1. A gas turbine engine, characterized in that the gas turbine engine comprises:

a compressor, a combustor and a turbine in serial flow relationship;

a heat exchanger having an inlet, an outlet, and an inner surface coated with a catalyst, the heat exchanger being located upstream of the compressor;

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.

2. The gas turbine engine of claim 1, wherein the distribution system delivers reformed hydrocarbon fuel to the gas turbine engine.

3. The gas turbine engine of claim 1, wherein the gas turbine engine comprises a superburner or an augmentor.

4. The gas turbine engine of claim 3, wherein the distribution system delivers reformed hydrocarbon fuel to the super burner or the augmentor.

5. The gas turbine engine of claim 3, wherein the afterburner or the augmentor comprises an afterburner fuel nozzle, and wherein the distribution system is configured to selectively deliver liquid hydrocarbon fuel or reformed hydrocarbon fuel to the fuel nozzle.

6. The gas turbine engine of claim 5, wherein the distribution system delivers a liquid hydrocarbon fuel to the combustor and a reformed gaseous hydrocarbon fuel to the super burner.

7. The gas turbine engine of claim 1, wherein the hydrocarbon fuel is aircraft Jet fuel, such as Jet-a, JP-4 and JP-8, gasoline, kerosene, rocket propulsion fuels, such as RPS1 and RPS2, diesel fuels, such as D2 and D4, blends, mixtures and combinations thereof.

8. The gas turbine engine of claim 1, wherein the oxygen source is an oxygenate such as steam, liquid water, or an oxygenated fuel such as ethanol or methanol.

9. The gas turbine engine of claim 1, wherein the system comprises a plurality of heat exchangers or branches of heat exchangers.

10. The gas turbine engine of claim 9, wherein the plurality of heat exchangers are arranged in series or in parallel.

Background

Many aircraft components, such as the internal components of a propulsion system including a gas turbine engine, are exposed to heat sources during operation. Particularly during high-speed flight operations, these components, as well as the exterior 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 aircraft skin, the leading edge of structures (such as wings, ridges, and control surfaces), and the engine inlet may also be particularly affected.

These heat sources may cause the surface and interior temperatures of such aircraft components to exceed their structural and/or operational capabilities, requiring expensive or special materials and changing operational characteristics.

Accordingly, it is desirable to provide systems and methods for removing heat from aircraft components that are reliable and durable in operational service and that are capable of removing heat from these components during high speed flight.

It is further desirable to collect heat for improving the propulsion efficiency of the aircraft.

Disclosure of Invention

In one aspect, a gas turbine engine includes: a compressor, a combustor and a turbine in serial flow relationship; a heat exchanger having an inlet, an outlet, and an inner surface coated with a catalyst, the heat exchanger being located upstream of the compressor; 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 the reformed hydrocarbon fuel from the heat exchanger.

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

Drawings

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

FIG. 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.

FIG. 2 is a schematic cross-sectional view of a tube of an exemplary heat exchanger as described herein.

FIG. 3 is a cross-sectional schematic view of an exemplary embodiment of a heat exchanger having an array of multiple tubular structures as described herein.

Fig. 4 is a schematic plan view of an exemplary embodiment of a heat exchanger having an array of multiple tubular structures as described herein, and showing inlet and outlet manifolds.

FIG. 5 is a cross-sectional schematic 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.

FIG. 6 is a schematic illustration of a gas turbine engine that utilizes a heat exchanger as described herein downstream of the last turbine stage to reform a hydrocarbon fuel.

FIG. 7 is a schematic view of a gas turbine engine similar to FIG. 6 utilizing a heat exchanger as described herein between turbine stages to reform a hydrocarbon fuel.

FIG. 8 is a schematic diagram illustrating a method of operating the exemplary embodiment of FIG. 5.

FIG. 9 is a cross-sectional schematic view of an exemplary embodiment of a gas turbine engine, similar to the embodiment of FIG. 5, suitable for use as an aircraft propulsion system but suitable for including a superburner system.

FIG. 10 is a schematic illustration of an exemplary embodiment of a heat exchanger system including a plurality of heat exchangers similar to the heat exchanger of FIG. 4, the system configured to operate with the gas turbine engine of FIG. 9.

Corresponding reference characters indicate corresponding parts throughout the several views. The examples set forth herein illustrate exemplary embodiments of the disclosure, and these examples should not be construed as limiting the scope of the disclosure in any way.

