Patent Grant
10927687
This disclosure generally relates to gas turbine engine blade containment systems, and more specifically relates to gas turbine engine blade containment systems having a casing made of a first material coaxially surrounded by a ring made of a second material.
Gas turbine engines are well known internal combustion engines typically used to provide thrust to an aircraft or to provide power for land-based operations. Generally speaking, a gas turbine engine includes a fan, a compressor, a combustor and a turbine arranged in a serial fashion. The fan draws in ambient air as it rotates and moves it to the compressor where the air is compressed or pressurized. The compressed air is then communicated to the combustor where it is mixed with fuel and ignited. The products of the combustion are hot gases which are then directed into the turbine. This causes the airfoils in the turbine to rotate, and as turbine is mounted on the same shaft, or shafts, as the compressor and fan, this causes the compressor and fan to rotate as well. Accordingly, once started, it can be seen that the operation of the engine is self-sustaining in that the combustion of more fuel causes more rotation of the turbine and in turn the compressor and the fan. Moreover, the rotation of the fan, which typically has a diameter many times that of the compressor and the turbine, causes the engine to generate thrust.
In order to increase operational efficiency, and thereby decrease fuel consumption, new gas turbine engine designs require the turbine section casings to operate at greater temperatures than ever before. This means that the casing of new gas turbines must be made of materials that are more resistant to heat than the legacy substances they are substituting. However, the casing must also have sufficient resilience to resist foreign object damage and contain blade liberation. Foreign object damage is a constant concern of the aerospace industry, as any impingement of ice, water, sand, dirt, animals (e.g., birds) and other foreign objects found in the air or on the ramps, taxiways and runways of airports can damage the turbine section of the engine and in turn detrimentally affect thrust generation.
One way to increase resistance to foreign object damage is by increasing the thickness of the casing. However, while increasing the thickness of the casing improves its resilience to foreign object damage, the increased thickness also increases the mass of the engine, and thereby offsets the efficiency gained by operating the engine at elevated temperatures. Thus, while certain gas turbine engine blade containment systems are known, improvements to the aforementioned areas of mass and resilience to foreign object damage are desired, all while maintaining the increased efficiency obtained when operating the engine at increased temperatures.
In accordance with one aspect of the present disclosure, a gas turbine engine blade containment system is disclosed. The containment system may include a generally cylindrical casing being made of a first material and a generally cylindrical ring made of a second material coaxially surrounding the casing. At least some portion of the ring may be metallurgically bonded to the casing.
In a refinement of the blade containment system, the containment system may further include a generally cylindrical second ring comprising a third material axially spaced apart from the ring, at least some portion of the second ring metallurgically bonded to the casing.
In another refinement of the blade containment system, the second material and third material are the same.
In another refinement of the blade containment system, the second material and third material are different.
In another refinement of the blade containment system, the containment system may further include a gap positioned between the casing and the ring.
In another refinement of the blade containment system, the containment system may further include a sheet metal core positioned between the casing and the ring.
In another refinement of the blade containment system, the containment system may further include a rib connected to at least some portion of the sheet metal core and extending through the ring.
In another refinement of the blade containment system, the containment system may further include metallic foam positioned between the casing and the ring.
In accordance with another aspect of the present disclosure, a gas turbine engine is disclosed. The gas turbine engine may have a fan section, a compressor section downstream from the fan section, a combustor section downstream from the compressor section and a turbine section downstream from the compressor section. The turbine section may include a blade containment system. The blade containment system may include a generally cylindrical casing being made of a first material and a generally cylindrical ring made of a second material coaxially surrounding the casing. At least some portion of the ring may be metallurgically bonded to the casing.
In a refinement of the gas turbine engine, the blade containment system may include a generally cylindrical second ring being made of a third material axially spaced apart from the ring, at least some portion of the second ring metallurgically bonded to the casing.
In another refinement of the gas turbine engine, the second material and third material are the same.
In another refinement of the gas turbine engine, the second material and third material are different.
In another refinement of the gas turbine engine, the blade containment system may further include a sheet metal core positioned between the casing and the ring.
In another refinement of the gas turbine engine, the blade containment system may further include a rib connected to at least some portion of the sheet metal core and extending through the ring.
