Gas turbine engine and method of operating the same

Abstract

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 reformed hydrocarbon fuel from the heat exchanger.

Description

Technical Field

[0001] The present invention relates to aircraft components such as gas turbine engines, and more particularly to those aircraft components that are exposed to high levels of heat during operation (such as during high-speed flight). Background Art

[0002] 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, due to the speed at which the aircraft travels through the atmosphere, these components as well as the external components of the aircraft itself are exposed to heat from skin friction. In addition to the air temperature entering the gas turbine engine, components such as the aircraft skin, the leading edges of structures (such as wings, spines, and control surfaces), and the engine inlet are also particularly affected.

[0003] These heat sources can cause the surface and internal temperatures of such aircraft components to exceed their structural and / or operating capabilities, thereby requiring expensive or special materials and altering operating characteristics.

[0004] Accordingly, it is desirable to provide systems and methods for removing heat from aircraft components that are reliable and durable in service and capable of removing heat from these components during high-speed flight.

[0005] It is further desirable to collect heat for use in improving the propulsion efficiency of the aircraft. Summary of the Invention

[0006] 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 reformed hydrocarbon fuel from the heat exchanger.

[0007] These and other features, aspects, and advantages of the present invention will be better understood with reference to the following description and the 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. Brief Description of the Drawings

[0008] A complete and enabling disclosure of the invention, including the best mode thereof, for the ordinary skilled person in the art, is set forth in the specification, reference being made to the accompanying drawings, in which:[[]]

[0009] Figure 1 is a perspective view of an exemplary embodiment of a high-speed aircraft adapted to implement the heat removal apparatus and method described herein.

[0010] Figure 2 is a cross-sectional view of a tube of an exemplary heat exchanger as described herein.

[0011] Figure 3 is a cross-sectional view of an exemplary embodiment of a heat exchanger having an array of a plurality of tubular structures as described herein.

[0012] Figure 4 is a plan view of an exemplary embodiment of a heat exchanger having an array of a plurality of tubular structures as described herein, and shows inlet and outlet manifolds.

[0013] Figure 5 is a cross-sectional view of an exemplary embodiment of a gas turbine engine adapted to be used as an aircraft propulsion system and incorporating a heat exchanger as described herein.

[0014] Figure 6 is a schematic view of a gas turbine engine that utilizes a heat exchanger as described herein downstream of a last turbine stage to reform a hydrocarbon fuel.

[0015] Figure 7 is similar to Figure 6 schematic view of a gas turbine engine that utilizes a heat exchanger as described herein between turbine stages to reform a hydrocarbon fuel.

[0016] Figure 8 is a schematic view showing a method of operating Figure 5 exemplary embodiment.

[0017] Figure 9 is a cross-sectional view of an exemplary embodiment of a gas turbine engine similar to Figure 5 embodiment adapted to be used as an aircraft propulsion system but adapted to include a scramjet system.

[0018] Figure 10 is a schematic view of an exemplary embodiment of a heat exchanger system including a plurality of heat exchangers including a heat exchanger similar to Figure 4 configured to operate with a Figure 9 gas turbine engine.

[0019] In several views, corresponding reference numerals indicate corresponding parts. The examples set forth herein illustrate exemplary embodiments of the present disclosure, and these examples should not be construed as limiting the scope of the present disclosure in any way. Detailed Description

[0020] 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. Identical or similar designations have been used in the drawings and the description to refer to identical or similar parts of the invention.

[0021] The following description is provided to enable a person skilled in the art to make and use the described embodiments expected to implement the invention. However, various modifications, equivalents, variations, and alternatives will still be apparent to a person skilled in the art. Any and all such modifications, variations, equivalents, and alternatives are intended to fall within the spirit and scope of the invention.

[0022] All directional references (e.g., radial, axial, proximal, distal, upper, lower, upward, downward, left, right, lateral, front, rear, top, bottom, up, down, vertical, horizontal, clockwise, counterclockwise, upstream, downstream, forward, backward, etc.) are for identification purposes only to assist the reader in understanding the invention and do not impose a limitation on the invention, particularly as to its position, orientation, or use. Unless otherwise noted, connection references (e.g., attached, coupled, connected, and joined) will be construed broadly and may include intermediate members between assemblies of elements and relative movement between elements. Thus, a connection reference does not necessarily infer that two elements are directly connected and in a fixed relationship to each other. Exemplary drawings are for illustrative purposes only, and the dimensions, positions, sequences, and relative sizes reflected in the attached drawings may vary.

