System for reducing thermal stress in the leading edge of a high-speed vehicle

Abstract

The present invention relates to a system for reducing thermal stress in the leading edge of a high-speed vehicle. A hypersonic vehicle includes one or more leading edge assemblies designed to manage the thermal loads experienced at the leading edge during high-speed or hypersonic operations. The leading edge assembly includes a plurality of structural layers and a plurality of compliant layers stacked alternately with each other to facilitate thermal expansion and movement between the plurality of structural layers while also providing a thermal break between the plurality of structural layers.

Description

Technical Field

[0001] The present subject matter generally relates to leading edge technology for use in high-speed vehicles, such as hypersonic vehicles. Background Art

[0002] High-speed vehicles often experience thermal management issues caused by the high heat loads experienced during high-speed operation (particularly where the free air flow impinges on the leading edge of the vehicle). For example, in applications involving hypersonic vehicles, the leading edges may include the nose, engine cowlings, and the leading edges of wings and stabilizers. Especially when these vehicles operate in the hypersonic range (e.g., Mach 5 or greater), the leading edges may be subject to very high heat loads (e.g., 500–1500 W / cm 2 ), as the incident air flow passes through the bow shock and settles at the vehicle surface, converting the kinetic energy of the gas into internal energy and greatly increasing its temperature. Full exposure to such thermal loads can cause component degradation and / or failure.

[0003] Improvements in materials and manufacturing techniques have enabled hypersonic vehicles to operate at higher speeds and temperatures. For example, materials have been developed to increase the temperatures that components can withstand while maintaining their structural integrity. In this regard, for example, nickel-based superalloys can be used at 800°C, single crystal materials can be used at 1200°C, and refractory metals may be required for even higher temperatures. In addition, a variety of cooling techniques have been developed to provide cooling to the leading edge of a hypersonic vehicle. However, corresponding advances in vehicle speed and duration of high-speed flight time have created a need to further improve the cooling capacity and high-temperature durability of the leading edge of a high-speed vehicle.

[0004] Additionally, because thermal loads are greatest at the stagnation point or leading edge of a hypersonic vehicle, large temperature gradients may be generated within the structure defining the leading edge. For example, the outermost layers of the leading edge may experience the highest heat and the greatest thermal expansion, while the innermost layers may experience relatively low heat and less thermal expansion. As a result, thermal stresses may be generated within the leading edge, which may cause component failure or degradation.

[0005] Therefore, it would be useful to improve hypersonic vehicle and propulsion technology. More specifically, it would be particularly beneficial to improve thermal management techniques for reducing thermal stresses within the leading edge of a hypersonic vehicle. Summary of the invention

[0006] 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.

[0007] In an exemplary embodiment of the present disclosure, a leading edge assembly for a hypersonic vehicle includes: an outer wall that tapers toward a front end; a tip portion that is joined to the front end of the outer wall and extends forward toward the leading edge, the tip portion including: an inner layer that is positioned at the rear end of the tip portion; an outer layer that defines the leading edge at the front end of the tip portion; and one or more compliant layers that are positioned between the inner layer and the outer layer to facilitate movement between the inner layer and the outer layer.

[0008] According to another exemplary embodiment, a leading edge assembly for a hypersonic vehicle includes: an outer wall that tapers toward a front end; a tip portion that is joined to the front end of the outer wall, the tip portion including a plurality of structural layers and a plurality of compliant layers that are alternately stacked with each other, wherein the plurality of compliant layers facilitate movement between the plurality of structural layers.

[0009] Technical Solution 1. A leading edge assembly for a hypersonic vehicle, the leading edge assembly comprising:

[0010] an outer wall that tapers at a front end of the outer wall;

[0011] a tip portion joined to the front end of the outer wall and extending forward toward the front edge, the tip portion comprising:

[0012] an inner layer positioned at a rear end of the tip portion;

[0013] an outer layer defining the leading edge at a forward end of the tip portion; and

[0014] One or more compliant layers are positioned between the inner layer and the outer layer for facilitating movement between the inner layer and the outer layer.

[0015] Technical Solution 2. The leading edge assembly according to any of the above technical solutions, characterized in that the tip portion further comprises:

[0016] 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.

[0017] Technical Solution 3. A leading edge assembly according to any of the preceding technical solutions, characterized in that the outer wall comprises a first wall segment and a second wall segment separated by a chamber, and wherein the inner layer, the middle layer and the outer layer are each curved and extend between the first wall segment and the second wall segment.

[0018] Technical Solution 4. The leading edge assembly according to any of the foregoing technical solutions is characterized in that the inner layer and the outer layer are made of different materials.

[0019] Technical Solution 5. The leading edge assembly according to any of the foregoing technical solutions is characterized in that the outer layer and the inner layer are made of metal, ceramic material or ceramic matrix composite material.

[0020] Technical Solution 6. The leading edge assembly according to any of the foregoing technical solutions is characterized in that the outer layer and the inner layer have different densities.

[0021] Technical Solution 7. The leading edge assembly according to any of the preceding technical solutions, characterized in that at least two of the one or more compliant layers are made of different materials.

[0022] Technical Solution 8. The leading edge assembly according to any of the foregoing technical solutions is characterized in that the inner layer and the outer layer have a first thermal expansion coefficient, and the one or more compliant layers in the compliant layer have a second thermal expansion coefficient, which is greater than the first thermal expansion coefficient.

[0023] Technical Solution 9. The leading edge assembly according to any of the preceding technical solutions is characterized in that at least one cooling channel is defined through the inner layer, the outer layer and the one or more compliant layers.

[0024] Technical Solution 10. The leading edge assembly according to any of the above technical solutions, characterized in that it further comprises:

[0025] A coolant supply is in fluid communication with the at least one cooling passage for selectively providing a coolant flow through the at least one cooling passage.

[0026] Technical Solution 11. The leading edge assembly according to any of the foregoing technical solutions is characterized in that the leading edge assembly is positioned on the wing, nose cone, engine fairing, engine inlet, fuselage or stabilizer of the hypersonic vehicle.

