Performance Trends of High-Bypass Civil Turbofans

Transcription

1 Original Author :Adam J. Head Primary Author :Adam J. Head Editor :Arvind Rao, Feijia Yin Revision : 1.0 Date : July 19, 2015 Doc no : LR-FPP

2 Authorization Signature Date Prepared Adam J. Head Approved Arvind Rao Doc nr: LR-FPP Revision: 1.0 Page 2/33

3 Summary Twenty four exhausted (unmixed) high-bypass civil turbofan powerplants performance was investigated with NLR s GSP and GasTurb 1. Simulations were made from a dual spool turbofan configuration in the GSP software. The selected engines were taken from three engine manufactures over a range of net thrust rating 80 to 490 kn (T-O). The mathematical models were constructed in order to perform an On-Design and Off-design analysis. These models were tuned with known/published On-Design and Off-Design data. Of particular interest were the polytropic efficiencies of the primary core components and their influence on the overall performance of the powerplant at cruise setting in the flight envelope. Trending could therefore be derived with FAA certification issued dates over a period of 50 years ( ). Graphs of component and system efficiencies of various families of engines were graphed. The primary points of interest are: Gas path parameters such as the TIT (CET), OPR, BPR, Max spool rotational speeds. Component parameters of the High Pressure Compressor (HPC), High Pressure Turbine (HPT), Low Pressure Turbine (LPT), Low Pressure Compressor (LPC) such as Polytropic efficiencies (ETA_poly). System performance parameters such as Thermal, Propulsive, Overall Efficiency (ETA_th, ETA_prop, ETA_overall). 1 0D thermodynamic Gas Turbine performance modeling software codes. Doc nr: LR-FPP Revision: 1.0 Page 3/33

4 Contents 1 Introduction Motivation and Assumptions Global Assumptions Research Question Selected Power Plants Performance Analysis CRP, Design Point and Off-design Analysis Cycle Reference Point (CRP) and Design Point Matching Design Point Matching Off-design Point Matching Summary Method of Performance Matching Introduction to the Modeling Process Method of Performance matching Generic Maps Reference engine Cruise setting Certification date Averaged HPC PR/Stage (Static) Limitations of Model Discussion of Results Graphs of selected performance parameters Polytropic Component Efficiencies System Performance Efficiencies System Gas Path Performance Parameters Critical evaluation of results Reason for increase over the last years Verification Parameters used for verification Verification/checks Verification/checks Improvements GSP BUGS Doc nr: LR-FPP Revision: 1.0 Page 4/33

5 10 Recommendations (Future Work) Suggestions for further development: Conclusions Annexes: References: Appendices: Station Numbering (Turbofan Exhaust) Equations Bypass Ratio Propulsive Efficiency: Thermal Efficiency: Doc nr: LR-FPP Revision: 1.0 Page 5/33

6 1 Introduction Engine Manufacturers such as Pratt & Whitney, General Electric and Rolls Royce have engaged in intense competitions for new engine contracts with either commercial or military applications over the years. This has resulted in steep improvements in technology levels with regards to compressor/combustor/turbine components the core engine turbomachinery. The civil market is of particular interest and within this sector exhausted High-Bypass civil turbofan powerplants warrant some observation. A performance analysis of powerplants with different thrust outputs and thus different aircraft application was conducted. Naturally the technology level of these turbomachienery components has increased over the decades. Performance input variables such as component efficiencies were guessed- and subsequently tuned and their influence over the system thermal, propulsive and ideal efficiency and combustor exit temperature and BPR recorded. Generic maps for the core gas path components were used for all powerplants in the study. The CRP for each engine was often chosen to be the same for a majority of the variants in the generic maps. However there were some engines where the CRP had to be changed to a different point on the map otherwise the results deviated to far from the TSFC or fuel flow at off-design conditions. The design point was defined at T-O setting in the flight envelope because this is where most available information is recorded. It is also the thermodynamic and rotational maximum of the gas cycle, [2]. Unknown values such as BPR, OPR and polytropic efficiencies of the HPC, LPC, HPT,LPT need to be tweaked to match known performance points which are scarce at the cruise aircraft rating, and becomes harder if component maps are unavailable. In the design point phase the component efficiencies/bpr/opr were chosen and tuned until the individual powerplants matched their rated thrust and TSFC at T-O. Subsequently in the OD phase a thrust controller was used to set the thrust constant at the required off-design setting in the flight envelope. At most, two off-design points (two flight settings) were cross-checked with available engine data in the Off-Design tuning process. The first was at cruise (as this is the setting of primary interest) and the second was Idle. Tweaking parameters such as the efficiency deterioration modifier in each component of every powerplant was tuned to match the TSFC at cruise. When/If the Idle OD point was matched, the same deterioration modifiers where used. The generic map CRP may also be shifted for each individual engine for reasons of tweaking to the TSFC or Fuel flow. The trends obtained will be useful for a complete predictive maintenance monitoring of the system, as explained in an article written by W. E. Forsthoffer [3]. Trending will enable the user to optimize run times between maintenance cycles amongst other advantages. The annexes of the report are mxl and.xlsx files which are the GSP performance models arranged in sets of families of the modeled engines. The Turbofan_Powerplants10.xlsx file contains over 700 civil turbofan variants. Within the excel spreadsheet engine characteristics are shown in the same synthetic way: thrust, specific fuel consumption, engine weight, bypass-ratio, overall pressure ratio, turbine entry temperature as with the reference [4]. Doc nr: LR-FPP Revision: 1.0 Page 6/33