Detailed Description

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

The following description is presented to enable any person skilled in the art to make and use the described embodiments, which are intended to be useful in the practice of the invention. Various modifications, equivalents, changes, and substitutions will still be 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.

All directional references (e.g., radial, axial, proximal, distal, upper, lower, upward, downward, left, right, lateral, forward, rearward, top, bottom, upward, downward, vertical, horizontal, clockwise, counterclockwise, upstream, downstream, forward, rearward, etc.) are only used for identification purposes to aid the reader's understanding of the present invention, and do not create limitations, particularly as to the position, orientation, or use of the invention. Unless otherwise specified, connection references (e.g., attached, coupled, connected, and engaged) are to be construed broadly and may include intermediate members between a collection of elements and relative movement between elements. As such, joinder references do not necessarily infer that two elements are directly connected and in fixed relation to each other. The exemplary drawings are for illustrative purposes only and the dimensions, locations, order and relative sizes reflected in the accompanying drawings may vary.

Unless specified otherwise, the terms "coupled," "secured," "attached," and the like refer to a direct coupling, securing, or attachment, as well as an indirect coupling, securing, or attachment through one or more intermediate components or features.

The singular forms "a", "an" and "the" include plural referents unless the context clearly dictates otherwise. Furthermore, the suffix "(s)" as used herein is generally intended to include both the singular and the plural of the term that it modifies, thereby including one or more of that term.

As used herein, the term "or" is not meant to be exclusive, and refers to the presence of at least one of the recited components (e.g., materials), as well as including instances where combinations of the recited components may be present, unless the context clearly indicates otherwise.

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

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

As used herein, the terms "may" and "may be" refer to the likelihood of occurring in a set of circumstances; having a particular property, characteristic or function; and/or by expressing capabilities associated with qualifying verbs, another verb can be qualified or by one or more of the possibilities. Thus, use of "may" and "may be" means that the modifier is apparently suitable, capable of, or adapted to the stated capacity, function, or usage, while taking into account that in some cases the modifier may sometimes not be suitable, capable of, or suitable. For example, in some cases, an event or capacity may be expected, while in other cases, no event or capacity may occur. This distinction is expressed by the terms "may" and "may be" (capture).

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

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

FIG. 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 aircraft designed to operate at above sonic speeds (i.e., above mach 1), and more particularly to aircraft designed to operate at hypersonic flight regimes above mach 5 (e.g., in the range of mach 5 to mach 10).

In the configuration shown in FIG. 1, 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 that functions as an aircraft propulsion system. An outer skin surface 16 covers at least a portion of the fuselage 11 and the wing 12. In high-speed flight, the movement of the aircraft 10 through the atmosphere causes heating of aircraft surfaces (e.g., the skin surface 16 on the exterior of the aircraft 10), and in particular, heat build-up in the region 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 atmospheric air is reduced to subsonic speeds, which increases the pressure within the gas turbine engine. Each of these locations may utilize a system for removing heat from an aircraft component as described herein. The system of interest is further described in commonly assigned, co-pending U.S. patent application serial No. 16/681,292, filed on 12.11.2019, the specification and drawings of which are incorporated in appendix 1 and appendix 2, respectively.

FIG. 2 is a cross-sectional schematic 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 comprises at least one tubular structure 21 having an inner surface 22 and an outer surface 23. As will be described below, the inner surface 22 includes a coating 24 and an interior space 25 located inside the inner surface 22 and the coating 24, through which interior space 25 fluid may pass.

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

As used herein, the terms "hydrocarbon cracking," "cracked hydrocarbon," or any anomaly thereof, refer to, but are not limited to, a process in which hydrocarbons are cracked in a facility to obtain a material having smaller molecules. The hydrocarbons may include ethane, heptane, liquid petroleum gas, naphtha, gas oils, substrates from atmospheric and vacuum distillation of crude oil, or any combination thereof.

As used herein, the term "coke" or any variant thereof refers to, but is not limited to, carbonaceous solids or liquids, or particles or macromolecules that form carbonaceous solids or liquids, derived from coal, petroleum, wood, hydrocarbons and other carbonaceous materials.

