Turbine engine cmas in situ mitigation

Abstract

In a method for using a gas turbine engine (720, 720A, 720B), the gas turbine engine has: a compressor section (840, 842); a combustor section (844) including a plurality of fuel nozzles (860); and a turbine section (846, 848) having one or more stages of blades (400). The method includes spraying a solution or suspension (23, 440) into the combustor section via the fuel nozzles or additional nozzles (26, 26A, 26B) of the gas turbine engine, if any. The solution or suspension has a binder selected from gadolinium and gadolinium compounds.

Description

TURBINE ENGINE CMAS IN SITU MITIGATIONCROSS-REFERENCE TO RELATED APPLICATION

[0001] Benefit is claimed of U.S. Patent Application No. 63525256, filed July 6, 2023, and entitled “Turbine Engine CMAS In Situ Mitigation”, the disclosure of which is incorporated by reference herein in its entirety as if set forth at length.BACKGROUND

[0002] The disclosure relates to gas turbine engines. More particularly, the disclosure relates to addressing calcium magnesium aluminosilicate (CMAS).

[0003] Gas turbine engines (used in propulsion and power applications and broadly inclusive of turbojets, turboprops, turbofans, turboshafts, industrial gas turbines, and the like) have combustor and turbine sections that operate at high temperature (the “hot section” of the engine).

[0004] Extremely fine dust, including calcium, magnesium, aluminum, and silicon oxides (CMAS), is a major problem for modern jet engines. It is mostly found in and above desert areas but can drift into other areas as well. CMAS is a problem because it is ingested by the engine and melts in the combustion chamber. The liquid CMAS deposits on the airfoils and walls (e.g., gaspath inner diameter (ID) and outer diameter (OD) surfaces) of the hot section of the engine and infiltrates the ceramic thermal barrier coatings that protect them. The ceramic coatings have either a columnar structure or controlled porosity that gives them strain tolerance to compensate for the different thermal expansion rates between the coating and the metallic substrate. The CMAS seeps into the gaps and porosity. When the engine is shut down, it cools, the CMAS can freeze, reducing or eliminating the strain tolerance of the coating. The resulting build-up of shear stresses as the metallic substate cools and contracts, and can cause the ceramic coating to spall after multiple cycles.

[0005] There have been various proposals to apply supplemental coating(s) to spalled regions so as to extend service life.

[0006] Examples are found in US Patent Publication 2022 / 0136095 Al (the ‘095 publication), May 5, 2022, “Reactive Phase Spray Formulation Coatings”, US Patent Publication 2021 / 0324201A1 (the ‘201 publication), October 21, 2021, “Consumable Coatings and Methods of Protecting A High Temperature Component from Dust Deposits”, and US Patent Publication 2021 / 0277523 Al (the ‘523 publication), September 9, 2021, “Coating Systems Including Infiltration Coatings and Reactive Phase Spray Formulation Coatings”.

[0007] CMAS is not merely relevant to engines producing thrust for aircraft but also helicopter engines, armored land vehicle (wheeled or tracked) engines, ships, and the like.

[0008] Separately, water injection has been used in gas turbine engines for purposes including increasing thrust and altering the emissions profile. Such engines have used various injection locations. US Patent 4395874 (the ‘874 patent), August 2, 1983, “Fuel Nozzles with Water Injection for Gas Turbine Engines” discloses a particular coaxial injection system.SUMMARY

[0009] One aspect of the disclosure involves a method for using a gas turbine engine, the gas turbine engine comprising: a compressor section; a combustor section including a plurality of fuel nozzles; and a turbine section having one or more stages of blades. The method comprises: spraying a solution or suspension into the combustor section via the fuel nozzles or additional nozzles of the gas turbine engine, if any. The solution or suspension comprises a binder.

[0010] In a further embodiment of any of the foregoing embodiments, additionally and / or alternatively, the binder comprises at least one metal or compound selected from the group consisting of lanthanum, cerium, praseodymium, neodymium, promethium, samarium, europium, gadolinium, terbium, dysprosium, holmium, erbium, thulium, ytterbium, lutetium, indium, scandium, yttrium, zirconium, hafnium, titanium, and mixtures thereof as well as compounds of such metals.

[0011] In a further embodiment of any of the foregoing embodiments, additionally and / or alternatively, the binder is selected from: gadolinium; and gadolinium compounds.

[0012] In a further embodiment of any of the foregoing embodiments, additionally and / or alternatively, the spraying is of said solution as an aqueous gadolinium nitrate solution.

[0013] In a further embodiment of any of the foregoing embodiments, additionally and / or alternatively, the spraying is of said suspension as a suspension of gadolinium or gadolinia in an aqueous and / or alcohol-based carrier.

[0014] In a further embodiment of any of the foregoing embodiments, additionally and / or alternatively, the solution or suspension coats the one or more stages of blades (e.g., coats the blade and vane airfoils).

[0015] In a further embodiment of any of the foregoing embodiments, additionally and / or alternatively, the one or more stages of blades alternate with one or more stages of vanes.

[0016] In a further embodiment of any of the foregoing embodiments, additionally and / or alternatively, the spraying is through the plurality of fuel nozzles.

[0017] In a further embodiment of any of the foregoing embodiments, additionally and / or alternatively, the spraying is through a circumferentially distributed plurality of spray nozzles coaxial with respective associated fuel nozzles.

[0018] In a further embodiment of any of the foregoing embodiments, additionally and / or alternatively, the spray nozzles are coaxial with respective associated fuel nozzles.

