Gearbox for aircraft gas turbine engine

Abstract

The present disclosure relates to gearboxes for aircraft gas turbine engines, particularly to arrangements for journal bearings in such gearboxes, and to related methods of operating such gearboxes and gas turbine engines. An exemplary embodiment includes a gearbox (30) for an aircraft gas turbine engine, the gearbox (30) comprising: a sun gear (28); a plurality of planet gears (32) surrounding and engaged with the sun gear (28); and a ring gear (38) surrounding and engaged with the plurality of planet gears (32), each of the plurality of planet gears (32) being rotatably mounted about a pin (41) of a planet carrier (51), the planet carrier having a journal bearing (42) having an inner sliding surface on the planet gear (32) and an outer sliding surface on the pin (41).

Description

Technical Field

[0001] This disclosure relates to gearboxes for gas turbine engines in aircraft, and more particularly to arrangements of journal bearings in such gearboxes, and to methods for operating such gearboxes and gas turbine engines. Background Technology

[0002] Gas turbine engines with large-diameter fans can incorporate a gearbox that connects the fan to the spindle of the engine core. The advantage of this is that both the fan and the engine core can be designed to operate efficiently even when the fan size is scaled up, as the fan's rotational speed is limited by the tangential velocity of its tip. The gearbox allows for a lower fan speed compared to the engine core's rotational speed, at the cost of additional weight and some efficiency losses within the gearbox. To maintain engine operating efficiency, the gearbox needs to be designed to minimize weight and maximize efficiency. Bearings are a source of losses within the gearbox and therefore need to be optimized to maximize its efficiency. Summary of the Invention

[0003] According to a first aspect, a gearbox for an aircraft gas turbine engine is provided, the gearbox comprising:

[0004] Sun gear;

[0005] Multiple planetary gears that surround and engage with the sun gear; and

[0006] A ring gear surrounds and engages with a plurality of planetary gears, each of which is rotatably mounted about a pin of a planetary gear carrier. The planetary gear carrier has a journal bearing having an internal sliding surface on the planetary gear and an external sliding surface on the pin.

[0007] The journal bearing has a surface coating on its internal or external sliding surface, which includes an alloy layer with aluminum or copper as the main component.

[0008] Ring gears can have a pitch circle diameter of approximately 550 mm or greater.

[0009] Each planetary bearing in the planetary bearing system may have a maximum specific operating load and a maximum operating sliding speed, wherein the maximum specific operating load multiplied by the maximum operating sliding speed is about 240 MPa m / s or greater. The maximum operating sliding speed may be about 30 m / s or greater, and optionally not greater than about 60 m / s. The maximum specific operating load may be about 7 MPa or greater.

[0010] The maximum specific operating load multiplied by the maximum operating sliding speed can be less than about 720 MPa m / s.

[0011] A surface coating can be applied to the outer sliding surface of each journal bearing.

[0012] The external sliding surface of each journal bearing may be located on a sleeve mounted around the corresponding pin.

[0013] The thickness of the surface coating can range from approximately 40 micrometers to approximately 200 micrometers.

[0014] The thickness of this layer can range from approximately 40 micrometers to approximately 100 micrometers.

[0015] A gas turbine engine for an aircraft may include: an engine core including a turbine, a compressor, and a spindle connecting the turbine to the compressor; a fan located upstream of the engine core, the fan including a plurality of blades; and a gearbox according to the first aspect, the gearbox being configured to receive input from the spindle and provide output drive to the fan so as to drive the fan at a lower rotational speed than the spindle.

[0016] In the case where the turbine is a first turbine, the compressor is a first compressor, and the spindle is a first spindle, the engine core may also include a second turbine, a second compressor, and a second spindle connecting the second turbine to the second compressor, the second turbine, the second compressor, and the second spindle being arranged to rotate at a higher rotational speed than the first spindle.

[0017] According to the second aspect, a method for operating a gas turbine engine is provided, the method comprising operating the engine under maximum takeoff conditions, wherein for each journal bearing in the gearbox, a specific load multiplied by a slip velocity is greater than approximately 240 MPa / s.

[0018] The specific load of each journal bearing multiplied by the sliding speed can be less than approximately 720 MPa m / s.

[0019] According to a third aspect, a gearbox for an aircraft gas turbine engine is provided, the gearbox comprising:

[0020] Sun gear;

[0021] Multiple planetary gears that surround and engage with the sun gear; and

[0022] A ring gear surrounds and engages with a plurality of planetary gears, each of which is rotatably mounted about a pin of a planetary gear carrier. The planetary gear carrier has a journal bearing having an internal sliding surface on the planetary gear and an external sliding surface on the pin.

[0023] The ratio of the length L of the inner and outer sliding surfaces to the diameter D of the journal bearing is between approximately 0.5 and 1.4.

[0024] Ring gears can have a pitch circle diameter of approximately 550 mm or greater.

[0025] In some examples, the L / D ratio can be between approximately 1.1 and 1.3.

[0026] Each planetary bearing in the planetary bearing may have a maximum specific operating load and a maximum operating sliding speed, wherein the maximum specific operating load multiplied by the maximum operating sliding speed is about 240 MPa m / s or greater.

[0027] The maximum specific operating load multiplied by the maximum operating sliding speed can be less than about 720 MPa m / s.

[0028] The pitch circle diameter of the gear ring may not exceed 1200 mm.

[0029] A gas turbine engine for an aircraft may include: an engine core including a turbine, a compressor, and a spindle connecting the turbine to the compressor; a fan located upstream of the engine core, the fan including a plurality of blades; and a gearbox according to a third aspect, the gearbox being configured to receive input from the spindle and provide output drive to the fan so as to drive the fan at a lower rotational speed than the spindle.

[0030] In the case where the turbine is a first turbine, the compressor is a first compressor, and the spindle is a first spindle, the engine core may also include a second turbine, a second compressor, and a second spindle connecting the second turbine to the second compressor, the second turbine, the second compressor, and the second spindle being arranged to rotate at a higher rotational speed than the first spindle.

[0031] According to the fourth aspect, a method for operating a gas turbine engine is provided, the method comprising operating the engine under maximum takeoff conditions, wherein for each journal bearing in the gearbox, a specific load multiplied by a slip velocity is greater than approximately 240 MPa / s.

[0032] The specific load of each journal bearing multiplied by the sliding speed can be less than approximately 720 MPa m / s.

[0033] According to a fifth aspect, a method for operating a gearbox for an aircraft gas turbine engine is provided, the gearbox comprising:

[0034] Sun gear;

[0035] Multiple planetary gears that surround and engage with the sun gear; and

[0036] A ring gear that surrounds and engages with the plurality of planetary gears, the ring gear having a pitch circle diameter of approximately 550 mm or greater.

[0037] Each of the plurality of planetary gears is rotatably mounted around a pin of a planetary gear carrier, the planetary gear carrier having a journal bearing having an inner sliding surface on the planetary gear and an outer sliding surface on the pin, the oil film between the inner sliding surface on the planetary gear and the outer sliding surface on the pin varying between a maximum thickness and a minimum thickness around the journal bearing.

[0038] The method involves operating the aircraft's gas turbine engine under maximum takeoff conditions such that the minimum thickness of the oil film differs from the average minimum oil film thickness among the plurality of planetary gears by no more than 8%.

[0039] The diameter D of each journal bearing can be between approximately 120 mm and approximately 200 mm.

[0040] The ratio of the length L of the inner and outer sliding surfaces to the diameter D of the journal bearing is between approximately 0.5 and approximately 1.4. The ratio L / D can be between approximately 1.1 and approximately 1.3.

[0041] The average minimum oil film thickness under maximum takeoff conditions can be between approximately 3.5 micrometers and 8 micrometers.

[0042] During operation of the gas turbine engine under maximum takeoff conditions, the eccentricity ratio of each journal bearing can be in the range of approximately 0.94 to 0.97.

[0043] For each journal bearing in the gearbox, a specific load multiplied by the sliding speed can be greater than approximately 240 MPa m / s.

[0044] The specific load of each journal bearing multiplied by the sliding speed can be less than approximately 720 MPa m / s.

[0045] According to a sixth aspect, a gearbox for an aircraft gas turbine engine is provided, the gearbox comprising:

[0046] Sun gear;

[0047] Multiple planetary gears that surround and engage with the sun gear; and

[0048] A ring gear that surrounds and engages with the plurality of planetary gears.

[0049] Each of the plurality of planetary gears is rotatably mounted around a pin of a planetary gear carrier, the planetary gear carrier having a journal bearing having an inner sliding surface on the planetary gear and an outer sliding surface on the pin, the oil film between the inner sliding surface on the planetary gear and the outer sliding surface on the pin varying between a maximum thickness and a minimum thickness around the journal bearing.

[0050] Furthermore, during the operation of the aircraft gas turbine engine under maximum takeoff conditions, the difference between the minimum oil film thickness and the average minimum oil film thickness among the plurality of planetary gears does not exceed 8%.

[0051] The various optional features described above regarding the fifth aspect can also be applied to the sixth aspect.

[0052] A gas turbine engine for an aircraft may include: an engine core including a turbine, a compressor, and a spindle connecting the turbine to the compressor; a fan located upstream of the engine core, the fan including a plurality of blades; and a gearbox according to the sixth aspect, the gearbox being configured to receive input from the spindle and provide output drive to the fan so as to drive the fan at a lower rotational speed than the spindle.

[0053] In the case where the turbine is a first turbine, the compressor is a first compressor, and the spindle is a first spindle, the engine core may also include a second turbine, a second compressor, and a second spindle connecting the second turbine to the second compressor, the second turbine, the second compressor, and the second spindle being arranged to rotate at a higher rotational speed than the first spindle.

[0054] According to the seventh aspect, a gearbox for an aircraft gas turbine engine is provided, the gearbox comprising:

[0055] Sun gear;

[0056] Multiple planetary gears that surround and engage with the sun gear; and

[0057] A ring gear surrounds and engages with a plurality of planetary gears, each of which is rotatably mounted about a pin of a planetary gear carrier. The planetary gear carrier has a journal bearing having an internal sliding surface on the planetary gear and an external sliding surface on the pin.

[0058] During the operation of the aircraft's gas turbine engine under maximum takeoff conditions, the specific operating load of each journal bearing multiplied by the operating slip velocity is approximately 300 MPa m / s or greater.

[0059] Ring gears can have a pitch circle diameter of approximately 550 mm or greater.

[0060] During the operation of the aircraft's gas turbine engine under maximum takeoff conditions, the specific operating load of each journal bearing multiplied by the operating slip velocity may not exceed approximately 720 MPa m / s.

[0061] During operation of the aircraft's gas turbine engine under maximum takeoff conditions, the sliding speed of each journal bearing can be greater than approximately 30 m / s or 35 m / s.

[0062] During operation of the aircraft's gas turbine engine under maximum takeoff conditions, the sliding speed of each journal bearing can be less than approximately 49 m / s, 47 m / s, 43 m / s, or 40 m / s.

[0063] During operation of an aircraft gas turbine engine under maximum takeoff conditions, the specific operating load on each journal bearing can be approximately 5 MPa or greater.