[0027] Technical Solution 12. A leading edge assembly for a hypersonic vehicle, the leading edge assembly comprising:

[0028] an outer wall that tapers at a front end of the outer wall;

[0029] A tip portion is joined to the front end of the outer wall, the tip portion comprising a plurality of structural layers and a plurality of compliant layers alternately stacked on each other, wherein the plurality of compliant layers facilitate movement between the plurality of structural layers.

[0030] Technical Solution 13. The leading edge assembly according to any of the preceding technical solutions is characterized in that the outer wall comprises a first wall segment and a second wall segment separated by a cavity, and wherein the plurality of structural layers are curved and extend between the first wall segment and the second wall segment.

[0031] Technical Solution 14. The leading edge assembly according to any of the foregoing technical solutions is characterized in that the multiple structural layers are made of different materials.

[0032] Technical Solution 15. The leading edge assembly according to any of the foregoing technical solutions is characterized in that the multiple structural layers are made of metal, ceramic material or ceramic matrix composite material.

[0033] Technical Solution 16. The leading edge assembly according to any of the preceding technical solutions is characterized in that at least two of the multiple structural layers have different densities.

[0034] Technical Solution 17. The leading edge assembly according to any of the preceding technical solutions is characterized in that at least two of the plurality of compliant layers are made of different materials.

[0035] Technical Solution 18. The leading edge assembly according to any of the foregoing technical solutions is characterized in that each of the multiple structural layers has a first thermal expansion coefficient, and each of the multiple compliant layers has a second thermal expansion coefficient, and the second thermal expansion coefficient is greater than the first thermal expansion coefficient.

[0036] Technical Solution 19. The leading edge assembly according to any of the foregoing technical solutions is characterized in that at least one cooling channel is defined through the multiple structural layers and the multiple compliant layers.

[0037] Technical Solution 20. The leading edge assembly according to any of the above technical solutions, further comprising:

[0038] A coolant supply is in fluid communication with the at least one cooling passage for selectively providing a coolant flow through the at least one cooling passage.

[0039] 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 present invention and, together with the description, serve to explain the principles of the present invention. BRIEF DESCRIPTION OF THE DRAWINGS

[0040] A complete and enabling disclosure of the invention, including the best mode thereof, to one of ordinary skill in the art is set forth in the specification with reference to the accompanying drawings.

[0041] Figure 1 is a schematic close-up cross-sectional view of the leading edge of a hypersonic vehicle according to an exemplary embodiment of the present disclosure.

[0042] Figure 2 A method of cooling a leading edge and facilitating magnetohydrodynamic power generation or control of a hypersonic vehicle according to an exemplary embodiment of the present subject matter.

[0043] Figure 3 is a schematic close-up cross-sectional view of a leading edge of a hypersonic vehicle according to another exemplary embodiment of the present disclosure.

[0044] Figure 4 is a schematic close-up cross-sectional view of a leading edge of a hypersonic vehicle according to another exemplary embodiment of the present disclosure.

[0045] Figure 5 is a schematic close-up cross-sectional view of a leading edge of a hypersonic vehicle according to another exemplary embodiment of the present disclosure.

[0046] Figure 6 is a schematic close-up cross-sectional view of a leading edge of a hypersonic vehicle according to another exemplary embodiment of the present disclosure.

[0047] Figure 7 is a schematic close-up cross-sectional view of an outer wall of a leading edge of a hypersonic vehicle according to another exemplary embodiment of the present disclosure.

[0048] 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

[0049] Reference will now be made in detail to the present embodiments of the present 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 designations in the drawings and description have been used to refer to the same or similar parts of the present invention.

[0050] The word "exemplary" is used herein to mean "serving as an example, instance, or illustration." Any implementation described herein as "exemplary" is not necessarily to be construed as preferred or advantageous over other implementations. Moreover, each example is provided by way of explanation of the invention rather than limitation of the invention. Indeed, it will be apparent to those skilled in the art that various modifications and variations may be made in the invention without departing from the scope of the invention. For example, features illustrated or described as part of one embodiment may be used with another embodiment to produce yet another embodiment. It is intended, therefore, that the invention encompasses such modifications and variations as come within the scope of the appended claims and their equivalents.

[0051] As used herein, the terms "first," "second," and "third" may be used interchangeably to distinguish one component from another and are not intended to indicate the position or importance of individual components. Unless the context clearly dictates otherwise, the singular forms "a," "an," and "the" include plural references. Unless otherwise specified herein, the terms "coupled," "fixed," "attached to," and the like refer to both direct coupling, fixing, or attachment and indirect coupling, fixing, or attachment through one or more intermediate components or features.

[0052] The terms "front" and "rear" refer to relative positions within a gas turbine engine or vehicle, and refer to the normal operating attitude of the gas turbine engine or vehicle. For example, with respect to a gas turbine engine, "front" refers to a position closer to the engine inlet, and "rear" refers to a position closer to the engine nozzle or exhaust. The terms "upstream" and "downstream" refer to relative directions relative to the flow of a fluid in a fluid passage. For example, "upstream" refers to the direction from which the fluid is flowing, and "downstream" refers to the direction to which the fluid is flowing.

[0053] Approximate language, as used throughout the specification and claims herein, is applicable to modify any quantitative representation that can be allowed to vary without causing a change in the basic function associated therewith. Therefore, a value modified by one or more terms such as "about", "approximately", and "substantially" will not be limited to the exact value specified. In at least some instances, the approximate language may correspond to the accuracy of an instrument used to measure the value, or the accuracy of a method or machine used to construct or manufacture a component and / or system. For example, approximate language may refer to being within a 10 percent margin.

[0054] Here and throughout the specification and claims, range limitations are combined and interchanged, and unless context or language indicates otherwise, such ranges are identified and include all sub-ranges contained therein. For example, all ranges disclosed herein are inclusive of the endpoints, and the endpoints are independently combinable with each other.

