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- h: 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 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 1 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 1 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 Th,outÞ CcðTc,out Tc,inÞ high-pressure compressor before introducing it into the " ¼ ¼ ¼ ð5Þ qmax CminðTh,in Tc,inÞ CminðTh,in 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. For this reason intercooling alone produces Fig. 3. Behavior of temperatures and capacity rates in counterflow heat more output power than a conventional cycle but increases exchanger. fuel consumption. Coupling with the regenerative cycle prevents use of extra fuel described above because the necessary heat is provided by enthalpy in the exhaust by assumed to be the same of that through the compressor, the heat exchanger. In this case, we can get the extra power while the specific heat is different due to the function of guaranteed by intercooling without spending extra fuel. temperature. The program accounts for the variability of In the combustion chamber, gas temperature rises to its specific heat with temperature. Thanks to the intercooled maximum level thanks to fuel combustion. We can write compression, the work to reach the same final pressure level the energy balance equation across the burner as: is less than for a simple compressor. m_ c T þ m_ H ¼ðm_ þ m_ Þc T ð9Þ With reference to Fig. 2, the compression can be divided a p(air) 05 f i b a f p(gas) 06 into two stages. The first stage, common to both compres- Where, Hi is the lower heating value and b is the combus- sors, compresses from point 1 to point 02, with a pressure tion efficiency. Combustion efficiency considers incomplete ratio of I. The second stage is different between the two combustion in a real burner. In this case, some of the fuel compressors; a conventional compressor continues its proc- reaction heat is not transferred to the gas to increase its tem- ess following the transformation 02–041, while a compres- perature. In the above equation, we neglect the enthalpy of sor with intercooling cools the air from point 02 to point the liquid fuel, because it does not cause serious error. The 03, with the intercooler subtracting the amount of heat constant pressure specific heat is computed separately for air Qcool given by and gas; in the first case it is considered as function of tem- perature only, while in the other case it is computed as func- Q ¼ m_ c ðT T Þð6Þ cool a p 02 03 tion of temperature and combustion humidity, computed as Then the compression continues from the point 03 to point function of the burner fuel/air ratio. Since fuel introduced 04. Indicating the common second-stage pressure ratio into the burner depends on the air temperature at the com- as II, the power adsorbed by the two compressors for the pressor exit T05, and on the amount of heat transferred in second stage is the regenerator, its value is obtained by numeric iteration. ð1Þ= Wconv ¼ m_ acpT02 II 1 ð7Þ As consequence of combustion, total pressure is reduced at the burner exit. We account for this by introducing the ð1Þ= Winterc ¼ m_ acpT03 II 1 ð8Þ pneumatic efficiency of the burner b, defined as Since the temperature T03 is lower than T02, the power p06 adsorbed by the second-stage compressor with intercooling b ¼ ð10Þ p05 is less than that adsorbed by the conventional design. This reduced compressor power absorption leads to a larger The high-pressure turbine drives the compressor, so enthalpy drop available to the power turbine and hence for assuming that no auxiliary power is extracted we can write propulsion. So intercooling reducing compression work, the power balance as WC ¼ WHPT or: gives more power at the propeller shaft. In a real case, we m_ acp(air)½ðT04 T03ÞþðT02 T01Þ must consider that the air passing through the intercooler ¼ðm þ m Þc ðT T ÞðÞ experiences a pressure drop. We can consider a pressure _ a _ f p(gas) 06 07 11 drop of about 3–6%. This means that the work to reach The power balance ignores air bleed from the compressor. the same final pressure will be a little higher with respect With this assumption, the mass flow rate through the com- to the ideal case described above; in any case lower with pressor end turbines only differs for fuel flow. From respect to the conventional case. Also, the air temperature Eq. (11) we can obtain the gas temperature level T07 at T03 at the heat exchanger exit will depend on the intercooler the high-pressure turbine exit (HPT) where the pressure efficiency. This will depend on many factors, especially the p07 is given by: May 2011 R. ANDRIANI et al.: Thermodynamic Analysis of a Turboprop Engine with Intercooling and Heat Recovery 47

