Thermodynamic Analysis of a Turboprop Engine with Intercooling and Heat Recovery

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Thermodynamic Analysis of a Turboprop Engine with Intercooling and Heat Recovery Trans. Japan Soc. Aero. Space Sci. Vol. 54, No. 183, pp. 44–50, 2011 Thermodynamic Analysis of a Turboprop Engine with Intercooling and Heat Recovery By Roberto ANDRIANI,1Þ Fausto GAMMA2Þ and Umberto GHEZZI1Þ 1ÞDepartment of Energy, Politecnico di Milano, Milan, Italy 2ÞDepartment of Mechanics and Aeronautics, Universita` ‘‘La Sapienza’’, Rome, Italy (Received January 19th, 2010) In many modern gas turbine engine ground-based power plants, used for power generation, several auxiliary systems are integrated with the main gas generator to improve the generated output power or reduce fuel consumption. Two of the most effective practices are regeneration and intercooling. The first recovers part of the enthalpy in exhaust gas to pre-heat air before introducing it into the combustion chamber. The second cools the air during compression to reduce the work, and consequently obtain more power at the output shaft. These techniques are not used in gas turbine engines for propulsion mainly due to the extra weight and size caused by the heat exchangers and more complex flow patterns that result. However, if we could overcome these difficulties by means of compact heat exchangers the same benefits obtained for the ground-based plants could be obtained for aero engines. In particular the turboprop engine seems to be the best suited to this purpose due to its smaller mass flow rate and gas path. A thermodynamic cycle analysis shows the advantages of introduction of regeneration and intercooling in a turboprop engine in terms of increased power and reduced fuel consumption. Key Words: Turboprop, Heat Regeneration, Intercooling, Efficiency Nomenclature 1. Introduction T: temperature Fuel saving is always of primary importance in develop- s: entropy ing aircraft engines. Today, it has become fundamental, both M: flight Mach number for fuel price and environmental reasons. It is obvious that cp: constant pressure specific heat burning less fuel cuts emissions. In jet engines, research fuel Ch: hot fluid capacity rate consumption has led to development of very high by-pass Cc: cold fluid capacity rate engines, with the aim of reducing fuel consumption and : specific heats ratio increasing propulsion efficiency. Not much has been done : diffuser efficiency on the side of thermal efficiency, which is fundamental for ": heat exchanger efficiency fuel saving, except for maximum turbine inlet temperature j: nozzle efficiency (TIT) and overall pressure ratio. However, technologies p: propeller efficiency such as regeneration and intercooling, change the thermal g: gear box efficiency cycle and reduce fuel consumption (regeneration) or in- 1–9) t: power turbine efficiency crease output power (intercooling). These techniques Wconv: power adsorbed by conventional II stage com- are well known and used widely in ground-based power pressor plants, but not in aircraft engines, because of the need for Winterc: power adsorbed by intercooled II stage compressor heat exchangers, which are usually large and heavy, and WHPT: power delivered by high pressure turbine the consequent complex air/gas flow pattern. However, WC: power adsorbed by compressor compact heat exchangers are now available and weight I: compressor first stage pressure ratio and size can be reduced to levels suitable for aircrafts. Start- II: compressor second stage pressure ratio ing from these considerations, we have studied regeneration Qfuel: heat introduced by combustion and intercooling in a gas turbine engine, to see how they Qregen: heat transferred by heat exchanger affect performance and efficiency. We developed a thermo- Qcool: heat subtracted by intercooler dynamic numerical program to compute the thermal cycle of HPT: high pressure turbine a gas turbine engine with regeneration and intercool- TIT: turbine inlet temperature ing.10–14) The program solves all the thermodynamic charac- Áh: enthalpy drop available after HPT teristics of the working fluid: pressure, temperature, density, : power turbine enthalpy drop fraction entropy, specific heat, etc. Once solved, we can determine the main performance, such as specific power, specific fuel consumption, thermal efficiency, propulsion efficiency, Ó 2011 The Japan Society for Aeronautical and Space Sciences May 2011 R. ANDRIANI et al.: Thermodynamic Analysis of a Turboprop Engine with Intercooling and Heat Recovery 45 Intercooler 06 T Q Power to Inlet fuel compressor fuel 07 Gear LPC HPC HPT Heat Power to Power Turbine exchanger Nozzle propeller c.c. 05 08 041 Heat from recuperator Qregen 04 09 Fig. 1. Scheme of the turboprop engine, with regenerator and intercooler, 02 used for numeric simulation. 