[0055] 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 speeds 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, but may instead be applied to applications involving other high-speed vehicles, projectiles, objects, etc. The description of the use of leading edge assemblies on hypersonic aircraft herein is merely exemplary, 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.

[0056] It is noteworthy that, as explained above, high-speed vehicles such as hypersonic vehicles typically experience extremely high temperatures and thermal gradients during high-speed or hypersonic operations. The temperature gradient caused by high heat flux is usually a more serious problem than the temperature itself. For example, the thermal conductivity of the structural material combines with the heat flux to set the temperature gradient within the material, and under high thermal loads, this gradient causes 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.

[0057] As explained above, the leading edge of such a high-speed vehicle is usually subject to the highest thermal load. For example, a hypersonic vehicle may include a plurality of leading edge assemblies (e.g., generally identified herein by reference numeral 300), which experience high thermal loads during hypersonic flight. In this regard, the leading edge assembly 300 may be arranged on the front end of an aircraft wing, a nose cone, a vertical stabilizer, an engine cowling of a propulsion engine, or other leading edges or surfaces of a hypersonic aircraft. According to an exemplary embodiment of the present theme, the leading edge assembly 300 includes features for alleviating the effects of such thermal loads (e.g., by taking away heat from the region).

[0058] It is noted that it is typically desirable to make the leading edge assembly 300 as sharp or pointed as possible, for example, in order to reduce drag on a hypersonic vehicle. Figure 1 , when the leading edge assembly 300 is formed with a small radius of curvature, extremely high temperatures and thermal gradients are experienced within the leading edge assembly 300 at its front portion or leading edge (also referred to herein as a stagnation line, stagnation point 302, or similar terms). In this regard, as the hypersonic vehicle travels through the air at hypersonic speeds, a free stream of air (e.g., identified herein by reference numeral 304) passes over and around the leading edge assembly 300, thereby generating large heat loads. Aspects of the present subject matter relate to thermal management techniques and features for cooling the leading edge assembly 300, with particular focus on the region proximate the stagnation point 302, where the most severe thermal management issues typically occur.

[0059] 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 technology may vary while remaining within the scope of this theme. For example, the leading edge assembly 300 described herein is limited to a radius between approximately 1 mm and 3 mm. However, according to alternative embodiments, the leading edge assembly may have any other suitable radius.

[0060] Cooling techniques and thermal management features are described herein as being used to cool portions of one or more parts of a hypersonic vehicle, such as the leading edge of a wing, a nose, a propulsion engine, or other parts of a hypersonic vehicle that experience large temperatures and thermal gradients. However, it should be appreciated that aspects of the present subject matter can 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 can be applied to any other hypersonic vehicle or any other technology or system with a component that is exposed to high temperatures and / or large temperature gradients.

[0061] 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 may be made to such techniques without departing from the scope of the present subject matter. Additionally, one or more 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 in isolation in order to clearly describe how each technique works, the described embodiments are merely examples intended for illustration and explanation purposes and are not intended to limit the scope of the present subject matter in any way.

[0062] In addition, 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 a hypersonic vehicle, such as leading edge assembly 300, to be integrally formed as a single, unitary component, or formed into any suitable number of subcomponents. As used herein, the term "additive manufacturing" or "additive manufacturing techniques or processes" generally refers to a manufacturing process in which successive layers of (one or more) materials are disposed one above the other to "build" a three-dimensional component layer by layer. Successive layers are typically fused together to form a unitary component, which may have a variety of integral subcomponents.

[0063] Although additive manufacturing techniques are described herein as enabling the manufacture of complex objects by typically building objects point by point, layer by layer in a vertical direction, other methods of manufacture are possible and within the scope of the present subject matter. For example, although the discussion herein relates to adding materials to form successive layers, those skilled in the art will recognize that the methods and structures disclosed herein may be practiced using any additive manufacturing technique or manufacturing technology. For example, embodiments of the present invention may use an additive layer process, a subtractive layer process, or a hybrid process.

[0064] 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 jetting and binder jetting, stereolithography (SLA), direct selective laser sintering (DSLS), electron beam sintering (EBS), electron beam melting (EBM), laser engineered net shaping (LENS), laser net shape manufacturing (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.

[0065] The additive manufacturing process described herein can be used to form a component using any suitable material. For example, the material can be metal, concrete, ceramic, epoxy resin or any other suitable material, which can be solid, liquid, powder, sheet, wire or any other suitable form or combination thereof. More specifically, according to exemplary embodiments of the present subject matter, the component of the additive manufacturing described herein can be formed in part, in its entirety or in some combination by materials including but not limited to the following: pure metal, nickel alloy, chromium alloy, titanium, titanium alloy, magnesium, magnesium alloy, aluminum, aluminum alloy and nickel-based or cobalt-based superalloy (for example, the name Inconel® superalloy available from Special Metals Corporation). These materials are examples of materials suitable for use in the additive manufacturing process described herein, and can be generally referred to as "incremental materials".

[0066] In addition, the additive manufacturing process disclosed herein allows a single component to be formed by a variety of materials. Therefore, the components described herein can be formed by any suitable mixture of the above-mentioned materials. For example, the component may include multiple layers, sections or parts formed using different materials, processes and / or on different additive manufacturing machines. In this way, components with different materials and material properties can be constructed to meet the needs of any specific application. In addition, although the components described herein are constructed entirely by additive manufacturing processes, it should be appreciated that in alternative embodiments, all or part of these components can be formed via casting, machining and / or any other suitable manufacturing process. In fact, any suitable combination of materials and manufacturing methods can be used to form these components.

[0067] Still reference Figure 1 , the leading edge assembly 300 will be described in more detail according to an exemplary embodiment of the present subject matter. Specifically, Figure 1A cross-sectional view of a leading edge assembly 300 is provided, which can be positioned at the leading edge (e.g., front end, leading end, upstream end, etc.) of any component of a hypersonic vehicle. For example, the leading edge assembly 300 can 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, the nose of an aircraft, the front end of a vertical stabilizer, etc.