As used herein, the term "perovskite material" or any synonym thereof refers to, but is not limited to, having ABO₃A perovskite structure and having the formula A_aB_bO_3-δAny material of (1), wherein 0.9<a≤1.2；0.9<b≤1.2；-0.5<δ<0.5; a comprises a first element selected from the group consisting of calcium (Ca), strontium (Sr), barium (Ba), lithium (Li), sodium (Na), potassium (K), rubidium (Rb), and any combination thereof, and optionally a second element selected from the group consisting of 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.

In some embodiments, the perovskite material may have the formula n (a)_aB_bO_3-δ) Wherein n is 2, 3, 4, 8, etc., and formula A_aB_bO_3-δIs a simplified form thereof. In some embodiments, in ABO₃In the perovskite structure, a cations are surrounded by twelve anions in cubic octahedral coordination, B cations are surrounded by six anions in octahedral coordination, and oxygen anions are coordinated by two B cations and four a cations. In some embodiments, ABO₃BO shared by corners for perovskite structures₆And (4) constructing an octahedron. In some embodiments, ABO₃The perovskite structure includes a deformed derivative. The deformation may be due to a rotation or tilt of a regular, rigid octahedron, or due to a deformed BO₆The presence of octahedra. In some embodiments, ABO₃The perovskite structure is cubic. In some embodiments, ABO₃The perovskite structure is hexagonal.

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).

In some embodiments, a comprises a combination of the first element and the second element. The second element may be a single element or a combination of elements selected from the group consisting of 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).

Likewise, B may be a single element or a combination of elements 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), and zirconium (Zr).

In some embodiments, the perovskite material comprises SrCeO₃，SrZr_0.3Ce_0.7O₃，BaMnO₃，BaCeO₃，BaZr_0.3Ce_0.7O₃，BaZr_0.3Ce_0.5Y_0.2O₃，BaZr_0.1Ce_0.7Y_0.2O₃，BaZrO₃，BaZr_0.7Ce_0.3O₃，BaCe_0.5Zr_0.5O₃，BaCe_0.9Y_0.1O₃，BaCe_0.85Y_0.15O₃Or BaCe_0.8Y_0.2O₃. For example, for SrCeO₃A is Sr, a ═ 1, B is Ce, B ═ 1, and δ is 0. For SrZr_0.3Ce_0.7O₃A is Sr, a ═ 1, B is a combination of Zr and Ce, B ═ 1, and 6 ═ 0. For BaMnO₃A is Ba, a ═ 1, B is Mn, B ═ 1, and 6 ═ 0. For BaCeO₃A is Ba, a ═ 1, B is Ce, B ═ 1, and 6 ═ 0. For BaZr_0.3Ce_0.7O₃A is Ba, a ═ 1, B is a combination of Zr and Ce, B ═ 1, and δ ═ 0. For BaZr_0.3Ce_0.5Y_0.2O₃A is Ba, a ═ 1, B is Zr, a combination of Ce and Y, B ═ 1, and δ ═ 0.

In some embodiments, the perovskite material comprises La_0.1Ba_0.9Ce_0.7Zr_0.2Y_0.1O₃，Ce_0.1Ba_0.9Ce_0.7Zr_0.2Y_0.1O_3.05，Ce_0.5Ba_0.5Ce_0.7Zr_0.2Y_0.1O_3.45，Y_0.1Ba_0.9Ce_0.7Zr_0.2Y_0.1O₃，Y_0.5Ba_0.5Ce_0.7Zr_0.2Y_0.1O_3.2，Bi_0.1Ba_0.9Ce_0.7Zr_0.2Y_0.1O₃，Bi_0.5Ba_0.5Ce_0.7Zr_0.2，Y_0.1O_3.2，Pr_0.1Ba_0.9Ce_0.7Zr_0.2Y_0.1O₃Or Pr_0.5Ba_0.5Ce_0.7Zr_0.2Y_0.1O_3.2. For La_0.1Ba_0.9Ce_0.7Zr_0.2Y_0.1O₃Where a is a combination of Ba and La, the first element is La, the second element is Ba, a ═ 1, B is Ce, Zr and Y, a combination of B, B ═ 1, and δ ═ 0. For Ce_0.1Ba_0.9Ce_0.7Zr_0.2Y_0.1O_3.05And Ce_0.5Ba_0.5Ce_0.7Zr_0.2Y_0.1O_3.45A 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-0.05 and-0.45, respectively. For Y_0.1Ba_0.9Ce_0.7Zr_0.2Y_0.1O₃And Y_0.5Ba_0.5Ce_0.7Zr_0.2Y_0.1O_3.2A is a combination of Y and Ba, the first element is Y, the second element is Ba, a ═ 1, B is Ce, a combination of Zr and Y, B ═ 1, and δ is 0 and-0.2, respectively. For Bi_0.1Ba_0.9Ce_0.7Zr_0.2Y_0.1O₃And Bi_0.5Ba_0.5Ce_0.7Zr_0.2Y_0.1O_3.2A is a combination of Bi and Ba, the first element is Bi, the second element is Ba, a ═ 1, B is Ce, a combination of Zr and Y, B ═ 1, and δ is 0 and-0.2, respectively. Also for Pr_0.1Ba_0.9Ce_0.7Zr_0.2Y_0.1O₃And Pr_0.5Ba_0.5Ce_0.7Zr_0.2Y_0.1O_3.2A is a combination of Pr and Ba, the first element is Pr, the second element is Ba, a ═ 1, B is Ce, a combination of Zr and Y, B ═ 1, and δ is 0 and-0, respectively.2。

