阅读说明：本技术 用于降低高速运载工具的前缘中的热应力的系统 (System for reducing thermal stress in the leading edge of a high speed vehicle ) 是由 W·D·格斯特勒 N·W·拉齐 B·M·拉什 P·R·苏布拉曼尼亚 于 2020-11-13 设计创作，主要内容包括：本发明涉及用于降低高速运载工具的前缘中的热应力的系统。一种高超音速飞行器包括一个或多个前缘组件,这些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历的热负荷。前缘组件包括彼此交替堆叠的多个结构层和多个顺应层,以便于多个结构层之间的热膨胀和移动,同时还在多个结构层之间提供热中断。(The invention relates to a system for reducing thermal stresses in the leading edge of a high speed vehicle. A hypersonic aircraft includes one or more leading edge assemblies designed to manage the thermal loads experienced at the leading edge during high speed or hypersonic operation. The leading edge assembly includes a plurality of structural layers and a plurality of compliant layers alternately stacked with one another to facilitate thermal expansion and movement between the plurality of structural layers while also providing thermal breaks between the plurality of structural layers.)

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

an outer wall tapered at a forward end thereof;

a tip portion joined to the forward end of the outer wall and extending forward toward a leading edge, the tip portion comprising:

an inner layer positioned at a rear end of the tip portion;

an outer layer defining the leading edge at a leading end of the tip portion; and

one or more compliant layers positioned between the inner layer and the outer layer for facilitating movement between the inner layer and the outer layer.

2. The leading edge assembly of claim 1, wherein the tip portion further comprises:

an intermediate layer positioned between the inner layer and the outer layer, and wherein the one or more compliant layers include a first compliant layer positioned between the inner layer and the intermediate layer and a second compliant layer positioned between the intermediate layer and the outer layer.

3. The leading edge assembly of claim 2, wherein the outer wall comprises a first wall section and a second wall section separated by a cavity, and wherein the inner layer, the intermediate layer, and the outer layer are each curved and extend between the first wall section and the second wall section.

4. The leading edge assembly of claim 1, wherein the inner layer and the outer layer are made of different materials.

5. The leading edge assembly of claim 1, wherein the outer layer and the inner layer are made of a metal, ceramic material, or ceramic matrix composite.

6. The leading edge assembly of claim 1, wherein the outer layer and the inner layer have different densities.

7. The leading edge assembly of claim 1, wherein at least two of the one or more compliant layers are made of different materials.

8. The leading edge assembly of claim 1, wherein the inner layer and the outer layer have a first coefficient of thermal expansion, and the one or more of the compliant layers have a second coefficient of thermal expansion that is greater than the first coefficient of thermal expansion.

9. The leading edge assembly of claim 1, wherein at least one cooling channel is defined through the inner layer, the outer layer, and the one or more compliant layers.

10. The leading edge assembly of claim 9, further comprising:

a coolant supply in fluid communication with the at least one cooling channel for selectively providing a flow of coolant through the at least one cooling channel.

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 experience thermal management issues (particularly at the leading edge where the free air stream impinges the vehicle) caused by the high thermal loads experienced during high speed operation. For example, in applications involving hypersonic aircraft, the leading edge may include the leading edges of the nose, engine fairings, and wings and stabilizers. Particularly when these vehicles are operated in the hypersonic range (e.g., Mach 5 or greater), the leading edge may be subjected to very high thermal loads (e.g., 500-²) Since the incident air flow passes through the bow shock and rests at the vehicle surface, the kinetic energy of the gas is converted into internal energy and its temperature is greatly increased. Complete exposure to such thermal loads can cause 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-based superalloy may be used for 800 ℃, a single crystal material may be used for 1200 ℃, and for even higher temperatures, refractory metals may be required. In addition, a variety of cooling techniques have been developed to provide cooling to the leading edge of hypersonic vehicles. However, corresponding advances in vehicle speed and duration of high speed flight time have created a need to further improve the cooling capability and high temperature durability of the leading edge of high speed vehicles.

In addition, because the thermal load is greatest at the stagnation point or leading edge of a hypersonic vehicle, large temperature gradients may be created within the structure defining the leading edge. For example, the outermost layer of the leading edge may experience the highest heat and greatest thermal expansion, while the innermost layer may experience relatively low heat and lesser thermal expansion. As a result, thermal stresses may be generated within the leading edge, which may cause the component to fail or deteriorate.

Accordingly, improvements in hypersonic aircraft and propulsion technology would be useful. More specifically, it would be particularly beneficial to improve thermal management techniques for reducing thermal stresses within the leading edge of hypersonic vehicles.

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 an exemplary embodiment of the present disclosure, a leading edge assembly for a hypersonic vehicle, the leading edge assembly comprising: an outer wall tapering to a forward end; a tip portion joined to the forward end of the outer wall and extending forward toward the leading edge, the tip portion comprising: an inner layer positioned at a rear end of the tip portion; an outer layer defining a leading edge at a leading end of the tip portion; and one or more compliant layers positioned between the inner and outer layers for facilitating movement between the inner and outer layers.

According to another exemplary embodiment, a leading edge assembly for a hypersonic vehicle, the leading edge assembly comprising: an outer wall tapering to a forward end; a tip portion joined to the forward end of the outer wall, the tip portion comprising a plurality of structural layers and a plurality of compliant layers alternately stacked with one another, wherein the plurality of compliant layers facilitate movement between the plurality of structural layers.

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

an outer wall tapered at a forward end thereof;

a tip portion joined to the forward end of the outer wall and extending forward toward a leading edge, the tip portion comprising:

an inner layer positioned at a rear end of the tip portion;

an outer layer defining the leading edge at a leading end of the tip portion; and

one or more compliant layers positioned between the inner layer and the outer layer for facilitating movement between the inner layer and the outer layer.

The leading edge assembly of any preceding claim, wherein the tip portion further comprises:

an intermediate layer positioned between the inner layer and the outer layer, and wherein the one or more compliant layers include a first compliant layer positioned between the inner layer and the intermediate layer and a second compliant layer positioned between the intermediate layer and the outer layer.

