Impingement/Effusion Cooled Combustor Liner

This invention relates generally to a gas turbine engine and more particularly to an improved low emission combustor for use with the gas turbine engine.

1. Background Art

High performance gas turbine engines require increased firing temperatures and increased compressor pressures. Coolant from the compressor section is directed through cooling passages in various components to enhance reliability and cycle life of individual components within the engine. For example, to improve fuel economy characteristics, engines are being operated at higher temperatures than the material physical property limits of which the engine components are constructed. These higher temperatures, if not compensated for, oxidize engine components, distort engine components and decrease component life. Cooling passages are used to direct a flow of air to such engine components to reduce the high temperature of the components and prolong component life by limiting the temperature to a level which is consistent with material properties of such components.

However, as the amount of coolant air is increased to cool the engine components the amount of air available for the combustion chamber is decreased. Thus, systems and methods of increasing cooling efficiency and reducing the amount of coolant used to cool the engine components must be utilized.

The present invention is directed to overcome one or more of the problems as set forth above.

2. Disclosure of the Invention

In one aspect of the present invention, a combustor is comprised of an interior liner defining an inlet end portion and an outlet end portion being spaced apart by an axial portion. The interior liner defines a combustion side and a cooling side having a plurality of effusion holes defined therein extending between the combustion side and the cooling side. The plurality of effusion holes are formed in a preestablished pattern defining a centroid. The combustor further includes an exterior liner defining an inlet end portion and an outlet end portion being spaced apart by an axial portion. The exterior liner defines a first surface and a second surface having a plurality of impingement holes defined therein extending between the first surface and the second surface at an angle of about 90 degrees. The plurality of impingement holes are formed in a preestablished pattern and at least a portion of the plurality of impingement holes in the exterior liner are positioned in radial alignment with the centroid of the preestablished pattern of the plurality of effusion holes in the interior liner.

FIG. 1 is a partially sectioned partial view of a gas turbine engine embodying the present invention;

FIG. 2 is an enlarged sectional side view of a combustion liner embodying the present invention;

FIG. 3 is an enlarged sectional view taken along line 3 of FIG. 2; and

FIG. 4 is an enlarged sectional view taken along line 4 of FIG. 2.

Referring to FIG. 1, a gas turbine engine 10 is shown but not in its entirety. The gas turbine engine 10 includes an air flow delivery system 12 for providing combustion air and for providing cooling air for cooling components of the engine 10. The engine 10 includes a turbine section 14, a combustor section 16 and a compressor section 18. The combustor section 16 and the compressor section 18 are operatively connected to the turbine section 14. In this application the combustor section 16 includes an annular combustion chamber 24 being positioned about a central axis 26 of the gas turbine engine 10. As an alternative this could include a plurality of can combustors without changing the essence of the invention. The annular combustion chamber 24 is operative positioned between the compressor section 18 and the turbine section 14. A plurality of fuel nozzles 34 (one shown) are positioned in an inlet end portion 36 of the annular combustion chamber 24. The turbine section 14 includes a first stage turbine 38 being centered about the central axis 26.

As best shown in FIG. 2, the annular combustion chamber 24 is enclosed by an inner liner portion 40 and an outer liner portion 42 being spaced apart a preestablished distance. The inner liner portion 40 is spaced from the central axis 26 a preestablished distance and has a generally cylindrical configuration. The inner liner portion 40 includes an outer thin sheet metal annularly shaped skin member or interior liner 44 and an inner thin sheet metal annularly shaped skin member or exterior liner 46 being generally spaced one from the other a preestablished distance which in this application ranges from about 6 mm to about 15 mm. The outer skin members 44 has an inlet end portion 48 and an outlet end portion 50 axially spaced one from the other by an axial portion 52. And, the inner skin member 46 has an inlet end portion 54 and an outlet end portion 56 axially spaced one from the other by an axial portion 58.