[0023] Unless otherwise noted, the terms "coupled," "fixed," "attached to," etc. refer to direct coupling, fixing, or attachment, as well as indirect coupling, fixing, or attachment through one or more intermediate components or features.

[0024] Unless the context clearly dictates otherwise, the singular forms "a," "an," and "the" include plural referents. In addition, the suffix "(s)" as used herein is generally intended to include both the singular and plural of the term it modifies, thereby including one or more of that term.

[0025] As used herein, unless the context clearly dictates otherwise, the term "or" does not mean exclusive and refers to the presence of at least one of the recited components (e.g., materials), as well as to instances that may include combinations in which the recited components may be present.

[0026] As used throughout this specification and the claims, approximating language is used to modify any quantitative representation that can vary without resulting in a change in the basic function associated therewith. Thus, values modified by terms such as "about," "approximately," and "substantially" are not limited to the precise values specified. In at least some instances, the approximating language may correspond to the precision of the instrument used to measure the value, or the precision of the method or machine used to construct or manufacture the components and / or systems. For example, the approximating language may refer to being within a range of 10%.

[0027] Here and throughout the specification and claims, range limitations are combined and interchanged, and such ranges are recognized and include all the sub-ranges contained therein unless the context or language indicates otherwise. For example, all ranges disclosed herein include the endpoints, and the endpoints are combined independently of each other.

[0028] As used herein, the terms "may" and "might" denote the possibility of occurring in a set of circumstances; having a particular property, characteristic, or function; and / or qualify another verb by expressing one or more of the ability, capacity, or likelihood associated with the qualifying verb. Thus, the use of "may" and "might" indicates that the modifier is apparently suitable, able, or adapted to the specified capacity, function, or usage, while taking into account that in some cases, the modifier may sometimes not be suitable, unable, or inappropriate. For example, in some cases, an event or capacity may be expected, while in other cases, the event or capacity may not occur. This distinction is captured by the terms "may" and "might."

[0029] References throughout the specification to "some embodiments" and the like refer to particular elements (e.g., features, structures, and / or characteristics) described in connection with the invention being included in at least one embodiment described herein and may or may not be present in other embodiments. Additionally, it should be understood that the described features of the invention may be combined in any suitable manner in various embodiments.

[0030] Aspects of the present invention are more fully explained with reference to the exemplary embodiments discussed below. It should be understood that, in general, the features of one embodiment may also be used in conjunction with the features of another embodiment, and these embodiments are not intended to limit the scope of the present invention.

[0031] Figure 1 is a perspective view of an exemplary embodiment of a high-speed aircraft 10 adapted to implement the thermal removal devices and methods described herein. As used herein, the term "high-speed aircraft" is intended to refer to an aircraft designed to operate at speeds higher than the speed of sound (i.e., higher than Mach 1), and more particularly to an aircraft designed to operate in a hypersonic flight regime at speeds higher than Mach 5 (e.g., in the range of 5 Mach to 10 Mach).

[0032] In Figure 1 the configuration shown, the exemplary hypersonic vehicle 10 includes a fuselage 11, wings 12, a vertical stabilizer 13, leading edges 14 of the wings 12 and the fuselage 11, and a gas turbine engine 15 that serves as the vehicle propulsion system. An outer skin surface 16 covers at least a portion of the fuselage 11 and the wings 12. During hypersonic flight, the movement of the vehicle 10 through the atmosphere causes heating of the vehicle surface (such as the skin surface 16 on the exterior of the vehicle 10), and in particular causes heat buildup in the regions of the leading edges 14 of the wings 12 and the fuselage 11. Hypersonic operation also causes the gas turbine engine 15 to experience high-temperature operation because the atmospheric air is reduced to subsonic speed, which increases the pressure within the gas turbine engine. Each of these locations can utilize a system for removing heat from the vehicle components as described herein. The system of interest is further described in the commonly assigned, co-pending U.S. Patent Application Serial No. 16 / 681,292, filed on November 12, 2019, the specification and drawings of which are incorporated herein by reference as Appendix 1 and Appendix 2, respectively.

