Aircraft Design 3 (2000) 17 } 31
Turbofan engine database as a preliminary design tool Charlie Svoboda* Department of Aerospace Engineering, The University of Kansas, 2004 Learned Hall, Lawrence, KS 66045, USA
Abstract
A large database of currently manufactured turbofan engines with a bypass ratio of at least 2.0 was compiled compiled in 1996. Key parameters parameters (dry weight, length, fan diameter, diameter, nacelle diameter, cruise thrust, air mass #ow, bypass ratio, total pressure ratio, take-o ! speci speci"c fuel consumption, and cruise speci "c fuel consumption) were plotted, most as a function of take-o ! thrust. The resulting plots are a rich source of basic information, which can be used to quickly de "ne an engine for use in a preliminary airplane design. The database is sorted by take-o ! thrust and can also be used to determine if an existing engine can be used in the proposed airplane. Relationships are suggested for use in preliminary design. 2000 Elsevier Science Science Ltd. All rights reserved.
1. Intro Introduct duction ion
While involved in the viability of re-engining a 747-400 with a di ! erential erential thrust application similar simil ar to the 3-X Jet Jet Concept [1], it was was necessary necessary to estimate estimate some basic basic engine engine parameters parameters for an engine with 108,000 108,000 lb of take-o! thrust [2]. At At that time, time, no turbofan turbofan engines engines were being produced produced in that thrust class. It was decided that data from existing engines should be examined for trends that could be used to provide reasonable estimates of basic parameters for a 115,600 lb take-o! thrust engine.
2. Methods Methods and results results
Engine data were assembled in a spreadsheet from Refs. [3}5] and sorted by take-o! thrust. thrust. It was decided to limit the scope of the engines to be surveyed to engines with a bypass ratio greater
* Corresponding author. Tel.: 001-785-864-4267.
E-mail address:
[email protected] (C. Svoboda).
1369-886 1369-8869/00 9/00/$ /$ - see front matter matter 2000 Elsevier Science Ltd. All rights reserved. PII: S 1 3 6 9 - 8 8 6 9 ( 9 9 ) 0 0 0 2 1 - X
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C. S voboda / Aircraft Design 3 (2000) 17 } 31
Nomenclature
¹
¹
mdot
P
P
Dia Dia Dia Leng =
=
SFC SFC M h
Desig Country Type Fan Compressor Combustion Turbine
take-o! thrust, lb cruise thrust, lb air mass #ow, lb/s bypass ratio, dimensionless fan pressure ratio, dimensionless Total pressure ratio, dimensionless diameter, in fan diameter, in total diameter, in length, in dry engine weight, lb nacelle weight, lb take-o! speci"c fuel consumption, lb/(lb3h) cruise speci"c fuel consumption, lb/(lb3h) cruise Mach number, dimensionless cruise pressure altitude, ft engine designation, dimensionless country of engine manufacture, dimensionless engine description, dimensionless fan description, dimensionless compressor description, dimensionless combustion description, dimensionless turbine description, dimensionless
than 2.0 since the engine of interest would be a high bypass ratio engine. The following relationships were plotted:
Dry Weight, = vs. Take-O! Thrust, ¹ , Length, Leng vs. Take-O! Thrust, ¹ , Fan Diameter, Dia , vs. Take-O! Thrust, ¹ , Nacelle Diameter, Dia , vs. Take-O! Thrust, ¹ , Cruise Thrust, ¹ , vs. Take-O! Thrust, ¹ , Air Mass Flow, mdot, vs. Take-O! Thrust, ¹ , Bypass Ratio, , vs. Take-O! Thrust, ¹ , Total Pressure Ratio, P , vs. Take-O! Thrust, ¹ , Take-O! Speci"c Fuel Consumption, SFC , vs. Take-O! Thrust, ¹ , Cruise Speci"c Fuel Consumption, SFC , vs. Take-O! Thrust, ¹ , Take-O! Speci"c Fuel Consumption, SFC , vs. Take-O! Thrust, ¹ .
