System and method for removing heat from aircraft components

Abstract

A system for removing heat from an aircraft component includes: an aircraft component; a heat exchanger proximate the aircraft component and having an inlet, an outlet, and an inner surface coated with a catalyst; a hydrocarbon fuel source in fluid communication with the inlet of the heat exchanger; an oxygen source in fluid communication with the inlet of the heat exchanger; and a distribution system for receiving reformed hydrocarbon fuel from the heat exchanger. A method of removing heat from an aircraft component includes the steps of providing a heat exchanger proximate the aircraft component, the heat exchanger in fluid communication with the hydrocarbon fuel source and a water source, and having an inner surface coated with a catalyst; introducing the hydrocarbon fuel into the heat exchanger; introducing the oxygen into the heat exchanger; contacting the hydrocarbon fuel with the catalyst; and cracking the hydrocarbon fuel to form the reformed hydrocarbon fuel and removing heat from the aircraft component.

Description

Technical Field

[0001] The present invention relates to aircraft components and, more particularly, to those aircraft components that are exposed to heat during operation, such as during high speed flight. Background Art

[0002] Many aircraft components are exposed to heat sources during operation, such as internal components of the propulsion system including gas turbine engines. Particularly during high-speed flight operations, these components, as well as external components of the aircraft itself, are exposed to heat from skin friction due to the speed at which the aircraft travels through the atmosphere. In addition to the temperature of the air entering the gas turbine engine, components such as the leading edge of the aircraft skin, structures such as wings, spines and control surfaces, and engine inlets may be particularly affected.

[0003] These heat sources may cause surface and internal temperatures of such aircraft components to exceed their structural and / or operating capabilities, requiring expensive or exotic materials and changes in operating characteristics.

[0004] It would therefore be desirable to provide a system and method for removing heat from aircraft components that is reliable and durable in operational service and is capable of removing heat from such components during high speed flight.

[0005] It would also be desirable to capture the heat for use in improving the propulsion efficiency of the aircraft. Summary of the invention

[0006] In one aspect, a system for removing heat from an aircraft component includes: an aircraft component; a heat exchanger proximate the aircraft component and having an inlet, an outlet, and an interior surface coated with a catalyst; a hydrocarbon fuel source in fluid communication with the inlet of the heat exchanger; an oxygen source in fluid communication with the inlet of the heat exchanger; and a distribution system for receiving reformed hydrocarbon fuel from the heat exchanger.

[0007] In another aspect, a method of removing heat from an aircraft component includes the steps of providing a heat exchanger proximate the aircraft component, the heat exchanger being in fluid communication with a hydrocarbon fuel source and a water source and having an interior surface coated with a catalyst; introducing a hydrocarbon fuel into the heat exchanger; introducing oxygen into the heat exchanger; contacting the hydrocarbon fuel with the catalyst; and cracking the hydrocarbon fuel to form a reformed hydrocarbon fuel and removing heat from the aircraft component.

[0008] On the other hand, a gas turbine engine includes: a compressor, a combustor, and a turbine in a series flow relationship; a heat exchanger having an inlet, an outlet, and an inner surface coated with a catalyst; a hydrocarbon fuel source connected to the inlet fluid of the heat exchanger; an oxygen source connected to the inlet fluid of the heat exchanger; and a distribution system for receiving reformed hydrocarbon fuel from the heat exchanger.

[0009] Technical Solution 1. A system for removing heat from an aircraft component, the system comprising:

[0010] Aircraft parts;

[0011] a heat exchanger proximate the aircraft component, the heat exchanger having an inlet, an outlet, and an inner surface coated with a catalyst;

[0012] a hydrocarbon fuel source in fluid communication with the inlet of the heat exchanger;

[0013] a source of oxygen in fluid communication with the inlet of the heat exchanger; and

[0014] A distribution system is provided for receiving the reformed hydrocarbon fuel from the heat exchanger.

[0015] Technical Solution 2. The system according to any of the preceding technical solutions, characterized in that the aircraft component is a skin surface, a leading edge, a gas turbine engine inlet or a waste heat recovery heat exchanger.

[0016] Technical Solution 3. A system according to any of the foregoing technical solutions, characterized in that the hydrocarbon fuel is aircraft jet fuel such as Jet-A, JP-4 and JP-8, gasoline, kerosene, rocket propellant fuel such as RPS1 and RPS2, diesel fuel such as D2 and D4, and blends, mixtures and combinations thereof.

[0017] Technical Solution 4. The system according to any of the foregoing technical solutions is characterized in that the oxygen source is an oxygen-containing substance such as steam, liquid water, or an oxidized fuel such as ethanol or methanol.

[0018] Technical Solution 5. The system according to any of the preceding technical solutions is characterized in that the system comprises a plurality of heat exchangers.

[0019] Technical Solution 6. The system according to any of the foregoing technical solutions is characterized in that the multiple heat exchangers are arranged in series or in parallel.

[0020] Technical Solution 7. The system according to any of the preceding technical solutions, characterized in that the distribution system delivers reformed fuel to the gas turbine engine.

[0021] Technical Solution 8. The system according to any of the preceding technical solutions, characterized in that the heat exchanger comprises a plurality of tubular structures.

[0022] Technical Solution 9. The system according to any of the preceding technical solutions, characterized in that the heat exchanger is formed integrally with the aircraft component.

[0023] Technical Solution 10. A method for removing heat from an aircraft component, the method comprising the following steps:

[0024] providing a heat exchanger proximate the aircraft component, the heat exchanger being in fluid communication with a hydrocarbon fuel source and a water source and having an interior surface coated with a catalyst;

[0025] introducing a hydrocarbon fuel into the heat exchanger;

[0026] introducing oxygen into the heat exchanger;

[0027] contacting the hydrocarbon fuel with the catalyst; and

[0028] The hydrocarbon fuel is cracked to form a reformed hydrocarbon fuel and to remove heat from the aircraft component.

[0029] Technical Solution 11. The method according to any of the above technical solutions is characterized in that the aircraft component is a skin surface, a leading edge, a gas turbine engine inlet or a waste heat recovery heat exchanger.

