Patent Application
20140290211
The present disclosure relates generally to turbofan engines, and more particularly to a balanced stage count in a turbofan engine having a high bypass ratio.
Turbine engines, such as those used in commercial aircraft, typically include a large fan on a fore end of the turbine engine gas path. Air drawn through the fan is either directed into the gas path of the turbine engine or provided to a bypass path that bypasses the turbine engine gas path. The ratio of air bypassing the turbine engine gas path to air entering the turbine engine gas path is referred to as the engine bypass ratio, or alternatively as the bypass ratio. As the turbine engine fan increases in size, the bypass ratio typically undergoes a corresponding increase.
In existing turbine engines, an increase in bypass ratio typically requires that the turbine portion of the turbine engine have a corresponding increase in stage count. That is, the higher the bypass ratio in existing turbine engines, the higher the number of low pressure turbine stages that are required for operation of the turbine engine. The increased number of low pressure turbine stages increases the ratio of low pressure turbine stages to low pressure compressor stages, and increases the weight of the engine.
A turbine engine according to an exemplary embodiment of this disclosure, among other possible things includes a fan, a compressor section having at least a first portion and a second portion, the first portion is at a high pressure relative to the second portion, a combustor in fluid communication with the compressor section, a turbine section in fluid communication with the combustor, the turbine section includes at least a first portion and a second portion and the first portion is at a high pressure relative to the second portion, each of the compressor section second portion and the turbine section second portion include a plurality of stages, a ratio of turbine section second portion stages to compressor section second portion stages is less than or equal to 1, and a fan bypass ratio of the turbine engine is greater than or equal to 11.
In a further embodiment of the foregoing turbine engine, a configuration complexity metric of the low pressure compressor and low pressure turbine=[1+N][1+[1/N×(SLPT)+N×(SLPC)]]/[N+(SLPC)/(SLPT)]/[2N] where, SLPT is the number of turbine second portion stages, SLPC is the number of compressor second portion stages, SLPC/SLPT is a reciprocal of the ratio of the number of turbine second portion stages to the number of compressor second portion stages and N is approximately 1.618034.
In a further embodiment of the foregoing turbine engine, the configuration complexity metric of the low pressure compressor and low pressure turbine is in the range of 2.63 to 4.27.
In a further embodiment of the foregoing turbine engine, the ratio of turbine section second portion stages to compressor section second portion stages is approximately 0.8.
In a further embodiment of the foregoing turbine engine, the turbine section second portion includes four stages and the compressor section second portion includes five stages.
In a further embodiment of the foregoing turbine engine, the turbine section second portion includes a number of stages in the range of 3 to 5 and the compressor section second portion includes a number of stages in the range of 5 to 7.
In a further embodiment of the foregoing turbine engine, the turbine engine has a fan bypass ratio in the range of 11 to 17.
In a further embodiment of the foregoing turbine engine, the turbine engine has a fan bypass ratio in the range of 11.6 to 15.
In a further embodiment of the foregoing turbine engine, said turbine engine has a fan bypass ratio of approximately 11.7.
A turbine engine according to an exemplary embodiment of this disclosure, among other possible things includes a fan, a compressor section having at least a first portion and a second portion, the first portion is at a high pressure relative to the second portion, a turbine section in fluid communication with the compressor, the turbine section includes at least a first portion and a second portion and the first portion is at a high pressure relative to the second portion, each of the compressor section second portion and the turbine section second portion include a plurality of stages, a core flow path defined at least by the compressor section and the turbine section, a bypass flow path bypassing the core flow path, a fan bypass ratio is defined as a ratio of air passing through the fan and entering the bypass flow path to air passing through the fan and entering the core flow path, a ratio of turbine section second portion stages to compressor section second portion stages is less than or equal to 1, and a fan bypass ratio of the turbine engine is greater than or equal to 11.
In a further embodiment of the foregoing turbine engine, a configuration complexity metric of the low pressure compressor and low pressure turbine=[1+N][1+[1/N×(SLPT)+N×(SLPC)]]/[N+(SLPC)/(SLPT)]/[2N] where, SLPT is the number of turbine second portion stages, SLPC is the number of compressor second portion stages, SLPC/SLPT is a reciprocal of the ratio of the number of turbine second portion stages to the number of compressor second portion stages, and N is approximately 1.618034.