[0064] During operation of the aircraft's gas turbine engine under maximum takeoff conditions, the specific operating load on each journal bearing can be less than approximately 20 MPa.

[0065] During operation of an aircraft gas turbine engine under maximum takeoff conditions, the specific operating load on each journal bearing can exceed approximately 10 MPa.

[0066] The diameter clearance of each journal bearing can be between approximately 1‰ and approximately 2‰. The diameter clearance can be between approximately 1.4‰ and approximately 1.6‰.

[0067] A gas turbine engine for an aircraft, the gas turbine engine comprising:

[0068] The engine core includes a turbine, a compressor, and a spindle that connects the turbine to the compressor;

[0069] A fan, located upstream of the engine core, comprising multiple fan blades; and

[0070] According to the gearbox in the seventh aspect, the gearbox is configured to receive input from the spindle and provide output drive for the fan so as to drive the fan at a lower rotational speed than the spindle.

[0071] In the case where the turbine is a first turbine, the compressor is a first compressor, and the spindle is a first spindle, the engine core may also include a second turbine, a second compressor, and a second spindle connecting the second turbine to the second compressor, the second turbine, the second compressor, and the second spindle being arranged to rotate at a higher rotational speed than the first spindle.

[0072] According to the eighth aspect, a method for operating a gas turbine engine of an aircraft is provided, the method comprising operating the gas turbine engine of the aircraft under maximum takeoff conditions such that a specific operating load of each journal bearing multiplied by an operating slip velocity is about 300 MPa m / s or greater.

[0073] The various optional features related to the seventh aspect can also be applied to the eighth aspect.

[0074] According to a ninth aspect, a gearbox for an aircraft gas turbine engine is provided, the gearbox comprising:

[0075] Sun gear;

[0076] Multiple planetary gears that surround and engage with the sun gear; and

[0077] A ring gear surrounds and engages with a plurality of planetary gears, each of which is rotatably mounted about a pin of a planetary gear carrier. The planetary gear carrier has a journal bearing having an internal sliding surface on the planetary gear and an external sliding surface on the pin.

[0078] Among them, when the aircraft's gas turbine engine is operating under maximum takeoff conditions, the eccentricity ratio of each journal bearing (defined as 1-2H) min / c, where H min It is the minimum oil film thickness between the internal and external sliding surfaces, and c is the diameter clearance of the journal bearing) greater than approximately 0.84.

[0079] Ring gears can have a pitch circle diameter of approximately 550 mm or greater.

[0080] When the aircraft's gas turbine engine is operating under maximum takeoff conditions, the eccentricity ratio of each journal bearing can be between approximately 0.94 and 0.97.

[0081] The diameter clearance of each journal bearing can be between approximately 1‰ and approximately 2‰. In some examples, the diameter clearance of each journal bearing can be between approximately 1.4‰ and 1.6‰.

[0082] When the gas turbine engine of an aircraft is operating under maximum takeoff conditions, the temperature of the oil flowing into each journal bearing may not exceed approximately 100°C.

[0083] When the gas turbine engine of an aircraft is operating under maximum takeoff conditions, the pressure of the oil flowing into each journal bearing can range from about 50 kPa to about 350 kPa.

[0084] The inefficiency of each journal bearing (defined as the percentage power loss of an aircraft gas turbine engine operating under maximum takeoff conditions) can be less than about 0.225%. In some examples, the inefficiency of each journal bearing can be no less than about 0.1%.

[0085] A gas turbine engine for an aircraft, the gas turbine engine comprising:

[0086] The engine core includes a turbine, a compressor, and a spindle that connects the turbine to the compressor;

[0087] A fan, located upstream of the engine core, comprising multiple fan blades; and

[0088] According to the gearbox of the ninth aspect, the gearbox is configured to receive input from the spindle and provide output drive for the fan so as to drive the fan at a lower rotational speed than the spindle.

[0089] In the case where the turbine is a first turbine, the compressor is a first compressor, and the spindle is a first spindle, the engine core may also include a second turbine, a second compressor, and a second spindle connecting the second turbine to the second compressor, the second turbine, the second compressor, and the second spindle being arranged to rotate at a higher rotational speed than the first spindle.

[0090] According to the tenth aspect, a method for operating a gas turbine engine is provided, the method comprising operating the aircraft gas turbine engine under maximum takeoff conditions such that the eccentricity ratio of each journal bearing (defined as 1-2H) is such that... min / c, where H min The minimum oil film thickness between the internal and external sliding surfaces (where c is the diameter clearance of the journal bearing) can be greater than approximately 0.84.

[0091] According to the eleventh aspect, a gearbox for an aircraft gas turbine engine is provided, the gearbox comprising:

[0092] Sun gear;

[0093] Multiple planetary gears that surround and engage with the sun gear; and

[0094] A ring gear surrounds and engages with a plurality of planetary gears, each of which is rotatably mounted around a pin of a planetary gear carrier. The planetary gear carrier has journal bearings having an internal sliding surface on the planetary gear and an external sliding surface on the pin. Each journal bearing includes an oil flow path through it from inlet to outlet.

[0095] Specifically, when the aircraft gas turbine engine is operating under maximum takeoff conditions, the temperature of the oil flowing through each oil flow path from the inlet to the outlet increases by 15°C to 30°C, and the temperature of the oil at the inlet is less than 105°C.

[0096] Ring gears can have a pitch circle diameter of approximately 550 mm or greater.

[0097] When the gas turbine engine of an aircraft is operating under maximum takeoff conditions, the temperature of the oil flowing through each oil flow path from the inlet to the outlet increases by 15°C to 25°C or 15°C to 20°C.

[0098] When the aircraft's gas turbine engine operates under maximum takeoff conditions, the specific oil flow rate through each oil flow path (defined as the oil flow rate through the oil flow path divided by the diameter and length of the corresponding journal bearing) can be less than approximately 2000 L / min. -1 m -2 .

[0099] In some examples, the gearbox can be a planetary gearbox, where the ring gear is connected to the output shaft and the sun gear is connected to the input shaft.

[0100] When the gas turbine engine of an aircraft operates under maximum takeoff conditions, the specific oil flow rate through each oil flow path (defined as the oil flow rate through the oil flow path divided by the diameter and length of the corresponding journal bearing) can be less than approximately 1000 L / min. -1 m -2 .

[0101] In some examples, the gearbox can be a stellar gearbox, where the output shaft is connected to the planet carrier, the planet carrier is connected to each planet gear, and the input shaft is connected to the sun gear.

[0102] When the gas turbine engine of an aircraft operates under maximum takeoff conditions, the specific oil flow rate through each oil flow path can be greater than approximately 400 L / min. -1 m -2 .

[0103] When the gas turbine engine of an aircraft is operating under maximum takeoff conditions, the specific operating load of each journal bearing multiplied by the operating slip speed can be approximately 250 MPa m / s or greater.

[0104] When the gas turbine engine of an aircraft is operating under maximum takeoff conditions, the specific operating load of each journal bearing multiplied by the operating slip speed can be up to about 450 MPa m / s.

[0105] When the gas turbine engine of an aircraft is operating under maximum takeoff conditions, the specific operating load of each journal bearing multiplied by the operating slip speed can be approximately 450 MPa m / s or greater.

[0106] When the gas turbine engine of an aircraft is operating under maximum takeoff conditions, the specific operating load of each journal bearing multiplied by the operating slip speed can be up to approximately 720 MPa m / s.

[0107] A gas turbine engine for an aircraft, the gas turbine engine comprising:

[0108] The engine core includes a turbine, a compressor, and a spindle that connects the turbine to the compressor;

[0109] A fan, located upstream of the engine core, comprising multiple fan blades; and

[0110] According to the gearbox of aspect eleven, the gearbox is configured to receive input from the spindle and provide output drive for the fan so as to drive the fan at a lower rotational speed than the spindle.

[0111] In the case where the turbine is a first turbine, the compressor is a first compressor, and the spindle is a first spindle, the engine core may also include a second turbine, a second compressor, and a second spindle connecting the second turbine to the second compressor, the second turbine, the second compressor, and the second spindle being arranged to rotate at a higher rotational speed than the first spindle.

[0112] According to the twelfth aspect, a method of operating a gas turbine engine is provided, the method comprising operating the aircraft gas turbine engine under maximum takeoff conditions such that the temperature of the oil flowing through each oil flow path from the inlet to the outlet increases by 15°C to 30°C and the temperature of the oil at the inlet is less than 105°C.

[0113] The optional features of aspect eleven can also be applied to the method of aspect twelfth.

[0114] According to the thirteenth aspect, a gearbox for an aircraft gas turbine engine is provided, the gearbox comprising:

[0115] Sun gear;

[0116] Multiple planetary gears that surround and engage with the sun gear; and

[0117] A ring gear surrounds and engages with a plurality of planetary gears, each of which is rotatably mounted about a pin of a planetary gear carrier. The planetary gear carrier has a journal bearing having an internal sliding surface on the planetary gear and an external sliding surface on the pin.

[0118] The diameter of each journal bearing divided by the pitch circle diameter of the corresponding planetary gear is less than approximately 55%.

[0119] Ring gears can have a pitch circle diameter of approximately 550 mm or greater.

[0120] The diameter of each journal bearing divided by the pitch circle diameter of the corresponding planetary gear can be greater than approximately 50%.

[0121] When the aircraft's gas turbine engine is operating under maximum takeoff conditions, the sliding speed of each journal bearing can be between approximately 30 m / s and approximately 40 m / s.

[0122] When the gas turbine engine of an aircraft is operating under maximum takeoff conditions, the specific operating load of each journal bearing multiplied by the operating slip speed can be approximately 400 MPa m / s or greater.

[0123] When the gas turbine engine of an aircraft is operating under maximum takeoff conditions, the specific operating load of each journal bearing multiplied by the operating slip speed can be up to approximately 720 MPa m / s.

[0124] The internal or external sliding surfaces of journal bearings may have a surface coating, which includes an alloy layer with aluminum or copper as the main component.

[0125] A gas turbine engine for an aircraft, the gas turbine engine comprising:

[0126] The engine core includes a turbine, a compressor, and a spindle that connects the turbine to the compressor;

[0127] A fan, located upstream of the engine core, comprising multiple fan blades; and

[0128] According to the gearbox of aspect thirteen, the gearbox is configured to receive input from the spindle and provide output drive for the fan so as to drive the fan at a lower rotational speed than the spindle.

[0129] In the case where the turbine is a first turbine, the compressor is a first compressor, and the spindle is a first spindle, the engine core may also include a second turbine, a second compressor, and a second spindle connecting the second turbine to the second compressor, the second turbine, the second compressor, and the second spindle being arranged to rotate at a higher rotational speed than the first spindle.

[0130] According to the fourteenth aspect, a method of operating a gas turbine engine is provided, the method comprising operating an engine core to drive a spindle, and providing output drive from a gearbox to a fan to drive the fan at a lower rotational speed than the spindle.

[0131] The optional features related to the thirteenth aspect are also applicable to the method of the fourteenth aspect.