[0030] Technical Solution 12. The method according to any of the aforementioned technical solutions is characterized in that the hydrocarbon fuel is aircraft jet fuel such as Jet-A, JP-4 and JP-8, gasoline, kerosene, rocket propellant fuel such as RPS1 and RPS2, diesel fuel such as D2 and D4, and their blends, mixtures and combinations.

[0031] Technical Solution 13. The method according to any of the foregoing technical solutions is characterized in that the oxygen source is an oxygen-containing substance such as steam, liquid water, or an oxidized fuel such as ethanol or methanol.

[0032] Technical Solution 14. The method according to any of the foregoing technical solutions is characterized in that the method includes multiple heat exchangers.

[0033] Technical Solution 15. The method according to any of the foregoing technical solutions is characterized in that the multiple heat exchangers are arranged in series or in parallel.

[0034] Technical Solution 16. The method according to any of the preceding technical solutions, characterized in that the distribution system delivers the reformed fuel to the gas turbine engine.

[0035] Technical Solution 17. The method according to any of the foregoing technical solutions is characterized in that the heat exchanger comprises a plurality of tubular structures.

[0036] Technical Solution 18. The method according to any of the above technical solutions is characterized in that the heat exchanger is formed integrally with the aircraft component.

[0037] Technical Solution 19. A gas turbine engine, comprising:

[0038] A compressor, combustor and turbine in series flow relationship;

[0039] a heat exchanger having an inlet, an outlet, and an inner surface coated with a catalyst;

[0040] a hydrocarbon fuel source in fluid communication with the inlet of the heat exchanger;

[0041] a source of oxygen in fluid communication with the inlet of the heat exchanger; and

[0042] A distribution system is provided for receiving the reformed hydrocarbon fuel from the heat exchanger.

[0043] Technical Solution 20. The gas turbine engine according to any of the foregoing technical solutions is characterized in that the heat exchanger is a waste heat recovery heat exchanger.

[0044] Technical Solution 21. A gas turbine engine according to any of the foregoing technical solutions, characterized in that the hydrocarbon fuel is aircraft jet fuel such as Jet-A, JP-4 and JP-8, gasoline, kerosene, rocket propellant fuel such as RPS1 and RPS2, diesel fuel such as D2 and D4, and blends, mixtures and combinations thereof.

[0045] Technical Solution 22. A gas turbine engine according to any of the foregoing technical solutions, characterized in that the oxygen source is an oxygen-containing substance such as steam, liquid water, or an oxidized fuel such as ethanol or methanol.

[0046] Technical Solution 23. A gas turbine engine according to any of the preceding technical solutions, characterized in that the system comprises a plurality of heat exchangers.

[0047] Technical Solution 24. A gas turbine engine according to any of the foregoing technical solutions, characterized in that the multiple heat exchangers are arranged in series or in parallel.

[0048] Technical Solution 25. The gas turbine engine according to any of the preceding technical solutions, characterized in that the distribution system delivers reformed fuel to the gas turbine engine.

[0049] Technical Solution 26. A gas turbine engine according to any of the foregoing technical solutions, characterized in that the heat exchanger comprises a plurality of tubular structures.

[0050] 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

[0051] A full and enabling disclosure of the present 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, in which:

[0052] Figure 1 is a perspective view of an exemplary embodiment of a high-speed aircraft suitable for implementing the heat removal apparatus and methods described herein.

[0053] Figure 2 is a schematic cross-sectional view of the tubes of an exemplary heat exchanger as described herein.

[0054] Figure 3 is a cross-sectional schematic diagram of an exemplary embodiment of a heat exchanger as described herein having an array of a plurality of tubular structures.

[0055] Figure 4 is a plan view schematic diagram of an exemplary embodiment of a heat exchanger as described herein having an array of a plurality of tubular structures and depicting inlet and outlet manifolds.

[0056] Figure 5 is a schematic cross-sectional view of an exemplary embodiment of a gas turbine engine suitable for use as an aircraft propulsion system and incorporating a heat exchanger as described herein.

[0057] Figure 6 is a schematic diagram of a gas turbine engine that utilizes a heat exchanger as described herein downstream of the last turbine stage to reform a hydrocarbon fuel.

[0058] Figure 7 is similar to Figure 6 Schematic diagram of a gas turbine engine that utilizes a heat exchanger as described herein between turbine stages to reform a hydrocarbon fuel.

[0059] Corresponding reference characters indicate corresponding parts throughout the several views.The exemplifications set out herein illustrate exemplary embodiments of the disclosure, and these exemplifications are not to be construed as limiting the scope of the disclosure in any way. DETAILED DESCRIPTION

[0060] Reference will now be made in detail to the present embodiments of the present invention, one or more examples of which are shown 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 are used to refer to the same or similar parts of the present invention.

[0061] The following description is provided to enable those skilled in the art to make and use the described embodiments contemplated for carrying out the present invention. However, various modifications, equivalents, variations and alternatives will remain apparent to those skilled in the art. Any and all such modifications, variations, equivalents and alternatives are intended to fall within the spirit and scope of the present invention.

[0062] All directional references (e.g., radial, axial, proximal, distal, upper, lower, upward, downward, left, right, lateral, front, rear, top, bottom, above, below, vertical, horizontal, clockwise, counterclockwise, upstream, downstream, forward, backward, etc.) are used only for identification purposes to help the reader understand the present invention and do not create limitations, especially limitations on the position, orientation, or use of the present invention. Connection references (e.g., attachment, connection, connection, and engagement) will be interpreted broadly and may include intermediate members between a series of elements and relative movement between elements unless otherwise indicated. As such, connection references do not necessarily infer that two elements are directly connected and are in a fixed relationship to each other. The exemplary drawings are for illustrative purposes only, and the dimensions, positions, orders, and relative sizes reflected in the drawings attached herein may vary.

[0063] The terms "coupled," "fixed," "attached to," and the like refer to both direct coupling, fixing, or attachment as well as indirect coupling, fixing, or attachment through one or more intermediate components or features, unless otherwise specified herein.

[0064] The singular forms "a", "an", and "the" include plural references unless the context clearly dictates otherwise. In addition, the suffix "(one or more)" used herein is generally intended to include both the singular and plural of the term that it modifies, thereby including one or more of the term.