The plots were examined to see if a rationale for the trends observed could be discerned. A total of 67 actual engines were included in the database and plotted in Figs. 1 }11. Engines 69 and 70 are possible future engine descriptions developed in the course of the previously mentioned
C. S voboda / Aircraft Design 3 (2000) 17 } 31
19
Fig. 1. Dry weight.
research and are not plotted. Engine 57 is an engine description based on engine 56, which was also used in the previously mentioned research. All engines are presented in Tables 1}3. Based on the data in Figs. 1}11 the following design trends are suggested: Dry weight Length
=
(lb)"250#0.175¹
Leng(in)"40#0.59 ¹
Fan diameter
Dia
Nacelle diameter Cruise thrust
Air mass # ow Bypass ratio
Take-O! SFC
(lb)"200#0.2¹
(&)"3.2#0.01 ¹ P
SFC
(3)
(lb),
(4)
(lb),
(5)
(lb),
Total pressure ratio
(lb),
(in)"5#0.39 ¹
mdot(lb/s)"0.032¹
(2)
(1)
(lb),
(in)"2#0.39 ¹
Dia
¹
(lb),
(6)
(lb),
(7)
(&)"11#0.082 ¹
(lb),
(lb/lb h)"0.49!0.0007 ¹
(lb),
(8) (9)
20
C. S voboda / Aircraft Design 3 (2000) 17 } 31
Fig. 2. Length.
Cruise SFC SFC (lb/lb h)"0.8!0.00096 ¹
Take-O! SFC
SFC
(lb),
(lb/lb h)"0.71!0.15 (&).
(10) (11)
3. Discussion and conclusions
Some of the engines in the database are derivatives of other engines in the database. No e! ort was made to eliminate these engines, nor were all derivative engines included. In Fig. 1, Dry Weight is plotted as a function of take-o! thrust. It must be understood that many currently produced engines have a de-rated take-o! thrust to extend engine life. In Fig. 1, this could cause some engines to appear to be above the trend. It is also helpful to remember that, while all the engines plotted are currently in production, all these engines are not recent designs. Engines designed 10 years ago may not be as e $cient as engines designed last year. It is also necessary to note that the bypass ratios of these engines vary from 2 to 9. This being said, the relationship between dry weight and take-o! thrust is fairly linear.
C. S voboda / Aircraft Design 3 (2000) 17 } 31
Fig. 3. Fan diameter.
Fig. 4. Nacelle diameter.
21
22
C. S voboda / Aircraft Design 3 (2000) 17 } 31
Fig. 5. Cruise thrust.
Fig. 6. Air mass #ow.
C. S voboda / Aircraft Design 3 (2000) 17 } 31
Fig. 7. Bypass ratio.
Fig. 8. Total pressure ratio.
23
24
C. S voboda / Aircraft Design 3 (2000) 17 } 31
Fig. 9. Take-o! speci"c fuel consumption.
Fig. 10. Cruise speci"c fuel consumption.
C. S voboda / Aircraft Design 3 (2000) 17 } 31
25
Fig. 11. Bypass ratio e! ect on take-o! speci"c fuel consumption.