1 Table 1. Efficiencies of components used in simulation. T07 HPT 1 p07 ¼ p06 ð12Þ Component Value T06 Intake efficiency 0.950 Where HPT is the HPT polytropic efficiency. Low pressure compressor polytropic efficiency 0.870 The propeller power is provided by the power turbine High pressure compressor polytropic efficiency 0.870 through a reduction gearbox. The enthalpy drop available High pressure turbine polytropic efficiency 0.900 for the power turbine is a fraction of the total drop, because Power turbine polytropic efficiency 0.900 part is given to the discharge nozzle. The global thrust Intercooler efficiency 0.6 horsepower of a turboprop is the sum of two terms due to Recuperator efficiency 0.6 the propeller and exhaust jet. Differentiation of the global Intercooler pressure drop (both sides) 5% thrust horsepower relation shows that the maximum value Recuperator pressure drop (both sides) 5% (assuming constant efficiencies of components) occurs when V 2 ¼ j 0 ð13Þ m_ f ð Þ2 EBSFC ¼ ð16Þ 2h pgt Pshaft þ Tjv1 Where, is the fraction of the ideal enthalpy drop h, avail- Together with SP and EBSFC, the code can compute the able after the HPT given to the power turbine. In the above thermal efficiency th, defined as: relation, , , , and represent the nozzle efficiency, j p g t Pshaft þ Ek(jet) propeller propulsive efficiency, gear efficiency, and power th ¼ ð17Þ m_ fHi turbine isentropic efficiency respectively, while V0 is the flight speed. In our simulation four different enthalpy distri- Where, Ek(jet) is the kinetic energy variation of the exhaust butions between the power turbine and nozzle are consid- jet and Hi is the lower heating value of the fuel. ered. Instead of fixing the value, as given by Eq. (13), we chose four values for pressure ratio in the nozzle: 1.1, 3. Results 1.2, 1.3, and 1.4. This choice was made to make calculations easier, because the presence of a heat exchanger after the Table 1 lists values of the efficiencies of the different power turbine does not allow direct use of Eq. (13). These engine components used in the simulation. nozzle pressure ratios correspond to values of about The graph simulate operation at 7,000 meters altitude and 0.10. Using these fixed values the exhaust nozzle is always 0.7 flight Mach number. Two turbine inlet temperatures completely expanded. The thrust provided by the nozzle Tj are considered, 1,300 K and 1,500 K, because they seem can be written as suitable for a new-generation engines at those flight condi- pffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffi tions. There are four curves for each engine, depending on T ¼ðm_ þ m_ Þ 2ð1 Þh m_ v ð14Þ j a f j a 1 the pressure ratio in the nozzle. The pressure ratios vary From Eq. (14) if is too great (at least 1) the thrust pro- from 1.1 to 1.4. For the engine with intercooling and regen- vided by the nozzle can become negative, depending on eration both at 1,300 K and 1,500 K, the efficiency of inter- the flight velocity v1. This explains why it is convenient cooling and regeneration is 0.6. This value is not high, and to leave part of the available enthalpy drop h to the nozzle, there are ground-application heat exchangers with higher especially if the flight velocity is considerable, to have efficiencies. However, there are two reasons for this choice. greater exhaust velocity than flight velocity. The first is to have a ‘‘baseline’’ for the advantages of the After solving the thermal cycle at different operating and introduction of the two technologies. If the results are good engine conditions, we can compute the main engine per- with low efficiencies they can only be better at higher formances and efficiencies. Two of the most important per- values. The second is because the size of a heat exchanger formance parameters are the specific power, and the specific mounted on an aircraft must necessarily be much smaller fuel consumption. The first is defined as the global power than used in ground-based plants. Since efficiency is related developed by the engine for unit air mass flow rate. It is in part to size of the exchanging surface, it seems unlikely the sum of the power at the propeller shaft and the propul- that very high efficiency levels can be reached. sive power of the exhaust jet, written as Figure 4 shows specific power (SP) for a turboprop 1 È engine with intercooling and regeneration (I-R engine) SP ¼ ðm_ a þ m_ fÞcpðT05 T06Þ and for a conventional turboprop (reference engine) as func- m_ a tion of compressor pressure ratio. Looking at the curves þ½ðm_ a þ m_ fÞv10 m_ av1v1gð15Þ of SP, the maxima correspond to nozzle pressure ratios Two different specific fuel consumption forms can be between 1.1 and 1.2 (the two lines nearly touch). This defined for the turboprop, depending on whether the contri- happens when the parameter (the enthalpy drop fraction bution of the exhaust jet is considered or not. In our case, of the power turbine) is about 0.90–0.95. In this condition, since part of the enthalpy drop is given to the nozzle, we the jet velocity and flight speed are almost the same. An do consider it and the specific fuel consumption expression increase in turbine inlet temperature of 200 K increases becomes: the maximum SP attained by both engines by about 30% 48 Trans. Japan Soc. Aero. Space Sci. Vol. 54, No. 183