[0019] In a further embodiment of any of the foregoing embodiments, additionally and / or alternatively, the gas turbine engine is a propulsion engine of an aircraft and the spraying is from a reservoir contained within the aircraft.

[0020] In a further embodiment of any of the foregoing embodiments, additionally and / or alternatively, the gas turbine engine is a propulsion engine of an aircraft; and the spraying is from an external reservoir connected to the aircraft via a hose.

[0021] In a further embodiment of any of the foregoing embodiments, additionally and / or alternatively, the hose is a single hose connected to a single port to treat multiple engines.

[0022] In a further embodiment of any of the foregoing embodiments, additionally and / or alternatively, the spraying comprises blending the binder and a solvent or carrier and spraying the blend.

[0023] In a further embodiment of any of the foregoing embodiments, additionally and / or alternatively, the method further comprises, after spraying the blend, a purge.

[0024] In a further embodiment of any of the foregoing embodiments, additionally and / or alternatively, the purge is for not more than half a duration of the spraying.

[0025] A further embodiment of any of the foregoing embodiments may additionally and / or alternatively include: shutting down the engine; monitoring a temperature of the engine; and responsive to the monitoring showing temperature decrease to a predetermined level, commencing the spraying.

[0026] A further embodiment of any of the foregoing embodiments may additionally and / or alternatively include driving rotation of a spool of the engine during the spraying.

[0027] In a further embodiment of any of the foregoing embodiments, additionally and / or alternatively, the driving comprises directing air from an auxiliary power unit (APU).

[0028] In a further embodiment of any of the foregoing embodiments, additionally and / or alternatively, the driving is at a speed of 10.0-100.0 RPM.

[0029] In a further embodiment of any of the foregoing embodiments, additionally and / or alternatively, the driving comprises directing air from a ground power unit (GPU or “huffer cart”).

[0030] In a further embodiment of any of the foregoing embodiments, additionally and / or alternatively, the solution is stored in the huffer cart or a further service cart and a hose connects the huffer cart or further service cart to the aircraft for delivering the solution.

[0031] A further aspect of the disclosure involves, an aircraft comprising: at least one gas turbine engine comprising: a compressor section; a combustor section including a plurality of fuel nozzles; a turbine section having one or more stages of blades; a fuel tank; and a fuel flowpath from the fuel tank to the plurality of fuel nozzles. The aircraft further comprises: a reservoir of a CMAS binder; and a binder flowpath from the reservoir to the fuel nozzles or additional fluid nozzles, if any in the combustor.

[0032] In a further embodiment of any of the foregoing embodiments, additionally and / or alternatively, the combustor comprises said additional fluid nozzles and the binder flowpath is from the reservoir to the additional fluid nozzles.

[0033] In a further embodiment of any of the foregoing embodiments, additionally and / or alternatively: the at least one gas turbine engine is a plurality of gas turbine engines; the reservoir is shared by the plurality of gas turbine engines; and binder flow is separately controllable for the plurality of gas turbine engines.

[0034] A further aspect of the disclosure involves a vehicle having a gas turbine engine, the engine comprising: a compressor section; a combustor section including a plurality of fuel nozzles; and a turbine section having one or more stages of blades, wherein: the combustor comprises additional fluid nozzles; and the additional fluid nozzles are coupled to a port.

[0035] In a further embodiment of any of the foregoing embodiments, additionally and / or alternatively: the additional fuel nozzles are concentric with or circumferentially alternating with the fuel nozzles; and the gas turbine engine has a temperature sensor coupled to a connector on the vehicle.

[0036] A further aspect of the disclosure involves a method for using the vehicle, the method comprising: connecting a fluid source to the port; delivering through the port a solution or suspension; and spraying into the combustor section via the additional nozzles of the gas turbine engine. The solution or suspension comprises a binder.

[0037] In a further embodiment of any of the foregoing embodiments, additionally and / or alternatively the binder comprises at least one metal or compound selected from the group consisting of lanthanum, cerium, praseodymium, neodymium, promethium, samarium, europium, gadolinium, terbium, dysprosium, holmium, erbium, thulium, ytterbium, lutetium, indium, scandium, yttrium, zirconium, hafnium, titanium, and mixtures thereof as well as compounds of such metals.

[0038] In a further embodiment of any of the foregoing embodiments, additionally and / or alternatively the binder is selected from: gadolinium; and gadolinium compounds.

[0039] A further aspect of the disclosure involves a method for using a gas turbine engine, the gas turbine engine comprising: a compressor section; a combustor section including a plurality of fuel nozzles; and a turbine section having one or more stages of blades. The method comprises: spraying a solution or suspension into the combustor section via a circumferentially distributed plurality of nozzles of the gas turbine engine. The solution or suspension comprises a binder.

[0040] In a further embodiment of any of the foregoing embodiments, additionally and / or alternatively the binder comprises at least one metal or compound selected from the group consisting of lanthanum, cerium, praseodymium, neodymium, promethium, samarium, europium, gadolinium, terbium, dysprosium, holmium, erbium, thulium, ytterbium, lutetium, indium, scandium, yttrium, zirconium, hafnium, titanium, and mixtures thereof as well as compounds of such metals.

[0041] In a further embodiment of any of the foregoing embodiments, additionally and / or alternatively the binder is selected from: gadolinium; and gadolinium compounds.

[0042] A further aspect of the disclosure involves an aircraft service cart comprising: a reservoir of a CMAS binder; a binder flowpath from the reservoir to an outlet; and means for driving a flow from the reservoir to the outlet.