Solution 3. the leading edge assembly of any preceding solution, wherein the outer wall comprises a first wall section and a second wall section separated by a chamber, and wherein the inner layer, the intermediate layer, and the outer layer are each curved and extend between the first wall section and the second wall section.

Claim 4. the leading edge assembly of any preceding claim, wherein the inner layer and the outer layer are made of different materials.

Claim 5. the leading edge assembly of any preceding claim, wherein the outer layer and the inner layer are made of a metal, ceramic material or ceramic matrix composite.

Claim 6. the leading edge assembly of any of the preceding claims, wherein the outer layer and the inner layer have different densities.

Solution 7. the leading edge assembly of any preceding solution, wherein at least two of the one or more compliant layers are made of different materials.

The leading edge assembly of any preceding claim, wherein the inner layer and the outer layer have a first coefficient of thermal expansion, and the one or more of the compliant layers have a second coefficient of thermal expansion that is greater than the first coefficient of thermal expansion.

The leading edge assembly of any preceding claim, wherein at least one cooling channel is defined through the inner layer, the outer layer, and the one or more compliant layers.

The leading edge assembly of any preceding claim, further comprising:

a coolant supply in fluid communication with the at least one cooling channel for selectively providing a flow of coolant through the at least one cooling channel.

Solution 11. the leading edge assembly of any preceding solution, wherein the leading edge assembly is positioned on a wing, nose cone, engine fairing, engine inlet, fuselage or stabilizer of the hypersonic vehicle.

Technical solution 12. a leading edge assembly for a hypersonic vehicle, the leading edge assembly comprising:

an outer wall tapered at a forward end thereof;

a tip portion joined to the forward end of the outer wall, the tip portion comprising a plurality of structural layers and a plurality of compliant layers alternately stacked with one another, wherein the plurality of compliant layers facilitate movement between the plurality of structural layers.

The leading edge assembly of any preceding claim, wherein the outer wall comprises a first wall section and a second wall section separated by a chamber, and wherein the plurality of structural layers are curved and extend between the first wall section and the second wall section.

The leading edge assembly of any preceding claim, wherein the plurality of structural layers are made of different materials.

Claim 15. the leading edge assembly of any preceding claim, wherein the plurality of structural layers are made of a metal, ceramic material, or ceramic matrix composite.

The leading edge assembly of any preceding claim, wherein at least two of the plurality of structural layers have different densities.

The leading edge assembly of any preceding claim, wherein at least two of the plurality of compliant layers are made of different materials.

The leading edge assembly of any preceding claim, wherein the plurality of structural layers each have a first coefficient of thermal expansion and the plurality of compliant layers each have a second coefficient of thermal expansion that is greater than the first coefficient of thermal expansion.

The leading edge assembly of any preceding claim, wherein at least one cooling channel is defined through the plurality of structural layers and the plurality of compliant layers.

The leading edge assembly of any preceding claim, further comprising:

a coolant supply in fluid communication with the at least one cooling channel for selectively providing a flow of coolant through the at least one cooling channel.

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 method of cooling a leading edge and facilitating magnetohydrodynamic generation or control of a hypersonic vehicle, according to an exemplary embodiment of the present subject matter.

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.

FIG. 7 is a close-up cross-sectional schematic view of an outer wall 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 will now be made in detail to the present embodiments of the invention, one or more examples of which are illustrated in the accompanying 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 explanation of the invention, 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 of the invention. For instance, features illustrated or described as part of one embodiment, can be used with another embodiment to yield a still further embodiment. It is therefore intended that the present invention cover such modifications and variations as 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 component from another component and are not intended to denote the position or importance of the individual components. The singular forms "a", "an" and "the" include plural referents unless the context clearly dictates otherwise. Unless otherwise specified herein, the terms "coupled," "fixed," "attached," and the like refer to both a direct coupling, fixing, or attachment, as well as an indirect coupling, fixing, or attachment 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 normal operating attitudes of the gas turbine engine or vehicle. For example, with respect to a gas turbine engine, "forward" refers to a location closer to an engine inlet, and "aft" refers to a location closer to an engine nozzle or exhaust outlet. The terms "upstream" and "downstream" refer to relative directions with respect to fluid flow in a fluid pathway. 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, may be 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," will not be limited to the precise value specified. In at least some examples, 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 within a 10 percent margin.

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 relate 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 greater, such as mach 5 and above. However, it should be appreciated that aspects of the present subject matter are not limited to hypersonic flight only, but may instead be applied to applications involving other high speed vehicles, projectiles, objects, and the like. The description herein of the leading edge assembly with respect to use on a hypersonic aircraft is merely an example, which is intended to facilitate explanation of aspects of the present subject matter. The present subject matter is not limited to such exemplary embodiments and applications.

Notably, as explained above, high speed vehicles, such as hypersonic aircraft, typically experience extremely high temperatures and thermal gradients during high speed or hypersonic operation. Temperature gradients caused by high heat fluxes are generally 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, this gradient results in mechanical stresses that cause plastic deformation or fracture of the material. The thermal load on the structural material should be reduced to maintain the structural integrity of the component.

As explained above, the leading edge of such high speed vehicles typically experiences 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 experience 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, a nose cone, a vertical stabilizer, a cowling for a propulsion 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 loads (e.g., by carrying heat away from the area).

Notably, it is typically 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 within leading edge assembly 300 at its leading or leading edge (also referred to herein as the stagnation line, stagnation point 302, or the like). In this regard, as the hypersonic vehicle travels through the air at hypersonic speed, a free stream of air (e.g., identified herein by reference numeral 304) passes over the leading edge assembly 300 and around the leading edge assembly 300, thereby creating a large thermal load. Aspects of the present subject matter relate to thermal management techniques and features for cooling the leading edge assembly 300, with particular attention being directed to 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 illustration 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 1 mm 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, such as the leading edge of a wing, the nose, the propulsion engines, or other parts of the hypersonic aircraft that experience large temperature and thermal gradients. However, it should be appreciated 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 the leading edge assembly 300 of a hypersonic vehicle are described herein, it should be appreciated that variations and modifications to such techniques may be made without departing from the scope of the present subject matter. Additionally, one or more of such techniques may be used in combination with each other to achieve improved cooling and thermal management. In this regard, although each cooling technique is described in isolation for the purpose of clearly describing how each technique works, 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 3D printing process. Using 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 subcomponents. As used herein, the terms "additive manufacturing" or "additive manufacturing technique or process" generally refer to a manufacturing process as follows: in which successive layers of material(s) are placed one on top of the other to "build" a three-dimensional structure layer by layer. The successive layers are typically fused together to form a unitary member, which may have a variety of integral subcomponents.

