Combustor liner panel with a multiple of heat transfer ribs for a gas turbine engine combustor
A liner panel for use in a combustor of a gas turbine engine, including a multiple of heat transfer ribs located in at least one discrete area of the liner panel. A wall assembly within a gas turbine engine including a support shell, a liner panel mounted to the support shell via a multiple of studs and a multiple of heat transfer ribs located in at least one discrete area on a cold side of the liner panel, each of the multiple of heat transfer ribs account for the multiple of studs.
1. A liner panel for a combustor of a gas turbine engine, the liner panel comprising:
a multiple of heat transfer ribs that extend from a cold side of the liner panel, each of the multiple of heat transfer ribs comprising a respective multiple of cooling holes that extend therethrough, each cooling hole of the respective multiple of cooling holes for each of the multiple of heat transfer ribs define a direction vector directed at an angle to penetrate a hot side and the cold side, the direction vector of each cooling hole aligned with the respective heat transfer rib, and wherein the multiple of heat transfer ribs avoids cold side structures.
2. The liner panel as recited in claim 1 , wherein the multiple of heat transfer ribs are located in response to a pressure field, the multiple of heat transfer ribs flow with respect to the pressure field, the higher the pressure, the greater the number of heat transfer ribs.
3. The liner panel as recited in claim 1 , wherein the multiple of heat transfer ribs are located in response to a pressure field on a hot side of the liner panel the multiple of heat transfer ribs flow with respect to the pressure field, the higher the pressure, the greater the number of heat transfer ribs.
4. The liner panel as recited in claim 1 , wherein the cold side structures comprise a mount stud.
5. The liner panel as recited in claim 1 , wherein the multiple of heat transfer ribs form a semi-circular pattern around a mount stud.
6. The liner panel as recited in claim 1 , wherein the multiple of heat transfer ribs form a converging pattern downstream of a first mount stud.
7. The liner panel as recited in claim 6 , wherein the converging pattern converges between two mount studs.
8. The liner panel as recited in claim 7 , wherein the converging pattern is located on both sides of a semi-circular pattern around a mount stud.
9. The liner panel as recited in claim 1 , wherein each of the multiple of heat transfer ribs are about 0.040 (1 mm) in width.
10. The liner panel as recited in claim 1 , wherein each of the multiple of heat transfer ribs are about 0.033 inches (0.84 mm) in height.
11. The liner panel as recited in claim 1 , wherein each of the multiple of heat transfer ribs are spaced about 0.040 (1 mm) center to center.
12. A wall assembly within a gas turbine engine comprising:
a support shell;
a liner panel mounted to the support shell via a multiple of studs that extend from a cold side of the liner panel; and
a multiple of heat transfer ribs located in at least one discrete area on a cold side of the liner panel, each of the multiple of heat transfer ribs comprising a respective multiple of cooling holes that extend therethrough, each cooling hole of the respective multiple of cooling holes for each of the multiple of heat transfer ribs define a direction vector directed at an angle to penetrate a hot side and the cold side, each cooling hole of the respective multiple of cooling holes aligned with the respective heat transfer rib, wherein the multiple of heat transfer ribs form a semi-circular pattern around a mount stud, and wherein the multiple of heat transfer ribs form a converging pattern downstream of the mount stud, the converging pattern converges between two downstream mount studs.
13. The assembly as recited in claim 12 , wherein the converging pattern is located adjacent to the semi-circular pattern.