As further shown in FIG. 2, the inner liner portion 40 further includes an inner inlet member 60 positioned at the inlet end portion 48 of the outer liner portion 44 being in communication with the compressor section 18 and being supported within the gas turbine engine 10 in a conventional manner. The outer skin member 44 defines a combustion side 62 and a cooling side 64 and has a preestablished configuration including a first end 66 being formed at the inlet end portion 48 and being attached to the inlet member 60. The inlet end portion 48 includes an axial portion 68 being connected to the inlet member 60 and a radial portion 70 extending from the axial portion 68. A straight portion 72 is connected to the radial portion 70 and forms a portion of the axial portion 52. An annular gallery 74 is formed between a portion of the straight portion 72, the radial portion 70 and a portion of the inlet member 60. A plurality of passages 76 extend through the radial portion 70 and communicate a flow of cooling air from the air flow delivery system 12 to the annular gallery 74. Spaced along the straight portion 72 at a preestablished distance and attached to the cooling side 64 is a plurality of stiffener members 78. A plurality of effusion cooling holes 80 are positioned in rows 82 along the straight portion 72. The rows 82 of the plurality of effusion cooling holes 80 are positioned axially along the straight portion 72 being spaced apart at a preestablished distance. The cooling holes 80 are spaced circumferentially along the rows 82 at preestablished intervals. The plurality of effusion cooling holes 80 are positioned in the outer skin member 44 at an angle of about 15 to 30 degrees and extend from the cooling side 64 through to the combustion side 62 and angle from the inlet end portion 48 toward the outlet end portion 50. A frustoconical or tapered portion 84 is connected to the straight portion 72 and forms the outlet end portion 50. The frustoconical portion 84 defines a cooling side 86 and a combustion side 88. Additional ones of the plurality of effusion cooling holes 80 are positioned in additional rows 82 along the frustoconical portion 84 and extend between the cooling side 86 and the combustion side 88 at an angle and angle from the inlet end portion 48 toward the outlet end portion 50. A transition portion 90 is connected to the frustoconical portion 84 and communicates with the turbine section 14. Further positioned in the frustoconical portion 84 is at least a row of dilution holes 92. The dilution hole 92 extends from the cooling side 86 through to the combustion hot side 88 at about a 90 degree angle. As best shown in FIG. 3, the spacing of the rows 82 and the positioning of the plurality of effusing cooling holes 80 along each of the rows 82 are arranged in a preestablished pattern 94 being generally defined as a diamond configuration having a centroid 96.

As further shown in FIG. 2, the inner skin member 46 of the inner liner portion 40 defines a first surface 100 being positioned adjacent the cooling side 64,86 and a second surface 102 being opposite the first surface 100. The inlet end portion 54 of the inner skin member 46 is attached to the straight portion 72 of the outer skin member 44 and has a configuration which spaces the outer and inner skin members 44,46 apart forming a first cooling cavity 106 therebetween. A straight portion 108 of the inner skin member 46 has a first end 110 and a second end 112. The first end 110 is connected to the first end portion 54 of the inner skin member 46 and has the first surface 100 spaced from the cooling side 64 a preestablished distance being generally equal along the entire axial distance of the straight portion 108 and forms a portion of the axial portion 52. The first cavity 106 is generally uniformly spaced apart a preestablished distance along an axial distance of the first cavity 106. The axial distance of the first cavity 106 being generally equal to the axial distance of the straight portion 108. A plurality of impingement holes 114 are positioned in a row 116 along the straight portion 108. The rows 116 of the plurality of impingement holes 114 are positioned axially along the straight portion 108 being spaced apart at a preestablished distance. The impingement holes 114 are spaced circumferentially along the rows 116 at preestablished intervals. The impingement holes 114 are positioned at generally a 90 degree angle to the first and second surfaces 100,102 of the inner skin member 46. The flow of cooling air from the air flow delivery system 12 is communicated to the first cooling cavity 106 through the plurality of impingement cooling holes 114. As best shown in FIG. 3, the spacing of the rows 116 and the positioning of the plurality of impingement holes 114 along each of the rows 116 are arranged in a preestablished pattern 118 being generally defined as a diamond configuration having a centroid 120. The plurality of holes 114 in the straight portion 108 of the inner member 46 are positioned in radial alignment with the centroid 96 of the preestablished pattern 94 of the plurality of holes 80 in the outer member 44. At the second end 112 of the straight portion 108, a plurality of spacer members 122 are intermittently positioned between the cooling side 64 of the outer skin member 44 and the first surface 100 of the inner skin member 46. Each of the spacer members 122 is attached to an annular member 124 in which the second end 112 of the straight portion 108 is positioned therein. Connected to the spacer members 122 and the annular sliding member 124 is an annular arcuate or tapered portion 126 at a first end 128 and has a second end 130 corresponding to the outlet end portion 56 connected to the transition portion 90. The annular arcuate portion 126 is spaced from the frustoconical portion 84 and forms a second cooling cavity 140. The spacing of the annular arcuate portion 126 from the frustoconical portion 84, in this application, is not necessarily evenly spaced along the second cooling cavity 140 between the first end 128 and the second end 130 of the annular arcuate portion 126. In this application, the spaced apart distance of the second cavity 140 is of a non-uniform spacing and the distance is smaller adjacent the second end 130. A plurality of non metering airflow inlet holes 142 are positioned in rows 144 and along the circumference of the rows 144 at predetermined locations. The plurality of non metering airflow inlet holes 142 are located closer to the first end 128 than to the second end 130 of the frustoconical portion 84. The flew of cooling air from the air flow delivery system 12 is communicated to the second cooling cavity 140 through the plurality of non metering airflow inlet holes 142. But, cooling airflow from the flow delivery system 12 is delivered to the first cooling cavity 106 and to the areas between the plurality of spacer members 122 by the impingement cooling holes 114. A support member 146 is attached to the annular arcuate portion 126 and supports the outlet end portion 50 of the outer skin member 44 by way of the transition portion 90 and the outlet end portion 56 of the inner skin member 46 in a conventional manner.

