Small gas turbine engine with multiple burn zones
A small gas turbine engine for use in an UAV such as a cruise missile, the gas turbine having a combustor forming a primary burn zone and a secondary burn zone, and in which fuel is injected into both the primary and the secondary burn zones by either a rotary cup injector or a plurality of fuel injector nozzles. The secondary burn zone with separate fuel injection allows for the diameter of the engine to be reduced in size but still allow for adequate power and efficiency to be reached for powering the vehicle. Air flow from the compressor is used to cool the combustor walls before being injected into the combustor, and to pass through and cool the guide nozzles and a main bearing located near the hot section of the combustor prior to being introduced into the combustor.
1. A gas turbine engine comprising:
a compressor connected to a rotor shaft;
a turbine connected to the rotor shaft;
an annular combustor positioned between the compressor and the turbine and encircling the rotor shaft;
the annular combustor being defined by an outer wall and an inner wall;
the annular combustor having an axial length greater than a distance between the outer wall and the inner wall;
the annular combustor forming a primary combustion zone in a forward end of the combustor and a secondary combustion zone immediately downstream from the primary combustion zone;
a first fuel injection means to inject a fuel into the primary combustion zone to produce a primary combustion swirl;
a second fuel injection means to inject a fuel into the secondary combustion zone to produce a secondary combustion swirl in an opposite direction than the primary combustion swirl direction;
the first and second fuel injection means injects both fuels from one of the outer or inner walls of the combustor, and the distance between the inner wall and the outer wall of the combustor is short enough so that the first and second fuels and air are impinged against the opposite wall to produce the swirl in the combustion chamber; and,
the first and second fuel injection means includes a rotary cup injector rotatably fixed to the rotor shaft, the rotary cup injector includes a front face and a rear face in which the fuel is supplied onto and then discharged as a thin film into the two combustion zones.
2. The gas turbine engine of claim 1 , and further comprising:
the front face of the rotary cup injector is parabolic in shape and the rear face is substantially flat and slightly inclined toward the front face.
3. The gas turbine engine of claim 1 , and further comprising:
the combustor includes the primary combustion zone and the secondary combustion zone and then a dilution zone located downstream from the two combustion zones and immediately upstream of a guide nozzle of the turbine, the dilution zone being thinner than the two combustion zones.
4. The gas turbine engine of claim 1 , and further comprising:
the forward end of the combustor is positioned underneath a diffuser of the compressor;
the combustor is positioned within an enclosure that defines a cooling air passage between the outer wall of the annular combustor and the enclosure wall;
the combustor outer wall at the primary and secondary combustion zones includes air supply holes to discharge air from the diffuser into the two combustion zones; and,
The cooling air passage between the enclosure and the outer wall is narrow such that a high velocity is developed to produce a high convection cooling of the outer wall.
5. The gas turbine engine of claim 3 , and further comprising:
the guide nozzle includes a cooling air channel extending through the nozzle; and,
means to divert cooling air from the diffuser through the cooling air channel in the guide nozzle and then into the secondary combustion zone.
6. The gas turbine engine of claim 5 , and further comprising:
a bearing to support the aft end of the rotor shaft positioned in the cooling air passage so that the diverted cooling air passing through the guide nozzle channel and into the secondary combustion zone also passes through the bearing to provide cooling for the bearing.
7. The gas turbine engine of claim 1 , and further comprising:
the first and second fuel injection means includes a first fuel injector positioned on an outer wall of the combustor to inject fuel into the primary combustion zone and a second fuel injector positioned on an outer wall of the combustor and downstream from the first fuel injector to inject fuel into the secondary combustion zone.
8. The gas turbine engine of claim 4 , and further comprising:
the annular combustor is enclosed within the enclosure such that all of the cooling air from the diffuser flowing over the outer wall to provide cooling for the combustor eventually flows into the combustor.
9. The gas turbine engine of claim 1 , and further comprising:
the rotary cup injector extends into the combustor through an opening in an inner wall of the combustor.
10. The gas turbine engine of claim 1 , and further comprising:
the first and second fuel injection means injects the fuel into a location where the two combustion zones meet.
11. The gas turbine engine of claim 1 , and further comprising:
the primary combustion zone and the secondary combustion zone having about the same height when measured from the outer wall and the inner wall.
12. A gas turbine engine comprising:
a compressor connected to a rotor shaft;
a turbine connected to the rotor shaft;
an annular combustor positioned between the compressor and the turbine and encircling the rotor shaft;
the annular combustor being defined by an outer wall and an inner wall;
the annular combustor being enclosed by an enclosure forming a cooling air passage around the combustor forward end and inner and outer walls, the cooling air passage being thin enough to produce a high rate of cooling air flow around the combustor;
a plurality of cooling air supply holes spaced around the combustor to discharge the cooling air into the combustor;
a first fuel injection means to inject a fuel into the primary combustion zone to produce a primary combustion swirl;
a second fuel injection means to inject a fuel into the secondary combustion zone to produce a secondary combustion swirl in an opposite direction than the primary combustion swirl direction;
the combustor is formed with a primary burn zone in the forward end and a secondary burn zone downstream from the primary burn zone; and,
the forward end of the combustor is positioned underneath a diffuser outlet.
13. The gas turbine engine of claim 12 , and further comprising:
the combustor includes a dilution zone located downstream from the secondary burn zone; and,
a guide nozzle downstream from the dilution zone, the guide nozzle including a cooling air passage to allow for cooling air to flow around the dilution zone and then into the combustor.
14. The gas turbine engine of claim 12 , and further comprising:
the enclosure provides for all of the air discharged from the diffuser and used to cool the combustor to be discharged into the combustor and burned with the fuel.
15. The gas turbine engine of claim 12 , and further comprising:
the first and second fuel injection means both inject fuel from the inner wall or the outer wall and at an axial location between the primary combustion and the secondary combustion zone.