[0068] As explained herein, during hypersonic flight operations, the leading edge assembly 300 may experience large thermal loads. As used herein, the terms "thermal loads" and the like are generally intended to refer to the high temperatures, temperature gradients, or heat fluxes experienced within components of a hypersonic or high-speed vehicle. According to an exemplary embodiment of the present subject matter, the leading edge assembly 300 is formed or provided with thermal regulation features or techniques for managing these thermal loads.

[0069] For example, as referenced below Figure 1 In more detail, the leading edge assembly 300 may include features for providing a coolant, such as a cooling fluid or cooling material, through the outer wall 320 of the leading edge assembly 300 to reduce the temperature of the outer wall 320 and / or temperature gradients experienced within the leading edge assembly 300 . Figures 2 to 5 Cooling techniques for leading edge assemblies 300 according to alternative embodiments are provided. It should be appreciated that the thermal regulation features and techniques described herein for each exemplary leading edge assembly 300 may be used alone, or in combination with any other leading edge techniques described herein, to regulate thermal loads on one or more leading edge assemblies 300 of a hypersonic vehicle, or on any other surface of any other component that experiences high thermal loads.

[0070] As explained above, the outer wall 320 and other components of the leading edge assembly 300 may be formed by any suitable material. According to an exemplary embodiment, such a material is selected to withstand the high thermal load experienced by the leading edge of the hypersonic vehicle. For example, the outer wall 320 may be constructed by at least one of aluminum, titanium, titanium aluminide, tungsten, tungsten alloy, nickel superalloy, refractory metal, single crystal metal, ceramic, ceramic matrix composite (CMC), ultra-high temperature ceramic (UHTC, including high melting point diboride, nitride, etc.) or carbon-carbon composite. In addition, or alternatively, the outer wall 320 may include composite materials such as silicon carbide (SiC), SiC composites, carbon fiber reinforced SiC matrix and other carbide matrix composites, composite materials with and without surface coatings, and / or high entropy alloys including refractory materials, platinum group metals, hafnium alloys, etc. However, it may still be desirable in some applications to provide additional cooling capabilities for thermal management of the high thermal load experienced by the leading edge assembly 300. Furthermore, as explained above, additive manufacturing techniques may be used to print the leading edge assembly 300 (eg, including the outer wall 320 ) as a single, unitary component and may facilitate improved cooling techniques and leading edge features.

[0071] As shown in the depicted embodiment, the outer wall 320 is generally formed by a first wall segment 322 and a second wall segment 324. More specifically, the first wall segment 322 and the second wall segment 324 each include an outer surface that together form an outer surface 326 and an inner surface that together form an inner surface 328. In addition, the first wall segment 322 and the second wall segment 324 can be angled relative to each other so that the leading edge assembly 300 tapers from the rear end 330 of the leading edge assembly 300 to the front 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 taller near the rear end 330 of the leading edge assembly 300 and narrows as it approaches the stagnation point 302. It is worth noting that the taper angle can vary depending on aerodynamics 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, for example, defining a more acute angle on one side.

[0072] As described above, for the illustrated embodiment, the outer surfaces 326 of the first wall segment 322 and the second wall segment 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 a portion 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 segment 322 and the second wall segment 324 and in fluid communication with the inner surface 328 of the leading edge portion 340. Additionally, according to an exemplary embodiment, the leading edge portion 340 defines at least one fluid channel 350 through the outer wall 320.

[0073] 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 forward end 332 of the leading edge assembly 300. Although Figure 1 , but 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, such as along the span of an aircraft wing, for example.

[0074] In practice, 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 is Figure 1 The circumferential direction C generally circumscribes the leading edge assembly 300 from the first wall segment 322 to the second wall segment 324, as shown in FIG. Figure 1 In addition, the leading edge assembly 300 can define a longitudinal direction L, which, according to exemplary embodiments described herein, corresponds to the longitudinal direction L of the hypersonic vehicle and is substantially parallel to the angle of attack defined between the leading edge assembly 300 and the main direction of the air flow 304.

[0075] According to the illustrated embodiment, the fluid passage 350 extends through the outer wall 320 along the radial direction R. However, it should be appreciated that according to alternative embodiments, the fluid passage 350 may pass through the outer wall 320 at an angle other than the radial direction R, may pass along the length of the first wall segment 322 and / or the second wall segment 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 the cooling fluid 360 to a desired location within or on the leading edge assembly 300.

[0076] In this regard, the leading edge assembly 300 may further include a coolant supply 362 for providing a flow of a cooling fluid 360 through the fluid passage 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. According to an embodiment of the present subject matter, as described in more detail below, the cooling fluid 360 may include at least one of air, water, steam, liquid metal, carbon dioxide, argon, helium, and the like.

[0077] It should be further 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, etc. Moreover, the present subject matter is not limited by the source of the 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 channels, evaporates to absorb thermal energy, and carries away thermal energy as it flows downstream.

[0078] According to an exemplary embodiment 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 through the leading edge assembly 300 as the cooling fluid 360 may include at least one of cesium (Cs, which has a boiling point of approximately 678° C.), potassium (K, which has a boiling point of approximately 774° C.), sodium (Na, which has a boiling point of approximately 883° C.), lithium (Li, which has a boiling point of approximately 1347° C.), indium (In, which has a boiling point of approximately 2000° C.), and / or gallium (Ga, which has a boiling point of approximately 2205° C.). Additionally, or alternatively, the cooling fluid 360 may include ammonia (NH3) or sulfur dioxide (SO2). According to an exemplary embodiment, the cooling fluid 360 may be a mixture or combination of such liquid metals or compositions.

[0079] In addition, the cooling fluid 360 can be selected based on its material properties such as ionization potential. The ionization potential of the cooling fluid 360 can be important (as explained below) for facilitating the use of the leading edge assembly 300 for magnetohydrodynamic control or power generation. For example, according to an exemplary embodiment, the liquid metal cooling fluid 360 can 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 so that its ionization potential is much lower than the ionization potential of air. For example, the ionization potential of O2 is approximately 12.6 eV, and the ionization potential of N2 is approximately 14.5 eV.