[0132] According to a fifteenth aspect, there is provided a gearbox for a gas turbine engine of an aircraft, the gearbox comprising:

[0133] a sun gear;

[0134] a plurality of planet gears, the plurality of planet gears surrounding and meshing with the sun gear; and

[0135] a ring gear, the ring gear surrounding and meshing with the plurality of planet gears, each planet gear of the plurality of planet gears being rotatably mounted about a pin of a planet carrier, the planet carrier having a journal bearing having an inner sliding surface on the planet gear and an outer sliding surface on the pin,

[0136] wherein, when the gas turbine engine of the aircraft is operating at maximum take-off conditions, the minimum oil film thickness H between the inner sliding surface and the outer sliding surface min is a function of the temperature T of the oil flowing into the journal bearing, such that H min < B-AT, where A is 0.139 μm / °C and B is 27.8 μm.

[0137] The ring gear may have a pitch circle diameter of about 550 mm or greater.

[0138] In some examples, H min may be greater than 2.3 μm.

[0139] In some examples, H min > B-AT, where A is 0.034 μm / °C and B is 6.4 μm.

[0140] In some examples, H min < B-AT, where A is 0.117 μm / °C and B is 22 μm.

[0141] T may be greater than about 60°C, optionally greater than about 100°C. T may be less than about 120°C.

[0142] A gas turbine engine for an aircraft, the gas turbine engine comprising:

[0143] an engine core, the engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor;

[0144] a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and

[0145] A gearbox according to the fifteenth aspect, the gearbox is configured to receive input from a spindle and provide output drive to a fan so as to drive the fan at a lower rotational speed than the spindle.

[0146] In the case where the turbine is a first turbine, the compressor is a first compressor, and the spindle is a first spindle, the engine core may further include a second turbine, a second compressor, and a second spindle connecting the second turbine to the second compressor, the second turbine, the second compressor and the second spindle being arranged to rotate at a higher rotational speed than the first spindle.

[0147] According to a sixteenth aspect, there is provided a method of operating a gas turbine engine, the method comprising operating an aircraft gas turbine engine at maximum take-off conditions, wherein the minimum oil film thickness H between the inner sliding surface and the outer sliding surface min is a function of the temperature T of the oil flowing into the journal bearing, such that H min <B-AT, where A is 0.139 μm / °C and B is 27.8 μm.

[0148] The optional features according to the fifteenth aspect may also be applicable to the method of the sixteenth aspect.

[0149] According to a seventeenth aspect, there is provided a gearbox for an aircraft gas turbine engine, the gearbox comprising:

[0150] a sun gear;

[0151] a plurality of planet gears surrounding and meshing with the sun gear; and

[0152] a ring gear surrounding and meshing with the plurality of planet gears, wherein each of the plurality of planet gears is rotatably mounted about a pin of a planet carrier, and the planet carrier has a journal bearing having an inner sliding surface on the planet gear and an outer sliding surface on the pin,

[0153] wherein, in the case where the aircraft gas turbine engine is operated at maximum take-off conditions, the eccentricity ratio E of each journal bearing is a function of the temperature T of the oil flowing into the journal bearing, such that E>AT+B, where A is 0.0015 / °C and B is 0.69.

[0154] The ring gear may have a pitch circle diameter of about 550 mm or more.

[0155] The eccentricity ratio may be defined as 1-2H min / c, where H min is the minimum oil film thickness between the inner sliding surface and the outer sliding surface and c is the diametral clearance of the journal bearing.

[0156] In some examples, E may be less than about 0.98.

[0157] In some examples, E < AT+B, where A is 0.00033 / °C and B is 0.94.

[0158] In some examples, E > AT+B, where A is 0.00083 / °C and B is 0.84.

[0159] T may be greater than about 60°C, optionally greater than about 100°C.

[0160] T may be less than about 120°C.

[0161] A gas turbine engine for an aircraft, the gas turbine engine comprising:

[0162] an engine core, the engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor;

[0163] a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and

[0164] A gearbox according to the seventeenth aspect, the gearbox being configured to receive input from the core shaft and provide output drive to the fan so as to drive the fan at a lower rotational speed than the core shaft.

[0165] Where the turbine is a first turbine, the compressor is a first compressor, and the core shaft is a first core shaft, the engine core may further comprise a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor, the second turbine, second compressor and second core shaft being arranged to rotate at a higher rotational speed than the first core shaft.

[0166] According to an eighteenth aspect, there is provided a method of operating a gas turbine engine, the method comprising operating the aircraft gas turbine engine at maximum take-off conditions, wherein the eccentricity ratio E of each journal bearing is a function of the temperature T of the oil flowing into the journal bearing such that E > AT+B, where A is 0.0015 / °C and B is 0.69.

[0167] The optional features of the seventeenth aspect may also be applied to the method of the eighteenth aspect.

[0168] According to a nineteenth aspect, there is provided a gearbox for an aircraft gas turbine engine, the gearbox comprising:

[0169] a sun gear;

[0170] a plurality of planet gears, the plurality of planet gears surrounding and meshing with the sun gear; and

[0171] A ring gear surrounds and engages with a plurality of planetary gears, each of which is rotatably mounted about a pin of a planetary gear carrier. The planetary gear carrier has a journal bearing having an internal sliding surface on the planetary gear and an external sliding surface on the pin.

[0172] In particular, when the aircraft's gas turbine engine is operating under maximum takeoff conditions, the Sommerfeld number of each journal bearing is greater than approximately 4.

[0173] Ring gears can have a pitch circle diameter of approximately 550 mm or greater.

[0174] The inefficiency of each journal bearing (defined as the percentage power loss under maximum takeoff conditions) can be less than approximately 0.225%.

[0175] The diameter clearance of each journal bearing can be between approximately 1‰ and 2‰. The diameter clearance of each journal bearing can be between approximately 1.4‰ and 1.6‰.

[0176] When the gas turbine engine of an aircraft is operating under maximum takeoff conditions, the temperature of the oil flowing into each journal bearing can be less than or equal to about 100°C.

[0177] When the gas turbine engine of an aircraft is operating under maximum takeoff conditions, the pressure of the oil flowing into each journal bearing under maximum takeoff conditions can be in the range of approximately 50 kPa to approximately 350 kPa.

[0178] A gas turbine engine for an aircraft, the gas turbine engine comprising:

[0179] The engine core includes a turbine, a compressor, and a spindle that connects the turbine to the compressor;

[0180] A fan, located upstream of the engine core, comprising multiple fan blades; and

[0181] According to the gearbox of the nineteenth aspect, the gearbox is configured to receive input from the spindle and provide output drive for the fan so as to drive the fan at a lower rotational speed than the spindle.

[0182] In the case where the turbine is a first turbine, the compressor is a first compressor, and the spindle is a first spindle, the engine core may also include a second turbine, a second compressor, and a second spindle connecting the second turbine to the second compressor, the second turbine, the second compressor, and the second spindle being arranged to rotate at a higher rotational speed than the first spindle.

[0183] According to the twentieth aspect, a method for operating a gas turbine engine of an aircraft is provided, the method comprising operating the gas turbine engine of an aircraft under maximum takeoff conditions, wherein the Sommerfeld number of each journal bearing is greater than about 4.

[0184] According to any of the above aspects, the sliding speed of each journal bearing under maximum takeoff conditions may be 30 m / s, 31 m / s, 32 m / s, 33 m / s, 34 m / s, 35 m / s, 36 m / s, 37 m / s, 38 m / s, 39 m / s, 40 m / s, 41 m / s, 42 m / s, 43 m / s, 44 m / s, 45 m / s, 46 m / s, 47 m / s, 48 ​​m / s, 49 m / s, or 50 m / s, or within a range defined by any two of the above values.

[0185] According to any of the foregoing aspects, the specific operating load of each journal bearing under maximum takeoff conditions may be 5 MPa, 6 MPa, 7 MPa, 8 MPa, 9 MPa, 10 MPa, 11 MPa, 12 MPa, 13 MPa, 14 MPa, 15 MPa, 16 MPa, 17 MPa, 18 MPa, 19 MPa, 20 MPa, 21 MPa, 22 MPa, 23 MPa, 24 MPa or 25 MPa or within the range defined by any two of the foregoing values.

[0186] In each of the above aspects, the gas turbine engine includes a turbine, a compressor, a spindle connecting the turbine to the compressor, and a fan located upstream of the engine core, the fan comprising multiple fan blades. The fan may have a size between approximately 5.5 × 10⁻⁶. 7 kg m 2 and 9×10 8 kg m 2 Between, another site is between approximately 7.4 × 10 7 kg m 2 and 7×10 8 kg m 2 Between, another site is between approximately 8.3 × 10 7 kg m 2 and 6.5×10 8 kg m 2 The moment of inertia between them. The same characteristics can also be applied to other aspects of the invention.

[0187] In each of the above aspects, the gearbox may have a gear ratio of 3.2 to 4.5, and optionally a gear ratio of 3.2 to 4.0. The gearbox may be in a star configuration.

[0188] In each of the above aspects, the gas turbine engine can have 70 N kg. -1Up to 90N kg -1 Specific thrust; and / or a bypass ratio of 12.5 to 18 or 13 to 16 under cruise conditions.

[0189] As described elsewhere herein, this disclosure relates to gas turbine engines. Such gas turbine engines may include an engine core comprising a turbine, a combustor, a compressor, and a spindle connecting the turbine to the compressor. Such gas turbine engines may include a fan (with fan blades) located upstream of the engine core.

[0190] A gearbox receives input from the spindle and outputs drive to the fan to drive the fan at a lower rotational speed than the spindle. The input to the gearbox can come directly from the spindle or indirectly, for example, via a spur shaft and / or gears. The spindle rigidly connects the turbine and compressor, causing them to rotate at the same speed (where the fan rotates at a lower speed).

[0191] The gas turbine engine described and / or claimed herein may have any suitable general architecture. For example, the gas turbine engine may have any desired number of shafts connecting the turbine and compressor, such as one shaft, two shafts, or three shafts. By way of example only, the turbine connected to the mandrel may be a first turbine, the compressor connected to the mandrel may be a first compressor, and the mandrel may be a first mandrel. The engine core may also include a second turbine, a second compressor, and a second mandrel connecting the second turbine to the second compressor. The second turbine, the second compressor, and the second mandrel may be arranged to rotate at a higher rotational speed than the first mandrel.

[0192] In such an arrangement, the second compressor may be axially positioned downstream of the first compressor. The second compressor may be arranged to receive flow from the first compressor (e.g., directly, or via a generally annular duct).

[0193] The gearbox can be arranged to be driven by a spindle configured (e.g., in use) to rotate at a minimum rotational speed (e.g., the first spindle in the example above). For example, the gearbox can be arranged to be driven only by a spindle configured (e.g., in use) to rotate at a minimum rotational speed (e.g., in the example above, only the first spindle, not the second spindle). Alternatively, the gearbox can be arranged to be driven by any one or more shafts, such as the first shaft and / or the second shaft in the example above.