In some embodiments, the perovskite material comprises BaZr_0.3Ce_0.7O₃。

As used herein, the term "conditioning material" or any anomaly thereof refers to any material that reduces the yield of carbon monoxide in hydrocarbon cracking. The conditioning material may comprise one material or a combination of materials. In some embodiments, the conditioning material comprises zirconia, doped zirconia, or any precursor or combination thereof.

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

In some embodiments, the process for cracking hydrocarbons operates at a temperature range of about 480 ℃ to about 600 ℃ in the presence of steam, the hydrocarbons comprise bottoms from atmospheric and vacuum distillation of crude oil, and the weight percent of steam is in the range of about 1 wt% to about 2 wt%.

The perovskite material may or may not chemically react with the tuning material. Thus, the inner surface may comprise a combination or reaction product of the perovskite material and the tuning material. In some embodiments, the inner surface comprises a perovskite material, a tuning material, and a combination of reaction products of the perovskite material and the tuning material.

The perovskite material and the tuning material may be applied in a coating to the device using different methods (e.g., air plasma spraying, slurry coating, sol-gel coating, and solution coating). In some embodiments, the perovskite material and the tuning material are coated using a slurry coating method.

The amount of conditioning material and perovskite material in the slurry may vary so long as a continuous, robust, carbon monoxide-reducing, and coking-resistant coating is formed, depending on the particular conditioning material and perovskite material used and the operating conditions of the coating. In some embodiments, the weight ratio of perovskite material to tuning material is 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.

The paste may further comprise an organic binder, an inorganic binder, a wetting agent, a solvent or any combination thereof to enhance the wetting ability of the paste, adjust the viscosity of the paste 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 percent of the conditioning 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%.

In some embodiments, the slurry comprises a perovskite material, a conditioning material, cerium oxide, yttrium oxide, glycerol, and polyvinyl alcohol (PVA).

The slurry may be applied to the equipment by various techniques, such as at least one of sponge 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 in the slurry). In some embodiments, the slurry is applied by filling and draining (i.e., filling the slurry in the article to be coated and then draining the slurry by, for example, gravity).

Additional description of cracking processes and systems may be found in the published patent texts, all of which are incorporated by reference: US9,499,747, WO2015105589a1, CA2821249a1, US20170260460, CA2932461a1, WO2015088671a1 and US 20170022428.

After the slurry is applied to the equipment, a sintering process may be performed. As used herein, the term "sintering" or any variant thereof refers to, but is not limited to, a method of heating a material in a sintering furnace or other heater apparatus. In some embodiments, the sintering temperature is in the range of about 850 ℃ to about 1700 ℃. In some embodiments, the sintering is at about 1000 ℃. In some embodiments, sintering is performed in an inert atmosphere (such as argon or nitrogen). In some embodiments, heat treatment in air is performed prior to sintering to form an oxide layer on the inner surface of the tube that improves the adhesion of the coating.

