber 374 and ensuring uniform melting and heat distribution therein. For example, according to the illustrated embodiment, conductive paths 392 form a grid-like structure extending throughout interior chamber 374. According to an exemplary embodiment, conductive path 392 may be formed from the same material as reservoir wall 372 or any other suitable conductive material. According to an exemplary embodiment, the thickness of the mesh may increase toward the rearward end 330 of the interior chamber 374, e.g., to increase the amount of thermal energy transferred to the thickest region of the interior chamber 374. It should be understood that the size, structure, geometry, and location of conductive paths 392 may vary while remaining within the scope of the present subject matter.

Accordingly, aspects of the present subject matter disclosed above present improved leading edge assemblies and methods of forming the same for cooling regions of hypersonic aircraft 100 subjected to ultra-high thermal loads. Notably, the leading edge assembly 300, including the outer wall 320, the reservoir wall 372, the capillary structure 364, the heat exchange fins 380, and other features, may be otherwise fabricated as a single, integral, unitary piece. Additionally, the additive manufacturing methods described herein facilitate forming the leading edge assembly 300 using any one or more suitable materials, and enable the formation of extremely complex heat exchange features with high surface areas to improve thermal contact between the materials. Furthermore, the use of working fluid 354 in vapor chamber 352 helps to transfer thermal energy from hotter regions to cooler regions of leading edge assembly 300, while phase change material 376 helps to absorb and manage the high thermal energy generated during hypersonic operation.

Further aspects of the invention are provided by the subject matter of the following clauses:

1. a leading edge assembly for a hypersonic vehicle, the leading edge assembly comprising: an outer wall tapering to a leading edge, wherein the outer wall at least partially defines a vapor chamber; a capillary structure positioned on an inner surface of the outer wall within the vapor chamber; and a thermal energy storage assembly positioned in thermal communication with the vapor chamber.

2. The leading edge assembly of any preceding claim, wherein the thermal energy storage assembly comprises: a reservoir wall; and a phase change material positioned within the reservoir wall.

3. The leading edge assembly of any preceding claim, wherein the phase change material comprises silicon or beryllium.

4. The leading edge assembly of any preceding claim, wherein the phase change material has a melting temperature greater than 1000 degrees celsius.

5. The leading edge assembly of any preceding claim, wherein the phase change material has a latent heat of fusion greater than 300 kJ/kg.

6. The leading edge assembly of any preceding claim, wherein the thermal energy storage assembly comprises a first chamber containing a first phase change material having a first melting temperature and a second chamber containing a second phase change material having a second melting temperature.

7. The leading edge assembly of any preceding claim, wherein the capillary structure is a wick, porous structure, or screen lining between the inner surface of the outer wall and the outer surface of the reservoir wall.

8. The leading edge assembly of any preceding claim, wherein the capillary structure comprises: at least one liquid bridge extending away from the reservoir wall towards the outer wall for providing a shorter path to the leading edge relative to the capillary structure.

9. The leading edge assembly of any preceding claim, wherein the outer wall, the reservoir wall, and the capillary structure are additively manufactured as a single monolithic component.

10. The leading edge assembly of any preceding claim, wherein the reservoir wall is a compliant containment structure that is capable of expanding or contracting depending on the state of the phase change material.

11. The leading edge assembly of any preceding claim, wherein the reservoir wall has an undulating profile.

12. The leading edge assembly of any preceding claim, wherein the reservoir wall extends substantially parallel to the outer wall.

13. The leading edge assembly of any preceding claim, further comprising: a plurality of heat exchange fins mounted on an outer surface of the reservoir wall.

14. The leading edge assembly of any preceding claim, wherein each heat exchange fin of the plurality of heat exchange fins extends in the airflow direction.

15. The leading edge assembly of any preceding claim, wherein the thermal energy storage assembly is positioned adjacent a rearward bulkhead of the leading edge assembly.

16. The leading edge assembly of any preceding claim, wherein the thermal energy storage assembly comprises:

one or more conductive paths extending through the interior chamber.

17. The leading edge assembly of any preceding claim, wherein the vapor chamber is loaded with lithium, sodium, or silver.

18. The leading edge assembly according to any of the preceding claims, wherein the outer wall is formed from a ceramic matrix composite, a carbon-carbon composite, or a refractory material.

19. A leading edge assembly for a hypersonic vehicle, the leading edge assembly comprising: an outer wall tapering to a leading edge, wherein the outer wall at least partially defines a vapor chamber; a capillary structure positioned on an inner surface of the outer wall within the vapor chamber; and a thermal energy storage assembly positioned within the vapor chamber, the thermal energy storage assembly comprising a reservoir wall containing a phase change material.

20. The leading edge assembly of any preceding claim, wherein the phase change material comprises silicon or beryllium.

This written description uses examples to disclose the invention, including the best mode, and also to enable any person skilled in the art to practice the invention, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the invention is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they include structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal languages of the claims.