Although additive manufacturing techniques are described herein to enable complex objects to be manufactured by building objects point-by-point, layer-by-layer, typically 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 recognize that the methods and structures disclosed herein may be practiced with any additive manufacturing technique or fabrication technique. For example, embodiments of the present invention may use a layering process, a subtractive process, or a hybrid process.

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 Net Shape (LENS), laser net shape 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, which may be in solid, liquid, powder, sheet, wire, or any other suitable form or combination thereof. More specifically, in accordance with exemplary embodiments of the present subject matter, the additively manufactured components described herein may be formed, in part, in whole or in some combination, from materials including, but not limited to: pure Metals, nickel alloys, chromium alloys, titanium alloys, magnesium alloys, aluminum alloys, and nickel-based or cobalt-based superalloys (e.g., a superalloy available from Special Metals Corporation under the name Inconel @). 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 process disclosed herein allows 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, sections, or portions 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 through an additive manufacturing process, it should be appreciated that in alternative embodiments, all or portions of these components may be formed via casting, machining, and/or any other suitable manufacturing process. In fact, any suitable combination of materials and manufacturing methods may be used to form these components.

Still referring to fig. 1, the leading edge assembly 300 will be described in more detail in accordance with exemplary embodiments of the present subject matter. 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., leading end, 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 inlet 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 leading end of a vertical stabilizer, etc.

As explained herein, the leading edge assembly 300 may experience a large thermal load during hypersonic flight operations. As used herein, the term "heat 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 a coolant, such as a cooling fluid or a cooling material, through the outer wall 320 of the leading edge assembly 300 to reduce the temperature of the outer wall 320 and/or the temperature gradient experienced within the leading edge assembly 300. FIGS. 2-5 provide cooling techniques for a 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 the thermal load on one or more leading edge assemblies 300 of a hypersonic vehicle, or any other surface of any other component experiencing a high thermal load.

As explained above, the outer wall 320 and other components of the leading edge assembly 300 may be formed of 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 high entropy alloys 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 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. Moreover, as explained 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 from a first wall section 322 and a second wall section 324. 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 the aft end 330 of the leading edge assembly 300 to the forward end 332 of the leading edge assembly 300 (e.g., which corresponds to the stagnation point 302). In other words, the leading edge assembly 300 is wider or higher near the trailing 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 not be symmetrical, e.g., defining a more acute 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. The leading edge portion 340 also defines at least portions of the outer surface 326 and the inner surface 328. For the depicted embodiment, the leading edge assembly 300 includes a cavity or chamber 342 positioned between the first wall section 322 and the second wall section 324 and in fluid communication with the inner surface 328 of the leading edge portion 340. Further, according to an exemplary embodiment, leading edge portion 340 defines at least one fluid channel 350 through outer wall 320.

More specifically, according to the illustrated embodiment, the flow channel 350 is a hole or aperture defined through the outer wall 320 such that the flow channel 350 extends from an inlet 352 positioned at the inner surface 328 to an outlet 354 positioned at the outer surface 326. Additionally, as illustrated, the outlet 354 is positioned at the stagnation point 302 or the leading end 332 of the leading edge assembly 300. Although a single fluid channel 350 is illustrated in the cross-section depicted in fig. 1, it should be appreciated that the fluid channel 350 may instead be an elongated slot extending along the leading edge or stagnation line of the leading edge assembly 300. According to still other embodiments, the leading edge portion 340 of the leading edge assembly 300 may include a plurality of fluid channels 350 spaced apart from one another along the length of the leading edge assembly 300 (e.g., such as along the span of an aircraft wing).

Indeed, the fluid passages 350 may be defined through the outer wall 320 in any suitable number, geometry, configuration, etc. For example, the leading edge assembly 300 may generally define a radial direction R and a circumferential direction C. The radial direction R extends outwardly from a center of curvature (not labeled) of the leading edge portion 340 in the cross-sectional plane illustrated in fig. 1. The circumferential direction C generally encircles the leading edge assembly 300 from the first wall section 322 to the second wall section 324, as shown in fig. 1. Additionally, the leading edge assembly 300 may define a longitudinal direction L that corresponds to the longitudinal direction L of the hypersonic aircraft and is substantially parallel to an angle of attack defined between the leading edge assembly 300 and the primary direction of airflow 304, according to exemplary embodiments described herein.

According to the illustrated embodiment, the fluid channel 350 extends through the outer wall 320 along the radial direction R. However, it should be appreciated that, according to alternative embodiments, the fluid passages 350 may pass through the outer wall 320 at angles other than the radial direction R, may pass along the length of the first wall section 322 and/or the second wall section 324, may terminate at other locations on both the inner surface 328 and the outer surface 326, may vary in cross-sectional size or profile, or may vary in any other suitable manner for providing a flow of cooling fluid 360 to a desired location within or on the leading edge assembly 300.

In this regard, the leading edge assembly 300 may further include a coolant supply 362 for providing a flow of cooling fluid 360 through the fluid channel 350 and out of the leading edge assembly 300. As used herein, the terms "cooling fluid," "coolant," and the like are generally intended to refer to any material that passes through the outer wall 320 to cool the outer surface 326 of the leading edge assembly 300. In accordance with embodiments of the present subject matter, as described in greater detail below, the cooling fluid 360 may include at least one of air, water, steam, liquid metal, carbon dioxide, argon, helium, and the like.