7 2 Motivation and Assumptions It may be interesting to know how the performance of various engines is trending with respect to each other over time. What is the behavior with respect to the ideal Brayton cycle efficiency and most importantly where are we heading. The technological development of gas turbine components is improving but often the exact numbers are difficult to devise. What overall cycle performance parameters are important and how their behavior has influenced the cycle efficiencies is hard to tie together. 2.1 Global Assumptions A few global assumptions will be stated first and then later in the report the local assumptions will be mentioned where necessary. Exhaust nozzle velocity loss coefficient and flow coefficient are constant and ideal for every modeled engine. As is the assumption in sources such as [5], Thus the losses in the nozzle inlet and the exhaust are assumed to be ideal, as is the case with most of the rated TSFC in Jane s aircraft. Real behavior of gases is implemented, Bleed and turbine cooling is sometimes implemented as well with max spool speeds. See [5, 6] for details, a maximum of 8% bleed is used (bled from end of compressor) whereby a max of 75% is used to cool the HPT (unless otherwise stated). There is NGV cooling (mixed before expansion in rotor) and additional rotor/nozzle blade cooling flows (mixed after rotor or stator). Please see respective models for details on numbers. Published DP values and OD point values can be stated for differently defined cruise settings amongst the selected engines. It is always the case that cruise is modeled the same for every engine. The cruise point will be defined at M=0.8 and H= 10,670m. This is the same for every engine. The corresponding data in annex turbofan_powerplants10 may differ for certain parameters but it is noted where possible. Eg. The TSFC is sometimes taken for M0.73 H10670m. Same generic maps same losses in every engine. The CRP is for the most part the same for every component in every engine. The design point (which is taken at T-O, as this is where the most information is found) will be different for every engine and thus the scaling will be different. The errors induced in the scaling will be reduced by the fact that the modeling is restricted to: o 2 shaft exhaust configurations and relatively same number of stages. The FAA certification date is the chosen certification date for the power plant of interest however there may be other certification dates which correspond to application on later aircraft. Other certification agencies such as JAA, EASA and CAA may differ. Where there are two or more aircraft applications the engine may be certified again for a later application. Therefore the latest certification is always taken in this case. There may be a reissue of a certification due to problems. The first certification date is always taken in that case. Measures to help the engineering judgment will come from keeping in mind T45 redline values, TIT temperature material limits and Max rotational speeds. Ambient conditions 60% humidity at ambient conditions. Doc nr: LR-FPP Revision: 1.0 Page 7/33

8 3 Research Question For the engine variants specified in Table 4-1 what are the system performance parameters (Overall Efficiency, Propulsive Efficiency, Thermal Efficiency, Overall Pressure Ratio, Bypass Ratio, Net Thrust to Weight Ratio) and component parameters (Polytropic HPC, LPC, HPT, LPT efficiency) at cruise setting in the flight envelope. Where do these performance parameters sit, in relation to each other, when plotted against the FAA certification date (entry into service). 4 Selected Power Plants The selected engines were: Table 4-1 The Selected Power plants and if Bleed/Turbine cooling was used. Engine Variant Bleed/Turbine Cooling Engine Variant Bleed/Turbine Cooling Engine Variant CF6-50C2B Yes GE90-76B Yes PW4052 No CF6-50 No GE90-94B No PW4060 No (Baseline) 2 CF6-80A Yes GE90-110B1 No PW4098 No CF6-80C2B1 Yes JT3D (Baseline) No TayRB.183- No 3Mk CF6-80E1A2 Yes JT3D-3B No V2527-A5 Yes CF34-3A1 No JT9D (Baseline) No V2527E-A5 No CF34-8C1 No JT9D-3A Yes CFM56-5B2 No JT9D-7Q3 No GP7270 No PW2040 No The engines which had bleed and turbine cooling implemented are listed above. For exact numbers and locations the reader is referred to [6] or the GSP mxl file (PowerPlantsExhaust.mxl). For closer tweaked results it is recommended that bleed and turbine cooling be implemented in all the turbofans. Implementing this in all the engine models would be impractical but recommended for future work. Bleed/Turbine Cooling 5 Performance Analysis The method of performance modeling will be one similar to the approach taken by [2]. However fewer operating points in the flight envelope will be matched and generic maps will be used in the reference engine as a basis for all 24 engines modeled. 5.1 CRP, Design Point and Off-design Analysis Information available at three points in flight envelope (including the design T-O point): T-O (Design Point): 2 The baseline is the original engine, original core. It was used as a calibration for the future variants because temperatures and pressures of some baseline engines were available. In principle the engine name/code doesn t exist on any aircraft application but has an extension to the name. Doc nr: LR-FPP Revision: 1.0 Page 8/33