In a further embodiment of the foregoing turbine engine, a configuration complexity metric of the low pressure compressor and low pressure turbine is in the range of 2.63 to 4.27.
In a further embodiment of the foregoing turbine engine, the ratio of turbine section second portion stages to compressor section second portion stages is approximately 0.8.
In a further embodiment of the foregoing turbine engine, the turbine section second portion includes four stages and the compressor section second portion includes five stages.
In a further embodiment of the foregoing turbine engine, the turbine section second portion includes a number of stages in the range of 3 to 5 and the compressor section second portion includes a number of stages in the range of 5 to 7.
In a further embodiment of the foregoing turbine engine, the turbine engine has a fan bypass ratio in the range of 11 to 17.
In a further embodiment of the foregoing turbine engine, the turbine engine has a fan bypass ratio in the range of 11.6 to 15.
In a further embodiment of the foregoing turbine engine, the turbine engine has a fan bypass ratio of approximately 11.7.
These and other features of the present invention can be best understood from the following specification and drawings, the following of which is a brief description.
The engine 20 generally includes a low speed spool 30 and a high speed spool 32 mounted for rotation about an engine central longitudinal axis A relative to an engine static structure 36 via several bearing systems 38. It should be understood that various bearing systems 38 at various locations may alternatively or additionally be provided.
The low speed spool 30 generally includes an inner shaft 40 that interconnects a fan 42, a low pressure compressor 44 and a low pressure turbine 46. The inner shaft 40 is connected to the fan 42 through a geared architecture 48 to drive the fan 42 at a lower speed than the low speed spool 30. The high speed spool 32 includes an outer shaft 50 that interconnects a high pressure compressor 52 and high pressure turbine 54. A combustor 56 is arranged between the high pressure compressor 52 and the high pressure turbine 54. A mid-turbine frame 57 of the engine static structure 36 is arranged generally between the high pressure turbine 54 and the low pressure turbine 46. The mid-turbine frame 57 further supports bearing systems 38 in the turbine section 28. The inner shaft 40 and the outer shaft 50 are concentric and rotate via bearing systems 38 about the engine central longitudinal axis A which is collinear with their longitudinal axes.
The core airflow C is compressed by the low pressure compressor 44 then the high pressure compressor 52, mixed and burned with fuel in the combustor 56, then expanded over the high pressure turbine 54 and low pressure turbine 46. The mid-turbine frame 57 includes airfoils 59 which are in the core airflow path C. The turbines 46, 54 rotationally drive the respective low speed spool 30 and high speed spool 32 in response to the expansion.
The engine 20 is in one example a high-bypass geared aircraft engine. In a further example, the engine 20 bypass ratio is greater than about eleven (11), with an example embodiment having a bypass ratio in the range of eleven (11) to seventeen (17), and a further example embodiment having a bypass ratio in the range of eleven and six tenths (11.6) to fifteen (15), and a further example embodiment being approximately eleven and seven tenths (11.7). The geared architecture 48 is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3 and the low pressure turbine 46 has a pressure ratio that is greater than about 5. In one disclosed embodiment, the engine 20 bypass ratio is greater than about eleven (11:1), the fan diameter is significantly larger than that of the low pressure compressor 44, and the low pressure turbine 46 has a pressure ratio that is greater than about 5:1. Low pressure turbine 46 pressure ratio is pressure measured prior to inlet of low pressure turbine 46 as related to the pressure at the outlet of the low pressure turbine 46 prior to an exhaust nozzle. The geared architecture 48 may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.5:1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other gas turbine engines including direct drive turbofans.
A significant amount of thrust is provided by the bypass flow due to the high bypass ratio. The fan section 22 of the engine 20 is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft, with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption (‘TSFC’)”—is the industry standard parameter of lbm of fuel being burned divided by lbf of thrust the engine produces at that minimum point. “Low fan pressure ratio” is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.45. “Low corrected fan tip speed” is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram ° R)/(518.7° R)]0.5. The “Low corrected fan tip speed” as disclosed herein according to one non-limiting embodiment is less than about 1150 ft/second.