[0043] In a further embodiment of any of the foregoing embodiments, additionally and / or alternatively the service cart is a ground start cart having at least one of: an electrical connection for starting an aircraft; and an air connection for starting an aircraft.

[0044] In a further embodiment of any of the foregoing embodiments, additionally and / or alternatively the binder is in solution or suspension.

[0045] The details of one or more embodiments are set forth in the accompanying drawings and the description below. Other features, objects, and advantages will be apparent from the description and drawings, and from the claims.BRIEF DESCRIPTION OF THE DRAWINGS

[0046] FIG. 1 is a schematic sectional view of a coating system.

[0047] FIG. 2 is a schematic sectional view of the coating system exposed to CMAS.

[0048] FIG. 3 is a schematic sectional view of the coating system during spalling.

[0049] FIG. 4 is a schematic sectional view of the coating system during restoration.

[0050] FIG. 5 is a schematic view of a CMAS mitigation system.

[0051] FIG. 6 is an underside schematic view of an aircraft with the CMAS mitigation system.

[0052] FIG. 7 is a front schematic view of an aircraft with the CMAS mitigation system.

[0053] FIG. 8 is a schematic sectional view of an aircraft engine for use with the CMAS mitigation system.

[0054] FIG 8 A is an enlarged view of a nozzle assembly of the engine of FIG. 8.

[0055] FIG. 9 is an end view of a nozzle assembly of the engine.

[0056] FIG. 10 is a schematic view of an aircraft with a partially external CMAS mitigation system.

[0057] Like reference numbers and designations in the various drawings indicate like elements.DETAILED DESCRIPTION

[0058] As is discussed below, a solution or suspension for mitigating CMAS may be sprayed in situ inside the engine. Some examples may use the fuel nozzles to spray the solution. Among examples using separate nozzles, some examples may use spray nozzles concentric with fuel nozzles in a similar structure to water injection on prior art engines. Alternatives may use other positioning such as circumferentially between the fuel nozzles.

[0059] In some examples, the solution / suspension acts to restore CMAS-binding material that has spalled off gaspath surfaces of hot section engine components. In some examples, the solution / suspension contains a binder and a solvent or carrier. For example, the binder may be a source of gadolinium (including compounds thereof) such as an aqueous gadolinium nitrate solution or a gadolinia suspension. More broadly, the binder may comprise, consist, or consist essentially of at least one metal selected from the group consisting of lanthanum, cerium, praseodymium, neodymium, promethium, samarium, europium, gadolinium, terbium, dysprosium, holmium, erbium, thulium, ytterbium, lutetium, indium, scandium, yttrium, zirconium, hafnium, titanium, and mixtures thereof as well as compounds of such metals. Among other examples are materials applied in the ‘095 publication, the ‘201 publication, and the ‘523 publication, the disclosures of which publications are incorporated by reference herein for materials disclosed therein as if set forth at length. More specifically, the metal may be a rare earth. Among compounds for solutions are nitrates and acetates. Example compounds for suspensions are oxides.

[0060] FIG. 1 shows an article 400 as a hot section component of a gas turbine engine. The component is exposed to the gaspath 520 of the engine and has a metallic substrate 402 and a baseline initial as-applied CMAS-resistant thermal barrier coating system 404. The coating may thus be on the gaspath-facing surface of a blade (airfoil exterior, platform OD surface and shroud (if any) ID surface), vane (airfoil exterior, platform (if any) OD surface and shroud ID surface), strut, and the like.

[0061] The example system schematically includes a bond coat (e.g., thermal sprayed MCrAlY) 406 atop a surface 408 of the substrate. A base layer 410 of the TBC is atop the bond coat. A CMAS-binding layer 412 is atop the base layer. The example base layer 410 is yttria stabilized zirconia (YSZ), namely 7YSZ. The example CMAS binding layer is gadolinium zirconate (GDZ or GZO), for example 59 GZO. This is illustrative. But typically, the binder will contain gadolinia. Additional aspects (not shown) may include artifacts such as a thermally grown oxide (TGO) layer or additional features such as a sealant or topcoat atop the binder.

[0062] The example base layer 410 and binding layer 412 are columnar crystalline structures deposited by electron beam physical vapor deposition (EBPVD). In service, CMAS 420 (FIG. 2) deposits atop the surface of the binding layer 412 and will have both intercolumnar and intracolumnar infiltration of the binding layer 412. Alternatives for one or both layers include thermal spray such as air plasma spray, solution plasma spray, or suspension plasma spray and may include non-columnar structures.

[0063] As noted above, the CMAS-infiltrated binding layer 412 material will have differential thermal expansion relative to the uninfiltrated base layer 410 and / or have less ability to accommodate differential thermal expansion of the substrate than does the base layer. This can lead to spalling off of pieces 430 (FIG. 3) of the deposits including the infiltrated binding layer 412 material exposing uninfiltrated binding layer material 412 below and / or base layer 410 material. After such spalling, the sprayed solution / suspension 440 (FIG. 4) may deposit (or redeposit) binder material 442 in place of original binding layer 412 material or prior restorative sprays.