Table 1 Engine parameters, Set 1
1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18 19 20 21
Designation (* )
Country (* )
Type (* )
SCF (lb/lbhr)
SCF (lb/lbhr)
M
h
(* )
(ft)
FJ44-1C FJ44-1A JT15D-5D AI-25 TFE731-3 PW545 TFE731-5 TFE731-5B PW300 TFE731-60 ATF3-6A PW306A CFE738 ALF502R-5 ALF502R-3A AE 3007 ALF 502L-2 DV-22 TF34-GE-100 TF34-GE-400A/B FJR710
USA USA Canada Ukraine USA Canada USA USA Canada USA USA Canada USA USA USA USA USA International USA USA Japan
3 2 2 2 2 2
0.456 0.475 0.550 0.570
0.750 0.758
0.70 0.70
30,000 36,090
0.795 0.835
0.48 0.80
19,685 40,000
0.771 0.756 0.675 0.679 0.830 0.679 0.640
0.80 0.80 0.80 0.80 0.80 0.80 0.80 0.70
40,000 40,000 40,000 40,000 40,000 40,000 40,000 25,000
0.680
0.75
25,000
shaft shaft shaft shaft shaft & geared front fan shaft
shaft & geared front fan shaft cross-compound shaft shaft
2 2 2 2
shaft geared shaft shaft geared shaft
2 shaft 2 shaft
0.436
2 shaft & geared front fan 2 3 2 2
0.503 0.394 0.372 0.408 0.408 0.330 0.428 0.370 0.370 0.363 0.374
26
C. S voboda / Aircraft Design 3 (2000) 17 } 31
Table 1 (Continued)
22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 37 38 39 40 41 42 43 44 45 46 47 48 49 50 51 52 53 54 55 56 57 58 59 60 61 62 63 64 65 66 67 68 69 70
Designation (* )
Country (* )
Type (* )
Tay 611 D-36 BR 710 Tay 651 D-436K PS-90A10 BR 715 CFM56-2B1 CFM56-3B2 D-30KU CFM56-7B26 PS-90A12 D-30KU-90 V2528-D5 V2500-A5 CFM56-5C4 PS-90A-76 535-C PW2037 CF-6D PW2040 D-100 TF39 535-E4 JT9D3A CFMXX CF6-80A3 RB211-524B D-18T F103-GE-101 PW4052 CF6-50C1 JT9D-59A JT9D-7R4H1 *CF6-80C2BIF Safety CF6-80C2DIF RB211-524H Trent 768 CF6-80E1A4 Trentc 775 GE90-76B Trent 875 PW4084 PW4090 GE90-92B Trent 890 De-Rated Cruise Cruise
UK Ukraine Germany UK Ukraine Russia Germany International International Russia International Russia Russia International International International Russia UK USA USA USA Russia USA UK USA International USA UK Ukraine USA USA USA USA USA USA CRS USA UK UK USA UK USA UK USA USA USA UK CRS CRS
2 3 2 2 3 2 2 2 2 2 2 2 2 2 2 2 2 3 2 2 2 2 2 3 2 2 2
shaft shaft shaft shaft shaft shaft shaft shaft shaft shaft shaft shaft shaft shaft shaft shaft shaft shaft shaft shaft shaft shaft shaft shaft shaft shaft shaft
3 shaft 2 2 2 2 2
shaft shaft shaft shaft shaft
2 3 3 2 3 2 3 2 2 2 3
shaft shaft shaft shaft shaft shaft shaft shaft shaft shaft shaft
SCF (lb/lbhr) 0.360
0.490
0.330 0.348 0.330 0.286 0.315
SCF (lb/lbhr)
M
h
(* )
(ft)
0.710 0.650 0.630 0.690 0.610 0.630 0.610 0.657 0.655 0.700
0.80 0.75 0.80 0.78 0.75 0.80 0.76 0.85 0.85 0.80 0.80 0.80 0.80 0.80 0.80 0.80 0.80 0.80 0.85 0.85 0.80 0.80
43,000 26,250 41,000 35,000 36,090 36,090 35,000 35,000 35,000 36,090 35,000 36,090 36,090 35,000 35,000 35,000 36,090 35,000 35,000 35,000 35,000 36,090
0.80 0.85 0.80 0.85 0.85 0.75
35,000 35,000 35,000 35,000 35,000 36,090
0.570 0.565
0.80 0.85 0.85 0.85 0.80 0.80 0.85 0.85 0.82
35,000 35,000 35,000 35,000 35,000 35,000 35,000 35,000 35,000