z= 7000 [m] M0= 0.7 Tmax=1300 [K] z= 7000 [m] M0= 0.7 Tmax=1500 [K]

500 500

p.r. = 1.1 475 R=0; E=0 475 R=0.6; E=0.6 p.r. = 1.2

450 p.r. = 1.3 450 p.r. = 1.4 425 425

400 400

375 375

350 350

p.r. = 1.1 Specific Power [kW/kg/s] 325 Specific Power [kW/kg/s] 325 p.r. = 1.2 R=0; E=0 p.r. = 1.3 300 300 R=0.6; E=0.6 p.r. = 1.4 275 275

250 250 6 9 12 15 18 21 24 27 30 33 36 39 6 9 12 15 18 21 24 27 30 33 36 39 Pressure ratio Pressure ratio

Fig. 4. Specific power as function of pressure ratio for the reference engine and I-R engine, at four nozzle pressure ratios.

z=7000 [m] M0=0.7 Tmax=1300 [K] z=7000 [m] M0=0.7 Tmax=1500 [K]

0.5 0.5

0.48 0.48

0.46 0.46

0.44 0.44

0.42 0.42

0.4 0.4

0.38 0.38 Thermal efficiency Thermal efficiency

0.36 0.36

p.r. = 1.1 p.r. = 1.1 0.34 0.34 p.r. = 1.2 R=0; E=0 p.r. = 1.2 R=0; E=0 R=0.6; E=0.6 p.r. = 1.3 R=0.6; E=0.6 p.r. = 1.3 0.32 0.32 p.r. = 1.4 p.r. = 1.4

0.3 0.3 6 9 12 15 18 21 24 27 30 33 36 39 6 9 12 15 18 21 24 27 30 33 36 39 pressure ratio pressure ratio

Fig. 5. Thermal efficiency as function of pressure ratio for reference engine and I-R engine at four nozzle pressure ratios.

(mean). The behavior of both engines at 1,300 K and values for thermal efficiency than the I-R engine, but at very 1,500 K of TIT is very similar. At low pressure ratio, the high pressure ratios (at about 33–36 at 1,300 K and more reference engine at any nozzle pressure ratio, has a higher than 39 at 1,500 K). At these pressure ratios the I-R engine SP than the I-R engine. However, as pressure ratio increases, SP is very low as shown in Fig. 4, and it is impossible to at values greater than 9–15, the reference engine SP de- make the engine work in such condition. The I-R engine creases, while the I-R engine curves continue rising, reach- instead reaches its maximum thermal efficiency at low pres- ing their SP maximum at about 21 and 27 for 1,300 K and sure ratios (12–15 at 1,300 K, about 18 at 1,500 K) and then 1,500 K, respectively. This happens because at low pressure decreases at higher values. However the curves are quite ratios, intercooling has little effect on reducing compression flat, so thermal efficiency remains high also for pressure work while there is a pressure drop in the intercooler. Com- ratios higher than those where the maximum is. In particu- paring the curves, the maximum SP of the I-R engine is lar, the maximum values of thermal efficiency and SP are increased with respect to the reference engine by about at the same pressure ratios. Also thermal efficiency has its 4% at 1,300 K, and more than 5.5% at 1,500 K. maximum at nozzle pressure ratios of about 1.2. Figure 5 shows the behavior of thermal efficiency for the Figure 6 shows the EBSFC curves for both engines. two engines as function of pressure ratio at 1,300 K and The specific fuel consumption of the I-R engine is much 1,500 K of TIT. The reference engine reaches slightly higher lower than the reference engine, especially at low pressure May 2011 R. ANDRIANI et al.: Thermodynamic Analysis of a Turboprop Engine with Intercooling and Heat Recovery 49