[0064] FIG. 5 shows a CMAS mitigation system 20 containing a source 22 of a CMAS-binding agent (binder) and a source 24 of a solvent or cleaner 25. The solvent or cleaner 25 may be used for purging. An example source 22 contains binder in a solution or suspension 23 with a solvent or carrier. The source 24 solvent or cleaner 25 may be of the same material used as the source 22 solvent or carrier or a different material. As is discussed further below, the solution / suspension 23 and solvent / cleaner 25 are sprayed through nozzles 26 into the engine to restore spalled binding layer 412 portions of ceramic barrier coatings on engine components. The example solution / suspension 23 and solvent / cleaner 25 are pumped by respective pumps 32 and 34 along flowpaths 36 and 38 for distribution via one or more valves 28 to be passed along one or more flowpaths 40 to the nozzles 26. An example flowpath 40 is a branching flowpath to respective nozzles associated with the fuel nozzles of the engine (e.g., coaxial with). For purposes of illustration and not limitation, FIG. 5 shows a two-engine configuration.

[0065] For multi-engine aircraft, there may be a provision for independent processing of the various engines for any of several reasons. FIG. 5 shows a two-engine configuration wherein the nozzles for the two engines are respectively shown as 26 A and 26B. First, the engines may cool at different rates so that solution / suspension flow should commence earlier on one engine than on another engine even if they are both being treated at generally the same time (e.g., after a given flight and possibly overlapping in time). In one example, theengines share sources of the solution / suspension and solvent / cleaner. However, branches to the engines may have separate pumps or valving structures.

[0066] In variations, one or both of the reservoirs may be pressurized by a pump or compressor upstream so that no separate pump intervenes between the reservoir and the engine, only the valve(s). This is particularly relevant to suspensions that might foul or otherwise damage pumps. Thus, for example, the reservoir vessel may be divided by a diaphragm (not shown) with the suspension on one side of the diaphragm and the pump or compressor pressurizing the headspace at the other side of the diaphragm (e.g., with air).

[0067] As is discussed further below, the various components may be controlled by a controller 42. An example controller 42 is a programmable logic controller (PLC). The controller may have various input and outputs. These inputs and outputs may include communications with the aircraft’s existing control system or other components. FIG. 5 shows an example input from temperature sensor(s) 44. In one example, there is such a temperature sensor 44 for each engine. Such temperature sensors 44 may be existing temperature sensors of a baseline engine to which the system 20 is added (e.g., exhaust gas temperature (EGT) sensors). Alternatively, they may be additional sensors. Example temperature sensors 44 are thermocouples. An example is one per engine.

[0068] Example sensors 44 that are not the baseline EGT sensors may be used to get more upstream temperature measurement than that offered by the EGT sensors. Alternatively, even if closer to the EGT sensors, additional sensors 44 may be used for independence or for more precise measurement in a temperature range of interest. Although a direct input of sensors 44 to PLC 42 is shown, alternative inputs may be via the aircraft’ s existing control system or an individual engine FADEC (e.g., via digital data bus).

[0069] One example of a binder solution 23 for source 22 is a binder comprising gadolinium nitrate and water as the solvent. The binder and solvent may be preblended and introduced to the tank or other reservoir of the source 22. An example mixture content is 0.10 to 50.0 g / L, more narrowly 0.50 to 50.0 g / L or 1.0 to 50g / L or 5.0 to 50g / L. However, higher concentrations may be desirable including up to 100 g / L or more. An example may be at least 20% or at least 30% saturated (e.g., 30% to 90% or 50% to 85%). A corresponding example solvent / cleaner 25 is also water. One example of valve(s) 28 is a single spool valve (e.g., servo or solenoid controlled) with positions corresponding to: 1) both flowpaths 36, 38 closed; 2) binder solution flowpath 36 open and solvent / cleaner flowpath 38 closed / blocked; and 3) binder solution flowpath 36 closed / blocked and solvent / cleaner flowpath 38 open. Such a single spool valve may replace multiple separate valves (e.g., separate valves alongeach leg of the various flowpaths). Staying sufficiently below 100% saturation may avoid possible pump and valve damage and clogging for systems not designed to handle particulate (see suspension discussion below). But a hybrid solution / suspension may yet be used in a similar fashion to a pure suspension.

[0070] An example binder suspension 23 is gadolinium and / or a gadolinium compound (e.g., an oxide such as gadolinium zirconate) particulate in a water-based or ethanol-based (e.g., water and / or ethanol) carrier with optional dispersant and other additives (if any). For such suspension, an example solvent / cleaner 25 is of the same base as the carrier but not necessarily with all additives (e.g., using pure water v. the water plus dispersant of the carrier). A similar spool valve may be used for the valve 28 similarly to that described above. Example solids content is about 25 weight percent solids, more broadly 5.0 weight percent to 35.0 weight percent or 15.0 weight percent to 30.0 weight percent. Alternative examples are 5.0 to 35.0 g / L, more narrowly, 15.0 to 30.0g / L. Example dispersant is an organic dispersant such as polyacrylic acid (PAA), polyethylene imine (PEI), and 2-phosphonobutane- 1,2,4- tricarboxylic acid (PBTCA). Example particulate may be in a size distribution similar to that used in the original application of layer 412 if sprayed. However smaller particles and / or larger and / or different distributions may be advantageous for filling gaps left by spalling. The binder, carrier, and optional dispersant may be preblended and introduced to the tank or other reservoir of the source 24.

[0071] More complex multi-valve systems may be appropriate, including valves that isolate pumps. In general, example valves are spool, ball, or poppet (e.g., servo or solenoid controlled). However, for the suspension embodiments, the valves passing the suspension may advantageously be poppet valves to limit wear from the particulate.

