Igniter cable conduit for gas turbine engine

1. A gas turbine engine comprising:

a compressor section configured to compress a gas;

a combustor section aft of the compressor section and configured to combust the gas; a fan section configured to produce a bypass flow;

an outer case having an inner surface and an inner case having an outer surface defining a bypass flow duct there between; and

a cable conduit, disposed within the bypass flow duct, comprising:

a head end, a sleeve having a foot end, a boot, and a split grommet having a first upper surface,

wherein the head end is coupled to the sleeve opposite the foot end,

wherein the boot comprises a first flange and is coupled to the foot end,

wherein the head end comprises a second flange having a second upper surface and a cutout penetrating into an interior volume of the sleeve,

wherein the split grommet is coupled about an inner diameter of the cutout, and

wherein the second flange is coupled at the second upper surface to the inner surface of the outer case and the first flange is coupled to the outer surface of the inner case.

2. The gas turbine engine of claim 1 , wherein the cable conduit further comprises a forward section and an aft section.

3. The gas turbine engine of claim 2 , wherein the forward section comprises a first perforation through a forward aerodynamic surface and the aft section comprises a second perforation through an aft aerodynamic surface, wherein the first perforation and the second perforation are in fluid communication with the bypass flow duct and the interior volume of the sleeve, wherein a portion of the bypass flow passes through the first perforation into the interior volume and exits the interior volume through the second perforation.

4. The gas turbine engine of claim 3 , wherein the boot comprises a circumferential channel.

5. The gas turbine engine of claim 4 , wherein the forward section further comprises forward sleeve having a first distal end and the aft section further comprises an aft sleeve having a second distal end, wherein the first distal end and the second distal end are disposed within the circumferential channel.

6. The gas turbine engine of claim 5 , wherein the boot tends to allow the forward section and the aft section to translate in response to a differential thermal growth between the outer case and the inner case.

7. The gas turbine engine of claim 6 , wherein the sleeve comprises a sleeve locking feature.

8. The gas turbine engine of claim 6 , wherein the split grommet comprises a mating surface comprising at least one of an extrusion, a finger, a cavity, a pocket, a bore, or an embedded stud.

9. The gas turbine engine of claim 6 , wherein the sleeve comprises at least one of a circular cross section, an elliptical cross section, an oblate cross section, an angular cross section, a teardrop cross section, or an airfoil cross section.