It should further be appreciated that the cooling fluid 360 is not limited to a particular state or phase of material. In this regard, although the terms "liquid" and "fluid" may be used to refer to the cooling fluid 360, the cooling fluid 360 may also include solid materials, vapors, and the like. Furthermore, the present subject matter is not limited by the source of cooling fluid 360. For example, the cooling fluid 360 may be pumped from a storage reservoir, may be contained within the leading edge assembly 300 as a solid material that liquefies at a critical temperature, permeates through the leading edge portion 340 or fluid channel, evaporates to absorb thermal energy, and carries it away as it flows downstream.

According to exemplary embodiments of the present subject matter, the cooling fluid 360 may be a solid or liquid metal. For example, according to an exemplary embodiment, the cooling fluid 360 is an alkali metal. Exemplary liquid metals that may be pumped as the cooling fluid 360 through the leading edge assembly 300 may include at least one of cesium (Cs, which has a boiling point of approximately 678℃.), potassium (K, which has a boiling point of approximately 774℃.), sodium (Na, which has a boiling point of approximately 883℃.), lithium (Li, which has a boiling point of approximately 1347℃.), indium (In, which has a boiling point of approximately 2000℃.), and/or gallium (Ga, which has a boiling point of approximately 2205℃.). Additionally, or alternatively, cooling fluid 360 may include ammonia (NH)₃) Or sulfur dioxide (SO)₂). According to an exemplary embodiment, the cooling fluid 360 may be a mixture or combination of such liquid metals or compositions.

Additionally, the cooling fluid 360 may be selected based on its material properties, such as ionization potential. The ionization potential of the cooling fluid 360 may be important to facilitate magnetohydrodynamic control or power generation using the leading edge assembly 300 (as explained below). For example, according to exemplary embodiments, the liquid metal cooling fluid 360 may have an ionization potential between about 1 and 10 electron volts (eV), between about 2 and 8 electron volts (eV), between about 4 and 6 electron volts (eV), or about 5 electron volts (eV). According to an exemplary embodiment, the cooling fluid 360 is selected such that its ionization potential is much lower than that of air. E.g. O₂Has an ionization potential of approximately 12.6 eV, and N₂Has an ionization potential of approximately 14.5 eV.

These materials can be used as "seeds" which are at temperatures well below that of airAnd (4) lower ionization. It should be appreciated that the terms "ionize" and the like are generally used herein to refer to both full ionization and partial ionization. For example, the term ionization may refer to partial ionization or weak ionization, e.g., where the density of ions is a fraction (such as 10) of the density of neutral species^-5). Additionally, it should be recognized that if the temperature is high enough (e.g., due to high mach numbers), not only the seed species, but any species present (including oxygen and nitrogen) may be ionized.

According to alternative embodiments, the cooling fluid 360 may be selected based on another material property, such as boiling temperature. According to an exemplary embodiment, the cooling fluid 360 may have a boiling temperature that is lower than the melting temperature of the outer wall 320. In this manner, the outer wall 320 or other portions of the leading edge assembly 300 may be impregnated with a solid material (e.g., a solid metal) that passively melts and flows as the cooling fluid 360. The molten metal cooling fluid 360 will evaporate and begin to cool the outer wall 320 before the outer wall 320 begins to melt. In this manner, evaporation of the cooling fluid 360 on the outer surface 326 may maintain the temperature of the outer wall 320 at the boiling or boiling temperature of the cooling fluid 360. If the boiling point is selected to be below the melting temperature of the outer wall 320, the structural integrity of the outer wall 320 may be maintained until the cooling fluid 360 has been depleted. For example, according to one embodiment, the liquid metal cooling fluid 360 has a boiling temperature greater than about 500 ℃, greater than about 700 ℃, greater than about 900 ℃, greater than about 1100 ℃, or about 750 ℃. Additionally, or alternatively, the boiling temperature of the liquid metal cooling fluid 360 may be less than about 2000 ℃, less than about 1500 ℃, less than about 1000 ℃, or less than about 800 ℃. For example, according to an exemplary embodiment, the cooling fluid 360 is lithium (Li), which boils at about 1342 ℃.

As used herein, the terms "boiling point", "boiling temperature", and the like refer to the temperature at which the table is made at an ambient pressure of 1 atmosphere. However, it should be appreciated that such boiling temperatures may differ under flight conditions (e.g., when ambient pressure and experienced pressure are different). Additionally, although the boiling point of the cooling fluid 360 is described as being below the melting temperature of the outer wall 320, it should be appreciated that it may be desirable to have the boiling point of the cooling fluid 360 even lower, e.g., corresponding to some predetermined structural temperature at which the integrity of the outer wall 320 begins to deteriorate or diminish.

It should be appreciated that the cooling fluid 360 described herein is merely exemplary and is not intended to limit the scope of the present subject matter. For example, the cooling fluid 360 may be selected to be a liquid or a solid, may have varying boiling points and ionization potentials, and may also exhibit other material properties. For example, it may be desirable to have a material with a relatively low mass density, particularly for applications involving aircraft. In addition, the heat of vaporization can be important in determining the volume and mass of the cooling fluid 360 required, e.g., lithium (Li) has a uniquely high specific heat of vaporization (per unit mass). According to an exemplary embodiment, cooling fluid 360 may be selected to melt at a relatively low melting temperature, for example, to minimize the complexity required of an on-board material handling system.

Additionally, or alternatively, the cooling fluid 360 may be a metallic phase change material. For example, the coolant may be a metal configured to change from a solid phase to a liquid or gas phase when exposed to temperatures generated during operation of the hypersonic propulsion engine during hypersonic flight operation. Additionally or alternatively, the coolant may be a metal configured to change from a liquid phase to a gas phase when exposed to temperatures generated during operation of the hypersonic propulsion engine during hypersonic flight operation. However, in other embodiments, other suitable coolants may be utilized.