[0072] Example pumps are positive displacement pumps (e.g., piston pumps or gear pumps). This has the advantage of relatively precise metering. For example, for pumps that pass binder solution (or the solvent / cleaner), gear pumps, piston pumps, or even roots type pumps may be used. Scroll pumps or vane pumps can be used, particularly when precise metering / control is not believed needed. For pumps 32 passing solid suspensions 23, such positive displacement and scroll / vane pumps would not be appropriate because of the abrasive nature of the suspension on the working elements. In such a case it would be appropriate to use a peristaltic pump or to use a compressor or air pump to pressurize the suspension reservoir to force the suspension through the lines.

[0073] FIG. 6 schematically shows an aircraft 700 including the CMAS mitigation system 20. The example aircraft 700 has a fuselage 702. The fuselage extends from a nose704 to a tail 706. A main wing 710 extends laterally outward from the fuselage on left and right sides. A horizontal stabilizer 712 and vertical stabilizer 714 are proximate the tail aft of the main wing. Control surface details (not shown) may be conventional. The example aircraft has engines 720A, 720B positioned in respective nacelles 722 on respective wing pylons 724. The aircraft also has an auxiliary power unit (APU) 730. An example APU: (1) provides electricity to power the aircraft when the main engines are off; and (2) compresses air for pneumatic functions including starting the main engines.

[0074] An example APU itself has a small gas turbine engine driving a load compressor (e.g., co-spooled with the APU’s gas turbine engine compressor and turbine and an electrical generator (e.g., via a gearbox). Typical APU is in the tail section of the fuselage. A controlled inlet flap or gate in the side of the fuselage serves as an air inlet passing air to the APU engine compressor and load compressor in parallel. The APU engine turbine typically exhausts out an outlet at the rear end of the fuselage.

[0075] FIG. 6 and 7 (not to scale) schematically shows CMAS mitigation system 20 largely shared by the engines. The main components of the system 20 are contained in a bay 740 in the fuselage (e.g., below a cargo deck of an airliner or cargo jet).

[0076] Other aircraft configurations exist including delta wing and delta canard configurations among other configurations lacking the traditional combination of vertical stabilizer and horizontal stabilizer. And as noted above, the teaching may be applicable to rotary wing aircraft, land vehicles, ships, and stationary power sources.

[0077] FIG. 8 shows an example gas turbine engine 720 (representing the engines 720A and 720B) as a two-spool turbofan engine. The engine 720 has an engine case 822 surrounding a centerline or central longitudinal axis 500. An example engine has a fan section 824 including a fan 826 within a fan case 828. The example engine includes an inlet 830 at an upstream end of the fan case receiving an inlet flow along an inlet flowpath 520. The fan 826 has one or more stages 832 of fan blades. Downstream of the fan blades, the flowpath 520 splits into an inboard portion 522 being a core flowpath and passing through a core of the engine and an outboard portion 524 being a bypass flowpath exiting an outlet 834 of the fan case.

[0078] The core flowpath 522 proceeds downstream to an engine outlet 836 through one or more compressor sections, a combustor, and one or more turbine sections. The example engine has two axial compressor sections and two axial turbine sections, although other configurations are equally applicable. From upstream to downstream there is a low pressure compressor section (LPC) 840, a high pressure compressor section (HPC) 842, a combustorsection 844, a high pressure turbine section (HPT) 846, and a low pressure turbine section (LPT) 848. Each of the LPC, HPC, HPT, and LPT comprises one or more stages of blades which may be interspersed with one or more stages of stator vanes.

[0079] In the example engine, the blade stages of the LPC and LPT are part of a low pressure spool mounted for rotation about the axis 500. The example low pressure spool includes a shaft (low pressure shaft) 850 which couples the blade stages of the LPT to those of the LPC and allows the LPT to drive rotation of the LPC. In the example engine, the shaft 850 also drives the fan. In the example implementation, the fan is driven via a transmission (not shown, e.g., a fan gear drive system such as an epicyclic transmission) to allow the fan to rotate at a lower speed than the low pressure shaft.

[0080] The example engine further includes a high pressure shaft 852 mounted for rotation about the axis 500 and coupling the blade stages of the HPT to those of the HPC to allow the HPT to drive rotation of the HPC. In the combustor 844, fuel is introduced to compressed air from the HPC and combusted to produce a high pressure gas which, in turn, is expanded in the turbine sections to extract energy and drive rotation of the respective turbine sections and their associated compressor sections (to provide the compressed air to the combustor) and fan.

[0081] PIG. 8 also shows the location of the temperature sensor 44 (possibly the EGT sensor if it can read in the correct range) in the HPT. Lor example, the existing EGT sensor may be calibrated for measuring temperatures well above 150°C. Thus, added sensor(s) 44 may be configured to operate well in a range of approximately 90°C to 200°C. An example positioning of the added sensor may be in a strut or vane within (PIG. 8) the HPT or adjacent to the HPT (e.g., in a mid-turbine frame between the two turbine sections).

[0082] PIGs. 8 A and 9 show two examples of the relationship between the nozzles 26 and the fuel nozzles. The fuel nozzles are connected to one or more fuel tanks 760 (e.g., wing (not shown) and / or fuselage tanks) via fuel lines and pumps as in a conventional manner. There may be an example one main fuel line 762A, 762B schematically forming a flowpath to each engine from the tank(s). PIG. 8A is an example based on the ‘874 patent. Specifically, it shows a possible relation between the two nozzles in a coaxial relationship (shared axis 501) where a manifold is added to form the nozzle 26 for the binder. The fuel nozzle is shown as 860 (having fuel outlet 861) and the swirler shown as 862. PIG. 9 shows circumferentially alternating nozzles 26 and 860. Despite these examples, the count of fuel nozzles and additional nozzles need not be the same. In general, an example number of the additional nozzles is four to thirty, more particularly, eight to twenty-four. Alternatively, the number offuel nozzles may be at least half the count of fuel nozzles. In general, the nozzles should be evenly circumferentially distributed about the engine axis / centerline 500.