[0065] As used herein, the term "or" does not mean exclusive, but rather means that at least one of the referenced element (eg, material) is present and includes instances where combinations of the referenced elements may be present unless the context clearly dictates otherwise.

[0066] As used throughout the specification and claims, approximate language is used to modify any quantitative representation that is permissible for variation without causing a change in the basic function to which it is associated. 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 cases, approximate language may correspond to the accuracy of an instrument used to measure a 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 within a 10% margin.

[0067] Here and throughout the specification and claims, range limitations are combined and interchanged, such ranges are identified and include all sub-ranges contained therein unless context or language indicates otherwise. For example, all ranges disclosed herein include the endpoints, and the endpoints are independently combinable with each other.

[0068] As used herein, the terms "may" and "may be" indicate the possibility of occurring under a set of circumstances; possessing a specified attribute, characteristic, or function; and / or qualifying another verb by expressing one or more capabilities, properties, or possibilities associated with the qualified verb. Thus, the use of "may" and "may be" indicates that the modified term is clearly suitable for, capable of reaching, or suitable for the indicated capacity, function, or usage, while taking into account that in some circumstances the modified term may sometimes be unsuitable, incapable of reaching, or inappropriate. For example, in some circumstances, an event or capacity may be expected, while in other circumstances, the event or capacity cannot occur. This distinction is captured by the terms "may" and "may be".

[0069] References throughout the specification to "some embodiments" or the like mean that a particular element (e.g., feature, structure, and / or characteristic) described in connection with the present invention is included in at least one embodiment described herein, and may or may not be present in other embodiments. Furthermore, it is to be understood that the described inventive features may be combined in any suitable manner in the various embodiments.

[0070] Various aspects of the invention are explained more fully with reference to the exemplary embodiments discussed below.It should be understood that, in general, features of one embodiment may also be used in combination with features of another embodiment, and that these embodiments are not intended to limit the scope of the invention.

[0071] Figure 1 1 is a perspective view of an exemplary embodiment of a high-speed aircraft 10 suitable for implementing the heat removal apparatus and methods described herein. As used herein, the term "high-speed aircraft" is intended to refer to an aircraft designed to operate above the speed of sound (i.e., above Mach 1), and more particularly to an aircraft designed to operate in a hypersonic flight regime above Mach 5, such as in the range of Mach 5 to Mach 10.

[0072] exist Figure 1 In the configuration shown, an exemplary high-speed aircraft 10 includes a fuselage 11, wings 12, vertical stabilizers 13, leading edges 14 of the wings 12 and fuselage 11, and a gas turbine engine 15 used as a propulsion system for the aircraft. An outer skin surface 16 covers at least portions of the fuselage 11 and wings 12. In high-speed flight, the movement of the aircraft 10 through the atmosphere causes heating of aircraft surfaces such as the skin surfaces 16 on the exterior of the aircraft 10, and particularly causes heat accumulation in the area of ​​the leading edges 14 of the wings 12 and fuselage 11. High-speed operation also causes the gas turbine engine 15 to experience high temperature operation because the atmospheric air is reduced to subsonic speeds, which increases the pressure within the gas turbine engine. Each of these locations can utilize a system for removing heat from aircraft components as described herein.

[0073] Figure 2 is a schematic cross-sectional view of a tube of an exemplary heat removal system 20 in the form of a heat exchanger as described herein. The heat removal system 20 includes at least one tubular structure 21 having an inner surface 22 and an outer surface 23. The inner surface 22 includes a coating 24 as will be described below, and an interior space 25 located inside the inner surface 22 and the coating 24 through which a fluid can pass.

[0074] The inner surface of the tubular structure 21 includes a coating 24 of the perovskite material and the tuning material, and the inner surface of the tubular structure 21 is accessible to a fluid, such as a hydrocarbon fuel, in an interior space 25 .

[0075] As used herein, the term "hydrocarbon cracking", "cracked hydrocarbons" or any variation thereof refers to, but is not limited to, a process in which hydrocarbons are cracked in a device to obtain materials with smaller molecules. The hydrocarbons may include ethane, heptane, liquid petroleum gas, naphtha, gas oil, bottom products from atmospheric and vacuum distillation of crude oil, or any combination thereof.

[0076] As used herein, the term "coke" or any variation thereof refers to, but is not limited to, a carbonaceous solid or liquid, or particles or macromolecules forming a carbonaceous solid or liquid, obtained from coal, petroleum, wood, hydrocarbons, and other carbonaceous materials.

[0077] As used herein, the term "perovskite material" or any variation thereof refers to, but is not limited to, a material having an ABO3 perovskite structure and having a molecular formula A a B b O 3-δAny material, where 0.9 < a ≤ 1.2; 0.9 < b ≤ 1.2; -0.5 < δ < 0.5; A includes a first element and optionally a second element, the first element is selected from calcium (Ca), strontium (Sr), barium (Ba), lithium (Li), sodium (Na), potassium (K), rubidium (Rb), and any combination thereof, the second element is selected from yttrium (Y), bismuth (Bi), lanthanum (La), cerium (Ce), praseodymium (Pr), neodymium (Nd), promethium (Pm), samarium (Sm), europium (Eu), gadolinium (Gd), terbium (Tb), dysprosium (Dy), holmium (Ho), erbium (Er), thulium (Tm), ytterbium (Yb), lutetium (Lu), and any combination thereof; and B is selected from silver (Ag), gold (Au), cadmium (Cd), cerium (Ce), cobalt (Co), chromium (Cr), copper (Cu), dysprosium (Dy), erbium (Er), europium (Eu), iron (Fe), gallium (Ga), gadolinium (Gd), hafnium (Hf), holmium (Ho), indium (In), iridium (Ir), lanthanum (La), lutetium (Lu), manganese (Mn), molybdenum (Mo), niobium (Nb), neodymium (Nd), nickel (Ni), osmium (Os), palladium (Pd), promethium (Pm), praseodymium (Pr), platinum (Pt), rhenium (Re), rhodium (Rh), ruthenium (Ru), antimony (Sb), scandium (Sc), samarium (Sm), tin (Sn), tantalum (Ta), terbium (Tb), technetium (Tc), titanium (Ti), thulium (Tm), vanadium (V), tungsten (W), yttrium (Y), ytterbium (Yb), zinc (Zn), zirconium (Zr), and any combination thereof.