In another refinement of the gas turbine engine, the blade containment system may further include a gap positioned between the casing and the ring.
In another refinement of the gas turbine engine, the blade containment system may further include metallic foam positioned between the casing and the ring.
In accordance with another aspect of the present disclosure, a method for manufacturing a gas turbine engine blade containment system is disclosed. The method may include the steps of providing a generally cylindrical casing being made of a first material, then depositing a second material onto the casing that coaxially surrounds the casing to form a ring, at least a portion of the ring metallurgically bonded to the casing.
In a refinement of the method for manufacturing a gas turbine engine blade containment system, the depositing process may be selected from the group consisting of electron beam melting, electron beam wire deposition, direct metal laser sintering, selective laser sintering, laser beam powder deposition, selective laser melting, cold metal transfer, gas dynamic cold spraying, high velocity oxy-fuel coating spraying (HVOF) and combinations thereof.
In another refinement of the method for manufacturing a gas turbine engine blade containment system, the first material may be selected from the group consisting of titanium, titanium alloy, nickel, nickel alloy, steel, aluminum, aluminum alloy and combinations thereof.
In another refinement of the method for manufacturing a gas turbine engine blade containment system, the second material may be selected from the group consisting of titanium, titanium alloy, nickel, nickel alloy, steel, aluminum, aluminum alloy and combinations thereof.
These and other aspects and features of the present disclosure will be more readily understood when read in conjunction with the accompanying drawings.
Referring now to the drawings, and with specific reference to
The engine 20 generally includes a low speed spool 30 and a high speed spool 32 mounted for rotation about an engine central longitudinal axis A relative to an engine static structure 36 via several bearing systems 38. It should be understood that various bearing systems 38 at various locations may alternatively or additionally be provided.
The low speed spool 30 generally includes an inner shaft 40 that interconnects a fan blade assembly 42, a low pressure (or first) compressor section 44 and a low pressure (or first) turbine section 46. The inner shaft 40 is connected to the fan blade assembly 42 through a geared architecture 48 to drive the fan assembly 42 at a lower speed than the low speed spool 30. The high speed spool 32 includes an outer shaft 50 that interconnects a high pressure (or second) compressor section 52 and high pressure (or second) turbine section 54. The outer shaft 50 is typically concentric with and radially outward from the inner shaft 50. A combustor 56 is arranged between the high pressure compressor 52 and the high pressure turbine 54. A mid-turbine frame 57 of the engine static structure 36 is arranged generally between the high pressure turbine 54 and the low pressure turbine 46. The mid-turbine frame 57 supports one or more bearing systems 38 in the turbine section 28. The inner shaft 40 and the outer shaft 50 are concentric and rotate via bearing systems 38 about the engine central longitudinal axis A, which is collinear with their longitudinal axes. As used herein, a “high pressure” compressor or turbine experiences a higher pressure than a corresponding “low pressure” compressor or turbine.
The core airflow C is compressed first by the low pressure compressor 44, and then by the high pressure compressor 52, before being mixed and burned with fuel in the combustor 56, and lastly expanded over the high pressure turbine 54 and low pressure turbine 46. The mid-turbine frame 57 includes airfoils 59 which are in the core airflow path. The turbines 46, 54 rotationally drive the respective low speed spool 30 and high speed spool 32 in response to the expansion.
The engine 20 in one example is a high-bypass geared aircraft engine. In a high-bypass engine a greater volume of air moves along a bypass flowpath B than through core airflow C. The ratio of the mass of air moving through bypass flowpath B to core airflow C is known as the bypass ratio. In a further example, the engine 20 bypass ratio is greater than about six (6), with an example embodiment being greater than ten (10), the geared architecture 48 is an epicyclic gear train, such as a star gear system or other gear system, with a gear reduction ratio of greater than about 2.3 and the low pressure turbine 46 has a pressure ratio that is greater than about 5. In one disclosed embodiment, the engine 20 bypass ratio is greater than about ten (10:1), the fan diameter is significantly larger than that of the low pressure compressor 44, and the low pressure turbine 46 has a pressure ratio that is greater than about 5:1. Low pressure turbine 46 pressure ratio is pressure measured prior to inlet of low pressure turbine 46 as related to the pressure at the outlet of the low pressure turbine 46 prior to an exhaust nozzle. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other gas turbine engines including direct drive turbofans.