[0194] The gearbox can be a reduction gearbox (because the output to the fan rotates at a lower rate than the input from the spindle). Any type of gearbox can be used. For example, the gearbox can be a "planetary" or "stellar" gearbox, as described in more detail elsewhere in this document. The gearbox can have any desired reduction ratio (defined as the rotational speed of the input shaft divided by the rotational speed of the output shaft), for example, greater than 2.5, for example, in the range of 3 to 4.2, or 3.2 to 3.8, for example, approximately or at least 3, 3.1, 3.2, 3.3, 3.4, 3.5, 3.6, 3.7, 3.8, 3.9, 4, 4.1, or 4.2. For example, the gear ratio can be between any two values ​​in the preceding sentence. By way of example only, the gearbox can be a "stellar" gearbox with a gear ratio in the range of 3.1 or 3.2 to 3.8. In some arrangements, the gear ratio can be outside these ranges.

[0195] In any gas turbine engine as described and / or claimed herein, the burner may be axially positioned downstream of the fan and one or more compressors. For example, if a second compressor is provided, the burner may be located directly downstream of the second compressor (e.g., at its outlet). In another example, if a second turbine is provided, the flow at the burner outlet may be directed to the inlet of the second turbine. This burner may be positioned upstream of one or more turbines.

[0196] The compressor, or each compressor (e.g., the first and second compressors as described above), may include any number of stages, such as multiple stages. Each stage may include a row of rotor blades and a row of stator blades, which may be variable stator blades (because the angle of incidence of the row of stator blades may be variable). The row of rotor blades and the row of stator blades may be axially offset from each other.

[0197] The turbine, or each turbine (e.g., the first and second turbines as described above), may include any number of stages, such as multiple stages. Each stage may include a row of rotor blades and a row of stator blades. The row of rotor blades and the row of stator blades may be axially offset from each other.

[0198] Each fan blade may be defined as having a radial span extending from the root (or hub) at a radially inner gas scrubbing position or a 0% span position to the tip at a 100% span position. The ratio of the radius of the fan blade at the hub to the radius of the fan blade at the tip may be less than (or approximately) any of the following: 0.4, 0.39, 0.38, 0.37, 0.36, 0.35, 0.34, 0.33, 0.32, 0.31, 0.3, 0.29, 0.28, 0.27, 0.26, or 0.25. The ratio of the radius of the fan blade at the hub to the radius of the fan blade at the tip may be within a range defined by any two values ​​in the preceding sentence (i.e., these values ​​may form an upper or lower limit), for example, in the range of 0.28 to 0.32. These ratios may generally be referred to as the hub-to-tip ratio. Both the radius at the hub and the radius at the tip may be measured at the leading edge (or axially foremost) portion of the blade. Of course, the hub-to-tip ratio refers to the gas-washing portion of the fan blades, that is, the portion radially outside any platform.

[0199] The radius of the fan can be measured between the engine centerline and the tip of the leading edge of the fan blades. The fan diameter (which may be only twice the fan radius) can be greater than (or approximately) any of the following: 220cm, 230cm, 240cm, 250cm (approximately 100 inches), 260cm, 270cm (approximately 105 inches), 280cm (approximately 110 inches), 290cm (approximately 115 inches), 300cm (approximately 120 inches), 310cm, 320cm (approximately 125 inches), 330cm (approximately 130 inches), 340cm (approximately 135 inches), 350cm, 360cm (approximately 140 inches), 370cm (approximately 145 inches), 380cm (approximately 150 inches), 390cm (approximately 155 inches), 400cm, 410cm (approximately 160 inches), or 420cm (approximately 165 inches). The fan diameter can be within the range defined by any two values ​​in the preceding sentence (i.e., these values ​​can form an upper or lower limit), for example, in the range of 240cm to 280cm or 330cm to 380cm.

[0200] The fan speed can vary during use. Generally, for fans with larger diameters, the speed is lower. By way of non-limiting example only, the fan speed under cruising conditions may be less than 2500 rpm, for example, less than 2300 rpm. By way of another non-limiting example only, for engines with fan diameters in the range of 220cm to 300cm (e.g., 240cm to 280cm or 250cm to 270cm), the fan speed under cruising conditions may be in the range of 1700rpm to 2500rpm, for example, in the range of 1800rpm to 2300rpm, or for example, in the range of 1900rpm to 2100rpm. By way of another non-limiting example only, for engines with fan diameters in the range of 330cm to 380cm, the fan speed under cruising conditions may be in the range of 1200rpm to 2000rpm, for example, in the range of 1300rpm to 1800rpm, or for example, in the range of 1400rpm to 1800rpm.

[0201] When using a gas turbine engine, the fan (with associated fan blades) rotates about an axis of rotation. This rotation causes the tips of the fan blades to move at a speed U. 尖端 Movement. The work done by the fan blades 13 in convection results in an enthalpy increase of dH in the flow. The fan tip load can be defined as dH / U 尖端 2 Where dH is the enthalpy rise across the fan (e.g., 1-D average enthalpy rise), and U 尖端 This is the (translational) velocity of the fan tip, for example at the leading edge of the tip (which can be defined as the fan tip radius at the leading edge multiplied by the angular velocity). The fan tip load under cruise conditions can be greater than (or approximately) any of the following: 0.28, 0.29, 0.30, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39, or 0.4 (all units in this paragraph are in J / kg). -1 K -1 / (ms -1 ) 2 The fan tip load can be within the range defined by any two values ​​in the preceding sentence (i.e., these values ​​can form an upper or lower limit), for example, in the range of 0.28 to 0.31 or 0.29 to 0.3.

[0202] The gas turbine engine according to this disclosure may have any desired bypass ratio, wherein the bypass ratio is defined as the ratio of the mass flow rate of the flow through the bypass duct to the mass flow rate of the flow through the core under cruise conditions. In some arrangements, the bypass ratio may be greater than (or approximately) any of the following: 10, 10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5, 15, 15.5, 16, 16.5, 17, 17.5, 18, 18.5, 19, 19.5, or 20. The bypass ratio may be within a range defined by any two values ​​in the preceding sentence (i.e., these values ​​may form an upper or lower limit), for example, in the range of 12 to 16, or 13 to 15, or 13 to 14. The bypass duct may be substantially annular. The bypass duct may be located radially outside the core engine. The radially outer surface of the bypass duct may be defined by a nacelle and / or fan casing.

[0203] The total pressure ratio of the gas turbine engine described and / or claimed herein can be defined as the ratio of the stagnation pressure upstream of the fan to the stagnation pressure at the outlet of the highest-pressure compressor (before entering the combustor). By way of non-limiting example, the total pressure ratio of the gas turbine engine described and / or claimed herein may be greater than (or approximately) any of the following: 35, 40, 45, 50, 55, 60, 65, 70, 75. The total pressure ratio may be within a range defined by any two values ​​in the preceding sentence (i.e., these values ​​may form an upper or lower limit), for example, in the range of 50 to 70.

[0204] The specific thrust of an engine can be defined as the net thrust of the engine divided by the total mass flow rate through the engine. Under cruise conditions, the specific thrust of the engine described and / or claimed herein may be less than (or approximately) any of the following: 110 Nkg -1 s, 105N kg -1 s, 100N kg -1 s, 95N kg -1 s, 90N kg -1 s, 85N kg -1 s or 80N kg -1 s. This specific thrust can be within a range defined by any two values ​​in the preceding sentence (i.e., these values ​​can form an upper or lower limit), for example, at 80 Nkg. -1 s to 100N kg -1 s, or 85N kg -1 s to 95N kg -1 Within the range of s. Compared to traditional gas turbine engines, this type of engine may be particularly efficient.

[0205] The gas turbine engine described and / or claimed herein may have any desired maximum thrust. By way of non-limiting example only, the gas turbine described and / or claimed herein may produce a maximum thrust of at least (or approximately) any of the following: 160 kN, 170 kN, 180 kN, 190 kN, 200 kN, 250 kN, 300 kN, 350 kN, 400 kN, 450 kN, 500 kN, or 550 kN. The maximum thrust may be within a range defined by any two values ​​in the preceding sentence (i.e., these values ​​may form an upper or lower limit). By way of example only, the gas turbine described and / or claimed herein may be capable of producing a maximum thrust in the range of 330 kN to 420 kN, for example, 350 kN to 400 kN. The thrust mentioned above can be the maximum net thrust under standard atmospheric conditions, at sea level, plus 15°C (ambient pressure 101.3 kPa, temperature 30°C), when the engine is stationary.

[0206] During operation, the temperature of the flow at the inlet of the high-pressure turbine may be particularly high. This temperature, referred to as TET, can be measured at the combustor outlet, for example, just upstream of the first turbine blade, which may be referred to as the nozzle guide vane. At cruising speed, this TET can be at least (or approximately) any of the following: 1400K, 1450K, 1500K, 1550K, 1600K, or 1650K. The cruising TET can be within the range defined by any two values ​​in the preceding sentence (i.e., these values ​​can form an upper or lower limit). The maximum TET of the engine during operation can be, for example, at least (or approximately) any of the following: 1700K, 1750K, 1800K, 1850K, 1900K, 1950K, or 2000K. The maximum TET can be within the range defined by any two values ​​in the preceding sentence (i.e., these values ​​can form an upper or lower limit), for example, in the range of 1800K to 1950K. Maximum TET can occur, for example, under high thrust conditions, such as under maximum takeoff (MTO) conditions.

[0207] The fan blades and / or the airfoil portions of the fan blades described and / or claimed herein may be made of any suitable material or combination of materials. For example, at least a portion of the fan blades and / or airfoil may be made at least partially of a composite material, such as a metal matrix composite and / or an organic matrix composite, such as carbon fiber. As another example, at least a portion of the fan blades and / or airfoil may be made at least partially of a metal, such as a titanium-based metal or an aluminum-based material (such as an aluminum-lithium alloy) or a steel-based material. The fan blade may include at least two regions made of different materials. For example, the fan blade may have a protective leading edge made of a material that is better resistant to impacts (e.g., from birds, ice, or other materials) than the rest of the blade. Such a leading edge may be made, for example, of titanium or a titanium-based alloy. Thus, by way of example only, the fan blade may have carbon fiber or an aluminum-based body (such as an aluminum-lithium alloy) with a titanium leading edge.

[0208] The fan described and / or claimed herein may include a central portion from which fan blades may extend, for example, radially. The fan blades may be attached to the central portion in any desired manner. For example, each fan blade may include a retainer that engages with a corresponding slot in a hub (or disc). By way of example only, such a retainer may be in the form of a dovetail, which may be inserted into and / or engage with a corresponding slot in the hub / disc to secure the fan blade to the hub / disc. In another example, the fan blade may be integrally formed with the central portion. Such an arrangement may be referred to as a blade disc or blade ring. Such blade discs or blade rings may be manufactured using any suitable method. For example, at least a portion of the fan blade may be machined from a block, and / or at least a portion of the fan blade may be attached to the hub / disc by welding (such as linear friction welding).

[0209] The gas turbine engines described and / or claimed herein may or may not be equipped with variable area nozzles (VANs). Such variable area nozzles allow the outlet area of ​​the bypass duct to vary during use. The general principles of this disclosure can be applied to engines with or without VANs.

[0210] The fan of the gas turbine described and / or claimed herein may have any desired number of fan blades, such as 14, 16, 18, 20, 22, 24 or 26 fan blades.