[0078] In some embodiments, the perovskite material may have the formula n(A a B b O 3-δ ), where n = 2, 3, 4, 8, etc., and the formula A a B b O 3-δ is its simplified form. In some embodiments, in the ABO3 perovskite structure, the A cations are surrounded by twelve anions in a cubic-octahedral coordination, the B cations are surrounded by six anions in an octahedral coordination, and the oxygen anions are coordinated by two B cations and four A cations. In some embodiments, the ABO3 perovskite structure is constructed from corner-sharing BO6 octahedra. In some embodiments, the ABO3 perovskite structure includes distorted derivatives. The distortion may be due to the rotation or tilt of regular, rigid octahedra, or due to the presence of distorted BO6 octahedra. In some embodiments, the ABO3 perovskite structure is cubic. In some embodiments, the ABO3 perovskite structure is hexagonal.

[0079] In some embodiments, A includes only the first element. The first element may be a single element or a combination of elements selected from calcium (Ca), strontium (Sr), barium (Ba), lithium (Li), sodium (Na), potassium (K) and rubidium (Rb).

[0080] In some embodiments, A includes a combination of a first element and a second element. The second element may be a single element or a combination of elements selected from yttrium (Y), bismuth (Bi), lanthanum (La), cerium (Ce), praseodymium (Pr), neodymium (Nd), promethium (Pm), samarium (Sm), europium (Eu), gadolinium (Gd), terbium (Tb), dysprosium (Dy), holmium (Ho), erbium (Er), thulium (Tm), ytterbium (Yb), and lutetium (Lu).

[0081] Similarly, B can be selected from silver (Ag), gold (Au), cadmium (Cd), cerium (Ce), cobalt (Co), chromium (Cr), copper (Cu), dysprosium (Dy), erbium (Er), europium (Eu), iron (Fe), gallium (Ga), gadolinium (Gd), hafnium (Hf), holmium (Ho), indium (In), iridium (Ir), lanthanum (La), lutetium (Lu), manganese (Mn), molybdenum (Mo), niobium (Nb), neodymium (Nd), nickel ( single element or a combination of elements of Ni, osmium (Os), palladium (Pd), promethium (Pm), praseodymium (Pr), platinum (Pt), rhenium (Re), rhodium (Rh), ruthenium (Ru), antimony (Sb), scandium (Sc), samarium (Sm), tin (Sn), tantalum (Ta), terbium (Tb), technetium (Tc), titanium (Ti), thulium (Tm), vanadium (V), tungsten (W), yttrium (Y), ytterbium (Yb), zinc (Zn) and zirconium (Zr).

[0082] In some embodiments, the perovskite material includes SrCeO3, SrZr 0.3 Ce 0.7 O3, BaMnO3, BaCeO3, BaZr 0.3 Ce 0.7 O3、BaZr 0.3 Ce 0.5 Y 0.2 O3、BaZr 0.1 Ce 0.7 Y 0.2 O3、BaZrO3、BaZr 0.7 Ce 0.3 O3、BaCe 0.5 Zr 0.5 O3、BaCe 0.9 Y 0.1 O3、BaCe 0.85 Y 0.15 O3 or BaCe 0.8 Y 0.2O3. For example, for SrCeO3, A is Sr, a=1, B is Ce, b=1, and δ=0. For SrZr 0.3 Ce 0.7 O3, A is Sr, a=1, B is a combination of Zr and Ce, b=1, and δ=0. For BaMnO3, A is Ba, a=1, B is Mn, b=1, and δ=0. For BaCeO3, A is Ba, a=1, B is Ce, b=1, and δ=0. For BaZr 0.3 Ce 0.7 O3, A is Ba, a=1, B is a combination of Zr and Ce, b=1, and δ=0. 0.3 Ce 0.5 Y 0.2 O3, A is Ba, a=1, B is a combination of Zr, Ce and Y, b=1, and δ=0.

[0083] In some embodiments, the perovskite material includes La 0.1 Ba 0.9 Ce 0.7 Zr 0.2 Y 0.1 O3、Ce 0.1 Ba 0.9 Ce 0.7 Zr 0.2 Y 0.1 O 3.05 、Ce 0.5 Ba 0.5 Ce 0.7 Zr 0.2 Y 0.1 O 3.45 , Y 0.1 Ba 0.9 Ce 0.7 Zr 0.2 Y 0.1 O3, Y 0.5 Ba 0.5 Ce 0.7 Zr 0.2 Y 0.1 O 3.2 、Bi 0.1 Ba 0.9 Ce 0.7 Zr 0.2 Y 0.1 O3、Bi 0.5 Ba 0.5 Ce 0.7 Zr 0.2 Y 0.1 O 3.2 , Pr 0.1 Ba 0.9 Ce 0.7 Zr0.2 Y 0.1 O3 or Pr 0.5 Ba 0.5 Ce 0.7 Zr 0.2 Y 0.1 O 3.2 For La 0.1 Ba 0.9 Ce 0.7 Zr 0.2 Y 0.1 O3, A is a combination of Ba and La, the first element is La, the second element is Ba, a=1, B is a combination of Ce, Zr and Y, b=1, and δ=0. For Ce 0.1 Ba 0.9 Ce 0.7 Zr 0.2 Y 0.1 O 3.05 and Ce 0.5 Ba 0.5 Ce 0.7 Zr 0.2 Y 0.1 O 3.45 , A is a combination of Ce and Ba, the first element is Ce, the second element is Ba, a=1, B is a combination of Ce, Zr and Y, b=1, and δ is equal to −0.05 and −0.45 respectively. 0.1 Ba 0.9 Ce 0.7 Zr 0.2 Y 0.1 O3 and Y 0.5 Ba 0.5 Ce 0.7 Zr 0.2 Y 0.1 O 3.2 , A is a combination of Y and Ba, the first element is Y, the second element is Ba, a=1, B is a combination of Ce, Zr and Y, b=1, and δ is equal to 0 and −0.2 respectively. 0.1 Ba 0.9 Ce 0.7 Zr 0.2 Y 0.1 O3 and Bi 0.5 Ba 0.5 Ce 0.7 Zr 0.2 Y 0.1 O 3.2 , A is a combination of Bi and Ba, the first element is Bi, the second element is Ba, a=1, B is a combination of Ce, Zr, and Y, b=1, and δ is equal to 0 and −0.2, respectively. Similarly, for Pr 0.1 Ba 0.9 Ce0.7 Zr 0.2 Y 0.1 O3 and Pr 0.5 Ba 0.5 Ce 0.7 Zr 0.2 Y 0.1 O 3.2 , A is a combination of Pr and Ba, the first element is Pr, the second element is Ba, a=1, B is a combination of Ce, Zr and Y, b=1, and δ is equal to 0 and −0.2 respectively.