[0083] System component materials and manufacture techniques and assembly techniques may be otherwise conventional. Due to presence in the combustor, the nozzles 26 should be of appropriate material (e.g., cast / machined or additively manufactured nickel- based or cobalt based alloy / superalloy).

[0084] An example operational sequence may occur after a landing. The aircraft may taxi or be towed to a gate or other parking location. The engine(s) are shut down there by the pilot, if not already shut down. The control 42 continuously or intermittently measures the relevant engine temperature(s) via the sensor(s) 44.

[0085] When the temperature of an engine (or all engines) drops to a threshold temperature, spraying may commence. The threshold temperature should be sufficiently high to evaporate the water and nitrogen and oxidize the gadolinium, but not so high that the solvent with flash to vapor and blow the gadolinium nitrate off the surface. An example advantageous temperature for application is in a range of 95°C to 150°C. However, it may be advantageous to commence in an upper portion of this range such as 120°C to 150°C.

[0086] An example pre-set application time is 1.0 minutes to 10 minutes, more particularly 2.0 to 5.0 minutes.

[0087] An example engine speed during application is 10.0 to 100.0 revolutions per minute (RPM) measured for the high pressure spool (N2), more narrowly 20.0 to 60.0 RPM.

[0088] An example delay between shutoff of the solution / suspension flow and commencement of the purge is 0 to 20.0 seconds. This time period is advantageous to limit chances of settling in the lines. However, in some cases a small delay may be advantageous. If the target gaspath surfaces are hotter than the nozzles and lines a short delay improves chances of the bonding of the binder to the target surfaces before the binder has an opportunity to foul the lines and nozzles. Thus a small delay (e.g., 5.0 to 20.0 seconds or 10.0 to 20 seconds) may limit any removal of binder from the gaspath surfaces while still allowing effective purge of binder from the lines and nozzles. For example, when using GdNOa, an immediate purge may remove it from the gaspath surfaces before it transforms into Gd Oa; whereas a short delay allows more to transform and thus persist while not having similar persistence in lines.

[0089] An example duration of the purge is shorter than the binder spraying, e.g., 10.0 seconds to 30.0 seconds or an example not more than 50% or 20% of the binder spray duration. This may be advantageous to merely purge lines but not remove all the depositedmaterial. An example engine speed during the purge is similar to that during binder application (e.g., ranges above).

[0090] In a particular example, when the post-shutdown threshold temperature is reached, several things may occur, if not already occurring. The control 42 cause rotation of the engine such as via the APU. Fuel is not delivered. The control 42 will cause the valve(s) to open the flowpath(s) 36 from the source 22 to the nozzles 26 and start the binder pump 32 (e.g., electronic command to the pump or applying power to the pump). The binder solution / suspension 23 flows for a predetermined time, to fully coat the HPT blade and vane airfoils and other relevant surfaces. During this process, the engine rotation helps both with downstream flow and distributing solution / suspension 23 that has landed on any given surface. Upon reaching the predetermined time, the control 42 causes the binder pump 32 to stop, and the valve(s) to close the binder flowpath 36. In this example, the rotating of the engine (e.g., via the APU) continues.

[0091] The control 42 will then signal the valve(s) to open the flowpath(s) 38 from the source 22 to the nozzles 26 and start the solvent / c leaner pump 34 to drive the solvent / cleaner to purge various lines / conduits and the nozzles so that binder does not harden in the lines and nozzles. After the purge is complete, the control 42 causes the solvent / cleaner pump 34 to stop and the valve(s)to close the flowpath 38. The control 42 then disengages the APU to stop the engine rotation.

[0092] Although some implementations may treat both / all engines in a multi-engine aircraft at the same time, other examples may treat them sequentially or not in conjunction at all. For example, one example of simultaneous use is when both engines have had their relevant temperature fall into the target range. However, there may be issues of pump capacity. In such a case, when one engine temperature has dropped into the target range, control 42 may initiate binder solution / suspension flow only to that engine and maintain such flow for the target duration even if the other engine's temperature falls into the target range. Depending upon the particular combination of valve(s) 28, it may be possible to switch over and pass binder solution / suspension to that other engine even during the purge of the prior engine (but in any event at least afterward). For such independent operation, a more complex single spool valve 28 may be used or multiple stages of valves and / or branches of fluid lines may be used.

[0093] As an alternative to an APU bleed, some aircraft use compressed air from a ground power unit (GPU, with air variants also known as an air start unit (ASU) or “huffer cart”) to start. As an alternative to an APU bleed or other compressed air start, some aircraftengines have electrical starters in the form of generator / starter units which generate electricity when the engine is operating. These may be used as starters powered by a generator of the APU. Alternatively, the ground power unit may be temporarily wired to the engine for such start purposes. Thus, APU electrical power or ground power unit compressed air or electrical power may be used for rotating the engine during the treatment instead of APU compressed air.