[0080] These materials can be used as "seeds" that ionize at temperatures much lower than air. It should be appreciated that the term "ionized" and the like are generally used herein to refer to both complete ionization and partial ionization. For example, the term ionized can refer to partial ionization or weak ionization, e.g., where the density of ions is a small fraction (e.g., 10 times) of the density of neutral species. -5 ). In addition, it will be appreciated that if the temperature is high enough (eg, due to a high Mach number), not only the seed material may be ionized, but also any material present (including oxygen and nitrogen).

[0081] According to alternative embodiments, the cooling fluid 360 may be selected based on another material property, such as a 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 way, 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 way, the evaporation of the cooling fluid 360 on the outer surface 326 can maintain the temperature of the outer wall 320 at the boiling point or boiling temperature of the cooling fluid 360. If the boiling point is selected to be lower than the melting temperature of the outer wall 320, the structural integrity of the outer wall 320 can be maintained until the cooling fluid 360 has been exhausted. For example, according to one embodiment, the liquid metal cooling fluid 360 has a boiling temperature greater than about 500° C., greater than about 700° C., greater than about 900° C., greater than about 1100° C., or about 750° C. Additionally, or alternatively, the boiling temperature of the liquid metal cooling fluid 360 may be less than about 2000° C., less than about 1500° C., less than about 1000° C., or less than about 800° C. For example, according to an exemplary embodiment, the cooling fluid 360 is lithium (Li), which boils at about 1342° C.

[0082] As used herein, the terms "boiling point," "boiling temperature," and the like refer to temperatures tabulated at an ambient pressure of 1 atmosphere. However, it should be appreciated that such boiling temperatures may be different under flight conditions (e.g., when the ambient pressure and the experienced pressure are different). Additionally, while 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 degrade or weaken.

[0083] It should be appreciated that the cooling fluid 360 described herein is exemplary only and is not intended to limit the scope of the present subject matter. For example, the cooling fluid 360 may be selected as a liquid or solid, may have a varying boiling point and ionization potential, 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 evaporation may be important for determining the volume and mass of the required cooling fluid 360, for example, lithium (Li) has a unique high specific heat of evaporation (per unit mass). According to an exemplary embodiment, the cooling fluid 360 may be selected to melt at a relatively low melting temperature, for example, to minimize the complexity required for an onboard material handling system.

[0084] 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 phase or a gas phase when exposed to temperatures generated during operation of the hypersonic propulsion engine during hypersonic flight operations. 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 operations. However, in other embodiments, other suitable coolants may be utilized.

[0085] like Figure 1 350 for selectively providing a flow of coolant 360 through the fluid channel 350. Specifically, according to the illustrated embodiment, the coolant supply source 362 includes a coolant reservoir 364, and a mechanical pump 366 is configured to pressurize or supply the cooling fluid 360 to the fluid channel 350. According to still other embodiments, the cooling fluid 360 may be stored in 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, the cooling fluid 360 may be delivered from any suitable location in any other suitable manner. For example, as illustrated, the coolant supply source 362 generally provides a flow of pressurized cooling fluid 360 to the chamber 342 in fluid communication with the fluid channel 350. However, it should be appreciated that according to alternative embodiments, the coolant supply source 362 may include one or more conduits directly fluidly coupled to the fluid channel 350. Furthermore, although 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, etc.

[0086] As explained above, according to an exemplary embodiment of the present subject matter, the flow of cooling fluid 360 can include liquid metal. Notably, when the liquid metal cooling fluid 360 is exposed to such high temperatures at the leading edge assembly 300, it can ionize after exiting the fluid passage 350 to produce an ionized vapor flow (e.g., as generally described by Figure 1 370). The ionized vapor flow 370 is transmitted downstream from the stagnation point 302 on the leading edge assembly 300. According to an exemplary embodiment, the ionized vapor flow 370 can facilitate a magnetohydrodynamic process for power generation and / or vehicle control. Specifically, according to an exemplary embodiment, the leading edge assembly 300 can include a magnetohydrodynamic generator 372 ("MHD generator") that interacts with the ionized vapor flow 370 to generate power or thrust for vehicle control.

[0087] In general, the MHD generator 372 may be any device suitable for interacting with the ionized vapor flow 370 to generate electricity or thrust for vehicle control. It will be appreciated that the MHD generator 372 may generally utilize a magnetohydrodynamic converter that utilizes a Brayton cycle to convert thermal and kinetic energy directly into electricity. For example, the MHD generator 372 may include one or more magnetic coils and / or electrodes positioned near or within the outer wall 320. According to an exemplary embodiment, the MHD generator 372 may use the coils to generate electricity when the ionized vapor flow 360 passes through the magnetic field generated by the coils. Additionally, or alternatively, the MHD generator 372 may energize the magnetic coils or electrodes to interact with the ionized vapor flow 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 a unique structure for implementing a control method for steering during high-speed flight, such as during hypersonic flight.

[0088] Specifically, Figure 2 As best illustrated in FIG. 4 , the method of operating the leading edge assembly 300 may include a magnetohydrodynamic control method. Specifically, the method 400 includes providing a coolant flow through at least one fluid channel defined by a leading edge of a hypersonic vehicle at step 410, 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 transmitted downstream as an ionized vapor flow. Once the ionized vapor flow is achieved, step 420 includes using a magnetohydrodynamic generator that interacts with the ionized vapor flow to generate power or thrust for vehicle control.

[0089] Additionally, or alternatively, if Figure 3 and Figure 4500. As best shown in FIG. 5 , 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 to the inner surface 328 of the leading edge portion 340 through the chamber 342, so that the flow of cooling fluid 360 can permeate 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 operations). In this regard, the fluid channel 350 may be defined through the porous structure that constitutes or defines the porous tip 500. When describing the figures herein, it should be recognized that the same reference numerals may be used to refer to the same or similar features between embodiments.