The outer liner portion 42 is spaced from the central axis 26 a preestablished distance, which in this application is a greater distance than the preestablished distance from the central axis 26 than that of the inner liner portion 40, and has a generally cylindrical configuration. The outer liner portion 42 includes an inner thin sheet metal annularly shaped skin member or interior liner 150 and an outer thin sheet metal annularly shaped skin member or exterior liner 152 being generally spaced one from the other a preestablished distance which in this application ranges from about 6 mm and about 15 mm. The inner skin member 150 has an inlet end portion 154 and an outlet end portion 156 axially spaced one from the other by an axial portion 158. And, the outer skin member 152 has an inlet end portion 160 and an outlet end portion 162 axially spaced one from the other by an axial portion 164.

The outer liner portion 42 further includes an outer inlet member 166 positioned at the inlet end portion 154 of the inner skin member 150 being in communication with the compressor section 18 and being supported within the gas turbine engine 10 in a conventional manner. The inner skin member 150 defines a combustion side 168 and a cooling side 170 and has a preestablished configuration including a first end 172 being formed at the inlet end portion 154 and being attached to the outer inlet member 166. The inlet end portion 154 includes an axial portion 174 being connected to the outer inlet member 166 and a radial portion 176 extending from the axial portion 174. A straight portion 178 is connected to the radial portion 176 and forms a portion of the axial portion 158. An annular gallery 180 is formed between a portion of the straight portion 178, the radial portion 176 and a portion of the outer inlet member 166. A plurality of passages 182 extend through the radial portion 176 and communicate a flow of cooling air from the air flow delivery system 12 to the annular gallery 180. Spaced along the straight portion 178 at a preestablished distance and attached to the cooling side 170 is a plurality of stiffener members 184. A plurality of effusion cooling holes 186 are positioned in rows 188 along the straight portion 178. The rows 188 of the plurality of effusion cooling holes 186 are positioned axially along the straight portion 178 being spaced apart at a preestablished distance. The cooling holes 186 are spaced circumferentially along the rows 188 at preestablished intervals. The plurality of effusion cooling holes 186 are positioned in the inner skin member 150 at an angle of about 15 to 20 degrees and extend from the cooling side 170 through to the combustion side 168 and angle from the inlet end portion 154 toward the outlet end portion 156. An inner conical or tapered portion 190 is connected to the straight portion 178 and forms the outlet end portion 156. The inner conical portion 190 defines a cooling side 192 and a combustion side 194. Additional ones of the plurality of effusion cooling holes 186 are positioned in additional rows 188 along the inner conical portion 190 and extend between the cooling side 192 and the combustion side 194 at an angle and angle from the inlet end portion 154 toward the outlet end portion 156. A transition portion 196 is connected to the inner conical portion 190 and communicates with the turbine section 14. Further positioned in the inner conical portion 190 is at least a row of dilution holes 198. The dilution hole 198 extend from the cooling side 192 through to the combustion side 194 at about a 90 degree. As best shown in FIG. 4, the spacing of the rows 188 and the positioning of the plurality of effusing cooling holes 186 along each of the rows 188 are arranged in a preestablished pattern 200 being generally defined as a diamond configuration having a centroid 202.