9 Known: Thrust, TSFC, max temperatures at certain points on stations in cycle, BPR, OPR, Fuel flow at T-O (ICAO data sheets) and Airflow Unknowns: ETA_LPC, ETA_HPC, ETA_ LPT, ETA_HPT Assumed: deltap_cc loss = 0.04, ETA_CC = 0.99, ETA_mech_N1 = 0.99, ETA_mech_N2= 0.99, Cruise (Off-Design Point): Known: TSFC, thrust (not for all engines) Unknowns: Airflow, ETA_LPC, ETA_HPC, ETA_ LPT, ETA_HPT, BPR, OPR, Fuel flow ect. Idle (Off-Design Point): Known: TSFC, thrust Unknowns: See Cruise point Cycle Reference Point (CRP) and Design Point Matching Design point tuning is done by adapting the unspecified data such as pressure losses, unknown efficiency etc. in order to match all GSP results to the specified engine performance parameter values of the specific operating point assigned at GSP design point. A design point (or cycle reference point) must be defined which serves as a reference point for subsequent off-design steady-state. Important is the location of the map design point that coincides with the engine design point and is used to scale the map for off-design simulation. Off-design tuning is subsequently performed by adapting map design points (and the maps themselves if necessary,) and off-design loss and efficiency relations Design Point Matching The design point (T-O in this study) must be defined which serves as a reference point for subsequent off-design steady-state. This point is matched to variables given at take-off (T-O) mentioned in section 5.1. The chosen map design point on the generic maps prior to scaling can be shifted for the individual engines when using the generic maps to improve results at off-design Why model the Design Point at T-O? Performance and component data is limited at the cruise OP, and so the best way to approach the problem will be setting the initial design point at T-O (This is where a predominant amount of info is, in sources such as: Jane's Aircraft, ICAO, and FAA US certification websites.) Reference points/design Points are commonly also defined at cruise (and other points) for civil aircraft. This unfortunately restricts us to the amount of parameters that we can match due to limitations of information at this aircraft rating, especially if component maps are scarce. The net thrust (influenced by ram drag) and TSFC are commonly known for Doc nr: LR-FPP Revision: 1.0 Page 9/33

10 different aircraft ratings and so seem to be appropriate measures to match at the design point and other OD points in the envelope Off-design Point Matching The cruise thrust is controlled and the fuel flow or TSFC matched using deterioration modifiers in the turbine and compressor components. 3 Changing the location of the Map Design Point also assists with tweaking the above variables as already mentioned. Figure 1 Chosen/fixed/controlled cruise thrusts In an article written by Svoboda [7] turbofan data taken at takeoff is recorded in a series of graphs. A relation derived for cruise thrust calculation: T CR (N) = T TO (N) This is used where thrust wasn t recorded at cruise Summary The cycle reference point is the same for most engines and the design point will be taken at T-O. A steady state off-design point will be simulated at the same cruise conditions for all engines. Matching for the net thrust and Tuning/optimizing for the TSFC through deterioration modifiers will be done at cruise. 3 Fuel Flow data or rather data at different levels of flight were obtained in the ICAO database. Aircraft Engine Emissions Databank Doc nr: LR-FPP Revision: 1.0 Page 10/33

11 6 Method of Performance Matching 6.1 Introduction to the Modeling Process The engines will be initially modeled with the performance parameters at T-O (design point). The first step of the modeling process will be to supply unknown design input values such as BPR, OPR, ETA_PolHPC, ETA_PolLPC, ETA_PolHPT, ETA_PolLPT. These values need to be guessed (if unknown) to match known performance data at T-O; such as thrust, TSFC and any temperatures and pressures along the gas path which may also be known. FAA data sheets can be used to gauge maximum spool speeds and material red line limits. 4, [6]. A variety of engines may also be configured for bleed and subsequently turbine cooling when the information was available 5, see Table 4-1, [5]. The second step is to model off-design behavior and subsequently tweak parameters at the cruise setting Method of Performance matching Method or Approach for Turbofan trend Modeling in detail: A reference engine was taken with standard generic maps (BIGFAN_PerfDeck.mxl from the standard library of GSP as an Example. Which uses the bigfanc.map, bigfand.map, bigbst.map, bighpc.map, turbimap.map, and biglpt.map) from the GSP library (Unmixed). The reference engine will be used to model the 24 engines specified. All the engines will thus have the same generic maps. The steady state off design points will be simulated at the same cruise conditions for all engines by thrust (after which there is perhaps need for tuning/optimising and/or matching for the TSFC). T-O will be defined as the same for all 24 engines. Therefore in summary the method should be a 3 step process. 1. Match the CRP/Design Point to the T-O power setting of each turbofan (The Design Point -- CRP of the cycle). 2. Set up a steady state (OD points) by controlling the thrust (thrust controller) at cruise/idle to match the thrust level. The third step will involve the tuning step to TSFC, as there will be deviations. 3. Tuning to TSFC at cruise power setting by using the deterioration modifiers/variables to the efficiencies of components. Eg. Change the efficiency by the same differences in each component if TSFC is too low then increase ETAs, to match the TSFC at cruise. Caution needs to be given because this is the only variable that is known in most cases, at this point in the flight envelope. Any original deviation will be a consequence of the actual maps not corresponding with the components. Or additionally incorrectly matching the OP parameters in the flight envelope. A note on matching the net (thrust): The net thrust is an uncommon parameter (or inappropriate rating) to use as a matching performance variable, It is rare to be directly concerned with since the pilot will 4 The maximum permissible temperatures are noted from the FAA: (T45) Interturbine temperature are noted and used also as a guiding parameter when sizing the engine, [5]. 5 Janes Aircraft was used to gather information about bleeding locations and turbine cooling. Verifying certification dates and technology level updates on the engine variants. It has been taken into consideration when sizing the turbine power plant. 6 There may be double ups of some variants in the excel spread sheet due to the fact that the application of same civil aircraft have different demands on certain parameters. Eg. Certification dates. Doc nr: LR-FPP Revision: 1.0 Page 11/33

