Gas turbine engine systems and related methods involving vane-blade count ratios greater than unity
Gas turbine engine systems and related methods involving vane-blade count ratios greater than unity are provided. In this regard, a representative turbine stage for a gas turbine engine includes a first set of rotatable blades operative to be positioned downstream of and adjacent to a first set of vanes, a number of blades of the first set of blades being less than a number of vanes of the first set of vanes.
1. A gas turbine engine comprising:
a compressor section;
a combustion section operative to receive compressed gas from the compressor section;
a turbine section operative to impart rotational energy to the compressor section, the turbine section having a first set of vanes positioned adjacent the combustion section and a first set of rotatable blades positioned downstream of and adjacent to the first set of vanes, a number of vanes of the first set of vanes exceeding a number of blades of the first set of blades;
wherein the engine is a turbofan;
the engine has multiple turbine stages; and
a first stage of the multiple turbine stages comprises the first set of vanes and the first set of blades.
2. The engine of claim 1 , further comprising:
a second set of vanes positioned downstream of and adjacent to the first set of rotatable blades.
3. The engine of claim 2 , wherein:
the first set of vanes and the second set of vanes are provisioned with cooling air holes operative to provide film-cooling of the first set of vanes and the second set of vanes; and
a volume of cooling air per unit area of the vanes of the first set of vanes is greater than a volume of cooling air per unit area of the vanes of the second set of vanes.
4. The engine of claim 2 , wherein the engine is operative to complete a combustion reaction upstream of the second set of vanes.
5. The engine of claim 4 , wherein the engine is operative such that cooling air provided to the first set of vanes reacts with combustion products from the combustion section in order to complete the combustion reaction.
6. The engine of claim 1 , wherein the engine exhibits a high-fuel-to-air ratio turbine design.
7. A method for designing a gas turbine engine comprising:
selecting a number (N) of vanes and a corresponding number (N) of blades for a first turbine stage of a gas turbine engine having multiple turbine stages;
increasing the number of the vanes by M;
decreasing the number of blades by at least M such that the number of vanes of the first turbine stage exceeds the number of blades of the first turbine stage; and
wherein the engine is a turbofan.
8. The method of claim 7 , further comprising designing the engine as a high-fuel-to-air ratio turbine.