0.565
0.82
35,000
0.557 0.537
0.83 0.80
35,000 35,000
0.520 0.557 0.543 0.543
0.80 0.83 0.85 0.85
35,000 35,000 35,000 35,000
0.582 0.664 0.575 0.575 0.567 0.595 0.646 0.582 0.563 0.540 0.598 0.624
0.344 0.360 0.399 0.311 0.390
0.316 0.329 0.322
0.620 0.570 0.537 0.631 0.628 0.576 0.578
0.332
0.274
C. S voboda / Aircraft Design 3 (2000) 17 } 31
27
Table 2 Engine parameters, Set 2
1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18 19 20 21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 37 38 39 40 41 42 43 44 45
¹
¹
(lb)
(lb)
1500 1900 3045 3307 3700 3876 4500 4750 4750 5000 5440 5700 5725 6790 6970 7200 7500 8532 9065 9275 11,243 13,850 14,330 14,800 15,400 18,078 20,283 22,000 22,000 22,000 24,250 26,400 26,455 26,455 28,000 30,000 34,000 35,275 37,400 38,350 40,000 40,900 41,887 43,000 43,100
600 506 *
976 817 915 986 1052 1113 1120 1055 1320 1464 2250
mdot (lb/s) 63 63 75 100 118 *
140 143 180 * * *
*
*
*
*
* *
309 333 338
2976
*
*
410 562 435 426
*
3439 4343 3600 4969 5040 6063 5480 5071 6063 5752 5752 7100 7716 8453 6500 9120 *
8377 *
8700
210 *
3527 2300
162
*
*
*
582 636 784 683 593 783 816 540 848 1065 1036 1142 1210 1307 1340 1581 1541 1150
P
P
(* )
(* )
(* )
3.28 3.28 3.30 2.10 2.80 4.00 3.34 3.48 4.50 4.40 2.80 4.50 5.30 5.70 5.71 5.00 5.00 5.00 6.42 6.20 6.50 3.04 5.60
*
*
1.50 1.70 *
* * * * * *
*
1.70 * * * *
Dia (in)
12.80 12.80 10.00 9.60 14.60
Dia (in) 19.7 19.7
* *
*
*
14.40 14.60 23.00 14.60 21.30 12.70 23.00 12.20 11.60 23.00 13.60
* *
28.2 27.3
29.7
30.7 31.7
*
*
*
*
41.7
*
*
38.5 *
*
*
1.50 1.50
20.00 21.00
*
*
*
49.0 52.0 49.0
*
*
3.07 6.20 3.76
*
*
*
6.00 4.90 2.42 5.10 5.05 2.44 4.70 4.60 6.40 4.50 4.40 6.00 4.40 6.00 8.10 8.00 4.30
*
* *
* * * * * *
1.70 * * * * *
1.70 1.42 * *
15.80 20.00 26.00 16.60 21.00 23.10 32.00 30.50 28.80 20.00 32.60 25.30 35.02 30.00 29.40 38.30 36.40 21.10 31.80 30.40 27.60 40.75 22.00 25.80
44.0 *
* *
48.0 44.8
54.1
*
*
55.1 58.0 68.3 60.0 57.3 61.0 65.8 57.3
* * * * *
63.5 63.0
*
*
72.3 74.8 73.9 78.5 86.4 78.5 95.9 96.2 74.1
* * * * * * * *
33.8
91.1
899
36.7
82.3 102.3 75.6 99.0
929 1125 1043 1325
56.8 106.5 58.6
1336 1581 1311 1543 1440 1478 2160 3135 2445 3600 3380 3197 4180 4660 4671 4301 5110 5216 5071 5291 5400 5200 4995 6503 7294 7196 10,155 7300 7716 7900 7189
38.2 43.0
43.5
*
*
*
34.2 32.0
*
54.1
*
(lb)
445 447 632 705 754 765
*
33.9 36.5
41.7
*
(lb)
*
*
*
=
*
*
=
41.9 40.3 61.0 78.5 59.8 68.0
*
* *
21.7 20.9
*
28.0 32.3 *
Dia Leng (in) (in)
52.0
62.0
61.4
61.4
*
84.8 94.1 84.8 102.5
100.0 100.0 93.0 94.7 136.6 134.0 94.7 136.6 168.5 142.0 95.7 93.0 224.0 98.7 188.8 224.4 126.0 126.0 103.0 195.4 118.5 146.8 188.0 146.8 100.0 117.9
1040
1495
7900 7500
28
C. S voboda / Aircraft Design 3 (2000) 17 } 31
Table 2 (Continued)
46 47 48 49 50 51 52 53 54 55 56 57 58 59 60 61 62 63 64 65 66 67 68 69 70
¹
¹
(lb)
(lb)
mdot (lb/s)
43,600 45,000 50,000 50,000 51,660 51,711 52,500 52,500 53,000 56,000 57,160 57,898 60,090 60,600 67,500 70,000 75,150 76,400 77,900 84,600 90,000 90,200 91,300 88,682 115,600
