After some literature review and from a turbomachinery course I identified the following limiting design parameters,
- Minimum size of last row of high pressure compressor ( about 10 mm ? )
- Maximum temperature that can be tolerated by the machine's disks (about 950 K)
- Maximum temperature allowed for effective turbine blade cooling
Are there any other? Is there a maximum allowable pressure for combustion? Is there a maximum pressure that can be handled by the turbine's outer walls?
Thank you very much in advance
Well you seems to have all the answers aleady. At the Cavendish laboratory University of cambridge, they have current issues on pressure ratio..
At high pressures (and therefore, high starting temperatures in the combustion chamber), you have to consider ionization effects which take energy and may prevent its being released in the turbine expansion affecting the efficiency which is the reason for high pressure in the first place. The standard T-S diagram does not include non equilibrium energy exchanges of ionization, so you would have to run dynamic calculations for a specific process.
Metallurgical limits of GT components is one factor that limits pressure ratio.Similarly, high temperatures are experienced with increasing pressure ratio which will necessitate intercoolers between successive stages. This adds to the complexity and hence weight of the GT which becomes an issue with aero engines.
Actually, the pressure ratio is determined by the compressor, not the combustor or the turbine. The compressor needs more stages for higher pressure ratios. Since temperature rises with pressure, the later stages may become too hot without cooling, so at really high pressure ratios one might have to cool these stages actively or use intercoolers (difficult on aeroengines). The exit area depends on the mass flow, it might be a limitation for small mass flow engines.
Combustion temperature rises with compressor exit temperature, so NOx emissions become more difficult to handle. Also wall cooling becomes more difficult. Since NOx is formed at temperatures higher than 1800 K, the combustor exit temperature cannot be much higher than this. At rising compressor exit temperature, the amount of heat release becomes smaller, limiting the fuel flow.
The turbine will be more difficult to cool at higher temperature, however a much higher turbine entry temperature than 1900 K is unlikely due to NOx limitation. The turbine discs can be cooled efficiently at those temperatures, blades can be cooled with film cooling and internally.
The overal pressure ratio in compressor and maximum temperaure rise in the combustion chamber improve the gas turbine (GT) performance. There is an optimum pressure ratio maximazing simple cycle efficiency at given tempearure rise, but is not the same value that gives the maximum net power of GT. The situation is different when we apply the regeneration system. The from flue gas is heating of air after the compressor in order to improve the efficiency of an engine and save the fuel. In this case an optimum pressure ratio is much smaller at the fixed maximum TIT. In any case the limiting value of TIT (up to ~1800 K) is limited by cooling technology and thermal/mechanical resistance of materials (now we are able to use composite materials: ceramics&metal = cermetals).
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