[0090] Specifically, Figure 3 and Figure 4 , the outer wall 320 of the leading edge assembly 300 includes a first wall segment 322 and a second wall segment 324, each extending from a front end 502 and a rear end (e.g., which may correspond to the rear end 330 of the leading edge assembly 300) along a longitudinal direction L. Notably, the front ends 502 of the first wall segment 322 and the second wall segment 324 may stop before the stagnation point 302. Additionally, the porous tip 500 may be coupled to the front ends 502 of the first wall segment 322 and the second wall segment 324. In this regard, the porous tip 500 may be generally defined as a curved region coupled to the first wall segment 322 and the second wall segment 324, which are illustrated as being substantially straight.

[0091] According to an exemplary embodiment, for example, the first wall segment 322, the second wall segment 324, and the tip portion 500 may be additively manufactured simultaneously as a single, integral, unitary component. However, as will be described in more 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 front end 332 of the leading edge assembly 300.

[0092] According to an exemplary embodiment, the leading edge portion 340, or more specifically the porous tip 500, may define a constant porosity for passing the cooling fluid 360. As used herein, the term "porosity" may be generally used to refer to a measure of voids or empty spaces within a material or structure. Thus, a structure having porosity has open channels, cells, or structures through which fluids can flow from one porous cell to another. For example, porosity may be used to refer to the fraction of the volume of the voids or open spaces divided by the total volume of the component. According to an exemplary embodiment, the porosity of the porous tip 500 may be greater than about 5%, 10%, 20%, 40%, or greater than or 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 rate of cooling fluid 360 , the mechanical properties of porous tip 500 , based on expected flight conditions, or based on any other suitable parameter.

[0093] It is noteworthy that, according to the illustrated embodiment, the porous tip may define a variable porosity, for example, in order to concentrate 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 front 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 the 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 as or different from the second porosity, but also less than the first porosity. For example, according to an exemplary embodiment, 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.

[0094] Similar to the embodiments described above, a coolant supply source 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 chamber 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 can be adjusted by manipulating the porosity within each region 510-514. According to the illustrated embodiment, it is desirable to have the highest porosity in the first porous region 510 to direct the largest amount of cooling fluid 360 toward the region with the greatest thermal impact (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 toward the stagnation point 302 .

[0095] It is noteworthy that the porous tip 500 is illustrated as having three different porous regions 510-514. However, it should be appreciated that by using the additive manufacturing techniques described herein, the porous tip 500 can have a gradually varying porosity, i.e., such that the porosity increases continuously and gradually from the front end 502 of the outer wall 320 to the highest porosity at the stagnation point 302. In this manner, the porous tip 500 can be envisioned as having 10, 20, 50, 100, or even more sub-regions, each of which has a gradually increasing porosity as it approaches the stagnation point 302. As illustrated, each of these sub-regions extends from the inner surface 328 to the outer surface 326 along the radial direction R, and has a substantially constant porosity along the radial direction R. In addition, according to alternative embodiments, one or more of these regions (such as the porous regions 510-514) can also vary in porosity along the radial direction R.

[0096] According to an exemplary embodiment, 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, reference Figure 4, the porous tip 500 may include a plurality of internal barriers 520 that separate one or more of the porous regions 510-514. In this regard, as illustrated, the internal barriers 520 are straight and solid walls that extend 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 internal barriers 520, the internal barriers 520 may extend through the outer wall 320 in other directions, may have varying thicknesses, etc. Other configurations are possible and within the scope of the present subject matter.

[0097] Now specifically refer to Figure 4 , the leading edge assembly 300 may include a nose shield 530, which is at least partially positioned on the porous tip 500 to restrict the flow of the cooling fluid 360 from the coolant reservoir 364 and / or the chamber 342. Specifically, the nose shield 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, the nose shield 530 may be constructed of a material that melts before the porous tip 500 and the outer wall 320 reach their melting temperatures. Notably, when the nose shield 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. Therefore, according to an exemplary embodiment, the predetermined critical temperature at which the nose shield 530 melts is below the temperature at which the structural integrity of the leading edge assembly 300 begins to degrade. Although the nasal mask 530 is illustrated as being used with a porous tip 500 having a variable porosity, it should be appreciated that these two features may be used together or independently of one another.

[0098] According to the illustrated embodiment, the nose shield 530 extends over the entire porous tip 500, e.g., from the front end 502 of the first wall segment 322 to the front end 502 of the second wall segment 324 along the circumferential direction C. Notably, according to an exemplary embodiment, the nose shield 530 is impermeable to the flow of the cooling fluid 360. Thus, the cooling fluid 360 is contained within the porous tip 500, the coolant reservoir 364, and / or the chamber 342 until the nose shield 530 melts, which corresponds to the time required to cool the leading edge assembly 300. According to alternative embodiments, the nose shield 530 may cover less than the entire portion of the porous tip 500. For example, the nose shield 530 may only cover the front end 332 of the first porous region 510. Alternatively, the nose shield 530 may only cover the second porous region 512 and the third porous region 514. According to still other embodiments, the thickness of the nose mask 530 (which may be measured substantially along the radial direction R) may vary depending on the circumferential position along the porous tip 500. According to still other embodiments, the nose mask 530 may be a plug positioned at the outlet 354 of the fluid passage 350 (e.g., as shown in FIG. Figure 1 ). In this regard, for example, the nose mask 530 can extend at least partially within the fluid passage 350. Other variations and modifications can be made to the nose mask 530 while remaining within the scope of the present subject matter. For example, it should be appreciated that the nose mask 530 can be made by filling the holes with different materials to plug the holes and form the nose mask 530, by attaching a separate cap element to the leading edge assembly 300, etc.