As best shown in fig. 1, a coolant supply 362 is in fluid communication with the fluid channel 350 for selectively providing a flow of coolant 360 through the fluid channel 350. Specifically, according to the illustrated embodiment, the coolant supply 362 includes a coolant reservoir 364, and the mechanical pump 366 is configured for pressurizing or supplying the cooling fluid 360 to the fluid channel 350. According to still other embodiments, the cooling fluid 360 may be stored within a pressurized tank for delivery to the fluid channel 350, or the leading edge assembly 300 may include a capillary system or another capillary structure for driving the cooling fluid 360. Alternatively, cooling fluid 360 may be delivered from any suitable location in any other suitable manner. For example, as illustrated, coolant supply 362 generally provides a flow of pressurized cooling fluid 360 into chamber 342 in fluid communication with fluid passage 350. However, it should be appreciated that, according to alternative embodiments, the coolant supply 362 may include one or more conduits fluidly coupled directly to the fluid channel 350. Further, while the cooling fluid 360 is illustrated as a single stream from a single coolant reservoir 364, it should be appreciated that one or more cooling fluids 360 may be provided from multiple coolant reservoirs 364, may include different cooling fluid(s) 360 at different pressures, and so forth.

As explained above, according to exemplary embodiments of the present subject matter, the flow of cooling fluid 360 may include a liquid metal. Notably, when the liquid metal cooling fluid 360 is exposed to such high temperatures at the leading edge assembly 300, it may ionize after exiting the fluid channel 350 to produce an ionized vapor stream (e.g., as generally identified by reference numeral 370 in fig. 1). The ionized vapor flow 370 is conveyed downstream over the leading edge assembly 300 from the stagnation point 302. According to an exemplary embodiment, this ionized vapor stream 370 may facilitate a magnetohydrodynamic process for power generation and/or vehicle control. Specifically, according to an exemplary embodiment, the leading edge assembly 300 may include a magnetohydrodynamic generator 372 ("MHD generator") that interacts with the ionized vapor stream 370 to generate electrical power or thrust for vehicle control.

In general, the MHD generator 372 may be any device suitable for interacting with the ionized vapor stream 370 to generate power or thrust for vehicle control. It will be appreciated that MHD generator 372 may generally utilize a magnetohydrodynamic converter that utilizes the brayton cycle to directly convert thermal and kinetic energy into electricity. For example, MHD generator 372 may include one or more magnetic coils and/or electrodes positioned near outer wall 320 or within outer wall 320. According to an exemplary embodiment, the MHD generator 372 may use the coils to generate electricity as the ionized vapor stream 360 passes through the magnetic field generated by the coils. Additionally, or alternatively, the MHD generator 372 may energize magnetic coils or electrodes to interact with the ionized vapor stream 370 in a manner that controls or adjusts flight by providing thrust to the outer wall 320. The MHD generator 372 may further include a controller and unique structure for implementing control methods for steering during high speed flight, such as during hypersonic flight.

Specifically, as best illustrated in FIG. 2, the method of operating the leading edge assembly 300 may include a magnetohydrodynamic control method. Specifically, the method 400 includes: at step 410, a coolant flow is provided through at least one fluid channel defined through a leading edge of the hypersonic vehicle, wherein the coolant flow includes liquid metal (or any other suitable seed material) that is ionized after exiting the at least one fluid channel and is conveyed downstream as an ionized vapor flow. Once the ionized vapor stream is achieved, step 420 includes using a magnetohydrodynamic generator that interacts with the ionized vapor stream to generate electricity or thrust for vehicle control.

Additionally, or alternatively, as best shown in fig. 3 and 4, the leading edge portion 340 may be configured as a porous leading edge portion, which is identified herein as a porous tip 500. According to an exemplary embodiment, the leading edge assembly 300 is configured to provide a flow of cooling fluid 360 through the cavity 342 to the inner surface 328 of the leading edge portion 340 such that the flow of cooling fluid 360 may penetrate through the porous tip 500 and cool the leading edge portion 340 during operation of the hypersonic vehicle or hypersonic propulsion engine (e.g., during hypersonic flight operation). In this regard, the fluid passage 350 may be defined through a porous structure that constitutes or defines the porous tip 500. In describing the drawings herein, it should be appreciated that the same reference numerals may be used to refer to the same or similar features between embodiments.

Specifically, as shown in fig. 3 and 4, the outer wall 320 of the leading edge assembly 300 includes a first wall section 322 and a second wall section 324, each wall section extending along the longitudinal direction L from the forward end 502 and the aft end (e.g., which may correspond to the aft end 330 of the leading edge assembly 300). Notably, the front ends 502 of the first and second wall sections 322, 324 may stop before the stagnation point 302. Additionally, the porous tip 500 may be joined to the forward ends 502 of the first and second wall sections 322, 324. In this regard, the porous tip 500 may generally be defined as a curved region coupled to the first and second wall sections 322, 324, which are illustrated as being substantially straight.

According to an exemplary embodiment, for example, first wall section 322, second wall section 324, and tip portion 500 may be simultaneously additively manufactured as a single, unitary, one-piece component. However, as will be described in greater detail below, the outer wall 320 may be substantially solid and airtight, while the porous tip 500 may include a porous structure for allowing the flow of the cooling fluid 360 to pass through the porous structure toward the forward end 332 of the leading edge assembly 300.

According to exemplary embodiments, the leading edge portion 340, or more specifically the porous tip 500, may define a constant porosity for passage of the cooling fluid 360. As used herein, the term "porosity" may be used generally to refer to a measure of voids or empty space within a material or structure. Thus, a structure having porosity has open channels, cells, or structures through which fluid can flow from one porous unit to another. For example, porosity may be used to refer to the fraction of the volume of voids or open spaces divided by the total volume of the member. According to exemplary embodiments, the porosity of the porous tip 500 may be greater than about 5%, 10%, 20%, 40%, or greater than even 50%. Additionally, or alternatively, the porosity of the porous tip 500 may be less than about 80%, 60%, 40%, 20%, or 5%. It should be appreciated that the porosity of the porous tip 500 may vary depending on the application while remaining within the scope of the present subject matter. For example, the porosity may vary based on the mass flow of the cooling fluid 360, the mechanical properties of the porous tip 500, based on expected flight conditions, or based on any other suitable parameter.