12 vary the thrust of the aircraft to accommodate the required thrust to maintain constant altitude and airspeed to meet with air traffic control requirements, [8] and [9]. The net thrust seems to be unavailable in most instances as its definition at cruise changes with defined power setting at T-O thrust. Eg. Varies from 18 to 25 % of T-O thrust. A more common parameter listed instead, is the TSFC which is a function of thrust and fuel flow. Therefore the primary variable of interest to be matched at cruise will be the optimizing variable, the TSFC Generic Maps Comparing Engines on the same generic maps for aircraft gas turbine performance calculations have some important observations that need to be made in terms of assumptions and limitations. These have been listed in the assumption page, section Reference engine Figure 2 GSP Tree structure and model configuration The reference engine will use generic maps (from the standard GSP library bigfanc.map, bigfand.map, bigbst.map, bighpc.map, turbimap.map, and biglpt.map). This reference engine will be used to model the 24 engines specified Table Cruise setting 7 Is generally defined to be 20-25% of the net rated thrust at T-O. There were a few engines which did not have the thrust at cruise specified. As explained in section 6.2. Therefore in the case where the cruise thrust was not specified 22% of the net rated T-O thrust was used as an estimate. 7 In this report cruise is defined at M=0.8 and H=10670m. The corresponding data in annex turbofan_powerplants may differ for certain parameters but it is noted where possible. Eg. The TSFC is sometimes taken for M0.73 H10670m. The TSFC at altitude is taken from both pg 272 [Roskam J.] and also Fuel Flow data or rather data at different levels of flight were obtained in the ICAO database. Aircraft Engine Emissions Databank. Doc nr: LR-FPP Revision: 1.0 Page 12/33

13 6.3.1 Certification date The certification date is the certification for the aircraft application stated in turbofan_powerplants10.xslm excel. 6.4 Averaged HPC PR/Stage (Static) An estimation of the average PR per stage is gives the designer an idea of the size of new engine technology. Possible dimensions can be derived with the subsequent formula. 1 HPC avg,per stage = PRHPC stages Equation 1 Table 2 Static average HPC PR/ stage Engine variant HPC PR HPC stages Average HPC PR/stage JT3D-Baseline JT3D-3B JT9D-Baseline PW JT9D-3A GE90-76B V2527E-A CF6-50C2B CF CF6-80A V2527-A CF6-80C2B CF34-3A JT9D-7Q PW PW CF6-80E1A CF34-8C GE90-94B CFM56-5B PW GP GE90-110B Limitations of Model Limitations Comparing Engines on the same generic maps for aircraft gas turbine performance calculations may be acceptable at least for trend analysis. However separate Doc nr: LR-FPP Revision: 1.0 Page 13/33

14 component maps for the respective power plants should be used to reduce scaling errors. We don't actually know the correct combinations of efficiencies of the LPC,HPC,LPT,HPT when specifying the design point. There is a general consensuses of what component should have higher efficiencies relative to each other perhaps. The values are true guess work that can only be further refined with the suggestions in the future work section below. The map will induce error but is much better than taking a CRP at cruise where we have even more unknowns. Matching only at one OD (cruise point/idle) is perhaps the fastest way at the moment, please see recommendations for further work. There are some data that helps to model the engine at T-O. A few examples are given below The maximum permissible temperatures are noted from the FAA airweb: (T45) Interturbine temperature are noted and used also as a guiding parameter when sizing the engine at T-O. The maximum shaft speeds N1 and N2 are noted. Measures to help the engineering judgment will come from keeping in mind T45 redline values, TIT temperature material limits and Max rotational speeds. 7 Discussion of Results 7.1 Graphs of selected performance parameters Polytropic Component Efficiencies Figure 3 Polytropic High Pressure Turbine component efficiency trends [5, 6, 10] Doc nr: LR-FPP Revision: 1.0 Page 14/33

15 Figure 4 Polytropic Low Pressure Turbine component efficiency trends [5, 6, 10] Figure 5 Polytropic Low Pressure Compressor component efficiency trends [5, 6, 10] Doc nr: LR-FPP Revision: 1.0 Page 15/33