In operation, a hydrocarbon, such as aircraft fuel, is supplied to the interior space 25 along with an oxygenate, such as steam, which may be provided in the form of liquid water (which evaporates in the presence of sufficient heat) or an oxygenated fuel, such as ethanol or methanol. The cracking of the hydrocarbon fuel that occurs within the interior space 25 is an endothermic reaction in which all carbon-carbon bonds are broken and hydrogen and methylene radicals are formed. This highly endothermic process takes heat away from the tubular structure 21 and its surroundings and thus cools the tubular structure 21 and its surroundings. Thus, the cracking of the hydrocarbons transforms the tubular structure 21 into the heat exchanger 20 and serves as a heat removal system that may 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 may eventually impede the flow of fuel and reduce the capacity of the heat exchanger 20. The reformed fuel exiting heat exchanger 20 may then be used as fuel for the propulsion of the aircraft.

Fig. 3 is a cross-sectional schematic view 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 tubular structure 21 having an interior space 25. The heat exchanger 20 is positioned proximate to the aircraft component 30, and it is desired to remove thermal energy (heat) from the aircraft component 30. The heat exchanger 20 may be joined to or integrally formed with the aircraft component by conventional or additive manufacturing techniques known in the art. Alternatively, the heat exchanger 20 may be used as an air-fuel heat exchanger, and thus to remove heat from the airflow flowing through the array of tubular structures 21.

Fig. 4 is a schematic plan view of an exemplary embodiment of a heat exchanger 20 as described herein, the heat exchanger 20 having an array of a plurality of tubular structures 21 and showing an inlet manifold 26 and an outlet manifold 28, respectively. Manifolds 26 and 28 fluidly couple the inlet and outlet ends of the tubular structure 21 to a common inlet 27 and outlet 29, respectively. The inlet 27 may in turn be fluidly coupled to a pipe or conduit that serves as a source of hydrocarbon fuel and/or steam (water). After cracking occurs within the tubular structure 21 of the heat exchanger 20, the outlet 29 may in turn be fluidly coupled to a pipe or conduit that is the destination or recipient of the reformed fuel.

In addition to multiple tubular structures sharing a common manifold, the heat exchangers described herein may actually comprise multiple heat exchangers that are separately manifolded. A control system including sensors, valves and/or electronically controlled actuators may control flow between and among the various tubular structures and/or the plurality of heat exchangers. This can provide flexibility for different operating conditions and can cycle between heat exchangers if it is necessary or desirable to take certain tubular structures or certain heat exchangers offline to remove any coke deposits that may accumulate during operation.

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

FIG. 6 is a schematic illustration of a gas turbine engine 15, the gas turbine engine 15 utilizing a heat exchanger 20 as described herein to reform a hydrocarbon fuel downstream of a last turbine stage after a high pressure turbine 35 and a low pressure turbine 36. The heat contained in the residual air stream leaving the low pressure turbine 36 provides the energy for cracking the fuel, which is then discharged to the atmosphere 37. As previously discussed, a reformate fuel may be supplied to the combustor 34 to fuel the gas turbine engine 15.

FIG. 7 is a schematic illustration of a gas turbine engine 15 similar to FIG. 6, the 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.

For any of the exemplary embodiments described herein, multiple heat removal systems may be employed in series or in parallel, and may share inlets and outlets, or may be plumbed with separate sources and destinations of hydrocarbon fuel and reformed fuel, respectively.

For parallel heat removal systems, all systems may be operated simultaneously, or some systems may be deactivated to restore and remove coke deposits or adjust heat removal capacity levels during various phases of aircraft operation.

FIG. 8 is a schematic diagram illustrating a method of operating the exemplary embodiment of FIG. 5. In FIG. 8, hot inlet air 33 enters the front of the gas turbine engine 15. The heat exchanger 20 may be a single element arranged annularly and/or symmetrically about the engine centerline 40, or may be a plurality of individual heat exchanger elements, the heat exchanger 20 receiving liquid fuel 41 through the inlet 27. The liquid fuel 41 then becomes vaporized fuel 42, and the vaporized fuel 42 is then mixed with an oxygen-containing medium (e.g., bleed air 44 from a gas turbine engine compressor or water from, for example, a storage tank within the heat exchanger 20). The heat exchanger 20 then reforms the fuel using the available heat from the hot inlet air 33 and the oxygen-containing medium such that the reformed fuel 43 then exits the heat exchanger 20 through the outlet 29. Conditioned air 45, which is cooler than hot inlet air 33, then flows to gas turbine engine compressor 32 in the embodiment of fig. 5 and 9, or to the atmosphere in the embodiment of fig. 6, or to low pressure turbine 37 in the embodiment of fig. 7. Depending on the installation location of the heat exchanger 20, the hot inlet air 33 in the embodiment of fig. 5 and 9 may also be exhausted from the high pressure turbine 35 in the embodiment of fig. 7.

