I canno't understand why higher pressure ratios are not applied to GTs in the 150 MW to 400 MW range. I assume that firing temperature is similar for both aeroderivatives (AD) and heavy duty (HD) turbines ( is that correct? ). In addition HD turbines have higher mass flow rates meaning that it should be easier to design the HPC for higher pressure ratios.
I know that HD GTs are normally designed for combined cycle operation and that lowers the optimum pressure ratio, but it doesnt justify pressure ratios around 10, when the optimum is around 30.
Could gas turbines with very high mass flow rates AND very high pressure ratios be designed in theory?
I have read the technical literature, but I can't find any relevant information. Any comment on the difference between AD and HD gas turbines would be very much apreciated.
Pressure ratio in AD turbines is selected to obtain the maximum cycle efficiency. Optimum pressure ratio increases with higher TIT, higher turbomachinery efficiency, better blade cooling, all features found in modern sophisticated AD GTs . High cycle efficiency means low fuel consumption: this is an advantage in aero propulsion, for several reasons: (I) lower amount of fuel to be transported, (ii) lower operating cost (jet fuel is more expensive than natural gas, (iii) AD gas turbines are operated for a large number of hours.
The pressure ratio that yields the maximum combined cycle efficiency of gas turbines operating in combined cycles is lower than in simple cycle (because bottoming steam cycle benefits of higher turbine outlet temperatures). Moreover, also the pressure ratio that yields the maximum specific power (kW/(kg/s of air) is lower than pressure ratio that yields the maximum simple cycle efficiency: Higher specific power means lower GT dimensions and cost,
These are the main reasons why modern AD turbines have pressure ratios >30, while HD gas turbines have pressure ratios in the range 15-20..
I think in AD turbines we need much thrust so we try to rise PR to have better thrust but in HD turbines, thrust force is useless and we only use it to rotate the shaft.
Increasing the gas turbine power to produce higher power is performed through increasing the firing temperature or by increasing the mass flow rate. All improvements to increase the power are turning around these two ways. Due to their limitations, the pressure ratio of both the air compressor and the turbine is limited. That is also why the air compressor stages is about 17 stages (the stage ratio is about 1.15, the total ratio is about 10) By increasing the firing temperature to increase the power will lead to increasing the compressor stages over 17 and hence increasing the turbine ratio. If couldn’t increase the firing temperature, due to material limitation, increasing the mass flow rate for the same stages is the available solution which will lead to massive designed gas turbine. The weight limitation which is pertaining to aeroderivative gas turbine leads to extended improvement studies , to increase the power through increasing firing temperature, in the field of blade material. As well as this is the reason , to compensate for lesser weight , that is running faster compared to HD industrial gas turbine , which is huge and running slower. For the available gas turbine without posing the possibility to improve through by the mention ways the designer go around increasing the mass by. for example (air compressor evaporation and inter stage wet compression , combustion injection, modifying seal labyrinth techniques, decreasing extracted air , modifying gas turbine blade angle,………) or via increasing the fringe temperature , for example by ( improving blade coating and cooling techniques,……….) or by decreasing the consumed gas turbine mechanical and accessory loads,………….).
Aeroengines are designed with focus on high power / weight ratio while NOx regulations are much less tight that for stationary gas turbines. A higher overall pressure ratio will increase the compressor exit temperature and the adiabatic flame temperature (if no intercooling is used), leading to more NOx emissions (which increase exponentially at temperatures higher than 1600-1800 K). Also , higher turbine entry temperatures require more expensive materials. In aeroengines material cost is often secondary to power and weight.
Also aeroengines feature a 2 stage or 3 stage compressor, which is easier to design for high overall pressure ratio than a single stage compressor typical for stationary gas turbines.
In principle, it is not difficult to design a high pressure gas turbine for big mass flow. A pure geometrical scaling is, however, not possible for large scaling ratio since centrifugal forces increase nonlinearly with size and material stress will not allow too large wheels running at the same circumferential speed.
In principle you might operate several high pressure compressors and turbines in parallel with joint combustor, however it might be simpler to just operate several complete gas turbines in parallel.
Pressure ratio in AD turbines is selected to obtain the maximum cycle efficiency. Optimum pressure ratio increases with higher TIT, higher turbomachinery efficiency, better blade cooling, all features found in modern sophisticated AD GTs . High cycle efficiency means low fuel consumption: this is an advantage in aero propulsion, for several reasons: (I) lower amount of fuel to be transported, (ii) lower operating cost (jet fuel is more expensive than natural gas, (iii) AD gas turbines are operated for a large number of hours.
The pressure ratio that yields the maximum combined cycle efficiency of gas turbines operating in combined cycles is lower than in simple cycle (because bottoming steam cycle benefits of higher turbine outlet temperatures). Moreover, also the pressure ratio that yields the maximum specific power (kW/(kg/s of air) is lower than pressure ratio that yields the maximum simple cycle efficiency: Higher specific power means lower GT dimensions and cost,
These are the main reasons why modern AD turbines have pressure ratios >30, while HD gas turbines have pressure ratios in the range 15-20..
Maybe I have missed the point, or it is too obvious: Aeroengines work at much lower intake pressure (20 kPa instead of 100 kPa) so they need more compression to reach comparable absolute pressures, kg/s, power per weight, and power per size.
Actually, the maximum specific power (kW/(kg/s) is achieved at lower pressure ratios. It is true that the power is proportional to the density of Air at compressor inlet, but increasingincreasing
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