10,200
1495
*
10,477 11,000 10,716 *
9400 10,800 11,950 12,250 11,330 12,042 11,330 11,813 11,500 *
11,500 17,500 13,000 13,965 *
18,400 13,000 24,084 24,084
*
1460 1513 1687 1476 1700 1484 1640 1695 1769 1730 1769 1604 1932 1926 *
3000 2482 2558 2550 3221 2720 3759 3759
P
P
(* )
(* )
(* )
5.17
*
*
*
4.60 4.50 5.60 4.31 5.00 4.40 4.90 4.80 5.06 5.05 5.05 4.30
* *
* * * * * *
*
1.70 * *
*
*
5.30
*
*
*
Dia (in)
Dia (in)
21.50
95.6
*
*
28.40 28.40 27.50 30.20 27.50 30.40 24.50 26.70 29.90 30.40 31.80 33.00
86.4
*
*
84.8
*
109.9 86.4
*
93.6 86.4
* *
97.0 97.0 106.0
* *
96.3
106.0
*
*
*
*
34.60
*
*
*
*
39.30
*
*
*
*
*
4.85 6.41 9.00 5.75 8.02 8.02
1.70
30.00 34.40 45.50 42.80 30.40 30.40
*
* *
1.70 1.63 1.63
97.2 94.1
*
*
* * * * *
=
=
(lb)
(lb)
128.2
8608
157.4 119.4 212.6 173.0 153.6 183.0 132.2 132.7 168.0 193.0 168.0 125.0 154.0 173.5 154.0 193.0 172.0 191.7 191.7 193.0 172.0 193.0 193.0
8420 9195 9039 8768
84.0
*
8.40
Dia Leng (in) (in)
86.3 97.4 96.0 97.4 123.0 110.0 93.6 93.6 123.0 110.0 142.0 142.0
112.5
110.0 134.0 120.0 120.0 134.0 165.8 165.8
10,842 9140 8885 9499 9500 9874 10,550 11,189 10,550 13,333 13,965 16,664 13,333 19,000 19,000
Table 3 Engine parameters, Set 3
Fan (* )
Compressor (* )
Combustion (* )
Turbine (* )
1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16
1S 1S 1S 3S 1S 1S
1S 1S 1S 8S 4S 2S
Annular radial out#ow Annular radial out#ow Annular reverse #ow Annular Annular, reverse #ow Annular, folded reverse #ow
1S HP axial, 2S LP axial 1S HP axial, 2S LP axial 1S HP, 2S LP 1S HP, 2S LP 1S HP, 3S LP 1S HP, 3S LP
axial axial axial axial axial IBR
LP axial, 1S HP centrigufal LP axial, 1S HP centrigufal centrifugal LP, 1S HP cent, 1S axial
1S axial
4S LP, 1S HP
Annular, reverse #ow
1S HP, 3S LP
1S axial 1S 1S, overhung 1S
4S LP, 1S HP 5S axial, 1S centrifugal 4S axial, 1S centrifugal 5S axial, 1S centrifugal
Annular, reverse #ow Annular Annular Annular
1S HP, 3S LP 1S H, 3S I, 2S L 2S axial, HP, 3S axial 2S HP, 3S LP
2S 1S
7S axial, 1S centrifugal 14S axial
Annular Annular
2S HP, 2S LP 2S HP axial, 2S LP axial
C. S voboda / Aircraft Design 3 (2000) 17 } 31
29
Table 3 (Continued)
Fan (* )
Compressor (* )
Combustion (* )
Turbine (* )
17 18 19 20 21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 37 38 39 40 41 42 43 44 45 46 47 48 49 50 51 52 53 54 55 56 57 58 59 60 61 62 63 64 65 66 67 68 69 70
2S
7S axial, 1S centrifugal
Annular
2S HP, 2S LP
1S 1S 1S 1S 1S 1S 1S 1S 1S axial 1S axial 1S 3S 1S axial 1S 3S 1S 1S 1S axial 1S 1S 1S 1S 1S 1S 1.5S 1S 1S
14S axial 14S axial 12S 3S LP, 12S HP 6S IP, 7S HP 10-S 3S LP, 12S HP 6S IP, 7S HP 12S HP 10S 3S axial LP, 9S HP 3S axial LP, 9S HP 11S HP 3S axial LP, 9S HP 12S HP 13S H
Annular Annular Smokeless annular Tubo-annular Annular Annular Tubo-annular Annular Annular
Annular Can-annular