Now turning to
The casing 60 may generally be cylindrical in shape and surround turbine section 28. Ring 62 may be generally cylindrical in shape and may coaxially surround casing 60. Ring 62 may be positioned along axis A such that it lies in a plane containing a turbine blade 66. Alternatively, ring 62 may be positioned along axis A such that it lies in a plane containing two or more turbine blades 66. Thus, ring 62 may have a varying width such that it lies in a plane containing only one blade 66, or it may have a width such that it lies in a plane containing two or more blades 66.
The casing 60 may be made of a first material including, but not limited to, metal and metal alloys. The metals from which casing 60 may be made include, but are not limited to, titanium, nickel and aluminum. While the following list is not exhaustive, the metal alloys from which casing 60 may be made include titanium alloys, nickel alloys, aluminum alloys and steel. Some examples of nickel alloys that may be utilized include, but are not limited to, WASPALOY®, INCONEL 718®, INCONEL 718+®, INCONEL 939®, and HAYNES 282®. Some examples of titanium alloys that may be utilized include Ti-6Al-4V and Ti-6Al-2Sn-4Zr-2Mo. Some examples of aluminum alloys that may be utilized include series 2000, 6000 and 7000 aluminum. Examples of steel that may be utilized include series 300 and 400 stainless steel, and precipitation hardenable stainless steel. As would be understood, the choice of the material for the casing may be determined based on the appropriate operating temperature of the gas turbine engine.
Now with reference to
However, this may not be the only bonding configuration that ring 62 may take. As demonstrated in
Further, in an alternative example, the blade containment system 64 of the present disclosure may further include a metallic foam 72 positioned between casing 60 and ring 62 as demonstrated in
Additionally, metallic foam 72 may be used in containment system 64 to increase the rate of heat transfer from casing 66, or even non-linearly dissipate energy across the blade containment system. Since the foam 72 may be used to alter the rate of energy dissipation across blade containment system 64, it may thus also be incorporated into the system 64 such that both foam 72 and gap 70 are positioned between the casing 60 and ring 62 (not shown). In this instance the foam 70 is positioned next to casing 60 thereby helping to increase the rate of heat transfer from casing 60 and while providing a differential rate of energy dissipation across containment system 64. The shape of the containment ring to accommodate either the metallic foam or the gap may be superior to a monolithic ring since the foam or the gap may increase the compliance of the ring assembly in the axial and radial directions. This applies to both pressure loads and thermal loads. This compliance reduces the local stress concentrations that could be associated with the edges/ends of a monolithic ring, and improves the life of the part while maintaining containment capability and reducing the weight of the part.
Referring now to
Moving onto
In an additional embodiment within the present disclosure, the containment system 64 may include one or more standups 80 as demonstrated in
While still referring to
Now while taking reference to the blade containment system discussed herein, and
The specific choice of material that is utilized to construct ring 62 may be based on the design of the containment system 64. Thus, when incorporating a metallic foam 72 into the containment system 64 (such as
With reference to
Now with reference to
While the foregoing describes a gas turbine engine 20 and a blade containment system 64, the present disclosure also recites methods for making a blade containment system for a gas turbine engine. An embodiment of the method is depicted in the flowchart of
In an additional, although not necessary, step 88 the casing may be prepared for the subsequent depositing a second material onto the casing step 90. In one example of this optional step, a sheet metal core, and any attached ribs if utilized, may be welded or brazed onto the casing, and then the second material may deposited onto the casing such that the sheet metal core is positioned between the casing and the ring, and any ribs extend through the ring. In an additional example, the metallic foam is laid onto or attached to the casing such that it is positioned between the casing and the ring once the second material is added onto the casing to form the ring. In another example, the stand-ups are welded or brazed onto the casing before the deposition of the second material to form the ring. In a further instance, the stand-ups are formed on the casing by one or more of the same manufacturing methods discussed below in regards to the deposition of the ring in step 88, before depositing the second material to form the ring. Then, the standups may be positioned intermediate the casing and the ring.