[0211] As used herein, cruise conditions may refer to the cruise conditions of an aircraft to which a gas turbine engine is attached. Such cruise conditions can generally be defined as conditions during the mid-cruise phase, such as the conditions experienced by the aircraft and / or engine at the midpoint between the peak of climb and the start of descent (in terms of time and / or distance).

[0212] By way of example only, the forward speed under cruise conditions can be any point in the range from Mach 0.7 to Mach 0.9, such as 0.75 to 0.85, 0.76 to 0.84, 0.77 to 0.83, 0.78 to 0.82, 0.79 to 0.81, approximately Mach 0.8, approximately Mach 0.85, or within the range of 0.8 to 0.85. Any single speed within these ranges can be considered cruise conditions. For some aircraft, cruise conditions may exceed these ranges, such as below Mach 0.7 or above Mach 0.9.

[0213] By way of example only, cruise conditions may correspond to standard atmospheric conditions at altitudes within the following ranges: 10,000 m to 15,000 m, for example, in the range of 10,000 m to 12,000 m, for example, in the range of 10,400 m to 11,600 m (approximately 38,000 feet), for example, in the range of 10,500 m to 11,500 m, for example, in the range of 10,600 m to 11,400 m, for example, in the range of 10,700 m (approximately 35,000 feet) to 11,300 m, for example, in the range of 10,800 m to 11,200 m, for example, in the range of 10,900 m to 11,100 m, for example, approximately 11,000 m. Cruise conditions may correspond to standard atmospheric conditions at any given altitude within these ranges.

[0214] By way of example only, cruise conditions could correspond to: a forward Mach number of 0.8; a pressure of 23,000 Pa; and a temperature of -55°C. Also by way of example only, cruise conditions could correspond to: a forward Mach number of 0.85; a pressure of 24,000 Pa; and a temperature of -54°C (which could be standard atmospheric conditions at 35,000 feet).

[0215] As used anywhere in this document, “cruising” or “cruising conditions” may refer to an aerodynamic design point (ADP). Such an aerodynamic design point (or ADP) may correspond to the conditions under which a fan is designed to operate (including, for example, one or more of Mach number, environmental conditions, and thrust requirements). For example, this might refer to the conditions under which a fan (or gas turbine engine) is designed to have optimal efficiency.

[0216] In use, the gas turbine engines described and / or claimed herein may operate under cruise conditions defined elsewhere herein. Such cruise conditions can be determined by the cruise conditions of the aircraft (e.g., intermediate cruise conditions), on which at least one (e.g., two or four) gas turbine engines may be mounted to provide propulsive thrust.

[0217] As used herein, maximum takeoff (MTO) conditions have a conventional meaning. MTO conditions can be defined as operating the engine at maximum takeoff thrust at the runway end under International Standard Atmosphere (ISA) sea-level pressure and temperature conditions +15°C. This is typically defined as an aircraft speed of approximately 0.25 Mn, or between approximately 0.24 Mn and 0.27 Mn. Therefore, the engine's maximum takeoff condition can be defined as operating the engine at maximum takeoff thrust (e.g., maximum throttle) under ISA sea-level pressure and temperature +15°C, where the fan inlet speed is 0.25 Mn.

[0218] Those skilled in the art will understand that, unless mutually exclusive, the features or parameters described with respect to any one of the foregoing aspects may be applied to any other aspect. Furthermore, unless mutually exclusive, any feature or parameter described herein may be applied to any aspect and / or combined with any other feature or parameter described herein. Attached Figure Description

[0219] The implementation scheme will now be described by way of example only, with reference to the accompanying drawings, in which:

[0220] Figure 1 This is a cross-sectional side view of a gas turbine engine;

[0221] Figure 2 This is a close-up cross-sectional side view of the upstream section of a gas turbine engine;

[0222] Figure 3 This is a partial cross-sectional view of a gearbox used in a gas turbine engine;

[0223] Figure 4 This is a schematic cross-sectional view of an exemplary planetary gear mounted on a pin of a planetary gear carrier with journal bearings;

[0224] Figure 5 yes Figure 4 A schematic longitudinal sectional view of an exemplary planetary gear;

[0225] Figure 6 This is a partial cross-sectional view of the outer surface of an exemplary planetary gear journal bearing;

[0226] Figure 7 This is a schematic cross-sectional view of an exemplary planetary gear, showing an enlarged view of the oil film thickness variation;

[0227] Figure 8 These are schematic graphs illustrating the variation of specific operating loads with sliding speed for several exemplary gearboxes;

[0228] Figure 9 This is a schematic graph showing the eccentricity ratio of several exemplary gearboxes as a function of percentage load.

[0229] Figure 10 This is a schematic graph illustrating the inefficiency of a series of exemplary gearboxes operating under maximum takeoff conditions as a function of the journal bearing eccentricity ratio.

[0230] Figure 11 This is a schematic graph showing the minimum oil film thickness as a function of inlet temperature for a series of exemplary gearboxes operating under maximum takeoff conditions.

[0231] Figure 12 This is a schematic graph showing the eccentricity ratio of a series of exemplary gearboxes operating under maximum takeoff conditions as a function of inlet oil temperature.

[0232] Figure 13 This is a schematic graph illustrating the inefficiency of a series of exemplary gearboxes operating under maximum takeoff conditions as a function of the Sommerfeld number; and

[0233] Figure 14 This is a schematic graph illustrating the variation of specific oil flow with PV in a series of exemplary gearboxes operating under maximum takeoff conditions. Detailed Implementation

[0234] Figure 1 A gas turbine engine 10 with a main axis of rotation 9 is shown. The engine 10 includes an intake 12 and a propulsion fan 23 that generates two airflows: a core airflow A and a bypass airflow B. The gas turbine engine 10 includes a core 11 that receives the core airflow A. The engine core 11 includes, in axial-flow series, a low-pressure compressor 14, a high-pressure compressor 15, a combustion device 16, a high-pressure turbine 17, a low-pressure turbine 19, and a core exhaust nozzle 20. A nacelle 21 surrounds the gas turbine engine 10 and defines a bypass duct 22 and a bypass exhaust nozzle 18. The bypass airflow B flows through the bypass duct 22. The fan 23 is attached to and driven by the low-pressure turbine 19 via a shaft 26 and a rotary gearbox 30.

[0235] In operation, the core airflow A is accelerated and compressed by the low-pressure compressor 14 and directed to the high-pressure compressor 15 for further compression. The compressed air discharged from the high-pressure compressor 15 is directed to the combustion device 16, where it is mixed with fuel and the mixture is burned. The resulting thermal combustion products then expand through the high-pressure turbine and low-pressure turbines 17, 19 before being discharged through nozzle 20, thereby driving the high-pressure turbine and low-pressure turbine to provide some propulsive thrust. The high-pressure turbine 17 drives the high-pressure compressor 15 via a suitable interconnecting shaft 27. The fan 23 typically provides most of the propulsive thrust. The rotary gearbox 30 is a reduction gearbox.

[0236] Figure 2An exemplary arrangement of a geared fan gas turbine engine 10 is shown. The low-pressure turbine 19 (see [reference]) Figure 1 A drive shaft 26 is connected to the sun gear or sun gear 28 of the planetary gear arrangement 30. A plurality of planet gears 32 mesh with the sun gear 28 radially outward, and are connected together by a planet carrier 34. The planet carrier 34 constrains the planet gears 32 to precess synchronously around the sun gear 28, while simultaneously causing each planet gear 32 to rotate about its own axis. The planet carrier 34 is connected to a fan 23 via a connecting rod 36 to drive the fan to rotate about the engine axis 9. A ring gear or ring gear 38 meshes with the planet gears 32 radially outward, and is connected to a fixed support structure 24 via a connecting rod 40.

[0237] It should be noted that the terms "low-pressure turbine" and "low-pressure compressor" as used herein may refer to the lowest-pressure turbine stage and the lowest-pressure compressor stage (i.e., excluding fan 23), and / or the turbine stage and compressor stage connected together by an interconnecting shaft 26 having the lowest rotational speed in the engine (i.e., excluding the gearbox output shaft driving fan 23). In some literature, the terms "low-pressure turbine" and "low-pressure compressor" mentioned herein may alternatively be referred to as "intermediate-pressure turbine" and "intermediate-pressure compressor." In the case of such alternative nomenclature, fan 23 may be referred to as the first or lowest-pressure compression stage.

[0238] exist Figure 3 The rotary gearbox 30 is shown in more detail by way of example. Each of the sun gear 28, planetary gear 32, and ring gear 38 includes teeth surrounding its periphery for meshing with other gears. However, for clarity, Figure 3 Only exemplary portions of the teeth are shown. Four planetary gears 32 are shown, but it will be apparent to those skilled in the art that more or fewer planetary gears 32 can be provided within the scope of the claimed invention. Practical applications of the planetary gearbox 30 typically include at least three planetary gears 32.

[0239] exist Figure 2 and Figure 3 The planetary gearbox 30 shown by way of example is a planetary type, in which the planet carrier 34 is connected to the output shaft via a connecting rod 36, and the ring gear 38 is fixed. However, any other suitable type of planetary gearbox 30 can be used. As another example, the planetary gearbox 30 can be a stellar arrangement, in which the planet carrier 34 remains fixed, allowing the ring gear (or gear ring) 38 to rotate. In such an arrangement, the fan 23 is driven by the ring gear 38. As yet another alternative example, the gearbox 30 can be a differential gearbox, in which both the ring gear 38 and the planet carrier 34 are allowed to rotate.

[0240] It should be understood that Figure 2 and Figure 3 The arrangement shown is merely exemplary, and various alternatives are within the scope of this disclosure. By way of example only, any suitable arrangement can be used to position the gearbox 30 within the engine 10 and / or for connecting the gearbox 30 to the engine 10. As another example, the connection between the gearbox 30 and other components of the engine 10 (such as the input shaft 26, output shaft, and mounting structure 24) (such as...) Figure 2 The connecting rods 36 and 40 in the example can have any desired level of stiffness or flexibility. In another example manner, any suitable arrangement of bearings between the rotating and stationary parts of the engine (e.g., between the input and output shafts from the gearbox and a stationary structure such as the gearbox housing) can be used, and this disclosure is not limited to... Figure 2 The exemplary arrangement is as follows. For example, in the case where the gearbox 30 has a stellar arrangement (as described above), those skilled in the art will readily understand that the arrangement of the output link, support link, and bearing positions is generally different from that of the gearbox 30. Figure 2 The arrangement structure is shown as an example.

[0241] Therefore, this disclosure extends to gas turbine engines having any arrangement of gearbox type (e.g., star or planetary gear), support structure, input and output shaft arrangement, and bearing location.

[0242] Optionally, the gearbox may drive additional and / or alternative components (e.g., a medium-pressure compressor and / or a booster compressor).