2S HP, 4S LP 2S HP, 4S LP 2S H, 4S L 2S HP, 3S LP 1S HP, 1S IP 2S HP, 2S LP 2S HP, 3S LP 1S HP, 1S IP, 4S LP 2S HP, 2S LP 3S LP 1S HP, 4S LP 1S HP, 4S LP 2S H, 4S L 1S HP, 4S LP 2S HP, 3S LP 2S HP, 3S LP
4S LP, 1-S HP 4S axial LP, 9S HP 2S LP, 13S HP 6S IP, 6S HP 4S LP, 12S HP 1S H, 16L 4S LP, 12S HP 4S L, 12S H 16S axial 6S IP, 6S HP 3S L, 11S H
Annular Dual annular Can-annular Annular Annular Annular Annular Fully annular Annular Annular Annular
2S 1S 2S 1S 2S 2S 2S 2S 2S 1S 2S
1S, 4S
3S LP, 14S HP
Annular
2S HP, 4S LP
1S
7S IP, 7S HP
Annular
1S-HP, 1S IP, 4S LP
1S 1S 1S 1S 1S, 4S
4S LP, 11S HP 1S H, 16L 4S LP, 14S HP
Annular Annular Annular Annular Annular
2S 2S 2S 2S 2S
1S, 4S 1S, overhung 1S 1S, 4S 1S 1S 1S 1S 1S 1S 1S
4S LP, 14S HP 7S IP, 6S HP 8S IP, 6S HP 4S LP, 14S HP 8S IP, 6S HP 3S LP, 10S HP 8S IP, 6S HP 4S LP, 11S HP 4S LP, 11S HP 3S LP, 10S HP 8S IP, 6S HP
Annular Annular
2S HP, 5S LP 1S-HP, 1S IP, 3S LP
Annular
2S HP, 5S LP
Double annular
2S HP, 6S LP
Annular Annular Double annular Annular
2S 2S 2S 1S
Can-annular
HP, 5 LP HP, 5S LP HP, 4S LP HP, 1S IP, 3S LP HP, 5S LP HP, 5S LP HP, 5S LP H, 6S L HP, 6S LP HP, 1S IP, 3S LP HP, 4S LP
HP, HP, HP, HP, HP,
HP, HP, HP, HP,
4S 5S 4S 4S 5S
4S 4S 6S 1S
LP LP LP LP LP
LP LP LP IP, 5S LP
30
C. S voboda / Aircraft Design 3 (2000) 17 } 31
In Fig. 2, Length is plotted as a function of take-o ! thrust. For engines with take-o! thrust less than 10,000 pounds, length increases signi"cantly with take-o! thrust. When the take-o! thrust is above 10,000 pounds, the length increases slowly with take-o! thrust. Length should not ever exceed 225 in. Length is determined by the number of compressor and turbine stages. The number of stages is an arbitrary design choice. In Fig. 3, Fan Diameter is plotted as a function of take-o ! thrust. The trend appears to be parabolic. This "ts the roughly linear relationship between air mass #ow and fan area that is required if the exit velocities of the engines being examined are similar. Fan diameter is a function of bypass ratio, which is an arbitrary design choice. A similar parabolic trend exists in the relationship between Nacelle diameter and take-o! thrust as seen in Fig. 4. In Fig. 5, cruise thrust is plotted as a function of take-o ! thrust. This relationship would be expected to be completely linear if none of the engines had a de-rated take-o! thrust and if all the engines had the same cruise altitude and velocity. Even with these variations, the relationship is reasonably linear. In Fig. 6, air mass # ow is plotted as a function of take-o! thrust. Since the fan exit velocities and turbine exit velocities for most engines with a bypass ratio greater than 2 are similar, it is expected that take-o! thrust of an engine would be proportional to the air mass #ow of the engine. This is con"rmed by the linear relationship of air mass #ow to take-o! thrust shown in Fig. 6. In Fig. 7, the bypass ratio is plotted as a function of take-o! thrust. It is notable that there is a lot of scatter. Bypass