z = 7000 [m] M0 = 0.7 Tmax=1300 [K] z = 7000 [m] M0 = 0.7 Tmax=1500 [K]

0.29 0.29

p.r. = 1.1 p.r. = 1.1 p.r. = 1.2 R=0; E=0 p.r. = 1.2 0.27 R=0; E=0 0.27 R=0.6; E=0.6 p.r. = 1.3 R=0.6; E=0.6 p.r. = 1.3 p.r. = 1.4 p.r. = 1.4 0.25 0.25

0.23 0.23

0.21 0.21 EBSFC [kg/h/kW] EBSFC [kg/h/kW]

0.19 0.19

0.17 0.17

0.15 0.15 6 9 12 15 18 21 24 27 30 33 36 39 6 9 12 15 18 21 24 27 30 33 36 39 Pressure ratio Pressure ratio

Fig. 6. Specific fuel consumption as function of pressure ratio for reference engine and I-R engine at four nozzle pressure ratios. ratios. As the pressure ratio increases, the two engines have z= 7000 [m] M0= 0.7 similar values, because the regeneration effect vanishes 500000 as pressure ratio increases, since the amount of heat trans- 450000 nozz.pr=1.1 ferred in the regenerator becomes smaller. If we consider nozz.pr=1.2 the case at 1,500 K the reference engine reaches the same 400000 nozz.pr=1.3 nozz.pr=1.4 value as EBSFC for the I-R at a very high pressure ratio 350000 (more than 30), while the I-R engine reaches its minimum Rig = 0.6 Eff = 0.6 at about 18. 300000 The EBSFC curves of the I-R engine have the same SP 250000 and thermal efficiency characteristic: they are almost con- 200000 stant over a wide range of pressure ratios. At 1,300 K, the Tmax=1500 [K] fuel consumption is between 0.19 and 0.21 over the whole 150000 Heat exchanged [j/kg] range of pressure ratios, while at 1,500 K, except the first 100000 point of the curve at pressure ratio of 6, the EBSFC value is about 0.18 at any pressure ratio. Considering SP in 50000 Tmax=1300 [K] Fig. 4, using the I-R engine we can choose a design pressure 0 ratio where SP is maximum and EBSFC is close to its mini- 6 9 12 15 18 21 24 27 30 33 36 39 mum. This is not possible with the reference engine because pressure ratio SP has a sharp maximum at low pressure ratios and EBSFC reaches a minimum at very high pressure ratios. In the Fig. 7. Heat transfer in heat exchanger at 1,300 K and 1,500 K turbine inlet temperature as function of pressure ratio for nozzle pressure ratios curves at 1,300 K, above some pressure ratio (from 33 to between 1.1 and 1.4. 36 depending on the nozzle pressure ratio) it is impossible to compute performances, because the heat exchange in the regenerator falls as pressure ratio rises, vanishing above 4. Discussion these values. Figure 7 shows the behavior of heat transfer in the heat exchanger at both TIT as function of pressure ratio. Simultaneous use of regeneration and staged intercooled Although at TIT 1,500 K it is possible to exchange heat compression process in a turboprop engine can give good in the regenerator at any pressure ratio and nozzle pressure results in terms of general performances. In particular, it is ratio, when TIT is reduced at 1,300 K this becomes impos- possible to reach high SP values without increasing specific sible. In fact at a pressure ratio of 39 no heat is exchanged fuel consumption. In fact, the increase in output power due regardless of the nozzle pressure ratio. At 36, the regenera- to reduced power adsorbed by the compressor thanks to tor works (although almost no heat is exchanged) only for intercooling is covered in terms of extra fuel by heat sub- a nozzle pressure ratio of 1.4. At 33, only nozzle pressure tracted by the regenerator from hot gas at the power turbine ratios of 1.4, 1.3, and 1.2 support heat exchange. The max- exit. It is very interesting to see how the