16 Figure 6 Polytropic Low Pressure Compressor component efficiency trends [5, 6, 10] Figure 3 through Figure 6 show the polytropic efficiencies of the high and low turbine and compressor components. The efficiencies are plotted against FAA entry into service. The spread and technology improvements are wider and extended along a bigger range for the compressor. Figure 7 Gas turbine engine Core Side Fan efficiency trends [5, 6, 10] Doc nr: LR-FPP Revision: 1.0 Page 16/33

17 7.1.2 System Performance Efficiencies Figure 8 Gas turbine engine Overall efficiency trends [5, 6, 10] Figure 9 Gas turbine engine thermal efficiency trends [5, 6, 10] Doc nr: LR-FPP Revision: 1.0 Page 17/33

18 Figure 10 Gas turbine engine propulsive efficiency trends [5, 6, 10] It may be important to note that not all engines are simulated at Mach 0.8. They range from 0.75 to This is because the reported SFC at cruise was made for the respective Mach numbers. The range of 0.7 to 0.85 is well with the recorded efficiency noted by sources such as, [11]. Where the ideal propulsive efficiency is around 0.8 at a Mach number of for large Fan Pressure Ratios, Figure 11 Gas turbine engine Brayton cycle efficiency trends [5, 6, 10] Doc nr: LR-FPP Revision: 1.0 Page 18/33

19 Figure 12 Gas turbine engine ideal turbofan thermal efficiency trends [5, 6, 10] Figure 13 Gas turbine engine cycle/ thermal efficiency trends [5, 6, 10] System performance efficiencies have also shown an increase with entry into service date. This is consistent with improvement in technology level and material technology over the past 50 years. Figure 13 attempts to illuminate the reader on the closing gap between what we can achieve and what is the theoretical max. As this gap closes we expect to see the component efficiencies plateau off to constant numbers. In other words the rate of increase will start to diminish to a fixed value. Doc nr: LR-FPP Revision: 1.0 Page 19/33

20 Figure 14 Gas turbine engine maximum thermal efficiency trends [5, 6, 10] System Gas Path Performance Parameters Figure 15 Gas turbine engine combustor exit temperature trends [5, 6, 10] Doc nr: LR-FPP Revision: 1.0 Page 20/33

21 Figure 16 Gas turbine engine overall pressure ratio trends [5, 6, 10] Figure 16 shows the history of aircraft engine pressure ratio versus entry into service, and it can be seen that there has been a large increase in cycle pressure ratio. Figure 17 Gas turbine engine bypass pressure ratio trends [5, 6, 10] Figure 17 shows the history of aircraft engine bypass pressure ratio versus entry into service, and it can be seen that there has been a large increase especially in the last 20 years. Doc nr: LR-FPP Revision: 1.0 Page 21/33

22 Figure 18 Gas turbine engine TSFC trends [5, 6, 10] Figure 18 shows the history of TSFC versus entry into service, and it can be seen that the TSFC has almost decreased to half its value in the last 50 years. Figure 19 Gas turbine engine net thrust to weight ratio trends [5, 6, 10] Doc nr: LR-FPP Revision: 1.0 Page 22/33

23 Figure 20 Static average HPC PR/ stage Equation 1 and thus Table 2 were used to create Figure Critical evaluation of results Numbers shown from Figures 1-14 seem consistent with today s technology level. Polytropic efficiencies for the LPT component in Figure 4 might be somewhat over predicted. Other figures/component performance parameters appear to provide realistic results and is consistent with open literature, [12]. 7.3 Reason for increase over the last years There are a few obvious reasons explaining the trends over the last 50 years. Some of them are: The use of single crystal alloys for turbine airfoils; Fan blade aerodynamics which transition from subsonic, through transonic, to supersonic across the span of a single airfoil; and, Extremely low volume heat release combustion chambers. New materials are still being developed today, eg. ceramics and ceramic coatings for airfoils. Performance benefits from such technology improvements can be seen in the trends over the years in three key cycle variables. Increases in maximum (turbine inlet) temperatures, pressure ratio, and turbofan bypass ratio are critical to system performance. Resulting in lighter, more fuel efficient engines. Trend curves show increasingly diminishing returns as engine performance approaches the max (ideal). However strong competition drives improved fuel consumption and lighter weight engines. Doc nr: LR-FPP Revision: 1.0 Page 23/33