[0033] Figure 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 includes at least one tubular structure 21 having an inner surface 22 and an outer surface 23. As will be described hereinafter, the inner surface 22 includes a coating 24 and an internal space 25 located on the inner side of the inner surface 22 and the coating 24 through which a fluid can pass.

[0034] The inner surface of the tubular structure 21 accessible to a fluid (such as a hydrocarbon fuel in the internal space 25) includes a coating 24 of a perovskite material and a modulating material.

[0035] As used herein, the term "hydrocarbon cracking", "cracking hydrocarbons", or any variant thereof refers to, but is not limited to, a process in which hydrocarbons are cracked in a device to obtain materials having smaller molecules. Hydrocarbons can include ethane, heptane, liquefied petroleum gas, naphtha, gas oil, bottoms from atmospheric and vacuum distillation of crude oil, or any combination thereof.

[0036] As used herein, the term "coke" or any variant thereof refers to, but is not limited to, a carbonaceous solid or liquid, or particles or macromolecules that form a carbonaceous solid or liquid, which are derived from coal, petroleum, wood, hydrocarbons, and other carbon-containing materials.

[0037] As used herein, the term "perovskite material" or any variant thereof refers to, but is not limited to, a material having an ABO3 perovskite structure and having the 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, and 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.

[0038] 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 cation is surrounded by twelve anions in a cubic octahedral coordination, the B cation is surrounded by six anions in an octahedral coordination, and the oxygen anion is 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 a deformed derivative. The deformation may be due to the rotation or tilting of regular, rigid octahedra, or due to the presence of deformed BO6 octahedra. In some embodiments, the ABO3 perovskite structure is cubic. In some embodiments, the ABO3 perovskite structure is hexagonal.

[0039] 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).

[0040] In some embodiments, A comprises 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).

[0041] Similarly, 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).

[0042] In some embodiments, the perovskite material comprises 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 0.2 O3. For example, for SrCeO3, A is Sr, a = 1, B is Ce, b = 1, and δ = 0. For SrZr 0.3 Ce 0.7For 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. For BaZr 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.

[0043] In some embodiments, the perovskite material comprises 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 Zr 0.2 Y 0.1 O3, or Pr 0.5 Ba 0.5 Ce 0.7 Zr0.2 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 -0.05 and -0.45 respectively. For Y 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 0 and -0.2 respectively. For Bi 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 0 and -0.2 respectively. Similarly, for Pr 0.1 Ba 0.9 Ce 0.7 Zr 0.2 Y 0.1 O3 and Pr 0.5 Ba 0.5 Ce 0.7Zr 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 0 and -0.2 respectively.

[0044] In some embodiments, the perovskite material comprises BaZr 0.3 Ce 0.7 O3.

[0045] As used herein, the term "modulating material" or any variant thereof refers to any material that reduces the carbon monoxide yield in hydrocarbon cracking. The modulating material can include a single material or a combination of multiple materials. In some embodiments, the modulating material includes zirconia, doped zirconia, or any precursor or combination thereof.

[0046] In some embodiments, the method for cracking hydrocarbons operates in the temperature range of about 700 °C to about 900 °C in the presence of steam, the weight ratio of steam to hydrocarbon is in the range of about 3:7 to about 7:3, and the hydrocarbon includes ethane, heptane, liquefied petroleum gas, naphtha, gas oil, or any combination thereof.

[0047] In some embodiments, the method for cracking hydrocarbons operates in the temperature range of about 480 °C to about 600 °C in the presence of steam, the hydrocarbon contains substrates from atmospheric and vacuum distillation of crude oil, and the weight percentage of steam is in the range of about 1 wt% to about 2 wt%.

[0048] The perovskite material may or may not chemically react with the modulating material. Thus, the inner surface may include a combination or reaction product of the perovskite material and the modulating material. In some embodiments, the inner surface includes a combination of the perovskite material, the modulating material, and the reaction product of the perovskite material and the modulating material.