FIG. 9 is a cross-sectional schematic view of an exemplary embodiment of a gas turbine engine 15, the gas turbine engine 15 being adapted for use as an aircraft propulsion system similar to the embodiment of FIG. 5 but adapted to include a superburner system 70. In the embodiment of FIG. 9, the overfire burner 70 is similar to an intensifier or afterburner, but is specifically designed for high speed and/or supersonic travel.

As with fig. 4, in fig. 9, the heat exchanger 20 is located in the inlet section 31 of the gas turbine engine 15. The liquid hydrocarbon fuel 41 is fed from the storage tank 52 through a supply line 58 to the first splitter 48, the first splitter 48 acting as a control valve, wherein a portion of the liquid hydrocarbon fuel 41 may be fed to a fuel nozzle 54 located within the combustor section 34 of the gas turbine engine 15. Liquid fuel 41 may flow from the first splitter 48 to the second splitter or control valve 50, where the liquid fuel 41 may be directed through an inlet 27 to the heat exchanger 20 via a supply line 58, or selectively to fuel nozzles 56 located behind the high-pressure turbine 35 and the low-pressure turbine 36, respectively, in a super-burner 70. A control valve 46 controls the flow of compressor discharge air or bleed air 44 to the reforming catalyst in the heat exchanger 20. Reformed (gaseous) fuel 60 exits heat exchanger 20 through outlet 29 and flows to a afterburner fuel nozzle 56 located in a afterburner 70.

The overfire fuel nozzle 56 may be configured for "dual fuel" operation, i.e., operation with either a liquid hydrocarbon fuel or a gaseous reformate fuel or both, as desired for the particular operating conditions present. The control valve or splitter may be operated by a control system that selectively directs liquid fuel to the gas turbine combustor and/or the overfire burner, and also controls the flow of reformed fuel from the heat exchanger to the overfire burner. The proportion of fuel flowing to the main gas turbine engine combustor and the afterburner can be adjusted as desired, and in high speed or hypersonic flight regimes, a split can be utilized, for example, approximately 5% to the main gas turbine engine combustor and 95% to the afterburner. A low pressure of e.g. 3-5bar may be experienced in the afterburner, which allows the afterburner fuel supply system to be operated also at a low pressure, thereby minimizing the risk of leakage. Higher pressures may be experienced in the gas turbine combustor, requiring liquid fuel to be supplied at higher pressures in the system.

FIG. 10 is a schematic illustration of an exemplary embodiment of a heat exchanger system 20, the heat exchanger system 20 including a plurality of heat exchangers 20A, 20B similar to the heat exchanger 20 of FIG. 4, the system configured to operate with the gas turbine engine 15 of FIG. 9. In the embodiment of FIG. 10, the plurality of heat exchangers or heat exchanger branches 20A and 20B may be supplied with liquid hydrocarbon fuel 41 and oxygen-containing compressor discharge air 44, respectively, via control valves 71, and similar control valves 71 may direct the reformed fuel to the gas turbine engine combustor fuel nozzles 54 or the overfire fuel nozzles 56 located at the turbine outlet. As previously mentioned, multiple (and possibly two or more) heat exchangers may be operated together, or one or more heat exchangers may be deactivated for cleaning or based on lower demand for reformed fuel.

Various 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 propulsion fuels (such as RPS1 and RPS2), diesel fuels (such as D2 and D4), blends thereof, mixtures thereof, and combinations thereof.

All publications, patents and patent applications cited herein, whether supra or infra, are hereby incorporated by reference in their entirety to the same extent as if each individual publication, patent or patent application were specifically and individually indicated to be incorporated by reference. It should be understood that all or portions of any patent, publication, or other disclosure material that is said to be incorporated by reference herein are incorporated herein only to the extent that the incorporated material does not conflict with existing definitions, statements, or other disclosure material 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 said to be incorporated by reference herein, but which conflicts with existing definitions, statements, or other disclosure material set forth herein will only be incorporated to the extent that no conflict arises between that incorporated material and the existing disclosure material.

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 context clearly dictates otherwise.