[0084] In some embodiments, the perovskite material includes BaZr 0.3 Ce 0.7 O3.

[0085] As used herein, the term "tuning material" or any variation thereof refers to any material that reduces the yield of carbon monoxide in hydrocarbon cracking. The tuning material may include one material or a combination of multiple materials. In some embodiments, the tuning material includes zirconium oxide, doped zirconium oxide, or any precursor or combination thereof.

[0086] In some embodiments, the process for cracking hydrocarbons is operated in the presence of steam at a temperature in the range from about 700°C to about 900°C, the weight ratio of steam to hydrocarbon is in the range from about 3:7 to about 7:3, and the hydrocarbon comprises ethane, heptane, liquefied petroleum gas, naphtha, gas oil, or any combination thereof.

[0087] In some embodiments, the process for cracking hydrocarbons comprises bottom products from atmospheric and vacuum distillation of crude oil and is operated in the presence of steam at a temperature in the range of from about 480°C to about 600°C, and the weight percentage of steam is in the range of from about 1 wt % to about 2 wt %.

[0088] The perovskite material may or may not chemically react with the tuning material. Thus, the inner surface may include a combination or reaction product of the perovskite material and the tuning material. In some embodiments, the inner surface includes a combination of the perovskite material, the tuning material, and a reaction product of the perovskite material and the tuning material.

[0089] The perovskite material and the tuning material can be applied in coatings to the device using different methods (eg, air plasma spraying, slurry coating, sol-gel coating, and solution coating). In some embodiments, the perovskite material and the tuning material are applied using a slurry coating method.

[0090] Depending on the specific tuning material and perovskite material used and the operating conditions of the coating, the amount of tuning material and perovskite material in the slurry can vary, as long as a continuous, strong, carbon monoxide reducing and anti-coking coating is formed. In some embodiments, the weight ratio of perovskite material to tuning material is from about 7:3 to about 7:93. In some embodiments, the weight of the perovskite material is equal to or less than the weight of the tuning material.

[0091] The slurry may further include an organic binder, an inorganic binder, a wetting agent, a solvent, or any combination thereof to enhance the slurry wetting ability, adjust the slurry viscosity, or obtain good green coating strength. When an organic binder, an inorganic binder, a wetting agent, a solvent, or any combination thereof is added to the slurry, the total weight percentage of the tuning material and the perovskite material in the slurry may be from about 10% to about 90%, or preferably from about 15% to about 70%, or more preferably from about 30% to about 55%.

[0092] In some embodiments, the paste includes a perovskite material, a tuning material, cerium oxide, yttrium oxide, glycerol, and polyvinyl alcohol (PVA).

[0093] The slurry can be applied to the equipment by different techniques, such as at least one of wiping, painting, centrifuging, spraying, filling and draining, and dipping. In some embodiments, the slurry is applied by dipping, i.e., dipping the part to be coated into the slurry. In some embodiments, the slurry is applied by filling and draining, i.e., filling the slurry into the article to be coated and then draining the slurry by, for example, gravity.

[0094] Additional descriptions of cracking processes and systems can be found in the published patent literature, all of which are incorporated herein by reference: US9,499,747, WO2015105589A1, CA2821249A1, US20170260460, CA2932461A1, WO2015088671A1, and US20170022428.

[0095] After the slurry is applied to the device, a sintering process may follow. As used herein, the term "sintering" or any variation thereof refers to, but is not limited to, a method of heating a material in a sintering furnace or other heater facility. In some embodiments, the sintering temperature is in the range of from about 850°C to about 1700°C. In some embodiments, the sintering temperature is about 1000°C. In some embodiments, sintering is performed in an inert atmosphere such as argon or nitrogen. In some embodiments, sintering is preceded by a heat treatment in air to form an oxide layer on the inner surface of the tube that improves the adhesion of the coating.

[0096] In operation, hydrocarbons such as aircraft fuels are supplied to the interior space 25 together with oxygen-containing substances such as steam, which can be provided in the form of liquid water or oxygenated fuels such as ethanol or methanol, which can evaporate in the presence of sufficient heat. The cracking of hydrocarbon fuels occurring in the interior space 25 is an endothermic reaction in which all carbon-carbon bonds are broken and hydrogen radicals and methylene radicals are formed. This highly endothermic process takes away heat from the tubular structure 21 and its surroundings, and thus cools the tubular structure 21 and its surroundings. Therefore, the cracking of hydrocarbons transforms the tubular structure 21 into a heat exchanger 20 and acts as a heat removal system that can be used to remove heat from the components 30 of the high-speed aircraft 10. The coating reduces or prevents the formation of coke in the interior space 25, which can eventually hinder the flow of fuel and reduce the capacity of the heat exchanger 20. The reformed fuel leaving the heat exchanger 20 can then be used as a fuel for aircraft propulsion.

[0097] Figure 3 2 is a cross-sectional schematic diagram of an exemplary embodiment of a heat removal system in the form of a heat exchanger 20 as described herein, having an array of a plurality of tubular structures 21, each having an interior space 25. The heat exchanger 20 is positioned in close proximity to an aircraft component 30 from which it is desired to remove thermal energy (heat). The heat exchanger 20 may be joined to the aircraft component by conventional or additive manufacturing techniques known in the art, or may be integrally formed with the aircraft component. Alternatively, the heat exchanger 20 may be used as an air-fuel heat exchanger, and thus used to remove heat from an air stream flowing through the array of tubular structures 21.