Notably, according to the illustrated embodiment, the porous tip may define a variable porosity, for example, in order to focus the cooling fluid 360 near the stagnation point 302. More specifically, the porous tip 500 may include a first porous region 510, a second porous region 512, and a third porous region 514. According to the illustrated embodiment, the first porous region 510 is positioned at the leading end 322 of the leading edge assembly 300 and includes the stagnation point 302. In contrast, each of the second porous region 512 and the third porous region 514 is positioned downstream of the first porous region 510 or stagnation point 302. According to an exemplary embodiment, the first porosity of the first porous region 510 is different from the second porosity of the second porous region 512. More specifically, according to an exemplary embodiment, the first porosity is greater than the second porosity. Similarly, the third porosity may be the same or different from the second porosity, but also less than the first porosity. For example, according to exemplary embodiments, the first porosity may be at least about 10% greater than the second and third porosities, such as at least about 25% greater, such as at least about 50% greater, such as at least about 100% greater, and up to about 1000% greater than the second and third porosities.

Similar to the embodiments described above, the coolant supply 362 may be in fluid communication with the chamber 342 and/or the porous tip 500 for selectively providing a flow of coolant 360 through the porous tip 500. Specifically, according to the illustrated embodiment, the inner surface 328 of the porous tip 500 is exposed to the pressurized cooling fluid 360 within the cavity 342. However, the amount of cooling fluid 360 flowing through the first porous region 510 may be greater than the amount of cooling fluid 360 flowing through one or both of the second porous region 512 and the third porous region 514. Notably, the ratio or amount of cooling fluid 360 flowing through each region of the porous tip 500 may be adjusted by manipulating the porosity within each region 510 and 514. According to the illustrated embodiment, it is desirable to have the highest porosity in the first porous region 510 to direct the maximum amount of cooling fluid 360 toward the region of greatest thermal influence (i.e., the stagnation point 302). Additionally, the higher porosity at the stagnation point 302 may help compensate for the fact that the pressure is highest at this point, which reduces the amount of cooling fluid 360 that naturally flows to the stagnation point 302.

Notably, the porous tip 500 is illustrated as having three distinct porous regions 510-514. However, it should be appreciated that by using the additive manufacturing techniques described herein, the porous tip 500 may have a gradually changing porosity, i.e., such that the porosity continuously and gradually increases from the leading end 502 of the outer wall 320 to the highest porosity at the stagnation point 302. In this manner, the porous tip 500 may be envisioned to have 10, 20, 50, 100, or even more sub-regions, each having a gradually increasing porosity as it approaches the stagnation point 302. As illustrated, each of these sub-regions extends along the radial direction R from the inner surface 328 to the outer surface 326 and has a substantially constant porosity along the radial direction R. Further, according to alternative embodiments, one or more of these regions (such as porous regions 510-514) may also vary in porosity along the radial direction R.

According to exemplary embodiments, the porous tip 500 may include additional features for directing the flow of the cooling fluid 360 to a desired location within the leading edge assembly 300. In this regard, for example, with reference to fig. 4, the porous tip 500 may include a plurality of internal barriers 520 that separate one or more of the porous regions 510 and 514. In this regard, as illustrated, the inner barrier 520 is a straight and solid wall that extends substantially along the radial direction R from the inner surface 328 to the outer surface 326. However, according to alternative embodiments, the porous tip 500 may have more inner barriers 520, the inner barriers 520 may extend through the outer wall 320 in other directions, may have varying thicknesses, and so on. Other configurations are possible and within the scope of the present subject matter.

Referring now specifically to FIG. 4, leading edge assembly 300 may include a nasal mask 530 positioned at least partially over porous tip 500 to restrict the flow of cooling fluid 360 from coolant reservoir 364 and/or chamber 342. Specifically, the nasal mask 530 prevents the cooling fluid 360 from escaping from the porous tip 500 or the leading edge portion 340, but may be made of a material that ablates or melts away when the leading edge assembly 300 is exposed to a predetermined critical temperature. In this regard, for example, nose cup 530 may be constructed of a material that melts before porous tip 500 and outer wall 320 reach their melting temperatures. Notably, when the nasal mask 530 melts, the cooling fluid 360 is released to cool or maintain the temperature of the leading edge assembly 300 at or below the melting point of the porous tip 500 and/or the outer wall 320. Thus, according to an exemplary embodiment, the predetermined critical temperature at which the nasal mask 530 melts is below a temperature at which structural integrity of the leading edge assembly 300 begins to degrade. Although nasal mask 530 is illustrated as being used with a porous tip 500 having variable porosity, it should be appreciated that these two features may be used together or independently of each other.

In accordance with the illustrated embodiment, nose cup 530 extends across porous tip 500, e.g., from forward end 502 of first wall section 322 to forward end 502 of second wall section 324 along circumferential direction C. Notably, according to an exemplary embodiment, the nasal mask 530 is impermeable to the flow of cooling fluid 360. Thus, cooling fluid 360 is contained within porous tip 500, coolant reservoir 364, and/or chamber 342 until nasal mask 530 melts, which corresponds to the time needed to cool leading edge assembly 300. According to alternative embodiments, nasal mask 530 may cover less than the entire portion of porous tip 500. For example, the nasal mask 530 may cover only the forward end 332 of the first porous region 510. Alternatively, the nasal mask 530 may cover only the second and third porous regions 512, 514. According to still other embodiments, the thickness of the nasal mask 530 (which may be measured substantially along the radial direction R) may vary depending on the circumferential location along the porous tip 500. According to still other embodiments, the nasal mask 530 may be a plug positioned at the outlet 354 of the fluid passage 350 (e.g., as shown in fig. 1). In this regard, for example, the nasal mask 530 may extend at least partially within the fluid passage 350. Other variations and modifications may be made to the nasal mask 530 while remaining within the scope of the present subject matter. For example, it should be appreciated that the nasal mask 530 may be made by filling the holes with a different material to plug the holes and form the nasal mask 530, by attaching a separate cap element to the leading edge assembly 300, and so forth.