Unless defined otherwise, all technical and scientific terms used herein have the same meaning as commonly understood by one of ordinary skill in the art to which this invention belongs. Although many methods and materials similar 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.

It should be noted that, as used in this disclosure, the terms "comprises," "comprising," and other derivatives from the root term "comprise" 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 preclude the presence or addition of one or more other features, elements, integers, steps, components, or groups thereof.

As required, detailed embodiments of the present invention are disclosed herein; however, it is to be understood that the disclosed embodiments are merely exemplary of the invention, which can be embodied in various forms. Therefore, specific structural and functional details disclosed herein are not to be interpreted as limiting, but merely as a basis for the claims and as a representative basis for teaching one skilled in the art to variously employ the present invention in virtually any appropriately detailed structure.

The various features, aspects, and advantages of the disclosure may also be embodied in any permutation of aspects of the disclosure, including but not limited to the following technical solutions defined in the enumerated aspects:

1. in one aspect, a gas turbine engine includes: a compressor, a combustor and a turbine in serial flow relationship; a heat exchanger having an inlet, an outlet, and an inner surface coated with a catalyst, the heat exchanger being located upstream of the compressor; 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 the reformed hydrocarbon fuel from the heat exchanger.

2. The gas turbine engine of any preceding item, wherein the distribution system delivers reformed hydrocarbon fuel to the gas turbine engine.

3. The gas turbine engine of any preceding item, wherein the gas turbine engine comprises a superburner or an augmentor.

4. The gas turbine engine of any preceding item, wherein the distribution system delivers reformed hydrocarbon fuel to the afterburner or the augmentor.

5. The gas turbine engine of any preceding claim, wherein the afterburner or the augmentor comprises an afterburner fuel nozzle, and wherein the distribution system is configured to selectively deliver liquid hydrocarbon fuel or reformed hydrocarbon fuel to the fuel nozzle.

6. The gas turbine engine of any preceding item, wherein the distribution system delivers liquid hydrocarbon fuel to the combustor and reformed gaseous hydrocarbon fuel to the super burner.

7. The gas turbine engine according to any preceding item, wherein the hydrocarbon fuel is an aircraft Jet fuel, such as Jet-a, JP-4 and JP-8, gasoline, kerosene, rocket propulsion fuels, such as RPS1 and RPS2, diesel fuels, such as D2 and D4, blends thereof, mixtures thereof, and combinations thereof.

8. The gas turbine engine according to any preceding item, wherein the oxygen source is an oxygenate, such as steam, liquid water, or an oxygenated fuel, such as ethanol or methanol.

9. The gas turbine engine of any preceding item, wherein the system comprises a plurality of heat exchangers or heat exchanger branches.

10. The gas turbine engine of any preceding item, wherein the plurality of heat exchangers are arranged in series or in parallel.

11. A method of operating a gas turbine engine having a compressor, a combustor and a turbine in serial flow relationship, the method comprising the steps of: providing a heat exchanger upstream of the compressor, the heat exchanger in fluid communication with a hydrocarbon fuel source and a water source and having an inner 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.

12. The method according to any of the preceding claims, wherein the hydrocarbon fuel is an aircraft Jet fuel, such as Jet-a, JP-4 and JP-8, gasoline, kerosene, rocket propulsion fuels, such as RPS1 and RPS2, diesel fuels, such as D2 and D4, blends thereof, mixtures thereof, and combinations thereof.

12. The method of any preceding claim, wherein the oxygen source is an oxygenate, such as steam, liquid water, or an oxygenated fuel, such as ethanol or methanol.

13. The method of any preceding claim, wherein the method comprises a plurality of heat exchangers.

14. The method of any preceding item, wherein the plurality of heat exchangers are arranged in series or in parallel.

15. The method of any preceding item, wherein the distribution system delivers the reformed fuel to the gas turbine engine.

16. The gas turbine engine of any one or more of the preceding items 1 to 10.

17. The method according to any one or more of the preceding items 11 to 15.

While this disclosure has been described as having exemplary embodiments, the present disclosure can be further modified within the spirit and scope of this disclosure. This application is therefore intended to cover any variations, uses, or adaptations of the disclosure using its general principles. Further, this application is intended to cover such departures from the present disclosure as come within known or customary practice in the art to which this invention pertains and which fall within the limits of the appended claims.