[0098] Figure 4 2 is a schematic plan view of an exemplary embodiment of a heat exchanger 20 described herein having an array of a plurality of tubular structures 21, and depicting an inlet manifold 26 and an outlet manifold 28, respectively. Manifolds 26 and 28 fluidly couple the inlet and outlet ends of the tubular structures 21 to a common inlet 27 and outlet 29, respectively. Inlet 27, in turn, may be fluidly coupled to a pipe or conduit that is a source of hydrocarbon fuel and / or steam (water). Outlet 29, in turn, may be fluidly coupled to a pipe or conduit that is a destination or recipient of reformed fuel after cracking occurs within the tubular structures 21 of the heat exchanger 20.

[0099] In addition to multiple tubular structures sharing a common manifold, a heat exchanger as described herein may actually include multiple heat exchangers having separate manifolds. A control system including sensors, valves, and / or electronically controlled actuators may control flow through each tubular structure and / or between multiple heat exchangers. This may provide flexibility for different operating conditions and for cycling between heat exchangers if it is necessary or desirable to take some tubular structures or some heat exchangers offline to remove any coke deposits that may have accumulated during operation.

[0100] Figure 5 1 is a schematic cross-sectional view of an exemplary embodiment of a gas turbine engine 15 suitable for use as an aircraft propulsion system for a high-speed aircraft 10 and incorporating a heat exchanger 20 as described herein. Figure 5 As shown, a heat exchanger 20 containing a catalyst coating 24 within a tubular structure 21 is located within an inlet section 31 of a gas turbine engine 15. The casing surrounding the inlet section serves as an aircraft component 30 from which it is desired to remove heat generated and transported by an incoming air flow 33 during high-speed flight. This reduces the temperature of the air before it enters a compressor section 32 of the gas turbine engine 15. The reformed fuel leaving the heat exchanger 20 may then be fed into a combustor section 34 of the gas turbine engine 15.

[0101] Figure 6 is a schematic diagram of a gas turbine engine 15 that utilizes a heat exchanger 20 as described herein downstream of the final turbine stage after a high pressure turbine 35 and a low pressure turbine 36 to reform a hydrocarbon fuel. Heat contained in the remaining air flow exiting the low pressure turbine 36 provides energy to crack the fuel and is then exhausted to the atmosphere 37. As previously discussed, the reformed fuel may be fed into the combustor 34 to fuel the gas turbine engine 15.

[0102] Figure 7 is similar to Figure 6 Schematic diagram of a gas turbine engine 15 utilizing a heat exchanger 20 as described herein between a high pressure turbine stage 35 and a low pressure turbine stage 36 to reform a hydrocarbon fuel.

[0103] For any of the exemplary embodiments described herein, multiple heat removal systems may be used in series or in parallel and may share inlets and outlets, or may be plumbed to separate sources and destinations of hydrocarbon fuel and reformate fuel.

[0104] For parallel heat removal systems, all systems may be operated simultaneously, or some systems may be deactivated to regenerate and remove coke deposits, or to adjust the level of heat removal capacity during various phases of aircraft operation.

[0105] A variety of hydrocarbon fuels may be used with the exemplary embodiments described herein, including aircraft jet fuels such as Jet-A, JP-4, and JP-8, gasoline, kerosene, rocket propellant fuels such as RPS1 and RPS2, diesel fuels such as D2 and D4, and blends, mixtures, and combinations thereof.

[0106] All publications, patents, and patent applications cited herein, whether above or below, are incorporated herein by reference in their entirety, to the extent that each individual publication, patent, or patent application is specifically and individually indicated as being incorporated herein by reference. It should be appreciated that any patent, publication, or other public material considered to be incorporated herein by reference in whole or in part is incorporated herein only to the extent that the incorporated material does not conflict with existing definitions, statements, or other public materials set forth in this disclosure. As such, and to the extent necessary, the disclosure as explicitly set forth herein supersedes any conflicting material incorporated herein by reference. Any material or portion thereof that is considered to be incorporated herein by reference but conflicts with existing definitions, statements, or other public materials set forth herein will be incorporated only to the extent that no conflict arises between the incorporated material and the existing public materials.

[0107] It must be noted that, as used in this specification and the appended claims, the singular forms "a," "an," and "the" include plural referents unless the content clearly dictates otherwise.

[0108] Unless otherwise defined, all technical and scientific terms used herein have the same meaning as commonly understood by one of ordinary skill in the art to which the invention belongs. Although many methods and materials similar to or equivalent to those described herein can be used in the practice of the present invention, materials and methods according to some embodiments are described herein.

[0109] It should be noted that when used in this disclosure, the terms "includes," "comprising," and other derivatives from the root term "comprising" are intended to be open-ended terms that specify the presence of any stated features, elements, integers, steps, or components, and are not intended to exclude the presence or addition of one or more other features, elements, integers, steps, components, or combinations thereof.

[0110] As required, detailed embodiments of the present invention are disclosed herein; however, it is to be understood that the disclosed embodiments are merely examples of the present invention, which can be embodied in various forms. Therefore, the specific structural and functional details disclosed herein will not be interpreted as limiting, but only as a basis for the claims, and as a representative basis for teaching those skilled in the art to use the present invention differently in virtually any suitable detailed structure.

[0111] The various features, aspects and advantages of the present disclosure may also be embodied in any arrangement of the aspects of the present disclosure, including but not limited to the following technical solutions as defined in the enumerated aspects:

[0112] 1. In one aspect, a system for removing heat from an aircraft component comprises: an aircraft component; a heat exchanger proximate the aircraft component and having an inlet, an outlet, and an inner surface coated with a catalyst; a hydrocarbon fuel source in fluid communication with the inlet of the heat exchanger; an oxygen source in fluid communication with the inlet of the heat exchanger; and a distribution system for receiving reformed hydrocarbon fuel from the heat exchanger.

[0113] 2. The system of aspect 1, wherein the aircraft component is a skin surface, a leading edge, a gas turbine engine inlet, or a waste heat recovery heat exchanger.