Referring now to fig. 5, the leading edge assembly 300 may include an aft bulkhead 532, the aft bulkhead 532 being positioned at the aft end 330 of the outer wall 320 and extending substantially perpendicular to the longitudinal direction L. As shown, chamber 342 is enclosed and defined between rear bulkhead 532, first wall section 322, second wall section 324, and nasal mask 530. Thus, chamber 342 may be a constant volume chamber or reservoir. For example, rear bulkhead 532 may be connected to the remainder of the hypersonic vehicle, may be a replaceable component, or the like. According to such an exemplary embodiment, chamber 342 may be filled with a cooling fluid 360. Notably, as the leading edge assembly 300 heats up during hypersonic operation, the nasal mask 530 may slowly melt while the pressure within the chamber 342 increases due to the increase in temperature of the cooling fluid 360. Once the temperature of the leading edge assembly 300 has reached the critical temperature, the nasal mask 530 melts away and the chamber 342 is sufficiently pressurized to drive a flow of cooling fluid 360 through the porous tip 500, thereby cooling the leading edge assembly 300.

According to still other embodiments, the porous tip 500 may be filled with a material that seeps out of the porous tip 500 when the nasal mask 530 melts or ablates away. More specifically, according to an exemplary embodiment, the porous tip 500 may be filled with a metallic cooling fluid 360 (e.g., in solid form), and the metallic cooling fluid 360 may have a relatively low melting point such that the metal filling the pores of the porous tip 500 is configured to melt during operation of the hypersonic propulsion engine or hypersonic aircraft during high temperature operations, such as hypersonic flight operations. Once the metallic cooling fluid 360 filling the pores of the porous tip 500 melts, the cooling fluid 360 may flow through the porous tip 500 in a manner similar to that described above with reference to FIG. 4.

Referring now specifically to fig. 6 and 7, a leading edge assembly 300 will be described in accordance with another exemplary embodiment of the present subject matter. As described below, the leading edge assembly 300 may be formed from one or more walls comprising a layered multifunctional material having a compliant interface for mitigating extreme thermal stresses experienced by hypersonic aircraft at high heat flux locations (e.g., near the stagnation point 302). As will be described below, each of the plurality of layers within the layered wall may be made of the same material (e.g., such as a metal or ceramic material) or may be made of different materials. Further, each layer may be tailored to meet specific needs for a given application, e.g., based on expected thermal loads, expected temperature gradients, the presence of localized cooling features, etc.

As explained above, the leading edge assembly 300 may include a variety of cooling techniques, such as techniques to provide transpiration cooling features that may include a cooling fluid 360 positioned within the cavity 342. Notably, in such embodiments, the greatest temperature may be experienced on the outer surface 326 near the stagnation point 302, and the lowest temperature may be experienced near the inner surface 328 that may be exposed to the cooling fluid 360. Thus, the large temperature gradients experienced across the outer wall 320 may cause significant thermal stresses (e.g., due to varying thermal expansion between layers or regions of the outer wall 320).

As shown in fig. 6, the leading edge assembly 300 may include a tip portion 550 (e.g., which may correspond to the leading edge portion 340), the tip portion 550 being joined to the forward end 502 of the outer wall 320 or otherwise positioned near the forward end 332 of the leading edge assembly 300. As shown, the tip portion 550 includes a plurality of structural layers 552 and a plurality of compliant layers 554 alternately stacked with one another. In this manner, the plurality of compliant layers 554 may facilitate some movement between the plurality of structural layers 552, and may also provide thermal breaks or insulating gaps between the plurality of structural layers 552. In accordance with the illustrated embodiment, a plurality of compliant layers 554 are embedded within the tip portion 550 between the plurality of structural layers 552 and are spaced apart along the thickness of the outer wall 320 or the tip portion 550.

Although layers 552, 554 are described herein as "structural" and "compliant," it should be recognized that these terms are used merely to distinguish various layers and are not intended to limit such layers to a particular material or material property. Additionally, it should be appreciated that the term "layer" of the outer wall 320 is used herein to generally refer to different areas within the outer wall 320. In this regard, each layer may be substantially isothermal or quasi-isothermal, e.g., such that each layer experiences substantially similar thermal loads. However, the term "layer" is not intended to limit the manner in which the leading edge assembly 300 is constructed, or otherwise dictate or require a precisely layered or laminar construction. For example, the "layers" of the leading edge assembly 300 may all be additively manufactured as a single, unitary piece using one or more materials having different densities, porosities, coefficients of thermal expansion, or other different material properties.

The plurality of structural layers 552 and the plurality of compliant layers 554 may be formed of any suitable material. For example, the layers 552, 554 may be made of metal, ceramic matrix composite, or any other suitable material described herein. In addition, according to an exemplary embodiment, the plurality of structural layers 552 may be formed of the same material and may have the same material properties. In contrast, according to alternative embodiments, the plurality of structural layers 552 may have different materials, compositions, or configurations for optimally managing the thermal load experienced by the leading edge assembly 300. Similarly, the plurality of compliant layers 554 may each be formed of the same material or different materials, which may or may not include the material used to form the plurality of structural layers 552.

For example, referring now specifically to fig. 7, a close-up view of the tip portion 550 will be described in accordance with an exemplary embodiment of the present subject matter. As shown, the tip portion 550 includes an inner layer 560 (i.e., one of the plurality of structural layers 552) positioned near a rear end of the tip portion 550. In this regard, inner layer 560 may at least partially define inner surface 328 of leading edge assembly 300 and cavity 342. In addition, the tip portion 550 includes an outer layer 562, the outer layer 562 defining a leading edge or stagnation point 302 at the leading end 332 of the leading edge assembly 300.

Between the inner layer 560 and the outer layer 562, the distal portion 550 may include one or more compliant layers 554. For example, compliant layer 554 may facilitate movement between inner layer 560 and outer layer 562. More specifically, as shown in fig. 7, the tip portion 550 further includes an intermediate layer 564 and two compliant layers 554. More specifically, a first compliant layer 566 (i.e., one of the plurality of compliant layers 554) is positioned between the inner layer 560 and the intermediate layer 564. Additionally, a second compliant layer 568 is positioned between the middle layer 564 and the outer layer 562. It should be appreciated that the tip portion 550 may include additional structural layers 552 and/or compliant layers 554.