24 Not to mention the continued development of engineering software codes. Compressible Fluid Dynamics (CFD) computer programs are now solving complex aerodynamic problems, heretofore unsolvable, for improved compressor and turbine airfoil designs; while ultra-high bypass ratio, geared-technology turbofans are on the drawing boards today. Therefore there is good reason to anticipate continued gains in propulsion gas turbine performance., [13, 14]. Figure 21 Taken from [13, 14] 8 Verification 8.1 Parameters used for verification Thrust, TSFC at various points in the envelope Table 8-1 Engine Variant and verified OD points Verified OD points on flight envelope Engine Variant Engine Variant Verified OD points on flight envelope Engine Variant Verified OD points on flight envelope CF6-50C2B Cruise/Idle GE90-76B Cruise PW4052 Cruise CF6-50 Cruise GE90-94B Cruise PW4060 Cruise (Baseline) 8 CF6-80A Cruise/Idle GE90-110B1 Cruise PW4098 Cruise CF6-80C2B1 Cruise JT3D (Baseline) Cruise TayRB.183- Cruise 3Mk CF6-80E1A2 Cruise JT3D-3B Cruise V2527-A5 Cruise CF34-3A1 Cruise JT9D (Baseline) Cruise V2527E-A5 Cruise CF34-8C1 Cruise JT9D-3A Cruise CFM56-5B2 Cruise JT9D-7Q3 Cruise GP7270 Cruise PW2040 Cruise There are two other points on the flight envelope that can be modeled (Approach and Climb out). In principle the more points matched and the closer they are to data supplied on the FAA data sheets, the more reliable the predicted performance results will be. Although this might prove difficult to do for every engine, for every point on the flight envelope, with the maps at hand. 8 The baseline is the original engine, original core. It was used as a calibration for the future variants because temperatures and pressures of some baseline engines were available. In principle the engine name/code doesn t exist on any aircraft application but has an extension to the name. Doc nr: LR-FPP Revision: 1.0 Page 24/33

25 8.2 Verification/checks The losses in the nozzle inlet and the exhaust are assumed to be ideal as is the case with most of the rated TSFC in Jane s aircraft. Which is another influencing parameter which concerns why the net thrust is quite high when calculating with GSP. In reality it will be smaller. Deviating results in design point matching can be attributed with the ideal assumption of the nozzle loss coefficients ect. Deviating results in off-design point performance parameter matches such as the TSFC or fuel flow can be attributed to scaling errors and map design point selection Verification/checks Net thrust, TSFC and fuel flow can be matched at the off design points and can also be used to verify the design point data at T-O. This needs to be further refined by matching more off design points in the flight envelope and can be done with future work. 9 Improvements 9.1 GSP BUGS. The ambient conditions in some cases (OD Idle case) revert to altitude of 10660m when changed to sea level conditions. If this is experienced then the user should delete the case and create a new one. 10 Recommendations (Future Work) 10.1 Suggestions for further development: A few suggestions will be made for extensions to the already developed performance method. This method will be prone to many errors when it comes to estimating the values for the component parameters. If the model is to be simulated with a higher fidelity, the certainty or accuracy can only be estimated when more operating points "operating performance points of interest" in the flight envelope can be matched and only when you match data such as thrust and/or TSFC at these various points can we be sure that the chosen unknowns (ETAs) are with given certainty. For 24 engines this is unrealistic to do, as it will take some time. However needs to be done to ensure accuracy. For each engine the map design beta and map design rotor speed values can be changed for each engine on the same generic maps. This will shift the reference point on the map. Therefore each engine will have a different scaling factor to its design point. To improve results this point should be shifted for the individual engines when using the generic maps. The ICAO exhaust emissions databank [15] was used to confirm the BPR, OPR, engine type and the rated output at T-O. It is noted that the atmospheric conditions are slightly different here than to the models in GSP. This needs to be addressed if it is esteemed that emissions data can be derived from this sheet and cross checked with GSP in the future. The engineer is cautioned that the rated power setting of cruise isn t explicitly listed for every engine probable reasons listed above section 6.2. However this gives a margin of where the fuel flow might sit when considering the flight at cruise and thus the TSFC and thus the net Doc nr: LR-FPP Revision: 1.0 Page 25/33

26 thrust. The fuel flows given in GSP would be within these margins if the models were defined at these atmospheric conditions. Please correct the H/C and fuel temperature selection/values. 11 Conclusions An initial iteration on unmixed high-bypass civil turbofans performance was investigated with respect to FAA certification date. Parameters such as: Overall system performance Component efficiencies Gas Path temperatures and pressures Where plotted and trends investigated. The primary results were: Gas path parameters such as TIT have increased Component parameters of the HPC, HPT, LPT, LPC such as ETA_poly have increased System performance parameters such as ETA_th, ETA_prop, ETA_overall have increased. While the TSFC has decreased. Doc nr: LR-FPP Revision: 1.0 Page 26/33

27 12 Annexes: 1. PowerPlantsExhaust.mxl 2. Turbine_Powerplants10.xlsx 9 9 The Turbofan_Powerplants.xlsm file contains over 700 civil turbofan variants. There may be double ups of some variants due to the fact that the application of civil aircraft has different demands on certain parameters. Eg. Certification date. This is based on an FAA certification however there may be other certification dates. Doc nr: LR-FPP Revision: 1.0 Page 27/33