[0114] 3. A system according to aspect 1 or 2, wherein the hydrocarbon fuel is aircraft jet fuel such as Jet-A, JP-4 and JP-8, gasoline, kerosene, rocket propellant fuel such as RPS1 and RPS2, diesel fuel such as D2 and D4, and blends, mixtures and combinations thereof.

[0115] 4. A system according to any preceding aspect, wherein the oxygen source is an oxygen-containing substance such as steam, liquid water, or an oxygenated fuel such as ethanol or methanol.

[0116] 5. A system according to any preceding aspect, wherein the system comprises a plurality of heat exchangers.

[0117] 6. A system according to aspect 5, wherein a plurality of heat exchangers are arranged in series or in parallel.

[0118] 7. The system of any preceding aspect, wherein the distribution system delivers reformed fuel to the gas turbine engine.

[0119] 8. The system of any preceding aspect, wherein the heat exchanger comprises a plurality of tubular structures.

[0120] 9. The system of any preceding aspect, wherein the heat exchanger is integrally formed with the aircraft component.

[0121] 10. In another aspect, a method of removing heat from an aircraft component includes the steps of providing a heat exchanger proximate the aircraft component, the heat exchanger being in fluid communication with a hydrocarbon fuel source and a water source and having an interior surface coated with a catalyst; introducing a hydrocarbon fuel into the heat exchanger; introducing oxygen into the heat exchanger; contacting the hydrocarbon fuel with the catalyst; and cracking the hydrocarbon fuel to form a reformed hydrocarbon fuel and removing heat from the aircraft component.

[0122] 11. The method of aspect 10, wherein the aircraft component is a skin surface, a leading edge, a gas turbine engine inlet, or a waste heat recovery heat exchanger.

[0123] 12. A method according to any preceding aspect, wherein the hydrocarbon fuel is aircraft jet fuel such as Jet-A, JP-4 and JP-8, gasoline, kerosene, rocket propellant fuel such as RPS1 and RPS2, diesel fuel such as D2 and D4, and blends, mixtures and combinations thereof.

[0124] 13. A method according to any preceding aspect, wherein the oxygen source is an oxygen-containing species such as steam, liquid water, or an oxygenated fuel such as ethanol or methanol.

[0125] 14. A method according to any preceding aspect, wherein the method comprises a plurality of heat exchangers.

[0126] 15. The method according to aspect 14, wherein a plurality of heat exchangers are arranged in series or in parallel.

[0127] 16. The method of any preceding aspect, wherein the distribution system delivers reformed fuel to the gas turbine engine.

[0128] 17. A method according to any preceding aspect, wherein the heat exchanger comprises a plurality of tubular structures.

[0129] 18. A method according to any preceding aspect, wherein the heat exchanger is integrally formed with the aircraft component.

[0130] 19. On the other hand, a gas turbine engine includes: a compressor, a combustor, and a turbine in a series flow relationship; a heat exchanger having an inlet, an outlet, and an inner surface coated with a catalyst; a hydrocarbon fuel source connected to the inlet fluid of the heat exchanger; an oxygen source connected to the inlet fluid of the heat exchanger; and a distribution system for receiving reformed hydrocarbon fuel from the heat exchanger.

[0131] 20. The gas turbine engine of aspect 19, wherein the heat exchanger is a waste heat recovery heat exchanger.

[0132] 21. A gas turbine engine according to any of the preceding aspects, wherein the hydrocarbon fuel is aircraft jet fuel such as Jet-A, JP-4 and JP-8, gasoline, kerosene, rocket propellant fuel such as RPS1 and RPS2, diesel fuel such as D2 and D4, and blends, mixtures and combinations thereof.

[0133] 22. A gas turbine engine according to any preceding aspect, wherein the oxygen source is an oxygen-containing substance such as steam, liquid water, or an oxygenated fuel such as ethanol or methanol.

[0134] 23. The gas turbine engine according to any preceding aspect, wherein the system comprises a plurality of heat exchangers.

[0135] 24. The gas turbine engine according to aspect 23, wherein a plurality of heat exchangers are arranged in series or in parallel.

[0136] 25. The gas turbine engine according to any preceding aspect, wherein the distribution system delivers reformed fuel to the gas turbine engine.

[0137] 26. A gas turbine engine according to any preceding aspect, wherein the heat exchanger comprises a plurality of tubular structures.

[0138] Although the present disclosure has been described as having exemplary embodiments, the present disclosure may be further modified in the spirit and scope of the present disclosure. Therefore, the present application is intended to cover any variation, use or adaptation of the present disclosure using the general principles of the present disclosure. In addition, the present application is intended to cover such deviations from the present disclosure as are within the known or customary practice of the field to which the present disclosure belongs and fall within the limits of the appended claims.

Claims

1. A system for removing heat from an aircraft component, the system comprising: Aircraft parts; a gas turbine engine having a high pressure turbine and a low pressure turbine in a series flow sequence with an exhaust air stream flowing from the low pressure turbine; a heat exchanger proximate the aircraft component, the heat exchanger having an inlet, an outlet, and a plurality of tubular structures fluidly coupled to the inlet and the outlet, wherein each of the plurality of tubular structures has an inner surface coated with a catalyst, wherein the plurality of tubular structures are disposed within the exhaust air flow from the low pressure turbine, and wherein the plurality of tubular structures are in thermal communication with the exhaust air flow; a hydrocarbon fuel source in fluid communication with the inlet of the heat exchanger, wherein the catalyst is for reforming a hydrocarbon fuel from the hydrocarbon fuel source; a source of oxygen in fluid communication with the inlet of the heat exchanger, wherein the inlet of the heat exchanger is fluidly coupled to both the source of hydrocarbon fuel and the source of oxygen, wherein hydrocarbon fuel and oxygen combine within the plurality of tubular structures resulting in cracking of hydrocarbons; and A distribution system is fluidly coupled to the outlet of the heat exchanger, wherein the distribution system receives the reformed hydrocarbon fuel from the heat exchanger, and wherein the distribution system includes an outlet fluidly coupled to a combustor of an aircraft.

2. The system according to claim 1, characterized in that The aircraft component is a skin surface, a leading edge, a gas turbine engine inlet or a waste heat recovery heat exchanger.