It should be appreciated that the layered structure used to define the tip portion 550 may be used in other areas of the leading edge assembly 300, elsewhere within a hypersonic aircraft, or in other applications where high thermal loads are anticipated. Additionally, although the structural layer 552 and the compliant layer 554 are illustrated as being curved between the first wall section 322 and the second wall section 324, these layers 552, 554 may take any other suitable shape, size, or location for facilitating improved management of thermal stresses.

According to an exemplary embodiment, one or more of the structural layers 552 and one or more of the compliant layers 554 may have different material densities. In this way, adjusting the material density may affect the stress induced within the material layer, may adjust the thermal conductivity of the layer, and the like. In addition, one or more of the structural layers 552 and one or more of the compliant layers 554 may have different coefficients of thermal expansion. In this manner, it may be desirable to form outer layer 562 to have a lower coefficient of thermal expansion than inner layer 560, e.g., to reduce the difference in expansion experienced between the two layers when exposed to different temperatures during hypersonic operation.

According to an exemplary embodiment, each of the plurality of compliant layers 554 may define a thickness 570. According to certain embodiments, the thickness 570 may be less than about 1 millimeter. The plurality of compliant layers 554 between the plurality of structural layers 552 may be effective to distribute heat at, for example, the stagnation point 302 along the outer surface 326 to reduce heat concentration at the stagnation point 302. The plurality of compliant layers 554 may be cavities having an interior volume defined by thickness 570 and may be filled with a fluid or material layer having a relatively high heat transfer coefficient, such as liquid sodium. Notably, in at least some example embodiments, the plurality of compliant layers 554 may define a smaller thickness 570, and the thickness of material between the plurality of compliant layers 554 may be less than or equal to about 1 millimeter.

It should be further appreciated that the layered tip portions described with respect to FIGS. 6 and 7 may be used with other techniques described herein to minimize the impact of the thermal loading of the leading edge assembly 300. For example, according to an exemplary embodiment, the tip portion 550 may define at least one cooling channel (e.g., such as the fluid channel 350 from FIG. 1) for providing a flow of cooling fluid 360 to the outer surface 326 of the leading edge assembly 300. The leading edge assembly 300 may further include a coolant supply, such as coolant supply 362, for providing a flow of cooling fluid 360 to the tip portion 550. Further, the tip portion 550 may be divided into sub-regions similar to the porous tip 500, with each region having a different number, type, material, and configuration of structural layers 552 and compliant layers 554. Other variations and modifications may be made while remaining within the scope of the present subject matter.

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 forward end; a tip portion joined to the forward end of the outer wall and extending forward toward the leading edge, the tip portion comprising: an inner layer positioned at a rear end of the tip portion; an outer layer defining a leading edge at a leading end of the tip portion; and one or more compliant layers positioned between the inner and outer layers for facilitating movement between the inner and outer layers.

2. The leading edge assembly according to any preceding clause, wherein the tip portion further comprises: an intermediate layer positioned between the inner layer and the outer layer, and wherein the one or more compliant layers comprise a first compliant layer positioned between the inner layer and the intermediate layer and a second compliant layer positioned between the intermediate layer and the outer layer.

3. The leading edge assembly according to any preceding clause, wherein the outer wall comprises a first wall section and a second wall section separated by a chamber, and wherein the inner layer, the intermediate layer, and the outer layer are each curved and extend between the first wall section and the second wall section.

4. The leading edge assembly according to any preceding clause, wherein the inner layer and the outer layer are made of different materials.

5. The leading edge assembly according to any preceding clause, wherein the outer layer and the inner layer are made of a metal, a ceramic material, or a ceramic matrix composite.

6. The leading edge assembly according to any preceding clause, wherein the outer layer and the inner layer have different densities.

7. The leading edge assembly according to any preceding clause, wherein at least two of the one or more compliant layers are made of different materials.

8. The leading edge assembly according to any preceding clause, wherein the inner and outer layers have a first coefficient of thermal expansion, and one or more of the compliant layers have a second coefficient of thermal expansion that is greater than the first coefficient of thermal expansion.

9. The leading edge assembly according to any preceding clause, wherein the at least one cooling channel is defined through the inner layer, the outer layer, and the one or more compliant layers.

10. The leading edge assembly according to any preceding clause, further comprising: a coolant supply in fluid communication with the at least one cooling channel for selectively providing a flow of coolant through the at least one cooling channel.

11. The leading edge assembly according to any preceding clause, wherein the leading edge assembly is positioned on a wing, nose cone, engine fairing, engine inlet, fuselage, or stabilizer of the hypersonic vehicle.

12. A leading edge assembly for a hypersonic vehicle, the leading edge assembly comprising: an outer wall tapering to a forward end; a tip portion joined to the forward end of the outer wall, the tip portion comprising a plurality of structural layers and a plurality of compliant layers alternately stacked with one another, wherein the plurality of compliant layers facilitate movement between the plurality of structural layers.

13. The leading edge assembly according to any preceding clause, wherein the outer wall comprises a first wall section and a second wall section separated by a chamber, and wherein the plurality of structural layers are curved and extend between the first wall section and the second wall section.

14. The leading edge assembly according to any preceding clause, wherein the plurality of structural layers are made of different materials.

15. The leading edge assembly according to any preceding clause, wherein the plurality of structural layers are made of a metal, ceramic material, or ceramic matrix composite material.

16. The leading edge assembly according to any preceding clause, wherein at least two of the plurality of structural plies have different densities.

17. The leading edge assembly according to any preceding clause, wherein at least two of the plurality of compliant layers are made of different materials.

18. The leading edge assembly according to any preceding clause, wherein the plurality of structural layers each have a first coefficient of thermal expansion and the plurality of compliant layers each have a second coefficient of thermal expansion, the second coefficient of thermal expansion being greater than the first coefficient of thermal expansion.

19. The leading edge assembly according to any preceding clause, wherein the at least one cooling channel is defined through the plurality of structural layers and the plurality of compliant layers.

20. The leading edge assembly according to any preceding clause, further comprising: a coolant supply in fluid communication with the at least one cooling channel for selectively providing a flow of coolant through the at least one cooling channel.

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.