ratio is a strong function of the type of application and not just the take-o ! thrust. Even for similar applications, the bypass ratio is an arbitrary design choice. It is also true that higher bypass ratio engines are a relatively recent phenomenon and that many older, lower bypass ratio engines are still in production. In Fig. 8, total pressure ratio is plotted as a function of take-o! thrust. Taking into account the large amount of scatter, there appears to be a parabolic relationship between the total pressure ratio and take-o! thrust. In engine design, a higher total pressure ratio usually yields to a lower speci"c fuel consumption, but also leads to more stages, more weight, more length, more complexity and a higher turbine temperature. These issues are often traded o! against each other, which could easily explain the high level of scatter for a given take-o ! thrust. In Fig. 9, take-o! speci"c fuel consumption is plotted as a function of take-o! thrust. Except for some scatter, probably due to di! erences in bypass ratio, the relationship appears to be parabolic. In Fig. 10, cruise speci"c fuel consumption is plotted as a function of take-o! thrust. This relationship also appears to be parabolic, with less scatter than the plot of take-o! speci"c fuel consumption. In Fig. 11, take-o! speci"c fuel consumption is plotted as a function of bypass ratio. This relationship also appears to be parabolic.
4. Summary
The "gures and formulas presented in this article are useful for making quick and reasonable estimates of various engine parameters when initially de"ning an engine for a new application. The major issues which need to be understood when these " gures are viewed are: (1) de-rating of engine
C. S voboda / Aircraft Design 3 (2000) 17 } 31
31
take-o! thrust, (2) age of engine design, and (3) variability of bypass ratio. Based on the data in Figs. 1}11 these design trends are suggested [see Eqs. (1)}(11)]. Afterword
The author is seeking missing data, especially nacelle weights, for the engines currently included in the current database. If you would like to help with that task, please contact me at Department of Aerospace Engineering, 2004 Learned Hall, Lawrence, KS 66045. The phone is (785) 864-4267. My current e-mail address is
[email protected].
References [1] Jones V. Comparative study of a 3X and conventional twin engine installation on a typical medium sized business jet, Master's Thesis, University of Kansas, Lawrence, KS, 1995. [2] Svoboda C. Comparative study of a 6X and a standard four engine installation on the boeing 747-400. Master 's Thesis, 2004 Learned Hall, University of Kansas, Lawrence, Kansas, 1998. [3] Jackson P. Jane's aero engines Aircraft. Jane 's Publishing Company, London, 1996 }1997. [4] Taylor, JWR. Jane's all the world aircraft. Jane 's Publishing Company, London, 1995 }1996. [5] Mattingly JD. Elements of gas turbine propulsion. New York: McGraw-Hill, 1996.