[0243] Other gas turbine engines to which this disclosure is applicable may have alternative configurations. For example, such engines may have an alternative number of compressors and / or turbines and / or an alternative number of interconnecting shafts. In another example, Figure 1 The gas turbine engine shown has split nozzles 18 and 20, meaning that the flow through the bypass duct 22 has its own nozzle 18, which is separate from and radially outside the core engine nozzle 20. However, this is not limiting, and any aspect of this disclosure can also be applied to engines in which the flow through the bypass duct 22 and the flow through the core 11 are mixed or combined before (or upstream of) a single nozzle, which may be referred to as a mixing nozzle. One or both nozzles (whether mixing or splitting) may have a fixed or variable area.

[0244] The geometry of the gas turbine engine 10 and its components are defined by a conventional shaft system, including the axial direction (aligned with the axis of rotation 9) and the radial direction (in... Figure 1 The direction from bottom to top) and the circumferential direction (perpendicular to) Figure 1 (Page in the view). The axial, radial, and circumferential directions are orthogonal to each other.

[0245] Figure 4 schematically shown Figure 3 An exemplary planetary gear 32 of the type shown in the exemplary revolute gearbox 30. The planetary gear 32 is mounted around a pin 41, and the inner surface of the planetary gear 32 and the outer surface of the pin 41 form a sliding surface of a journal bearing 42, thereby allowing the planetary gear 32 to rotate relative to the pin 41. In use, the sliding surface is lubricated with oil to allow the planetary gear 32 and the pin 41 to rotate smoothly relative to each other. For clarity, Figure 4 The gap between the sliding surfaces is shown magnified in the image.

[0246] Figure 4 In the example, the sliding surface of the journal bearing 42 is shown as the inner surface of the planetary gear 32 and the outer surface of the pin 41. In an alternative example, a sleeve may be provided around the pin 41, the outer surface of which provides the inner sliding surface of the journal bearing 42, and the sleeve may be fixed to the outer surface of the pin 41, for example, by an interference fit. Alternatively or otherwise, a bushing (or bushing) may be provided, the inner surface of which provides the outer sliding surface of the journal bearing 42, and the bushing may be fixed to the inner surface of the planetary gear 32, for example, by an interference fit between the bushing and the gear 32. The advantage of the sliding surface of the journal bearing 42 being formed by the pin 41 and the planetary gear 32 themselves is that the tolerances of the journal bearing 42 can be more tightly controlled, while the advantage of using one or both of the sleeve and bushing is that the journal bearing can be repaired more easily by replacing one or both components when worn.

[0247] The planetary gear 32 is defined by an inner surface diameter 43 and an outer pitch circle diameter 44, the inner surface diameter of which may also be defined as the diameter of the journal bearing 42. The planetary gear 32 includes a plurality of teeth 45 extending around the outer circumference of the gear 32. The total number of teeth 45 may vary depending on the specific application. Figure 4 The number shown. Teeth 45 can be arranged in the form of a spur gear or a helical gear, that is, parallel to or at an angle to the axis of rotation of gear 32. The helical gear arrangement is more common because it allows the gear teeth 45 of planetary gear 32 to engage with the ring gear 38 and the sun gear 28 when the gears rotate relative to each other. Figure 3 The transition between corresponding teeth on the surface is smoother.

[0248] Figure 5 schematically shown Figure 4 The axial cross section of planetary gear 32 and pin 41. Figure 4 The cross section shown is along Figure 5The line A-A' shown is cut off. In this example, pin 41 is mounted to planet carrier 34 by extending through the thickness of planet carrier 34. Pin 41 can be fixed to planet carrier 34 by welding, bolting, or other means of securing pin 41 and planet carrier 34 to prevent relative movement between pin 41 and planet carrier 34 during use. In operation, force is transmitted between pin 41 and planet carrier 34 mainly by shear force on pin 41 transverse to axis 51 of pin 41, which also results in a bending moment applied to pin 41 along axis 51. In the planetary gearbox arrangement, where planet carrier 34 is fixed relative to engine frame, the net force on planet gear 32 acts in a direction tangential to the diameter of the center of planet gear 32. In the planetary gearbox arrangement, where the outer ring gear 38 ( Figure 3 The planetary gear 32 is fixed, and due to the additional centripetal force component required to keep the planetary gear 32 rotating around the sun gear 28, the net force on the planetary gear 32 is tilted toward the center of the sun gear, and this centripetal force is a function of the rotational speed of the planet carrier 34. The advantage of configuring the gearbox in a star arrangement structure is that the load on the pins is reduced when the gearbox operates at high speeds.

[0249] Planetary Gear 32 in Figure 5 The diagram shows a journal bearing portion 52 having a length L smaller than the total width 54 of the planetary gear 32. The length L of the journal bearing 42 can be selected and optimized according to the load experienced during gearbox operation, and the ratio between the journal bearing length L and the diameter D of the journal bearing 42 can be determined by the outer sliding surface (corresponding to...). Figure 5 The average diameter of the planetary gear 32 in the example shown (either the inner surface, the inner sliding surface (corresponding to the outer surface of the pin 41), or both) is defined. In implementation, this is referred to as the diameter clearance c (distance c / 2 in...). Figure 4 The difference between the two diameters (as shown in the figure) is small, typically in the range of less than 0.5% of either diameter. For an exemplary diameter range between 120 mm and 200 mm, the difference could be between about 0.1% and 0.3%, that is, between about 120 μm and about 600 μm, where a typical diameter gap is about 150 μm.

[0250] In some examples, the length 52 of the journal bearing 42 may be equal to or greater than the total width of the planetary gear 32.

[0251] In a specific example, the ratio L / D of the length L of the journal bearing 42 to its diameter D can range from about 0.5 to 1.4, optionally between about 1.1 and 1.3. A lower L / D ratio reduces misalignment of the gear 32 relative to the pin 41 in part by reducing the bending moment applied to the pin, thereby keeping the pin 41 more parallel to the gear 32. However, the L / D ratio should be kept above about 0.5, or optionally about 1.1, to avoid certain loads on the journal bearing becoming too high and adversely affecting the bearing's life.

[0252] Figure 6 An exemplary structure of a surface coating 61 that can be applied to any sliding surface of a journal bearing 42 is schematically shown. The material 62 below can be a pin 41 or a ring gear 32, or in an alternative example, a sleeve or bushing of the type described above. The total thickness of the surface coating 61 can be in the range of thicknesses between 40 micrometers and 200 micrometers, with a specific exemplary thickness in the range of approximately 100 micrometers.

[0253] Although the surface coating 61 can be applied to either surface of the journal bearing 42, applying the coating 61 to the outer surface of the pin 41 is practically preferred due to the practical limitations of deposition methods for the inner surface. Common deposition methods such as physical vapor deposition (PVD) are more suitable for applying the coating to the outer surface than the inner surface. Other techniques such as casting are more suitable for applying the coating to the inner surface, but casting is generally less suitable for producing coatings with the thickness range defined above and the tolerances required for the journal bearing.

[0254] An exemplary surface coating 61 may include three layers 61a-61c. A first layer 61a is deposited, having a coefficient of thermal expansion between that of the underlying material 62 and the second layer 61b. In the case of steel as the underlying material, the first layer 61a may be, for example, a copper-based alloy. The second layer 61b typically forms the thickest layer in the surface coating 61, having a thickness between approximately 50% and 95% of the total thickness of the surface coating 61, and may be composed of a copper- or aluminum-based alloy, i.e., a metallic alloy having copper or aluminum as a major component; an example is lead bronze, i.e., an alloy of copper, lead, and tin. This alloy is chosen to have a lower hardness compared to the material forming the other surface of the journal bearing, so that any particles not filtered out of the oil can instead embed in the second layer 61b, thereby reducing their ability to abrade the surface of the journal bearing.

[0255] The third layer 61c can be much thinner than the first layer 61a and the second layer 61b and is made of a material with a lower hardness than the second layer 61b (e.g., a lead-based alloy). The third layer is used to reduce friction between the surfaces of the journal bearing, especially when starting from a resting position where no oil layer has yet accumulated between the surfaces. The third layer 61c can, for example, have a thickness between 1 micrometer and 10 micrometers.

[0256] In a specific example, the thickness of the first layer 61a may be between 10 micrometers and 20 micrometers, the thickness of the second layer may be between about 40 micrometers and 100 micrometers, and the thickness of the third layer may be between about 1 micrometer and 15 micrometers, or between 1 micrometer and 10 micrometers, or between 5 micrometers and 15 micrometers, or between 10 micrometers and 15 micrometers.

[0257] Bearing materials have been developed to offer an optimal trade-off between "hard / strong" and "soft / flexible" properties for specific applications in gearboxes of gas turbine engines. The "hard" properties address the requirements for contact wear resistance, fatigue, and load-bearing capacity. The "soft" properties facilitate surface compatibility (compatibility with opposing surfaces), conformability, and embeddability. This has been found to help ensure continued operation under imperfect conditions. Furthermore, the arrangement of this invention has been developed to address environmental factors such as corrosion resistance and oxidation resistance.

[0258] In one specific example, coating 61b can be an aluminum-tin-copper alloy (e.g., SAE 783). In another specific example, a lead bronze alloy (such as SAE 49) can be used for coating 61b. Such alloys have been found suitable for harsh duty cycles, such as where the maximum specific operating load multiplied by the maximum operating slip speed is about 240 MPa m / s or greater. Using a thin overlay coating 61c (e.g., up to about 12 μm) such as SAE 194 lead-indium can further enhance the soft properties at the operating surface without compromising the load-bearing capacity of the underlying material.

[0259] Figure 7 A schematic cross-sectional view of gear 32 is shown, which is mounted around pin 41, thereby forming a journal bearing 42 between the outer surface of pin 41 and the inner surface of gear 32. The difference between the inner diameter of gear 32 and the outer diameter of pin 41 (i.e., the diametrical clearance) is magnified to show the variation in oil film thickness produced when the gearbox operates. The oil film thickness is lower on a portion 71 of the journal bearing, where the load F is transmitted between pin 41 and gear 32, for example, in the direction indicated by arrow 72. A specific load on journal bearing 42 is defined by the load F applied to the area defined by the length L and diameter D of bearing 42, i.e., the specific load is F / LD, typically expressed in MPa or N / mm². 2 measure.

[0260] The oil flow path through journal bearing 42 passes through the center hole 73 of pin 41, through inlet channel 74, and into the gap between pin 41 and gear 32. The oil flows around the journal bearing, is dragged across the minimum clearance by the relative rotation between pin 41 and gear 32, and exits via the edge of bearing 42. The oil is cooled and recirculated via a return oil pump (not shown). The oil flowing into the journal bearing can be pressurized to between approximately 50 kPa and 350 kPa (0.5 bar to 35 bar). A minimum oil pressure is required to supply enough oil to the bearing so that the area under force is covered by the oil supply. Higher pressures will tend to force a larger amount of oil through the bearing, but have a diminishing effect on lubrication and cooling of the bearing because a larger amount of oil will tend to travel through a wider portion of the gap between pin 41 and gear 32 rather than through the minimum clearance portion 71. Higher oil pressures will tend to reduce the temperature difference between the inlet and outlet oil flows, making it more difficult to extract heat and thus requiring a larger radiator. Therefore, optimal oil flow pressure and temperature difference are often required to minimize the weight associated with the oil pump and radiator. Therefore, the pressure and temperature difference defined in this paper were chosen to provide the necessary lubrication, but with a sufficiently high temperature difference to allow a sufficiently lightweight heat exchanger to remove heat. Given the importance of weight to the total fuel consumption of the aircraft powering the engine, a low-weight heat exchanger may be a particularly important consideration for gearboxes intended for use in gas turbines within aircraft.