[0099] Reference now Figure 5 , the leading edge assembly 300 may include a rear baffle 532 positioned at the rear end 330 of the outer wall 320 and extending substantially perpendicular to the longitudinal direction L. As shown, the chamber 342 is enclosed and defined between the rear baffle 532, the first wall segment 322, the second wall segment 324, and the nose mask 530. Therefore, the chamber 342 can be a chamber or reservoir of constant volume. For example, the rear baffle 532 can be connected to the rest of the hypersonic vehicle, can be a replaceable component, etc. According to such an exemplary embodiment, the chamber 342 can be filled with a cooling fluid 360. Notably, as the leading edge assembly 300 heats up during hypersonic operation, the nose mask 530 can slowly melt while the pressure within the chamber 342 increases due to the increased temperature of the cooling fluid 360. Once the temperature of the leading edge assembly 300 has reached the critical temperature, the nose shield 530 melts away and the chamber 342 is pressurized sufficiently to drive the flow of the cooling fluid 360 through the porous tip 500 , thereby cooling the leading edge assembly 300 .

[0100] According to still other embodiments, porous tip 500 may be filled with material that seeps out of porous tip 500 when nose mask 530 melts or ablates away. More specifically, according to an exemplary embodiment, porous tip 500 may be filled with metal cooling fluid 360 (e.g., in solid form), which may have a relatively low melting point so that the metal filling the holes of porous tip 500 is configured to melt during operation of a hypersonic propulsion engine or hypersonic aircraft during high temperature operations (such as hypersonic flight operations). Once the metal cooling fluid 360 filling the holes of porous tip 500 is melted, cooling fluid 360 may be heated in the same manner as described above with reference to Figure 4 Flow through porous tip 500 in a similar manner as described.

[0101] Now specifically refer to Figure 6 and Figure 7, a leading edge assembly 300 will be described according to another exemplary embodiment of the present subject matter. As described below, the leading edge assembly 300 can be formed of one or more walls that include a layered multifunctional material having a compliant interface that is used to mitigate the extreme thermal stresses experienced by the hypersonic vehicle at a high heat flux location (e.g., near the stagnation point 302). As will be described below, each of the multiple layers within the layered wall can be made of the same material (e.g., such as a metal or ceramic material) or can be made of a different material. In addition, each layer can be customized to meet specific requirements for a given application, for example, based on expected heat loads, expected temperature gradients, the presence of localized cooling features, etc.

[0102] As explained above, the leading edge assembly 300 may include a variety of cooling techniques, such as techniques to provide a radiative cooling feature, which may include a cooling fluid 360 positioned within the cavity 342. Notably, in such an embodiment, the greatest temperatures may be experienced on the outer surface 326 near the stagnation point 302, and the lowest temperatures 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).

[0103] like Figure 6 , the leading edge assembly 300 may include a tip portion 550 (e.g., which may correspond to the leading edge portion 340) that is joined to the front end 502 of the outer wall 320 or otherwise positioned near the front 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 that are alternately stacked with each other. In this manner, the plurality of compliant layers 554 may facilitate some movement between the plurality of structural layers 552 and may also provide a thermal break or insulating gap between the plurality of structural layers 552. According to the illustrated embodiment, the 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.

[0104] Although layers 552, 554 are described herein as "structural" and "compliant," it should be appreciated that these terms are used merely to distinguish between various layers and are not intended to limit such layers to specific materials or material properties. In addition, 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, for example, 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 to otherwise indicate or require a precisely layered or layered construction. For example, the "layers" of the leading edge assembly 300 may all be additively manufactured as a single, integral piece using one or more materials having different densities, porosities, coefficients of thermal expansion, or other different material properties.

[0105] 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, ceramic matrix composite, or any other suitable material described herein. Additionally, 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 an alternative embodiment, the plurality of structural layers 552 may have different materials, compositions, or configurations for optimally managing the thermal loads 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 materials used to form the plurality of structural layers 552.

[0106] For example, now specifically refer to Figure 7 , a close-up view of a tip portion 550 will be described according to 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 the rear end of the tip portion 550. In this regard, the inner layer 560 can at least partially define the inner surface 328 and the cavity 342 of the leading edge assembly 300. In addition, the tip portion 550 includes an outer layer 562 that defines the leading edge or stagnation point 302 at the front end 332 of the leading edge assembly 300.

[0107] Between the inner layer 560 and the outer layer 562, the distal portion 550 may include one or more compliant layers 554. For example, the compliant layer 554 may facilitate movement between the inner layer 560 and the outer layer 562. More specifically, as shown in FIG. Figure 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 intermediate 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.

[0108] 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 vehicle, or in other applications where high thermal loads are expected. Additionally, while the structural layer 552 and the compliant layer 554 are illustrated as being curved between the first wall segment 322 and the second wall segment 324, these layers 552, 554 may take any other suitable shape, size, or position for facilitating improved management of thermal stresses.

[0109] 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 stresses induced within the material layer, the thermal conductivity of the layer may be adjusted, etc. Additionally, 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 way, it may be desirable to form the outer layer 562 to have a lower coefficient of thermal expansion than the inner layer 560, for example, to reduce the difference in expansion experienced between the two layers when exposed to different temperatures during hypersonic operation.

[0110] 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 for distributing 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 the 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 certain exemplary embodiments, the plurality of compliant layers 554 may define a smaller thickness 570, and the thickness of the material between the plurality of compliant layers 554 may be less than or equal to about 1 millimeter.

[0111] It should be further recognized that Figure 6 and Figure 7 The described layered tip portion may be used with other techniques described herein to minimize the effects of 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 from Figure 1350) 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 source, such as coolant supply source 362, for providing a flow of cooling fluid 360 to the tip portion 550. In addition, the tip portion 550 may be divided into sub-regions similar to the porous tip 500, wherein each region has 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.