3. The system according to claim 1, characterized in that The hydrocarbon fuels are aircraft jet fuel, gasoline, kerosene, rocket propellant fuel, diesel fuel, and blends, mixtures and combinations thereof.

4. The system according to claim 3, characterized in that The aircraft jet fuel is selected from at least one of Jet-A, JP-4 and JP-8.

5. The system according to claim 3, characterized in that The rocket propellant is selected from at least one of RPS1 and RPS2.

6. The system according to claim 3, characterized in that The diesel fuel is selected from at least one of D2 and D4.

7. The system according to claim 1, characterized in that The oxygen is an oxygen-containing species or an oxygenated fuel.

8. The system according to claim 7, characterized in that The oxygen-containing substance is selected from at least one of steam and liquid water.

9. The system according to claim 7, characterized in that The oxygenated fuel is selected from at least one of ethanol and methanol.

10. The system according to claim 1, characterized in that The system includes a plurality of heat exchangers.

11. The system according to claim 10, characterized in that The plurality of heat exchangers are arranged in series or in parallel.

12. The system according to claim 1, characterized in that The distribution system delivers the reformed fuel to the gas turbine engine.

13. The system according to claim 1, characterized in that The heat exchanger is integrally formed with the aircraft component.

14. A method of removing heat from an aircraft component, the method comprising the steps of: providing a heat exchanger proximate the aircraft component, the heat exchanger having an inlet, an outlet, and a plurality of tubular structures fluidly coupled to the inlet and the outlet, wherein each tubular structure of the plurality of tubular structures has an inner surface coated with a catalyst, wherein the plurality of tubular structures are disposed within an exhaust air flow from a low pressure turbine of a gas turbine engine of the aircraft, and wherein the plurality of tubular structures are in thermal communication with the exhaust air flow; introducing a hydrocarbon fuel into the heat exchanger via an inlet of the heat exchanger; introducing oxygen into the heat exchanger via the inlet of the heat exchanger; contacting the hydrocarbon fuel with the catalyst; and The hydrocarbon fuel is cracked to form a reformed hydrocarbon fuel and to remove heat from the aircraft component.

15. The method according to claim 14, characterized in that The aircraft component is a skin surface, a leading edge, a gas turbine engine inlet or a waste heat recovery heat exchanger.

16. The method according to claim 14, characterized in that The hydrocarbon fuels are aircraft jet fuel, gasoline, kerosene, rocket propellant fuel, diesel fuel, and blends, mixtures and combinations thereof.

17. The method according to claim 16, characterized in that The aircraft jet fuel is selected from at least one of Jet-A, JP-4 and JP-8.

18. The method according to claim 16, characterized in that The rocket propellant is selected from at least one of RPS1 and RPS2.

19. The method according to claim 16, characterized in that The diesel fuel is selected from at least one of D2 and D4.

20. The method according to claim 14, characterized in that The oxygen is an oxygen-containing species or an oxygenated fuel.

21. The method according to claim 20, characterized in that The oxygen-containing substance is selected from at least one of steam and liquid water.

22. The method according to claim 20, characterized in that The oxygenated fuel is selected from at least one of ethanol and methanol.

23. The method according to claim 14, characterized in that The method includes a plurality of heat exchangers.

24. The method according to claim 23, characterized in that The plurality of heat exchangers are arranged in series or in parallel.

25. The method according to claim 14, characterized in that The distribution system delivers the reformed fuel to the gas turbine engine.

26. The method according to claim 14, characterized in that The heat exchanger is integrally formed with the aircraft component.

27. A gas turbine engine, comprising: a compressor, a combustor, and a turbine in series flow relationship, the turbine having a high pressure turbine and a low pressure turbine in series flow order, wherein an exhaust air stream flows from the low pressure turbine; a heat exchanger having an inlet, an outlet, and a plurality of tubular structures fluidly coupled to the inlet and the outlet, wherein each of the plurality of tubular structures has an inner surface coated with a catalyst, wherein the plurality of tubular structures are disposed within the exhaust air flow from the low pressure turbine, and wherein the plurality of tubular structures are in thermal communication with the exhaust air flow; a hydrocarbon fuel source in fluid communication with the inlet of the heat exchanger, wherein the catalyst is for reforming a hydrocarbon fuel from the hydrocarbon fuel source; a source of oxygen in fluid communication with the inlet of the heat exchanger, wherein the inlet of the heat exchanger is fluidly coupled to both the source of hydrocarbon fuel and the source of oxygen, wherein hydrocarbon fuel and oxygen combine within the plurality of tubular structures resulting in cracking of hydrocarbons; and A distribution system is fluidly coupled to the outlet of the heat exchanger, wherein the distribution system receives the reformed hydrocarbon fuel from the heat exchanger, and wherein the distribution system includes an outlet fluidly coupled to the combustor of the gas turbine engine.

28. The gas turbine engine according to claim 27, characterized in that The heat exchanger is a waste heat recovery heat exchanger.

29. The gas turbine engine according to claim 27, characterized in that The hydrocarbon fuels are aircraft jet fuel, gasoline, kerosene, rocket propellant fuel, diesel fuel, and blends, mixtures and combinations thereof.

30. The gas turbine engine according to claim 29, characterized in that The aircraft jet fuel is selected from at least one of Jet-A, JP-4 and JP-8.

31. The gas turbine engine according to claim 29, characterized in that The rocket propellant is selected from at least one of RPS1 and RPS2.

32. The gas turbine engine according to claim 29, characterized in that The diesel fuel is selected from at least one of D2 and D4.

33. The gas turbine engine according to claim 27, characterized in that The oxygen is an oxygen-containing species or an oxygenated fuel.

34. The gas turbine engine according to claim 33, characterized in that The oxygen-containing substance is selected from at least one of steam and liquid water.

35. The gas turbine engine according to claim 33, characterized in that The oxygenated fuel is selected from at least one of ethanol and methanol.

36. The gas turbine engine according to claim 27, characterized in that The system includes a plurality of heat exchangers.

37. The gas turbine engine according to claim 36, characterized in that The plurality of heat exchangers are arranged in series or in parallel.

38. The gas turbine engine according to claim 27, characterized in that The distribution system delivers reformed fuel to the gas turbine engine.

Images