reference and I-R imum of pressure ratio where all four cases exchange heat is engines behave with respect to operating conditions. From 30, but with very low heat transfer values. Fig. 4 (SP), the reference engine has maximum of SP at 50 Trans. Japan Soc. Aero. Space Sci. Vol. 54, No. 183 low pressure ratios (about 12 at 1,300 K and about 15 at . The size and weight of the heat exchangers, and the 1,500 K) then it falls as pressure ratio rises. Figure 6 complexity of the consequent flow pattern, are critical (EBSFC) shows that the reference engine reaches low spe- problems to be solved for the introduction of these cific fuel consumption at high pressure ratios (about 24– techniques to aircraft engines. 27 at 1,300 K, about 30–33 at 1,500 K). At low pressure ratios where SP is maximum, fuel consumption of the refer- References ence engine is still high. This means that to change the en- gine working conditions from maximum power to minimum 1) Boggia, S. and Rud, K.: Intercooled and Recuperated Aeroengine, fuel consumption, we must change the pressure ratio using a DGLR Paper 2004-179, 2004. variable-geometry compressor or a bypass for some com- 2) Boggia, S. and Rud, K.: Intercooled and Recuperated Engine Concept, AIAA Paper 2005-4192, 2005. pressor stages at maximum power. From the same graph 3) McDonald, C. F.: Low Cost Recuperator Concept for Microturbine the I-R engine should also need variable geometry (varying Applications, ASME Paper 2000-GT-167, 2000. in opposite direction because maximum power is at higher 4) McDonald, C. F., Massardo, A. F., Rodgers, C. and Stone, A.: Recu- pressure ratios with respect to the minimum specific fuel perated Gas Turbine Aeroengines, Part II: Engine Design Studies Following Early Development Testing, Aircraft Engineering and consumption) to meet at any operating condition. However, Aerospace Technology, 80 (2008), pp. 280–294. the I-R engine curves are much flatter than the reference 5) McDonald, C. F. and Wilson, D. G.: The Utilization of Regenerative engine. In fact, at 1,300 K for TIT in the pressure ratio range and Recuperative Cycles for High Efficiency Gas Turbines in the 21st from 12 to 30, the maximum variation in SP is about 5%, Century, J. Appl. Thermal Engineering, 16 (1996), pp. 635–653. 6) McDonald, C. F. and Langworthy, R. A.: Advanced Regenerative and in the same range, the EBSFC variation is similar. At Gas Turbine for Lightweight and High Performance, ASME Paper 1,500 K for TIT at pressure ratios from 15 to 39, SP varies 1971-GT-67, 1971. less than 5% and EBSFC variation less than 4%. This is 7) European Commission Research Directorate’s Fifth Framework important for an aero engine. If we consider that turboprop Program. 8) Kentfield, J. A. C.: Regenerative Turbofans: A Comparison with operating conditions change often during a mission (take Non-regenerative Units, J. Aircraft, 12 (1975), pp. 174–182. off, climb, fast cruise, low cruise, etc.) the possibility of 9) Miura, Y. and Sakurai, C.: Advanced Turbofan Concept with Regen- always being at the best performance point over a wide erator and Intercooled Compressor, AIAA Paper 84-1270, 1984. range of pressure ratios is a considerable gain. It seems that 10) Andriani, R., Gamma, F. and Ghezzi, U.: Regeneration and Intercool- ing in Gas Turbine Engines for Propulsion Systems, 44th AIAA Aero- the I-R turboprop engine works at any operating condition at space Sciences Meeting and Exhibit, Hartford, CT, July 21–23, 2008. its maximum SP and near its minimum EBSFC. Off-design 11) Andriani, R. and Ghezzi, U.: Heat Recovery in Reciprocating Engine, analysis is needed for confirmation but the first results of this 35th AIAA/ASME/SAE Joint Propulsion Conference and Exhibit, study follow this direction. It is also interesting to note that Los Angeles, CA. Paper 99-2664, July 20–24, 1999. 