Existing turbine engine models, such as direct drive turbine engines, increase the bypass ratio of the turbine engine by increasing the fan size, thereby increasing the amount of air that is drawn through the gas path of the turbine engine. The large fan size necessitates an increased number of low pressure turbine stages in order to drive the fan at sufficient speeds. The additional turbine stages result in a heavier turbine engine where the number of low pressure turbine stages exceeds the number of low pressure compressor stages.
The gas path 102 extends through the turbine engine 20 and into the low pressure turbine 46, as shown in
The number of low pressure compressor stages 130 is identical to the number or low pressure compressor rotors 110. The number of low pressure turbine stages 132 is identical to the number of low pressure turbine rotors 112. The turbine engine 20 incorporating the low pressure compressor 44 and the low pressure turbine 46 illustrated in
Thus, the example turbine engine 20 including the low pressure compressor portion 44 and the low pressure turbine portion 46 of
In yet further alternate turbine engine configurations, the ratio of low pressure turbine stages 132, 232 to low pressure compressor stages 130, 230 can be anywhere in the range of 0.3 to about 0.9. In other words, alternate configurations can include ratios ranging from 1:3 or 2:6 or 3:9 to 4:5 or 6:7 or 7:8.
A set of examples of the number of low pressure compressor 44 stages and low pressure turbine 46 stages of the example gas turbine engine 20 is defined below in Table 1. Table 1 includes the combinations of the number of low pressure compressor 44 stages and low pressure turbine 46 stages, the ratio of low pressure turbine 46 stages to low pressure compressor 44 stages, the reciprocal of the ratio of low pressure turbine 46 stages to low pressure compressor 44 stages, the difference between the number of low pressure compressor 44 stages and low pressure turbine 46 stages, the sum of the number of low pressure compressor 44 stages and low pressure turbine 46 stages and a measure of the configuration complexity of the low pressure compressor 44 and low pressure turbine 46 in terms of a configuration complexity metric. The configuration complexity metric is defined as
where, SLPT is the number of low pressure turbine 46 stages, SLPC is the number of low pressure compressor 44 stages, SLPC/SLPT is the reciprocal of the ratio of the number of low pressure turbine 46 stages to the number of low pressure compressor 44 stages, and N=1.618034, approximately. N also is known in mathematics as the “golden number” due to the relationship, N×[N−1]=1.
The configuration complexity metric includes a weighted summation of the number of low pressure 44 compressor stages and the number of low pressure turbine 46 stages where the weighting factors are the golden number and the reciprocal of the golden number.
A balanced stage count has one more low pressure compressor stage than low pressure turbine stage, expressed mathematically as (SLPT)−(SLPT)=1. The above equation is of the form [1/(SLPT)]×[(SLPC)−1]=1. The configuration complexity metric balances the complexity of all low pressure compressor stages against the complexity of all low pressure turbine stages by applying weighting factors based on the golden number, N. The weighted sum of the stage counts of the low pressure compressor and low pressure turbine is defined as the sum of the stage count of the low pressure turbine, (SLPT), multiplied by the reciprocal of the golden number, 1/N, plus the stage count of the low pressure compressor, (SLPC), multiplied by the golden number, N. The simplest configuration of low pressure compressor and low pressure turbine has (SLPC)=1 and (SLPT)=1 and, therefore, (SLPT)/(SLPC)=1 (one). In one example, the configurations of interest have (SLPT)/(SLPC) less than or equal to 1 (one). The lowest value of the configuration complexity metric is set equal to 1 (one) by applying the factor {[1+N]/[2×N]}; see Table 1.
In alternate example configurations the low pressure compressor 44 and the low pressure turbine 46 have different values for the configuration complexity metric ranging from N×N=2.634, approximately, to N×N×N=4.262, approximately. While the ratio of the number of low pressure turbine 46 stages to the number of low pressure compressor 44 stages for various configurations (such as 1:2 and 2:4) may be the same but the configuration complexity metric is different, as the configuration complexity metric depends on the actual stage counts. No two distinct configurations comprising one low pressure compressor 44 and one low pressure turbine 46 and a second low pressure compressor 44 and a second low pressure turbine 46 have both the same ratio of the number of low pressure turbine 46 stages to the number of low pressure compressor 44 stages and the same configuration complexity metric.
It is further understood that any of the above described concepts can be used alone or in combination with any or all of the other above described concepts. Although various embodiments of this invention have been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of this invention. For that reason, the following claims should be studied to determine the true scope and content of this invention.