10 03 1 Qcool Heat from intercooler a global overall efficiency, etc. The code was written in s FORTRAN and implemented on a PC. Some simplifying Fig. 2. Thermal cycle of turboprop engine with regeneration of exhaust assumptions were made: both the air and gas flow through heat and staged-intercooled compression. the engine behave as perfect gas, following the perfect gas state equation; no auxiliary power is extracted by the turbine; no air bleeding from the compressor; turbine blades inlet at state 1 and then through the first compression stage are not cooled; complete expansion of the exhaust in the to state 02, where ‘‘0’’ indicates stagnation conditions. The nozzle; and component efficiencies are kept constant by stagnation (or total) conditions of the flow at state a are com- varying the operating conditions. Although these assump- puted as tions may seem unrealistic, they provide significant results À 1 T ¼ T 1 þ M2 ð1Þ without complicating the calculations too much. Moreover, 0a a 2 we do not intend to design a real engine, but instead to À1 À 1 2 provide a tool able to examine some engine configurations. p0a ¼ p0 1 þ M ð2Þ The results obtained in terms of absolute efficiencies and 2 performances are in good agreement with effective engines. Where, is the ratio of the specific heats of air, and M is the flight Mach number. The stagnation temperature is not 2. Equations and Analysis varied at the exit of the inlet, while the stagnation pressure is p ¼ p ð3Þ A turboprop engine was chosen for numeric simulation 01 0a for two reasons. 1. Turboprop and turboshaft engines have where is the diffuser efficiency. a smaller mass flow rate through their core compared to tur- The air temperature at the exit of the first stage of the bojet or turbofan engines. This means that size and weight of compressor is given by: heat exchangers can be smaller compared to jet engines; 2. 1 À1 The exhaust pressure from the power turbine is much lower T02 ¼ T0aðIÞ cI ð4Þ than in jet engines, consequently they cannot provide great Where, I is the compressor first stage pressure ratio, and cI power (in a turboprop engine the power provided by exhaust is its polytropic efficiency. thrust is about 10% of propeller power). This means that Before entering the second compression stage to reach the heat recovery from the exhaust in a turboprop engine can maximum pressure level, the air passes through the inter- be considered almost as a net gain. The scheme of the engine cooler where is cooled by external air. The amount of heat used for this simulation is shown in Fig. 1.15–20) exchanged between the two fluids depends, among other Unlike a conventional turboprop, there is an intercooler factors, on the temperature difference and efficiency of the and heat exchanger. The first cools the air between the heat exchanger.21) With reference to Fig. 3, the efficiency two compression stages. The cooling flow is external air. " of the intercooler is defined as follows: The heat exchanger (regenerator) pre-heats air from the q ChðTh,in À Th,outÞ CcðTc,out À Tc,inÞ high-pressure compressor before introducing it into the " ¼ ¼ ¼ ð5Þ qmax CminðTh,in À Tc,inÞ CminðTh,in À Tc,inÞ combustor, extracting the enthalpy from the exhaust. The gas flow pattern in the regenerator is more complex than where, q is the actual amount of heat exchanged, while qmax in the intercooler and this complexity with consequent pres- is the maximum amount of heat exchangeable between the sure drop is a main difficulty to overcome to introduce this two fluids in an ideal counterflow arranged heat exchanger technology into a propulsion system. of infinite surface. Ch is the hot fluid capacity rate and is The engine scheme in Fig. 1 is shown as a temperature- the product of the hot fluid mass flow rate m_ and its constant entropy (T-s) diagram in Fig. 2. pressure specific heat cp, Cc is the cold fluid capacity rate, Point ‘‘a’’ in Fig. 2 represents the external air. The air and Cmin is the smaller of the Ch and Cc magnitudes. In flow is then compressed by the flight velocity through the the intercooler, the mass flow rate of the external air is 46 Trans. Japan Soc. Aero. Space Sci. Vol. 54, No. 183 T size of the heat exchanger. Considering a compact heat exchanger with size and weight limited by the aircraft, we 21,22) Th,in chose a value of 0.6. Before entering the main combus- tor, air from the second-stage compressor is pre-heated in the heat exchanger (regenerator) where its temperature rises Ch from T04 to T05 by absorbing heat from hot gas at the power turbine exit. In this case, the added fuel is that required to increase temperature only from T05 to T06 because to Tc,out hot side increase from the compressor exit to burner inlet (T05) the enthalpy in the exhaust at the power turbine exit (T09)is Cc Th,out used. Intercooling, reducing the temperature at the second- cold side stage compressor exit, increases the amount of fuel needed Tc,in section 1 section 2 to reach the maximum cycle temperature with respect to the simple case.
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