[0261] The dimensional and positional accuracy of pins 41 and gears 32 in the gearbox will affect how the oil film thickness varies, as well as the viscosity and temperature of the oil. To maintain a uniform oil temperature across each journal bearing, a symmetrical oil inlet path can be provided in the gearbox, and the aeration chamber for mixing the oil before it enters the gearbox can be large enough to allow oil of uniform temperature to be fed into the gearbox at different inlet points. Therefore, for example, when the engine is operating under cruising conditions, the temperature variation between the oils fed into each journal bearing can be no more than 1 degree Celsius. The oil pressure variation between the journal bearings is also preferably uniform, but this will generally have a smaller effect than the temperature variation, as an increase in pressure above the minimum required will tend to simply result in more oil flowing through the bearing, thus having minimal impact on operation.

[0262] Operating oil film thickness, i.e., the thickness of the oil film in each journal bearing during engine operation, can be defined as a proportion of the journal bearing diameter. During operation, for example under the gearbox load at its highest MTO condition, the minimum operating oil film thickness of each journal bearing can be less than about 8 micrometers, with journal bearing diameters between about 120 mm and 200 mm, and optionally greater than about 3.5 micrometers. The journal bearing clearance can typically be between about 1‰ and 3‰ (0.1% and 0.3%) of the journal bearing diameter, for example, about 1.5‰ (0.15%). As mentioned above, the journal bearing diameter can be defined as the diameter of the internal sliding surface of the planetary gear. Again, for example under MTO conditions, the variation between the minimum operating thicknesses of each journal bearing can be less than about 8% of the average minimum oil film thickness. For example, if the average minimum oil film thickness is about 6 micrometers, the maximum difference between the minimum oil film thicknesses across all journal bearings will be about + / - about 0.5 micrometers.

[0263] like Figure 7 The diagram schematically illustrates that the operating oil film thickness will vary between the minimum thickness at the maximum load point and the maximum thickness at a point diametrically opposite the maximum load point for each journal bearing. The maximum load point will tend to follow a linear path along the length of the journal bearing parallel to its axis of rotation. The ratio between the maximum and minimum oil film thickness will be highest under maximum takeoff conditions and approximately 1 under idle conditions, i.e., no significant load is transmitted through the gearbox.

[0264] Figure 8 Exemplary graphs are shown illustrating the variation of a specific operating load (y-axis, MPa) with sliding velocity (x-axis, m / s) on the journal bearings of a series of exemplary gearboxes of the type disclosed herein, operating under maximum takeoff conditions. The specific load on the journal bearing is as defined above. The sliding velocity of the journal bearing is defined as the relative tangential velocity between the inner and outer surfaces of the journal bearing. Dashed lines 81a, 81b, and 81c represent constant values ​​of the product of the specific operating load, which may be referred to as PV (the product of pressure and velocity), and the sliding velocity at 200 MPa m / s, 400 MPa m / s, and 600 MPa m / s, respectively.

[0265] Generally speaking, under higher specific loads, sliding speeds, or higher PV values, surface coatings comprising an alloy layer with aluminum or copper as the main component can be used, for example, forming a surface coating such as... Figure 6 The second layer 61b is shown. Copper as the main component is preferably used for journal bearings with higher diameters, for example, those with a diameter greater than 120 mm.

[0266] The maximum specific operating load of each journal bearing in the gearbox can be greater than 5 MPa, or greater than any one of 6 MPa, 7 MPa, 8 MPa, 9 MPa, 10 MPa, 11 MPa, 12 MPa, 13 MPa, 14 MPa, 15 MPa, 16 MPa, or 17 MPa. The maximum sliding speed of the journal bearing can be determined by... Figure 8 The curves 81a-81c shown correspond to the specified sliding speeds. In specific examples, the maximum specific operating load can be approximately 13 MPa or approximately 18 MPa, and in each case, the sliding speed is approximately 42 m / s or 38 m / s, respectively, as shown below. Figure 8 Data points 82 and 83 are shown above. In general, under maximum takeoff conditions, a specific load can range from approximately 10 MPa to 20 MPa, and the skid speed can range from approximately 35 m / s to 45 m / s. These ranges are particularly suitable for planetary gearbox arrangements.

[0267] Points 82 and 83 represent specific pressure and slip velocity values ​​for journal bearings in two exemplary planetary gearboxes under maximum takeoff conditions, where the journal bearing diameters are approximately 155 mm and 140 mm, respectively, the journal bearing L / D ratios are approximately 1.11 and 1.24, respectively, and the diameter clearance is approximately 1.5‰ for both. The PV values ​​for points 82 and 83 under maximum takeoff conditions are approximately 560 MPa m / s and 650 MPa m / s, respectively.

[0268] Figure 8 Points 84, 85, and 86 represent specific pressure and slip velocity values ​​for journal bearings in three exemplary star gearboxes of different sizes under maximum takeoff conditions, with journal bearing diameters of approximately 100 mm, 120 mm, and 180 mm, and journal bearing L / D ratios of approximately 1.45, 1.35, and 1.13, respectively, and a diameter clearance of approximately 1.5‰ for each. The PV values ​​for points 84, 85, and 86 under maximum takeoff conditions are approximately 325 MPa m / s, 335 MPa m / s, and 370 MPa m / s, respectively. In general, specific loads for such examples under maximum takeoff conditions can range from approximately 5 MPa to 10 MPa, and slip velocities can range from approximately 45 m / s to 55 m / s. The PV value can be within a range having a lower limit of any one of 200 MPa m / s, 220 MPa m / s, 240 MPa m / s, 260 MPa m / s, 230 MPa m / s or 300 MPa m / s and an upper limit of any one of 310 MPa m / s, 330 MPa m / s, 350 MPa m / s, 370 MPa m / s, 390 MPa m / s, 410 MPa m / s or 430 MPa m / s.

[0269] Therefore, in another general aspect, under maximum takeoff conditions, the specific load of the above example can be in the general range of about 5 MPa to 20 MPa and the skid speed is in the range of about 30 m / s or 35 m / s to 50 m / s or 55 m / s.

[0270] The higher specific loads on the journal bearings (points 82, 83) of the planetary gearbox partially reflect the additional radial load on each journal bearing due to the rotation of each planetary gear around the central sun gear, whereas the planetary gears (points 84, 85, 86) in the sun gearbox do not rotate around the central sun gear.

[0271] The y-axis extension of a specific load on each of data points 82-86 represents the variation of a specific load within a range of + / -10% torque load variation around the nominal torque load under maximum takeoff conditions.

[0272] The upper limit of PV can be about 720 MPa m / s, while the lower limit can be about 240 MPa m / s or 300 MPa m / s. Alternatively, the upper limit can be defined by an upper limit of one or both of the sliding speed and the specific operating load, for example, an upper limit of about 45 m / s, 50 m / s, 55 m / s or 60 m / s for the sliding speed and an upper limit of about 10 MPa, 20 MPa or 30 MPa for the specific operating load. The lower limit can be defined by a sliding speed of about 30 m / s, 35 m / s, 40 m / s or 45 m / s, or by a specific load of about 5 MPa or 10 MPa, and other aspects specified herein.

[0273] During the operation of a gas turbine engine, such as under MTO conditions, the eccentricity ratio of the journal bearing is defined as 1-2H. min / c, where H min Minimum oil film thickness (e.g.) Figure 7 (as shown) and c is the diameter gap (as shown) Figure 4 As shown, gear 32 and pin 41 are arranged concentrically. Figure 9 The diagram illustrates the variation of the eccentricity ratio (y-axis) of a series of exemplary gearboxes as a percentage of the journal bearing design load (x-axis). The first exemplary star gearbox 91, the second exemplary star gearbox 92, and the third exemplary star gearbox 93 (corresponding to...) Figure 8 The same gearbox design with data points 84, 85, and 86 exhibits a variation in eccentricity ratio between approximately 0.2 and 0.3 between 90% and 110% of the design load, and an eccentricity ratio between approximately 0.79 and 0.91 within this design load range. The first exemplary planetary gearbox and the second exemplary planetary gearbox (corresponding to...) have higher absolute design loads. Figure 8The eccentricity ratios 94 and 95 of the gearbox designs with data points 82 and 83 are between approximately 0.94 and 0.97 within a similar design load range, where the eccentricity ratio varies between approximately 0.03 and approximately 0.05 within this range.

[0274] The diameter clearance c can be in the range of approximately 1‰ to 2‰, i.e., between approximately 0.1% and 0.2%. A smaller diameter clearance will tend to increase the area over which pressure is distributed between the inner and outer surfaces of the journal bearing, but this will combine with a narrower path, forcing oil through the bearing to flow through the narrower path as the bearing rotates, thus limiting the flow rate of oil through the bearing and ultimately causing the bearing to seize as the diameter clearance decreases further. A larger diameter clearance will tend to decrease the area over which pressure is distributed, but will also make it easier for oil to travel through the bearing. Therefore, an optimal balance between the factors is required, especially for higher eccentricity ratios between approximately 0.94 and 0.97, which can be between approximately 1‰ and 2‰, and optionally between approximately 1.4‰ and 1.6‰.

[0275] Figure 10 This is a schematic graph illustrating the variation of inefficiency as a function of the eccentricity ratio for the aforementioned series of gearboxes. The exemplary solar gearbox tends to have higher inefficiencies, ranging between approximately 0.17% and 0.3%, while the exemplary planetary gearbox has inefficiencies between approximately 0.1% and 0.16%. The eccentricity ratio falls within the range of values ​​described above. The variation of inefficiency with eccentricity follows... Figure 10 The general trend shown in Figure 1001 is that a higher eccentricity ratio leads to higher efficiency, i.e., lower inefficiency. The range of the eccentricity ratio can be as described above, while the range of inefficiency can be less than about 0.225%, and can be between about 0.225% and about 0.1%. Further increasing the eccentricity ratio will tend to increase the risk of journal bearing seizure because the minimum thickness of the oil film becomes too small to maintain the oil film at the required load range to separate the pins and gears.

[0276] Figure 11 The H of the above-described exemplary stellar gearbox and planetary gearbox operating under maximum takeoff conditions is shown. min A series of trend lines (unit: μm) showing the variation of oil inlet temperature (unit: °C). The star gearbox (lines 1101, 1102, 1103) often exhibits high Hg within the temperature range. min The value and trend line have a steep gradient, while planetary gearboxes (lines 1104, 1105) tend to have a lower H value. min The values ​​are relatively shallow and have shallow gradients. Each trend line tends to follow H. min=B-AT, a function of the form T, where T is the temperature (in °C) and A and B are constants characteristic of a specific gearbox design. The minimum oil film thickness H for each gearbox under maximum takeoff conditions, excluding one of the star gearboxes. min Within the region having the upper limit defined by line 1106, where A is 0.139 μm / ℃ and B is 27.8 μm. H min The minimum oil film thickness can be approximately 2.3 μm; below this minimum, the oil film may be insufficient to prevent journal bearing seizing. Two additional lines, 1107 and 1108, define additional upper and lower limits, respectively, with line 1107 defined by A = 0.117 μm / ℃ and B = 22 μm, and line 1108 defined by A = 0.034 μm / ℃ and B = 6.4 μm. The overall range of inlet temperature can be between 60℃ and 120℃, with a selectable range greater than approximately 100℃ and less than approximately 120℃. Oil viscosity increases at lower temperatures, thus reducing lubrication efficiency, while the resulting lower oil viscosity at higher temperatures can lead to an excessively small minimum oil film thickness.