[0049] The perovskite material and the modulating material can be applied to the coating of the device using different methods (such as air plasma spraying, slurry coating, sol-gel coating, and solution coating). In some embodiments, the perovskite material and the modulating material are coated using the slurry coating method.

[0050] As long as a continuous, robust, carbon monoxide-reducing, and coking-resistant coating is formed, the amounts of the modulating material and the perovskite material in the slurry can vary, depending on the specific modulating material and perovskite material used and the operating conditions of the coating. In some embodiments, the weight ratio of the perovskite material to the modulating 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 modulating material.

[0051] The slurry 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 slurry, adjust the viscosity of the slurry, 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 regulating 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%.

[0052] In some embodiments, the slurry comprises a perovskite material, a regulating material, cerium oxide, yttrium oxide, glycerol, and polyvinyl alcohol (PVA).

[0053] The slurry can be applied to the device by different techniques (such as at least one of sponge wiping, painting, centrifugation, 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).

[0054] Additional descriptions of the cracking method and system can be found in the published patent texts, all of which are incorporated by reference: US9,499,747, WO2015105589A1, CA2821249A1, US20170260460, CA2932461A1, WO2015088671A1, and US20170022428.

[0055] After the slurry is applied to the device, a sintering treatment can be carried out. 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 heating device. In some embodiments, the sintering temperature ranges from about 850 °C to about 1700 °C. In some embodiments, the sintering is at about 1000 °C. In some embodiments, the sintering is carried out in an inert atmosphere (such as argon or nitrogen). In some embodiments, a heat treatment is carried out in air before sintering to form an oxide layer on the inner surface of the tube, which improves the adhesion of the coating.

[0056] In operation, a hydrocarbon (such as aircraft fuel) is supplied to an internal space 25 together with an oxygen-containing substance (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 internal 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 removes heat from the tubular structure 21 and its surroundings and thus cools the tubular structure 21 and its surroundings. Accordingly, the cracking of the hydrocarbon transforms the tubular structure 21 into a heat exchanger 20 and serves as a heat removal system that can be used to remove heat from a component 30 of a high-speed aircraft 10. The coating reduces or prevents the formation of coke within the internal space 25, which could ultimately impede the flow of fuel and reduce the capacity of the heat exchanger 20. The reformed fuel exiting the heat exchanger 20 can then be used as fuel for an aircraft thruster.

[0057] Figure 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, the heat removal system having an array of a plurality of tubular structures 21, each tubular structure 21 having an internal space 25. The heat exchanger 20 is positioned adjacent to an aircraft component 30 from which it is desired to remove thermal energy (heat). 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 be used to remove heat from an air stream flowing through the array of tubular structures 21.

[0058] Figure 4 is a plan schematic 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. The manifolds 26 and 28 fluidly couple the inlet ends and the outlet ends of the tubular structures 21 to a common inlet 27 and an outlet 29, respectively. The inlet 27 may in turn be fluidly coupled to a pipe or conduit serving as a source of hydrocarbon fuel and / or steam (water). After cracking occurs within the tubular structures 21 of the heat exchanger 20, the outlet 29 may in turn be fluidly coupled to a pipe or conduit serving as a destination or recipient of the reformed fuel.

[0059] In addition to the multiple tubular structures sharing a common manifold, the heat exchangers described herein can actually include multiple heat exchangers with individual manifolds. A control system including sensors, valves, and / or electronically controlled actuators can control the flow through and among the individual tubular structures and / or the multiple heat exchangers. This can provide flexibility for different operating conditions and can cycle between the 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.

[0060] Figure 5 is a cross-sectional schematic view of an exemplary embodiment of a gas turbine engine 15 that is adapted to be used as an aircraft propulsion system for a high-speed aircraft 10 and incorporates a heat exchanger 20 described herein. As Figure 5 shown, the heat exchanger 20 containing a catalyst coating 24 within a tubular structure 21 is located within the inlet section 31 of the gas turbine engine 15. The outer casing surrounding the inlet section serves as an aircraft component 30 that is desired to remove heat from the heat generated by and transferred by the incoming air flow 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. Then, the reformed fuel exiting the heat exchanger 20 can be supplied into the burner section 34 of the gas turbine engine 15.