28 13 References: 1. Cumpsty, N., Jet Propulsion. 2003: Cambridge University Press. 2. Shakariyants, S.A., Generic Methods for Aero-Engine Exhaust Emission Prediction, in LR. 2008, TUD: Delft. 3. Forsthoffer, W.E., Benefits of trending compressor, generator and turbine performance, in International Turbomachinery. 2013, Tri-sen: USA. p Roux, E., Turbofan and Turbojet engines database handbook. 2007: Éditions Élodie Roux. 5. Various. Jane's Aero-Engines. Jane's Defense & Security Intelligence & Analysis 2013 [cited 2013; Available from: 6. Transportation, U.S.D.o. Certificate Data Sheets /06/2013]; Available from: ameset?openpage. 7. Svoboda, C., Turbofan engine database as a preliminary design tool. Pergamon, Aircraft Design, Khan, K. Jet Propulsion/Performance [cited 2013; Available from: 9. Roskam, J., Airplane Aerodynamics and Performance. 1997: Design, Analyis and Research Corporation 10. Visser, W., et al., Gas Turbine Simulation Program 2013, NLR and TUD: The Netherlands. 11. Avellan, R., On the Design of Energy Efficient Aero Engines Some Recent Innovations, in Applied Mechanics. 2011, Chalmers University of Technology: Goteborg, Sweden 12. Miller, H., Benefiting From Efficiency Improvements to Gas Compression, in The cost effectiveness of Technology. 2003: NY, Olean. 13. Cantwell, B. Aircraft and Rocket Propulsion AA283. Chapter 10: Summary Review and Future Trends 2013 [cited 2013 July]; Available from: Cantwell, B., Aircraft and Rocket Propulsion AA283, in The GE90- An Introduction 2013, Stanford: Standford University. 15. Agency, E.A.S. ICAO Aircraft Engine Emissions Databank 2013 [cited 2013; Available from: Davis, D.M., Performance Prediction and Simulation of Gas Turbine Engine Operation for Aircraft, Marine, Vehicular, and Power Generation, in NATO Technical Report Doc nr: LR-FPP Revision: 1.0 Page 28/33

29 14 Appendices: 14.1 Station Numbering (Turbofan Exhaust) The station numbering is the international one. If a booster compressor were installed on the LP shaft downstream of the fan then station 23 would be downstream of the booster stages. Figure 14-1 International Gas Turbine station numbering [1] Note that at stations 1 and 9 static pressure deviates from ambient pressure. At stations 0 and static pressure is equal to ambient pressure. Station 0 is always best to use to determine v0 since only aircraft airspeed must be specified or measured. Table 14-1 Station details Station Location 0 Far Upstream or Freestream 1 Inlet or diffuser entry 2 Inlet or diffuser exit, fan entry 13 Air side Fan Exit, duct entry 2.1 Core Side Fan exit, Low-Pressure compressor entry 2.3/2.5 Low-Pressure compressor exit High-Pressure compressor entry 3 High-Pressure compressor exit 3.1 Burner entry 4 Burner exit Nozzle vanes entry High-pressure turbine entry for pi_th definition Doc nr: LR-FPP Revision: 1.0 Page 29/33

30 4.1 Nozzle vanes exit 4.4 High-pressure turbine entry for tau_th definition 4.5/4.9 Low-pressure turbine entry 6A 5 Low-pressure turbine exit 6 Core stream mixer entry 16 Fan bypass stream mixer entry Mixer exit afterburner entry 7 core exhaust nozzle entry afterburner exit 8 exhaust nozzle throat 9 core exhaust nozzle exit 17 Bypass exhaust nozzle entry 18 Bypass exhaust nozzle throat entry 19 Bypass exhaust nozzle exit 14.2 Equations The thrust can be defined in the usual way and for an unmixed turbofan engine it is: F x = F + A 19 (P 19 P 0 ) + A 9 (P 9 P 0 ) = (W core + W fuel ) (V 9 V 0 ) + W bypass (V 19 V 0 ) + A 19 (P 19 P 0 ) + A 9 (P 9 P 0 ) Bypass Ratio The bypass ratio is defined as the mass flow of air passing outside the core divided by the mass flow through the core. Affects the appearance, size and weight of the engine. The RB211, CF6 and JT9D generation engines had BPR of around 5. Assume a BPR of at least this will be employed for future new Large Aircraft. [1], pg 69. BPR = m b/fan m c For civil engines a major requirement is low fuel consumption, equivalent to high overall efficiency (considering the energy available from the combustion process). Aerodynamic factor * thermal factor. η 0 = η p η th Or from Gasturb (another definition): Is the ratio of useful work done in overcoming the drag of the airplane to the energy content of the fuel. η 0 = F V 0 W f LHV Equation 14-1 Doc nr: LR-FPP Revision: 1.0 Page 30/33