[0112] Further aspects of the invention are provided by the subject matter of the following clauses:

[0113] 1. A leading edge assembly for a hypersonic vehicle, the leading edge assembly comprising: an outer wall that tapers toward a front end; a tip portion that is joined to a front end of the outer wall and extends forward toward the leading edge, the tip portion comprising: an inner layer that is positioned at a rear end of the tip portion; an outer layer that defines a leading edge at a front end of the tip portion; and one or more compliant layers that are positioned between the inner layer and the outer layer to facilitate movement between the inner layer and the outer layer.

[0114] 2. A 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.

[0115] 3. A leading edge assembly according to any preceding clause, wherein the outer wall comprises a first wall segment and a second wall segment 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 segment and the second wall segment.

[0116] 4. A leading edge assembly according to any preceding clause, wherein the inner layer and the outer layer are made of different materials.

[0117] 5. A leading edge assembly according to any preceding clause, wherein the outer layer and the inner layer are made of metal, ceramic material or ceramic matrix composite material.

[0118] 6. A leading edge assembly according to any preceding clause, wherein the outer layer and the inner layer have different densities.

[0119] 7. A leading edge assembly according to any preceding clause, wherein at least two of the one or more compliant layers are made of different materials.

[0120] 8. A 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, the second coefficient of thermal expansion being greater than the first coefficient of thermal expansion.

[0121] 9. A leading edge assembly according to any preceding clause, wherein at least one cooling passage is defined through the inner layer, the outer layer and the one or more compliant layers.

[0122] 10. A leading edge assembly according to any preceding clause, further comprising: a coolant supply in fluid communication with the at least one cooling passage for selectively providing a coolant flow through the at least one cooling passage.

[0123] 11. A leading edge assembly according to any preceding clause, wherein the leading edge assembly is positioned on a wing, nose cone, engine cowling, engine inlet, fuselage or stabilizer of a hypersonic vehicle.

[0124] 12. A leading edge assembly for a hypersonic vehicle, the leading edge assembly comprising: an outer wall that tapers toward a front end; a tip portion that is joined to the front end of the outer wall, the tip portion comprising a plurality of structural layers and a plurality of compliant layers that are alternately stacked with each other, wherein the plurality of compliant layers facilitate movement between the plurality of structural layers.

[0125] 13. A 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.

[0126] 14. A leading edge assembly according to any preceding clause, wherein the plurality of structural layers are made of different materials.

[0127] 15. A leading edge assembly according to any preceding clause, wherein the plurality of structural layers are made of metal, ceramic material or ceramic matrix composite material.

[0128] 16. A leading edge assembly according to any preceding clause, wherein at least two of the plurality of structural layers have different densities.

[0129] 17. A leading edge assembly according to any preceding clause, wherein at least two of the plurality of compliant layers are made of different materials.

[0130] 18. A 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.

[0131] 19. A leading edge assembly according to any preceding clause, wherein at least one cooling channel is defined through the plurality of structural layers and the plurality of compliant layers.

[0132] 20. A leading edge assembly according to any preceding clause, further comprising: a coolant supply in fluid communication with the at least one cooling passage for selectively providing a coolant flow through the at least one cooling passage.

[0133] 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 that do not differ substantially from the literal language of the claims.

Claims

1. A leading edge assembly for a hypersonic vehicle, the leading edge assembly include: an outer wall that tapers at a front end of the outer wall; a tip portion joined to the front end of the outer wall and extending forward toward the front 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 forward 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, and The inner layer and the outer layer have a first coefficient of thermal expansion, and the one or more compliant layers have a second coefficient of thermal expansion, which is greater than the first coefficient of thermal expansion.

2. The leading edge assembly according to claim 1, It is characterized in that The distal 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 according to claim 2, It is characterized in that The outer wall includes a first wall segment and a second wall segment 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 segment and the second wall segment.

4. The leading edge assembly according to claim 1, It is characterized in that The inner layer and the outer layer are made of different materials.

5. The leading edge assembly according to claim 1, It is characterized in that The outer layer and the inner layer are made of metal, ceramic material or ceramic matrix composite material.

6. The leading edge assembly according to claim 1, It is characterized in that The outer layer and the inner layer have different densities.

7. The leading edge assembly according to claim 1, It is characterized in that At least two of the one or more compliant layers are made of different materials.

8. The leading edge assembly according to claim 1, It is characterized in that At least one cooling channel is defined through the inner layer, the outer layer, and the one or more compliant layers.

9. The leading edge assembly according to claim 8, It is characterized in that Further including: A coolant supply is in fluid communication with the at least one cooling passage for selectively providing a coolant flow through the at least one cooling passage.

10. The leading edge assembly according to claim 1, It is characterized in that The leading edge assembly is positioned on a wing, nose cone, engine cowling, engine inlet, fuselage, or stabilizer of the hypersonic vehicle.

11. A leading edge assembly for a hypersonic vehicle, the leading edge assembly include: an outer wall that tapers at a front end of the outer wall; A tip portion joined to the front end of the outer wall, the tip portion comprising a plurality of structural layers and a plurality of compliant layers alternately stacked with each other, wherein the plurality of compliant layers facilitate movement between the plurality of structural layers, and 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.

12. The leading edge assembly according to claim 11, It is characterized in that The outer wall includes a first wall segment and a second wall segment separated by a cavity, and wherein the plurality of structural layers are curved and extend between the first wall segment and the second wall segment.

13. The leading edge assembly according to claim 11, It is characterized in that The plurality of structural layers are made of different materials.

14. The leading edge assembly according to claim 11, It is characterized in that The plurality of structural layers are made of metal, ceramic material or ceramic matrix composite material.

15. The leading edge assembly according to claim 11, It is characterized in that At least two of the plurality of structural layers have different densities.

16. The leading edge assembly according to claim 11, It is characterized in that At least two of the plurality of compliant layers are made of different materials.

17. The leading edge assembly according to claim 11, It is characterized in that At least one cooling channel is defined through the plurality of structural layers and the plurality of compliant layers.

18. The leading edge assembly according to claim 17, It is characterized in that Further including: A coolant supply is in fluid communication with the at least one cooling passage for selectively providing a coolant flow through the at least one cooling passage.

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