Subsequent the depositing step 90, additional optional steps may also be incorporated into the method of manufacturing a blade containment system. If the casing provided is wrought manufactured, then subsequent to the deposition step 90, the blade containment system may be heat treated in step 92 to improve the metallurgical bond formed between the casing and the ring during the deposition step. Following heat treatment step 92, the containment system may be machined to final specifications in step 94.
If the casing provided is a cast manufactured, then the containment system may be treated with hot isostatic pressing (HIP) in the optional step 96. As further shown in
The depositing the second material onto the casing to form the ring step 88 may be undertaken with a variety of methods. In one example, electron beam melting may be utilized. Electron beam melting is a process where a metal or metal alloy powder is melted with an electron beam in a layer-by-layer additive process. Thus, a layer of metal or metal alloy (e.g., titanium or titanium alloy) may be deposited onto the casing, and then an electron beam is utilized to melt the material to the casing. In another example, electron beam wire deposition may be utilized to deposit the ring onto the casing. In electron beam wire deposition a wire made of metal or metal alloy (e.g., titanium, titanium alloys, nickel or nickel alloy) is continuously fed into the path of an electron beam that is in close physical proximity to the casing, thereby melting the material onto the casing. The beam and wire combination may be moved in the X, Y or Z planes to create the ring in an additive process.
In a further example, direct metal laser sintering may be used to deposit the ring onto the casing. In direct metal layer sintering, a metal or metal alloy powder is sintered together with a laser beam in a layer-by-layer additive process. Thus, a layer of metal or metal alloy (e.g., titanium, titanium alloy, nickel or nickel alloy) may be deposited onto the casing, and then the laser is utilized to coalesce and sinter together the metal or metal alloy powders. In a close cousin to direct metal laser, selective laser sintering may also be used to deposit the ring. In selective laser sintering, a granular bed of metal or metal alloy powder (e.g., titanium, titanium alloy, nickel or nickel alloy), and then a laser is used to selectively fuse the powder together. Then, another layer of powder is added and the process is repeated. Selective laser melting may also be used. In selective maser melting, which is similar to selective laser sintering, a granular bed of metal or metal alloy powder (e.g., titanium, titanium alloy, nickel or nickel alloy), is laid down then then a laser is used to melt, rather than fuse, the powder granules together. Then, another layer of powder is added and the process is repeated. In another cousin to direct metal laser sintering, laser beam powder deposition may be used. Laser beam powder deposition is a process in which metal powder is injected into the focused beam of a high-power laser under tightly controlled atmospheric conditions. The focused laser beam melts the surface of the target material and generates a small molten pool of based material. The powder delivered to this spot is absorbed into the melt pool, thus generating a deposit thereby building a feature.
Next, cold metal transfer may be used to construct the ring. The cold metal transfer process is an engineered cousin of arc welding in which a filler metal repeatedly moves into and out of a weld pool. Moving the filler metal into the weld pool extinguishes the arc between the filler metal and the substrate, while moving the filler metal out of the pool regenerates the arc thereby melting more fill material for addition to the melt pool. Cold metal transfer is a gas metal arc welding (GMAW) process where the filler metal moves into and out of the weld pool, repeatedly forming and extinguishing the arc.
Additionally, gas dynamic cold spraying may be used to deposit the ring onto the casing. In gas dynamic cold spraying, metal or metal alloy powders are entrained in a high velocity gas stream, both of which then impinge the surface to be worked on. During impact with the surface the power particles undergo a plastic deformation and the particles adhere to the surface without any melting of the particles. Then, another layer of powder may be added to the previous layer. Heat treating may be used to improve the metallurgical bond between the powder and the casing, and between the powder particles themselves.
Finally, high velocity oxy-fuel coating spraying (HVOF) may be utilized to deposit the ring onto the casing. Like gas dynamic cold spraying, metal or metal alloy powders are entrained in a high velocity gas stream. However, in HVOF the gas stream is at an elevated temperature around the melting temperature of the powder. Thus, unlike gas dynamic cold spraying, the powder partially melts in the gas stream, and thus when the powder particles impinge the surface of the material to be worked upon, not only do the particles adhere due to plastic deformation, they also melt and fuse with the surface and with each other. Then, another layer of powder may be added to the previous layer.