31 Propulsive Efficiency: Propulsive efficiency is highest when jet velocity equals flight velocity; however, in this case thrust is zero. With a turbojet at subsonic flight conditions the exhaust jet velocity is very much higher than the flight velocity of the aircraft. The high kinetic energy which the exhaust jet has relative to the air is a loss and this results in poor propulsive efficiency and finally in high thrust specific fuel consumption even if all the component efficiencies are high. When propulsive efficiency is evaluated for an unmixed flow turbofan then the core stream and the bypass stream must be considered: η prop = P thrust P prop When the propulsive efficiency is evaluated for an unmixed flow turbofan then the core stream (index 9) and the bypass stream (index 19) must be considered. (Gasturb help: Propulsive Efficiency). The propulsive efficiency is (REF: Performance Prediction and Simulation of Gas turbines [16]) and [13]. For a choked nozzle. η pr = Power to vehicle TV 0 + [W core (V 9 V 0 ) 2 2 ΔKE air Power to vehicle TV 0 + W Fan (V 19 V 0 ) 2 ] 2 ΔKE fuel + [W fuel (V 9 V 0 ) 2 2 If both exhausts are fully expanded so that P 9 = P 0 and P 19 =P 0. Then the propulsive efficiency becomes: W fuel (V 0) 2 2 ] η p = (W fan (V 19 V 0 ) + W air (V 9 V 0 ) + W fuel V 9 )V 0 V 2 2 W 9 V 0 V 2 2 Core + W 19 V 0 2 Air W 2 fuelv 9 Please note that: The effective jet velocity, is defined as the jet velocity which, would give the same thrust as the actual case, if it were fully expanded. F = m a(v e V 0 ) + A e (p e p a ) = m a(v je V 0 ) The Equation for effective jet velocity is thus V je = V e + A e(p e p a ) m a Thus using the notation in the document V eff,9 = V 9 + A 9(p 9 p 0 ) m c V eff,19 = V 19 + A 19(p 19 p 0 ) m air Doc nr: LR-FPP Revision: 1.0 Page 31/33

32 The fuel mass flow is much less than the air mass flow, m f m a and is usually neglected in the above Thermal Efficiency: The thermal efficiency characterizes the net energy output extracted (shaft work) from the engine divided by the available thermal energy (fuel). OR is defined as increase of the kinetic energy of the gas stream passing through the engine by the amount of heat employed η th = w out Q in η th = thermal efficiecny w out = net power output of engine Q in = rate of heat input η th = w net m fuel LHV For an ideal turbofan in terms of temperatures and pressures the thermal efficiency can be written as: Brayton Cycle for Power Gneration: the Ideal joule The brayton cycle efficiency is: η B = 1 ( 1 π γ 1 γ ) With the ideal Joule process, and assuming constant isentropic exponent, γ, the thermal efficiency of the cycle is only dependent on cycle pressure ratio P3/P2 and does not change with T4. Highest pressure ratio yields the best thermal efficiency. We must be careful when comparing cycles/configurations to their theoretical max. For power generation purposes it is ok to compare to the brayton cycle efficiency. What happens when we switch to thrust generation? It may be more appropriate to compare to the form incorporating the flight velocity. Although the Brayton is still the CYCLE max! Brayton Cycle for Jet Propulsion: the Ideal Turbofan The ideal turbofan thermal efficiency is: Or 1 η th idealturbofan = 1 ( ) τ r τ c η th idealturbofan = 1 ( Q rejected during the cycle Q input during the cycle ) Doc nr: LR-FPP Revision: 1.0 Page 32/33

33 Notice that for the ideal turbofan with the heat rejected by the fan stream is zero. Therefore the thermal efficiency of the ideal turbofan is independent of the parameters of the fan stream. 1 η th idealturbofan = 1 ( T t0 TT3 ) T 0 TT2 Or more commonly expressed adiabatic-isentropic form in terms of the PR η th idealturbofan = 1 ( 1 T t0 T 0 [ P t3 P t2 ] γ 1) γ 1 η th idealturbofan = 1 (1 + γ 1 γ 1 2 (M 2 ))π γ Note: A little note on the Brayton cycle efficiency and the ideal turbofan efficiency: In the ideal cycle approximation, if the Mach number is very small the thermal and Brayton efficiencies are identical. Equivalently, if the stagnation temperature of the free stream entering the inlet is the same as the stagnation temperature well upstream of the system then you are most likely stationary. Note that, characteristically for a Brayton process, the thermal efficiency is determined entirely by the inlet compression process. The ideal turbofan efficiency represents the best we can do at this Mach number. In fact the final design is what we would call the ideal turbofan. The ideal cycle will be the basis for comparison with other engine cycles but it is not a practically useful design. The non-ideal turbofan efficiency is (For a choked nozzle): For a turbofan with a core and bypass stream thermal efficiency is: η th = Power to vehicle TV 0 + [m core (V 9 V 0 ) 2 2 ΔKE air + m Fan (V 19 V 0 ) 2 ] 2 m f Hv f ΔKE fuel + [m fuel (V 9 V 0 ) 2 2 m fuel (V 0) 2 2 ] If both exhausts are fully expanded so that P 9 = P 0 and P 19 =P 0. Then the thermal efficiency becomes: η p = W Core V 9 2 V V + W V 0 Air W 2 fuelv 9 m f Hv f The fan s isentropic efficiency is η isen fan = T 13is T 02 T 13 T 02 Doc nr: LR-FPP Revision: 1.0 Page 33/33

34 The performance of each component is defined in terms of the stagnation pressure and the temperature entering and leaving the component. A widely accepted notation is: M 0 = flight mach number Doc nr: LR-FPP Revision: 1.0 Page 34/33

35 Station 19- The fan nozzle exit. The temperature/pressure parameters across the fan nozzle are τ 1n = T t19 T t17 π = P t19 P t17 Doc nr: LR-FPP Revision: 1.0 Page 35/33

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