The present disclosure relates to a gas turbine engine, and more particularly, to a first stage vane assembly of a high pressure turbine of the engine.
Gas turbine engines are rotary-type combustion turbine engines built around a power core made up of a compressor, combustor and turbine, arranged in flow series with an upstream inlet and downstream exhaust. The compressor section compresses air from the inlet, which is mixed with fuel in the combustor and ignited to generate hot combustion gas. The turbine section extracts energy from the expanding combustion gas, and drives the compressor section via a common shaft. Expanded combustion products are exhausted downstream, and energy is delivered in the form of rotational energy in the shaft, reactive thrust from the exhaust, or both.
The turbine section typically includes alternating rows of turbine vanes and turbine blades. The turbine vanes are stationary and function to direct the hot combustion gases that exit the combustor section. The vanes and blades each project from respective platforms that when assembled form vane and blade rings. Airfoils of the vane and blade rings are designed with pitch-to-chord ratios that are generally dependent on a wide variety of engine characteristics and operating parameters. Achieving the optimal pitch-to-chord ratio is desirable to optimize engine efficiency and performance. Moreover, achieving higher pitch-to-chord ratios may reduce the number of required airfoils and/or reduce cooling requirements thereby improving engine efficiency and reducing engine manufacturing and maintenance costs.
USPTO Backgrounds
USPTO Backgrounds