[0061] Figure 6 is a schematic view of a gas turbine engine 15 that utilizes a heat exchanger 20 described herein to reform a hydrocarbon fuel downstream of the last turbine stage behind a high-pressure turbine 35 and a low-pressure turbine 36. The heat contained in the residual air flow exiting the low-pressure turbine 36 provides the energy to crack the fuel, which is then discharged into the atmosphere 37. As previously discussed, the reformed fuel can be supplied to the burner 34 to fuel the gas turbine engine 15.

[0062] Figure 7 is similar to Figure 6 and is a schematic view of a gas turbine engine 15 that utilizes a heat exchanger 20 described herein to reform a hydrocarbon fuel between a high-pressure turbine stage 35 and a low-pressure turbine stage 36.

[0063] For any of the exemplary embodiments described herein, multiple heat removal systems can be employed in series or in parallel, and the inlets and outlets can be shared, or they can be separately plumbed to separate sources and destinations for the hydrocarbon fuel and the reformed fuel.

[0064] For parallel heat removal systems, all systems can operate simultaneously, or some systems can be deactivated to recover and remove coke deposits or adjust the level of heat removal capacity during various stages of aircraft operation.

[0065] Figure 8 is a schematic diagram of a method showing an exemplary embodiment of an operation. In Figure 5 , hot inlet air 33 enters the front of a gas turbine engine 15. The heat exchanger 20 can be a single element arranged annularly and / or symmetrically around the engine centerline 40, or can be a plurality of separate heat exchanger elements. The heat exchanger 20 receives liquid fuel 41 through an inlet 27. Then, the liquid fuel 41 becomes vaporized fuel 42, and then the vaporized fuel 42 is mixed with an oxygen-containing medium (such as bleed air 44 from a gas turbine engine compressor or water from a storage tank, for example, within the heat exchanger 20). Then, the heat exchanger 20 uses the available heat from the hot inlet air 33 and the oxygen-containing medium to reform the fuel, such that the reformed fuel 43 then exits the heat exchanger 20 through an outlet 29. The conditioned air 45 having a temperature lower than the hot inlet air 33 then flows in Figure 8 and Figure 5 and Figure 9 in an embodiment to the gas turbine engine compressor 32, or in Figure 6 in an embodiment to the atmosphere, or in Figure 7 in an embodiment to the low-pressure turbine 37. Depending on the installation location of the heat exchanger 20, the hot inlet air 33 in Figure 5 and 9 in an embodiment can also be discharged from the high-pressure turbine 35 in Figure 7 in an embodiment.

[0066] Figure 9 is a cross-sectional schematic diagram of an exemplary embodiment of a gas turbine engine 15, which is similar to Figure 5 in an embodiment suitable for use as an aircraft propulsion system but adapted to include a scramjet system 70. In Figure 9 in an embodiment, the scramjet 70 is similar to a booster or afterburner, but is specifically designed for high-speed and / or supersonic travel.

[0067] As with Figure 4 , in Figure 9In this case, the heat exchanger 20 is located in the inlet section 31 of the gas turbine engine 15. The liquid hydrocarbon fuel 41 is supplied from the storage tank 52 through the supply line 58 to the first diverter 48, which serves as a control valve, and a portion of the liquid hydrocarbon fuel 41 can be supplied to the fuel nozzle 54 located within the burner section 34 of the gas turbine engine 15. The liquid fuel 41 can flow from the first diverter 48 to the second diverter or control valve 50, where the liquid fuel 41 can be directed through the inlet 27 via the supply line 58 to the heat exchanger 20, or selectively directed to the fuel nozzles 56 behind the high-pressure turbine 35 and the low-pressure turbine 36 located in the scramjet 70, respectively. The control valve 46 controls the flow of the compressor discharge air or bleed air 44 to the reforming catalyst in the heat exchanger 20. The reformed (gaseous) fuel 60 exits the heat exchanger 20 through the outlet 29 and flows to the scramjet fuel nozzle 56 located in the scramjet 70.