[0094] Relatedly, among further variations in general configuration, some implementations may place the mitigation system largely external to the aircraft. This may be particularly relevant where weight and / or space considerations are critical. Weight may be particularly relevant in smaller commuter type aircraft. Space may be particularly relevant in smaller military aircraft such as fighter aircraft and durable uncrewed vehicles (UCV). For example, the aircraft may include one or more external ports mating to one or more fluid couplings of the system and plumbing from the external port to the nozzles. This may involve valves to switch between engines such as discussed above. Similarly, the aircraft may include an external communications port so that the mitigation system may communicate with the aircraft for receiving information including the temperature data and potentially for controlling rotation of the engine if the engine is to be rotated by an onboard source. External components of the mitigation system may be contained in a service cart. The service cart may include a power source (e.g., a small electrical generator such as a diesel generator) or may be plugged into external electrical power.

[0095] In some implementations, such as wherein there is no APU, the external mitigation system may be paired with or integrated into a ground power unit (GPU, with air variants also known as an air start unit (ASU) or “huffer cart”). In examples of separate carts (e.g., separate from the GPU), the cart containing the external components of the mitigation system 20 may also have communication connections to the GPU to control use of the GPU to rotate the engine. Depending upon the aircraft, this may be via compressed air from the GPU or electrical power from the GPU. Nevertheless, such cart implementations may also be used for aircraft having APU (e.g., sometimes GPU / ASU are used for economic reasons).

[0096] In an example of a retrofit of such an aircraft for use with external inputs of the binder solution / suspension and solvent / cleaner, the aircraft may be fitted with modified engines having the additional nozzles 26 (if present) or additional connections to the fuel nozzles 860 (e.g., connecting in to the fuel line for each engine with a valved connection). The additional temperature sensor(s) 44 may be added (if used) or additional wiring may be coupled to an existing temperature sensor (e.g., in parallel with existing wiring). Exampleexternal connections on the fuselage include a single fluid connection and a multiconductor power / data connection. In some implementations, fluid and power / data connections may be coupled into a combined connector. In some implementations, the aircraft may include valve(s) for selectively coupling the fluid port to individual engines for a multi-engine aircraft. Alternative implementations may have separate fluid ports for the separate engines. After landing, and upon parking, the service cart may be connected to the ports and the process may proceed as discussed above for the onboard system.

[0097] FIG. 10 schematically shows an aircraft 900 with a largely external CMAS mitigation system 300. For purposes of illustration, the aircraft is shown as an underside view of one particular aircraft. The CMAS mitigation system is shown including a separate cart 302 (e.g., wheeled) but is schematically shown not in an actual orientation relative to the aircraft.

[0098] FIG. 10 schematically shows an aircraft modified for use with a partially external mitigation system 300. The example aircraft 900 has a fuselage 902. The fuselage extends from a nose 904 to a tail 906. A main wing 910 extends laterally outward from the fuselage on left and right sides. A horizontal stabilizer 912 and vertical stabilizer(s) (two in the example) 914 are proximate the tail aft of the main wing. FIG. 10 also shows a fuel tank 960 and fuel lines 962A, 962B for the respective engines (schematically represented by 720 in FIG. 8 although likely proportioned differently). A service cart 302 (e.g., wheeled / castered) has a number of external connections for connecting to the airplane. The example cart is notional and not all connections need be present and additional connections may be present. The example cart has at least one fluid outlet port / coupling / fitting 322 for connection via a conduit (e.g., hose 324) to a fluid inlet port / coupling / fitting 326 on the aircraft fuselage. The example fitting 326 is at a small on-board equipment bay 920 where the added on-board equipment may be located including various fluid connections, electronic / electrical connections and the like. One or more additional external ports / connectors 332 on the cart connect electrical / electronic power and / or communication to the aircraft via wiring 334 coupled to one or more corresponding connectors 336 on the fuselage. Additionally, for air-started engines as opposed to electrically-started engines, an air outlet port / fitting 352 on the cart connects a compressed air source 350 on the cart (e.g., a compressor shown) to a conventional port or ports 356 on the aircraft via a hose 354 with associated fittings. FIG. 10 also shows the cart having an electrical connection 340 to external power. However, as noted above, the cart may additionally or alternatively have an on-board diesel generator.

[0099] The use of “first”, “second”, and the like in the following claims is for differentiation within the claim only and does not necessarily indicate relative or absolute importance or temporal order. Similarly, the identification in a claim of one element as “first” (or the like) does not preclude such “first” element from identifying an element that is referred to as “second” (or the like) in another claim or in the description.

[0100] One or more embodiments have been described. Nevertheless, it will be understood that various modifications may be made. For example, when applied to an existing baseline configuration, details of such baseline may influence details of particular implementations. Accordingly, other embodiments are within the scope of the following claims.

Claims

CLAIMSWhat is claimed is:

1. A method for using a gas turbine engine (720, 720A, 720B), the gas turbine engine comprising: a compressor section (840, 842); a combustor section (844) including a plurality of fuel nozzles (860); and a turbine section (846, 848) having one or more stages of blades (400), the method comprising: spraying a solution or suspension (23,440) into the combustor section via the fuel nozzles or additional nozzles (26, 26A, 26B) of the gas turbine engine, if any, the solution or suspension comprising a binder.

2. The method of claim 1 wherein: the binder comprises at least one metal or compound selected from the group consisting of lanthanum, cerium, praseodymium, neodymium, promethium, samarium, europium, gadolinium, terbium, dysprosium, holmium, erbium, thulium, ytterbium, lutetium, indium, scandium, yttrium, zirconium, hafnium, titanium, and mixtures thereof as well as compounds of such metals.

