fig1 shows a turbomachine , in this example a turbojet 10 , having a pylon 12 and fastened by the pylon 12 to an airplane wing . the turbojet 10 is described in greater detail with reference to fig2 , 3 , and 4 . fig2 shows the turbojet 10 in perspective and in section on plane iia of fig1 . the turbojet 10 comprises a nozzle 14 ( or first nozzle 14 ) extending along an axial direction a corresponding to the axis of rotation of the turbine 15 . the turbojet also has a second nozzle 16 for the secondary stream and presenting an outlet 16 a . the second nozzle 16 is coaxial with the first nozzle 14 , being outside the first nozzle 14 . the turbojet 10 is a bypass turbojet in which the first nozzle 14 conveys a hot gas stream gc ( or primary stream of hot gas ) while the second nozzle 16 conveys a cold gas stream gf ( or secondary stream of cold gas ). more particularly , the hot gas stream gc flows inside the first nozzle 14 , while the cold gas stream gf flows between the first nozzle 14 and the second nozzle 16 . the downstream axial end 14 a of the first nozzle 14 forms the exhaust of the first nozzle 14 . in this example , the turbojet 10 also has an ejection cone 17 . naturally , in a variant , the turbojet need not have an ejection cone . in this example , the first nozzle 14 also has an intermediate outlet 14 b for ventilation gas gv . the ventilation gas gv is used to ventilate components of the turbomachine without contributing directly to its propulsive effect . naturally , in a variant , the first nozzle need not have an intermediate outlet . the pylon 12 extends in a radial direction r from a base 12 a . more particularly , in this example , the base 12 a of the pylon 12 includes a rear mount 22 for mounting a turbine casing 25 on the pylon , as shown in fig2 b . the rear mount 22 has a plurality of links 22 a distributed in azimuth , and in this example it has three links 22 a . a fairing of the rear mount 12 b protects the rear mount 22 and provides aerodynamic continuity for the cold gas stream gf . in general manner , the mounting of the body 10 a of the turbojet 10 on the pylon 12 is shown in fig5 , in which there can be seen the rear mount 22 to which the turbine casing 25 is mounted and a front mount 23 to which an intermediate casing is mounted that is located axially between the fan 27 and a low pressure compressor 28 . more precisely , an internal shroud 26 of the intermediate casing is mounted on the front mount 23 . it should be observed that the body 10 a of the turbomachine 10 also comprises , from upstream to downstream : a high pressure compressor 29 , a combustion chamber 30 , a high pressure turbine 31 , the turbine casing 25 in this non - limiting example housing a low pressure turbine and containing an annular flange fastened to the rear mount 22 . the turbojet 10 has a pylon cooler device 18 that , in this example , comprises two auxiliary gas ducts 20 arranged in azimuth on either side of the pylon 12 , being incorporated in the rear mount fairing 12 b . each duct 20 has a gas inlet 20 a taking gas from the cold gas stream gf , and a gas outlet 20 b leading to the vicinity of the base 12 a of the pylon 12 , radially between the exhaust 14 a and the base 12 a ( cf . fig4 ). in other words , in this example , the gas outlets 20 b open out radially “ under ” the base 12 a ( the term radially “ under ” should be understood as meaning that the base is on the outside relative to the outlet , or that the outlets are on the inside relative to the base , and substantially in radial alignment ). in this example , the two outlets 20 b extend in azimuth over substantially the entire length in azimuth z of the base 12 a of the pylon 12 , and more particularly over about 80 % of the length in azimuth z of the base 12 a of the pylon 12 . naturally , in a variant , the two outlets 20 b could extend in azimuth over the entire length in azimuth z of the base 12 a of the pylon 12 , i . e . over 100 % thereof . furthermore , the gas outlets 20 b are arranged at the same axial position as the exhaust 14 a ( cf . fig3 ). the gas outlets 20 b are also arranged downstream from the intermediate outlet 14 b for ventilation gas gv , in line with the intermediate outlet . this configuration presents the advantage that the stream of cooling gas gr creates a cooling film under the base 12 a while contributing to the overall thrust of the turbojet 10 , the cooling gas stream gr flowing substantially axially . as shown in fig2 , the gas inlet 20 a faces upstream and directly receives a fraction of the cold gas stream gf . the gas inlet 20 a is arranged in the upstream end 13 of the fairing 12 b of the rear mount 22 . the cross - section of each gas duct 20 varies going from the gas inlet 20 a to the gas outlet 20 b . specifically , the shape of the section of the duct 20 at its inlet 20 a is substantially triangular , whereas at the gas outlet 20 b , the shape of the section of the duct is substantially in the shape of a portion of a ring ( or a shape that is substantially rectangular ). the cooling gas stream gr leaving the gas outlets 20 b forms a film of gas over the downstream portion of the base 12 a of the pylon 12 , thereby preventing the hot gas stream gc coming from the exhaust 14 a ( i . e . downstream from the exhaust ) from coming into contact with the pylon 12 and its base 12 a . furthermore , the gas stream gr cools the pylon 12 and its base 12 a or keeps them at a temperature that is acceptable . although the present invention is described with reference to specific embodiments , it is clear that modifications and changes could be undertaken on those embodiments without going beyond the general scope of the invention as defined by the claims . in particular , individual characteristics of various embodiments as shown and / or mentioned above may be combined in additional embodiments . consequently , the description and the drawings should be considered in a sense that is illustrative rather than restrictive .