[0068] The scramjet fuel nozzle 56 can be configured for "dual fuel" operation, i.e., to operate with liquid hydrocarbon fuel or gaseous reformed fuel or both, depending on the particular operating conditions desired. The control valves or diverters can be operated by a control system that selectively directs the liquid fuel to the gas turbine burner and / or the scramjet, and also controls the flow of the reformed fuel from the heat exchanger to the scramjet. The ratio of the fuel flowing to the main gas turbine engine burner and the scramjet can be adjusted as needed, and in high-speed or hypersonic flight regimes, diversion can be utilized, for example, about 5% diversion to the main gas turbine engine burner and 95% diversion to the scramjet. A low pressure, such as 3 - 5 bar (bar), may be experienced in the scramjet, which allows the scramjet fuel supply system to also operate at low pressure, thus minimizing the risk of leakage. A higher pressure may be experienced in the gas turbine burner, thus requiring the liquid fuel to be supplied at a higher pressure in this system.

[0069] Figure 10 is a schematic diagram of an exemplary embodiment of the heat exchanger system 20, which includes a plurality of heat exchangers 20A, 20B similar to the Figure 4 heat exchanger 20, and the system is configured to operate with the Figure 9 gas turbine engine 15. In the Figure 10 embodiment, the liquid hydrocarbon fuel 41 and the oxygen-containing compressor discharge air 44 can be supplied to the plurality of heat exchangers or heat exchanger branches 20A and 20B, respectively, via the control valve 71, and a similar control valve 71 can direct the reformed fuel to the gas turbine engine burner fuel nozzle 54 or the scramjet fuel nozzle 56 located at the turbine outlet. As described above, multiple (two or more) heat exchangers can operate together, or one or more heat exchangers can be deactivated for cleaning or based on a lower demand for the reformed fuel.

[0070] A variety of hydrocarbon fuels can 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), and their blends, mixtures, and combinations thereof.

[0071] All publications, patents, and patent applications cited herein, whether above or below, are hereby incorporated by reference in their entirety to the same extent as if each individual publication, patent, or patent application was specifically and individually indicated to be incorporated by reference. It should be understood that any patent, publication, or other publicly available material incorporated by reference herein is incorporated only to the extent that the incorporated material does not conflict with the existing definitions, statements, or other publicly available material set forth in this disclosure. Thus, and to the extent necessary, the disclosure set forth herein supersedes any conflicting material incorporated by reference herein. Any material or portion thereof that is considered to be incorporated by reference herein but conflicts with the existing definitions, statements, or other publicly available material set forth herein will be incorporated only to the extent that there is no conflict between the incorporated material and the existing publicly available material.

[0072] It must be noted that the singular forms "a," "an," and "the" as used in this specification and the appended claims include plural referents unless the context clearly dictates otherwise.

[0073] 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 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, the materials and methods according to some embodiments are described herein.

[0074] It should be noted that when used in this disclosure, the terms "comprising," "including," and other derivatives from the root term "comprising" are intended to be open-ended terms that specify the presence of any stated feature, element, integer, step, or component, and are not intended to exclude the presence or addition of one or more other features, elements, integers, steps, components, or groups thereof.

[0075] As needed, detailed embodiments of the present invention are disclosed herein; however, it should be understood that the disclosed embodiments are merely examples of the present invention, which can be implemented in various forms. Thus, the specific structural and functional details disclosed herein should not be construed as limiting, but merely as a basis for the claims and as a representative basis for teaching one of ordinary skill in the art to apply the present invention in virtually any appropriate detailed structure.

[0076] The various features, aspects, and advantages of the present disclosure may also be embodied in any arrangement of aspects of the present disclosure, including but not limited to the following technical solutions defined in the recited aspects:

[0077] 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 reformed hydrocarbon fuel from the heat exchanger.

[0078] 2. The gas turbine engine according to any preceding item, wherein the distribution system delivers the reformed hydrocarbon fuel to the gas turbine engine.

[0079] 3. The gas turbine engine according to any preceding item, wherein the gas turbine engine includes a scramjet or a booster.

[0080] 4. The gas turbine engine according to any preceding item, wherein the distribution system delivers the reformed hydrocarbon fuel to the scramjet or the booster.

[0081] 5. The gas turbine engine according to any preceding item, wherein the scramjet or the booster includes a scramjet fuel nozzle, and wherein the distribution system is configured to selectively deliver liquid hydrocarbon fuel or reformed hydrocarbon fuel to the fuel nozzle.