What now follows are four examples demonstrating methods of making a blade containment system for a gas turbine engine in accordance with the method disclosed herein. In a first example of a method of making a blade containment system for a gas turbine engine, a casing made of wrought HAYNES 282® is provided. Next, the blade containment ability of the wrought casing is enhanced by depositing a ring of INCONEL 718® coaxially around the outside of the casing via the additive manufacturing process electron beam wire deposition. Finally, the metallurgical bond between the casing and the ring may be improved by a hybrid heat treatment process wherein the system is heated at a preselected temperature, followed by being cooled at a rate of air cooled or faster, and ending with a precipitation heat treating at a chosen temperature.
In an additional example of the method of making a blade containment system for a gas turbine engine, a casing made of cast HAYNES 282® is provided. Next, the blade containment ability of the cast casing is enhanced by depositing a ring of INCONEL 718® coaxially around the outside of the casing via the additive manufacturing process high velocity oxy-fuel coating spraying (HVOF). Next, the casing and ring are subjected to Hot Isostatic Pressing (HIP) to close pore porosity and enhance the metallurgical bond between the HAYNES 282® and INCONEL 718®. Finally, the metallurgical bond between the casing and the ring may be further improved by a heat treatment process wherein the system is heated to a preferred temperature, cooled at a rate of air cool or faster and followed by precipitation heat treatment at a selected temperature.
In another example of the method for making a blade containment system for a gas turbine engine, a casing made of cast WASPALOY® is provided. Then, the blade containing capacity of the casing is improved by depositing a ring of INCONEL 718® coaxially around the casing by the additive manufacturing process laser powder deposition. Subsequently, the casing and the added ring are treated by HIP to close porosity and promote the metallurgical bond formed between the WASPALOY® and INCONEL 718®. Finally, the casing and system is heat treated to the WASPALOY® process wherein the system is heated to a preferred temperature, followed by cooling at a rate of air cooled or faster, followed with a stabilization heat treatment at a selected temperature and cooled at a rate of air cooled or faster, and rounded out with precipitation heat treatment at a selected temperature.
In a further example of the method of making a blade containment system for a gas turbine engine, a cast INCONEL 939® casing is provided. Then a ring of INCONEL 718® is deposited onto the casing so that it coaxially surrounds the casing with the electron beam wire additive manufacturing process. Next, the casing and the added ring are treated by HIP to close porosity and promote the metallurgical bond formed between the INCONEL 939® and INCONEL 718®. Finally, the casing and system is heat treated to the INCONEL 939® process whereby the system is heated to a preselected temperature, followed by cooling at rate of air cooled or faster, followed with a stabilization heat treatment at selected temperature and cooled at a rate of air cooled or faster, and ending with a precipitation heat treatment at a chosen temperature.
In operation, the blade containment system of the present disclosure can find applicability in many industrial settings including, but not limited to, gas turbine engines for use in aircraft. More specifically, various gas turbine engine blade containment systems are disclosed. The blade containment systems disclosed herein have a generally cylindrical casing being made of a first material, and a generally cylindrical ring being made of a second material coaxially surrounding the casing, at least some portion of the ring metallurgically bonded to the casing. The casing made of a first material allows the engine to operate at higher temperature thereby increasing efficiency, while the ring made of a second material coaxially surrounding the casing increases the engine's blade containment capability for a given system weight.
The above description is meant to be representative only, and thus modifications may be made to the embodiments described herein without departing from the scope of the disclosure. Thus, these modifications fall within the scope of present disclosure and are intended to fall within the appended claims.
This Application is a divisional of U.S. patent application Ser. No. 14/794,576 filed on Jul. 8, 2015, which is non-provisional patent application claiming priority under 35 USC § 119(e) to U.S. Provisional Patent Application Ser. No. 62/036,677 filed on Aug. 13, 2014, the contents each of which are incorporated herein by reference thereto.
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| Number | Date | Country | |
|---|---|---|---|
| 20190106996 A1 | Apr 2019 | US |
| Number | Date | Country | |
|---|---|---|---|
| 62036677 | Aug 2014 | US |
| Number | Date | Country | |
|---|---|---|---|
| Parent | 14794576 | Jul 2015 | US |
| Child | 16207937 | US |