The outer skin member 152 of the outer liner portion 42 defines a first surface 210 being positioned adjacent the cooling side 170 and a second surface 212 being opposite the first surface 210. The inlet end portion 160 of the outer skin member 152 is attached to the straight portion 178 of the inner skin member 150 and has a configuration which spaces the inner and outer skin members 150,152 apart forming a first cooling cavity 216 therebetween. A straight portion 218 of the outer skin member 152 has a first end 220 and a second end 222. The first end 220 is connected to the inlet end portion 160 of the inner skin member 150 and has the first surface 210 spaced from the cooling side 192 a preestablished distance being generally equal along the entire axial distance of the straight portion 218 and forms a portion of the axial portion 164. The first cavity 216 being generally uniformly spaced apart a preestablished distance along an axial distance of the first cavity 216. The axial distance of the first cavity 216 being generally equal to the axial distance of the straight portion 218. A plurality of impingement holes 224 are positioned in a row 226 along the straight portion 218. The rows 226 of the plurality of impingement holes 224 are positioned axially along the straight portion 218 being spaced apart at a preestablished distance. The impingement holes 224 are spaced circumferentially along the rows 226 at preestablished intervals. The impingement holes 224 are positioned at generally a 90 degree angle to the first and second surfaces 210,212 of the outer skin member 152. The flow of cooling air from the air flow delivery system 12 is communicated to the first cooling cavity 216 through the plurality of impingement cooling holes 224. As best shown in FIG. 4, the spacing of the rows 226 and the positioning of the plurality of impingement holes 224 along each of the rows 226 are arranged in a preestablished pattern 228 being generally defined as a diamond configuration having a centroid 230. The plurality of holes 224 in straight portion 218 of the outer member 152 are positioned in radial alignment with the centroid 202 of the preestablished pattern 200 of the plurality of holes 186 in the inner member 150. At the second end 222 of the straight portion 218, a plurality of spacer members 232 are intermittently positioned between the cooling side 170 of the inner skin member 150 and the first surface 210 of the outer skin member 152. Each of the spacer members 232 are attached to an annular sliding member 234 in which the second end 222 of the straight portion 218 is slidably positioned. Connected to the spacer members 232 and the annular sliding member 234 is an outer conical or tapered portion 236 at a first end 238 and has a second end 240 corresponding to the outlet end portion 162 connected to the transition portion 196. The outer conical portion 236 is spaced from the inner conical portion 190 and forms a second cooling cavity 250. The spacing of the outer conical portion 236 from the inner conical portion 190 in this application is not necessarily evenly spaced along the second cooling cavity 250 between the first end 238 and the second end 240 of the outer conical portion 236. In this application, the spaced apart distance of the second cavity 250 is of a non-uniform spacing and the distance is smaller adjacent the second end 240. A plurality of non metering access holes 252 are positioned in rows 254 and along the circumference of the rows 254 at predetermined locations. The plurality of non metering access holes 252 are located closer to the first end 238 than to the second end 240 of the outer conical portion 236. The flow of cooling air from the air delivery system 12 is communicated to the second cooling cavity 250 through the plurality of non metering access holes 252. But, cooling airflow from the flow delivery system 12 is delivered to the first cooling cavity 216 and to the area between the plurality of spacer members 232 by the impingement cooling holes 224. A support member 256 is attached to the outer conical portion 236 and supports the outlet end portion 156 of the inner skin member 150 by way of the transition portion 196 and the outlet end portion 162 of the outer skin member 152 in a conventional manner.

Thus, the primary advantages of the improved combustor liner portions 24 is in the efficient use of the compressed cooling air. Since less cooling airflow per unit length of combustor wall, inner liner portion 40 and outer liner portion 42, is used there is a substantial reduction of CO emissions. The inner skin members 46 and outer skin member 152 of the inner liner and outer liner portions 40,42 respectively have a lower heat rejection to the gas turbine engine 10. The combination of the impingement and effusion cooling and the location of the plurality of impingement cooling holes 114,224 relative to the plurality of effusion cooling holes 80,186 allows the combustion chamber 24 to be subject to a very high heat flux as a result of high heat transfer rates conveyed by radiation and convection arising from the burning of fuel to be consistent with the design life expectancy of the combustor and its material properties. Thus, the improved impingement and effusion cooled combustor increases efficiency, reduces emissions and increases or maintains component life.

Other aspects, objects and advantages of this invention can be obtained from a study of the drawings, the disclosure and the appended claims.