[0277] Figure 12 A series of trend lines are shown illustrating the eccentricity ratio E of the journal bearings of various exemplary gearboxes operating under maximum takeoff conditions in each case, as a function of inlet oil temperature. Solar gearboxes (lines 1201, 1202, 1203) tend to have lower E values ​​and steeper gradients across the entire temperature range, while planetary gearboxes (lines 1204, 1205) tend to have higher E values ​​and shallower gradients. In each case, the trend lines tend to follow a function of the form E = AT + B, where T is the inlet oil temperature and A and B are constants. The maximum value of E is approximately 0.98; above this maximum, the oil film may be too small to sustain lubrication of the journal bearings. The eccentricity ratio E may be higher than trend line 1206 defined by A = 0.0015 / ℃ and B = 0.69, or alternatively higher than trend line 1207 defined by A = 0.00083 / ℃ and B = 0.84, and lower than trend line 1208 defined by A = 0.00033 / ℃ and B = 0.94. For Figure 11 In the example, the overall range of the inlet temperature can be between 60°C and 120°C, with the selectable range being greater than about 100°C and less than about 120°C.

[0278] The Sommerfeld number S of a journal bearing is defined as:

[0279]

[0280] Where d is the outer diameter of pin 41 ( Figure 7c is the diameter clearance, μ is the absolute viscosity of the lubricant, N is the relative rotational speed of the journal bearing (in revolutions per second), P is the load applied to the entire projected bearing area, i.e., F / LD, where L is the length of the journal bearing and D is the diameter.

[0281] Higher PV values ​​that result in lower efficiency values ​​will tend to increase the Sommerfeld number of journal bearings. Figure 13 The general relationship between inefficiency and Sommerfeld number for the gearboxes in the aforementioned range, given by trend line 1301, is illustrated. Exemplary planetary gearboxes often have journal bearings with low Sommerfeld numbers, typically in the range of about 1 to 9 under maximum takeoff conditions. Two planetary gearboxes with higher PV values, higher eccentricity, and lower inefficiency have journal bearings that tend to have higher Sommerfeld numbers under varying oil temperatures and loads, typically in the range of about 4 to 21. Under more optimized oil temperature conditions, the Sommerfeld numbers of these planetary gearbox journal bearings tend to be between about 10 and 16. In general, the Sommerfeld number per journal bearing can be greater than about 4, with an inefficiency of about 0.225% or less under maximum takeoff conditions. As mentioned above, the minimum inefficiency can be about 0.1%. The maximum Sommerfeld number can be about 21.

[0282] Figure 14 This diagram illustrates the relationship between a specific oil flow rate and PV (in MPa m / s), where the specific oil flow rate is the ratio of the journal bearing area (LD, in m²). 2 The oil flow rate (in liters per minute) is defined as follows. Region 1401, defined between upper and lower limits 1402 and 1403, encompasses the journal bearings of each of the five gearboxes mentioned above. Three of the planetary gearboxes are located at the left-hand end of region 1401, typically below approximately 450 MPa m / s and above 240 MPa m / s, while the two planetary gearboxes are located at the right-hand end, typically in the range between approximately 450 MPa m / s and 720 MPa m / s. The general relationship shows that a higher PV value is associated with a lower specific oil flow rate, thus requiring a lower pressure oil flow rate for a given rated power, and resulting in a reduction in the overall weight of associated equipment such as oil pumps and radiators. Therefore, a higher PV value is particularly advantageous for gas turbine engines used in aircraft.

[0283] It should be understood that the present invention is not limited to the embodiments described above, and various modifications and improvements can be made without departing from the concepts described herein. Unless mutually exclusive, any feature may be used alone or in combination with any other feature, and this disclosure extends to and includes all combinations and sub-combinations of one or more features described herein.

Claims

1. A method of operating a gearbox (30) for an aircraft gas turbine engine, said gearbox (30) comprising: Sun gear (28); Multiple planetary gears (32) surround and engage with the sun gear (28); and A ring gear (38) surrounds and engages with the plurality of planetary gears (32), the ring gear (38) having a pitch circle diameter of 550 mm or greater. Each of the plurality of planetary gears (32) is rotatably mounted about a pin (41) of a planetary gear carrier (51), the planetary gear carrier having a journal bearing (42) having an inner sliding surface on the planetary gear (32) and an outer sliding surface on the pin (41), the oil film between the inner sliding surface on the planetary gear (32) and the outer sliding surface on the pin (41) varying between a maximum thickness and a minimum thickness around the journal bearing (42). The method includes operating the aircraft gas turbine engine under maximum takeoff conditions such that the minimum thickness of the oil film differs from the average minimum oil film thickness of the plurality of planetary gears (32) by no more than 8%.

2. The method according to claim 1, wherein: The diameter D of each journal bearing (42) is between 120 mm and 200 mm; and / or The length L of the inner sliding surface and the outer sliding surface of each journal bearing (42) is between 0.5 and 1.4 of the diameter D; and / or The L / D ratio is between 1.1 and 1.

3.

3. The method according to claim 1 or claim 2, wherein the average minimum oil film thickness under maximum takeoff conditions is between 3.5 micrometers and 8 micrometers.

4. The method according to claim 1 or claim 2, wherein during operation of the gas turbine engine under maximum takeoff conditions, the eccentricity ratio of each journal bearing is in the range of 0.94 to 0.

97.

5. The method according to claim 1 or claim 2, wherein for each journal bearing in the gearbox, a specific load multiplied by a sliding speed is greater than 240 MPa m / s.

6. The method of claim 5, wherein for each journal bearing in the gearbox, the specific load multiplied by the sliding speed is less than 720 MPa m / s.

7. The method of claim 1 or claim 2, wherein the gas turbine engine comprises a turbine, a compressor, a spindle connecting the turbine to the compressor, and a fan located upstream of the engine core, the fan comprising a plurality of fan blades having a diameter between 5.5 × 10⁻⁶. 7 kg m 2 and 9×10 8 kg m 2 The moment of inertia between them.

8. The method of claim 7, wherein the fan has a diameter between 7.4 × 10⁻⁶. 7 kg m 2 and 7×10 8 kg m 2 The moment of inertia between them.

9. The method of claim 8, wherein the fan has a diameter between 8.3 × 10⁻⁶. 7 kg m 2 and 6.5×10 8 kg m 2 The moment of inertia between them.

10. A gearbox (30) for an aircraft gas turbine engine, the gearbox (30) comprising: Sun gear (28); Multiple planetary gears (32) surround and engage with the sun gear (28); and A ring gear (38) surrounds and engages with the plurality of planetary gears (32), the ring gear (38) having a pitch circle diameter of 550 mm or greater. Each of the plurality of planetary gears (32) is rotatably mounted about a pin (41) of a planetary gear carrier (51), the planetary gear carrier having a journal bearing (42) having an inner sliding surface on the planetary gear (32) and an outer sliding surface on the pin (41), the oil film between the inner sliding surface on the planetary gear (32) and the outer sliding surface on the pin (41) varying between a maximum thickness and a minimum thickness around the journal bearing (42). Furthermore, during the operation of the aircraft gas turbine engine under maximum takeoff conditions, the difference between the minimum thickness of the oil film and the average minimum oil film thickness between the plurality of planetary gears (32) does not exceed 8%.

11. The gearbox (30) according to claim 10, wherein: The diameter D of each journal bearing (42) is between 120 mm and 200 mm; and / or The length L of the inner sliding surface and the outer sliding surface of each journal bearing (42) is between 0.5 and 1.4 of the diameter D; and / or The L / D ratio is between 1.1 and 1.

3.

12. The gearbox (30) according to claim 10 or claim 11, wherein the average minimum oil film thickness is between 3.5 micrometers and 8 micrometers during operation of the aircraft gas turbine engine under maximum takeoff conditions.

13. The gearbox (30) according to claim 10 or claim 11, wherein during operation of the aircraft gas turbine engine under maximum takeoff conditions, the eccentricity ratio of each journal bearing is in the range of 0.94 to 0.

97.

14. The gearbox (30) according to claim 10 or claim 11, wherein during operation of the aircraft gas turbine engine under maximum takeoff conditions, for each journal bearing in the gearbox, a specific load multiplied by a sliding speed is greater than 240 MPa m / s.

15. The gearbox (30) of claim 14, wherein during operation of the aircraft gas turbine engine under maximum takeoff conditions, for each journal bearing in the gearbox, a specific load multiplied by a sliding speed is less than 720 MPa m / s.

16. The gearbox (30) according to claim 10 or claim 11, wherein: The gearbox has a gear ratio of 3.2 to 4.5 or 3.2 to 4.0; and / or The gearbox is in a star configuration.

17. A gas turbine engine (10) for an aircraft, the gas turbine engine comprising: An engine core (11) includes a turbine (19), a compressor (14) and a spindle (26) connecting the turbine to the compressor. A fan (23) is located upstream of the engine core and includes multiple fan blades; and According to claim 10 or claim 11, the gearbox (30) is configured to receive input from the spindle (26) and provide output drive for the fan to drive the fan at a lower rotational speed than the spindle.

18. The gas turbine engine according to claim 17, wherein: The turbine is a first turbine (19), the compressor is a first compressor (14), and the spindle is a first spindle (26). The engine core also includes a second turbine (17), a second compressor (15), and a second spindle (27) connecting the second turbine to the second compressor; and The second turbine, the second compressor, and the second spindle are arranged to rotate at a higher rotational speed than the first spindle.

19. The gas turbine engine (10) according to claim 17, wherein the gas turbine engine comprises: Specific thrust of 70 N / kg to 90 N / kg, wherein specific thrust is defined as the net thrust of the gas turbine engine divided by the total mass flow rate through the gas turbine engine; A bypass ratio of 12.5 to 18 under cruise conditions; and / or The fan has a size between 5.5 × 10 7 kg m 2 and 9×10 8 kg m 2 The moment of inertia between them.

20. The gas turbine engine (10) according to claim 19, wherein the bypass ratio under cruise conditions is in the range of 13 to 16.

21. The gas turbine engine (10) according to claim 19, wherein the fan has a diameter between 7.4 × 10⁻⁶. 7 kg m 2 and 7×10 8 kg m 2 The moment of inertia between them.

22. The gas turbine engine (10) according to claim 19, wherein the fan has a diameter between 8.3 × 10⁻⁶. 7 kg m 2 and 6.5×10 8 kg m 2 The moment of inertia between them.

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