3. The method of claim 1 wherein: the binder is selected from: gadolinium; and gadolinium compounds.

4. The method of claim 1 wherein: the spraying is of said solution as an aqueous gadolinium nitrate solution.

5. The method of claim 1 wherein: the spraying is of said suspension as a suspension of gadolinium or gadolinia in an aqueous and / or alcohol-based carrier.

6. The method of claim 1 wherein: the solution or suspension coats the one or more stages of blades.

7. The method of claim 1 wherein: the one or more stages of blades alternate with one or more stages of vanes.

8. The method of claim 1 wherein: the spraying is through the plurality of fuel nozzles.

9. The method of claim 1 wherein: the spraying is through a circumferentially distributed plurality of spray nozzles (26, 26A, 26B).

10. The method of claim 1 wherein: the spray nozzles are coaxial with respective associated fuel nozzles.

11. The method of claim 1 wherein: the gas turbine engine is a propulsion engine of an aircraft (700); and the spraying is from a reservoir (22, 24) contained within the aircraft.

12. The method of claim 1 wherein: the gas turbine engine is a propulsion engine of an aircraft (900); and the spraying is from an external reservoir (22, 24) connected to the aircraft via a hose (324).

13. The method of claim 12 wherein: the hose is a single hose connected to a single port (326) to treat multiple engines.

14. The method of claim 1 wherein: the spraying comprises blending the binder and a solvent or carrier and spraying the blend.

15. The method of claim 14 further comprising: after spraying the blend, a purge.

16. The method of claim 15 wherein: the purge is for not more than half a duration of the spraying.

17. The method of claim 1 further comprising: shutting down the engine; monitoring a temperature of the engine; and responsive to the monitoring showing temperature decrease to a predetermined level, commencing the spraying.

18. The method of claim 1 further comprising: driving rotation of a spool of the engine during the spraying.

19. The method of claim 18 wherein: the driving comprises directing air from an auxiliary power unit (APU) (730).

20. The method of claim 19 wherein: the driving is at a speed of 10.0-100.0 RPM.

21. The method of claim 18 wherein: the driving comprises directing air from a ground power unit (GPU or “huffer cart”) (300).

22. The method of claim 21 wherein: the solution is stored in the huffer cart or a further service cart; and a hose (324) connects the huffer cart or further service cart to the aircraft for delivering the solution.

23. An aircraft (700) comprising: at least one gas turbine engine (720, 720A, 720B) comprising: a compressor section (840, 842); a combustor section (844) including a plurality of fuel nozzles (860); a turbine section (846, 848) having one or more stages of blades (400); a fuel tank (760); and a fuel flowpath from the fuel tank to the plurality of fuel nozzles, wherein the aircraft further comprises: a reservoir of a CMAS binder; anda binder flowpath from the reservoir to the fuel nozzles or additional fluid nozzles, if any in the combustor.

24. The aircraft of claim 23 comprising: the combustor comprises said additional fluid nozzles; and the binder flowpath is from the reservoir to the additional fluid nozzles.

25. The aircraft of claim 23 wherein: the at least one gas turbine engine is a plurality of gas turbine engines; the reservoir is shared by the plurality of gas turbine engines; and binder flow is separately controllable for the plurality of gas turbine engines.

26. A vehicle (900) having a gas turbine engine (720, 720A, 720B), the engine comprising: a compressor section (840, 842); a combustor section (844) including a plurality of fuel nozzles (860); and a turbine section (846, 848) having one or more stages of blades (400), wherein: the combustor comprises additional fluid nozzles (26, 26A, 26B); and the additional fluid nozzles are coupled to a port (326).

27. The vehicle of claim 26 wherein: the additional fuel nozzles are concentric with or circumferentially alternating with the fuel nozzles; and the gas turbine engine has a temperature sensor coupled to a connector on the vehicle.

28. A method for using the vehicle of claim 26, the method comprising: connecting a fluid source (22, 24) to the port; delivering through the port a solution or suspension (23, 440); and spraying into the combustor section via the additional nozzles of the gas turbine engine, the solution or suspension comprising a binder selected from: gadolinium; and gadolinium compounds.

29. A method for using a gas turbine engine (720, 720A, 720B), the gas turbine engine comprising: a compressor section (840, 842); a combustor section (844) including a plurality of fuel nozzles (860); and a turbine section (846, 848) having one or more stages of blades (400), the method comprising: spraying a solution or suspension (23, 440) into the combustor section via a circumferentially distributed plurality of nozzles of the gas turbine engine, the solution or suspension comprising a binder.

30. The method of claim 29 wherein: the binder comprises at least one metal or compound selected from the group consisting of lanthanum, cerium, praseodymium, neodymium, promethium, samarium, europium, gadolinium, terbium, dysprosium, holmium, erbium, thulium, ytterbium, lutetium, indium, scandium, yttrium, zirconium, hafnium, titanium, and mixtures thereof as well as compounds of such metals.

31. The method of claim 29 wherein: the binder is selected from: gadolinium; and gadolinium compounds.

32. An aircraft service cart (300) comprising: a reservoir (22) of a CMAS binder; a binder flowpath from the reservoir to an outlet (322); and means (32) for driving a flow from the reservoir to the outlet.

33. The service cart of claim 32 being a ground start cart having at least one of: an electrical connection (332) for starting an aircraft; and an air connection (352) for starting an aircraft.

34. The service cart of claim 32 wherein: the binder is in solution or suspension.

5. The service cart of claim 32 further comprising: a reservoir of a purge liquid (24); and a pump (34) for pumping the purge liquid to the outlet.

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