12) Andriani, R. and Ghezzi, U.: Influence of Heat Recovery and Inter- the intercooler allows exchange of more heat in the regener- cooling on Turboprop Engine Behaviour, Int. J. Turbo Jet-Engines, ator. This is a sort of system self-balancing. As the pressure 25 (2008), pp. 259–267. ratio rises, the heat removed by the intercooler increases, 13) Andriani, R. and Ghezzi, U.: Main Performances of a Turboprop which increases the power output (due to reduced compres- Engine With Heat Exchange and Intercooling, 2nd EUCASS— European Conference for Aerospace Sciences, Bruxelles, 1–6 July, sion work), but increases fuel consumption, because air at 2007. the compressor exit is colder. However, the extra needed 14) Andriani, R. and Ghezzi, U.: Off-Design Analysis of a Jet Engine heat is not provided by extra fuel but is from the regenerator, With Heat Recovery, 38th AIAA Aerospace Sciences Meeting and so the overall effect increased power at almost constant Exhibit, Reno, NV., Paper 00-0743, Jan 10–13, 2000. 15) Ascher, H. Shapiro: The Dynamics and Thermodynamics of Compres- specific fuel consumption. sible Fluid Flow, Vol. 1, The Ronald Press Company, New York, 1953. 5. Conclusion 16) Oates, G. C.: Aerothermodynamics of Gas Turbine and Rocket Propul- sion, AIAA, Education Series, New York, 1984. 17) Mattingly, J. D.: Elements of Gas Turbine Propulsion, Mc Graw-Hill The authors conducted a numerical study regarding the International Editions, Mechanical Engineering Series, Singapore, possibility of introducing intercooling and regeneration to 1996. turboprop engines. The parametric study considered a turbo- 18) Liew, K. H., Urip, E., Yang, S. L. and Siow, Y. K.: A Complete prop engine at different turbine inlet temperatures and con- Parametric Cycle Analysis of a Turbofan with Interstage Turbine Burner, 41st AIAA Aerospace Sciences Meeting and Exhibit, Reno, ditions of constant-speed cruising. The main results are NV, January, 2003, AIAA-2003-0685. summarized as follows: 19) Liew, K. H., Urip, E. and Yang, S. L.: Parametric Cycle Analysis of a . The introduction of both intercooling and regeneration Turbofan Engine with an Interstage Turbine Burner, J. Propul. Power, has a positive effect on the performance and specific 21, 3 (2005), pp. 546–551. 20) Sirignano, W. A. and Liu, F.: Performance Increase for Gas-Turbine fuel consumption of turboprop engines. In particular, Engines Through Combustion Inside the Turbine, J. Propul. Power, it enables the actualization of higher specific power 15, 1 (1999), pp. 111–118. together with lower specific fuel consumption. 21) Kays, W. M. and London, A. L.: Compact Heat Exchangers, 2nd ed., . The efficiency of the heat exchangers for both inter- McGraw-Hill, New York, 1964. 22) Pellischek, G. and Kumpf, B.: Compact Heat Exchanger Technology cooling and regeneration is very important to obtain for Aero Engines, 10th International Symposium on Air Breathing improvements in performance: values that are too Engine (X ISABE), Nottingham, September 1–6, 1991. low do not lead to the desired results.