[0082] 6. The gas turbine engine according to any preceding item, wherein the distribution system delivers liquid hydrocarbon fuel to the combustor and reformed gaseous hydrocarbon fuel to the scramjet.

[0083] 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 fuel such as RPS1 and RPS2, diesel fuel such as D2 and D4, and their blends, mixtures, and combinations thereof.

[0084] 8. The gas turbine engine according to any preceding item, wherein the oxygen source is an oxygen-containing substance such as steam, liquid water, or an oxygen-containing fuel such as ethanol or methanol.

[0085] 9. The gas turbine engine according to any preceding item, wherein the system includes a plurality of heat exchangers or heat exchanger branches.

[0086] 10. The gas turbine engine according to any preceding item, wherein the plurality of heat exchangers are arranged in series or in parallel.

[0087] 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 being in fluid communication with a 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 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.

[0088] 12. The method 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 fuel such as RPS1 and RPS2, diesel fuel such as D2 and D4, and blends, mixtures, and combinations thereof.

[0089] 12. The method according to any preceding item, wherein the oxygen source is an oxygen-containing substance such as steam, liquid water, or an oxygen-containing fuel such as ethanol or methanol.

[0090] 13. The method according to any preceding item, wherein the method includes a plurality of heat exchangers.

[0091] 14. The method according to any preceding item, wherein the plurality of heat exchangers are arranged in series or in parallel.

[0092] 15. The method according to any preceding item, wherein a distribution system delivers the reformed fuel to the gas turbine engine.

[0093] 16. A gas turbine engine according to any one or more of the preceding items 1 to 10.

[0094] 17. The method according to any one or more of the preceding items 11 to 15.

[0095] Although the present disclosure has been described as having exemplary embodiments, the present disclosure may be further modified within the spirit and scope of the present disclosure. Accordingly, this application is intended to cover any variations, uses, or adaptations of the present disclosure using its general principles. Additionally, 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 fall within the limits of the appended claims.

Claims

1. A gas turbine engine, characterized in that, The 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 reformed hydrocarbon fuel from the heat exchanger; a scramjet or an augmenter, wherein the scramjet or the augmenter includes a scramjet fuel nozzle, and wherein the distribution system is configured to selectively deliver liquid hydrocarbon fuel or reformed hydrocarbon fuel to the fuel nozzle.

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

3. The gas turbine engine according to claim 1, characterized in that, Wherein, the distribution system delivers reformed hydrocarbon fuel to the scramjet or the augmenter.

4. The gas turbine engine according to claim 1, characterized in that, Wherein, the distribution system delivers liquid hydrocarbon fuel to the combustor and reformed gaseous hydrocarbon fuel to the scramjet.

5. The gas turbine engine according to claim 1, characterized in that, Wherein, the oxygen source is an oxygen-containing substance.

6. The gas turbine engine according to claim 1, characterized in that Wherein, the system includes a plurality of heat exchangers or heat exchanger branches.

7. The gas turbine engine according to claim 6, characterized in that, Wherein, the plurality of heat exchangers are arranged in series or in parallel.

8. A method of operating a gas turbine engine having a compressor, a combustor, and a turbine in serial flow relationship, characterized in that, The method includes the steps of: providing a heat exchanger upstream of the compressor, the heat exchanger being in fluid communication with a hydrocarbon fuel source and a water source and having an inner surface coated with a catalyst; introducing hydrocarbon fuel into the heat exchanger; introducing an oxygen source into the heat exchanger; contacting the hydrocarbon fuel with the catalyst; and cracking the hydrocarbon fuel to form reformed hydrocarbon fuel and removing heat from the aircraft component; delivering the reformed hydrocarbon fuel to a scramjet or an augmenter of the gas turbine engine, wherein the scramjet or the augmenter includes a scramjet fuel nozzle, and wherein a distribution system is configured to selectively deliver liquid hydrocarbon fuel or reformed hydrocarbon fuel to the fuel nozzle.

9. The method according to claim 8, wherein Wherein, the oxygen source is an oxygen-containing substance.

10. The method according to claim 8, characterized in that Wherein, the method includes a plurality of heat exchangers.

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

12. The method according to claim 8, wherein Wherein, a distribution system delivers reformed fuel to the gas turbine engine.

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