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NASA TECHNICAL MEMORANDUM
ANALYSIS OF TURBOFAN PROPULSION SYSTEM WEIGHT
AND DIMENSIGNS
Mark H. Waters and Edward T. Schairer
Ames Research Center Moffett Field, California 9403C
NASA TM X- 73,199
January 1977
RECEIVED - . N A C ~ csl cnrll lrv ' ' J
A - X - ANaLYSZS OF i U n 0 O E ~ h ti77-190b9 EhOYULSXCI S YSZEtl hbIGhL AND T I f i E L S I C h S (NASA) bS p HC hO*/BE A01 CSCL 21E
https://ntrs.nasa.gov/search.jsp?R=19770012125 2018-03-25T12:10:30+00:00Z
* F a Jr by tho National Tochnid lntornution h i c o , Springfirld, Virginic 22181
1. Ro@ort No. 2. Govmmnt k s a i m No. NASA TM X-73,199
4. Titk end Subtitla ANALYSIS OF TURROFAN PROPULSION SYSTEM WEIGHT AND DIMENSIONS
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Mark H. Waters and Edward T. Schairer
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Ames Research Center Moffett Field, Calif. 94035
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Natio a1 Aeronautics and Space Administration Washington, D. C. 20546
16. supplamauuy Nota
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6. R.pwt Date
6. ~artwm~nq O~ganizat~on Code
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Technical Memorandum
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An important factor often overlooked in aircraft preliminary design studies is the tradeoff chat exists in propulsion system weight and size. This is an even more important factor when there is an emphasis on reduced aircraft noise. Engines designed specifically for low noise may have signif- icant penalties in propulsion system weight and size. The objective of this paper is to summarize weight and dimensional relationships that can be used in aircraft preliminary design studies. These relationships must be rela- tively simple to prove useful to the preliminary designer, but they must be sufficiently detailed to provide meaningful design tradeoffs. All weight and dimensional relationships will be developed from data bases of existing and conceptual turbofan engines. The total propulsion system will be considered including both engine and nacelle, and all estimating relations will stem from physical principles, not statistical correlations.
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Propulsion weight Turbine engines Transport aircraft
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FPR
OPR
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Wff
SYMBOLS
area, m2 (it2)
bypass ratio
specific heat at constant pressure, joules BTU kg O K (lb OR)
diameter, m (ft)
fuel-air mass ratio
fan pressure ratio
acceleration of gravity, 9.9 m/sec2 (32.2 ft/sec2)
enthalpy change, J ! ( % ) hub diameter, m (ft)
constant, 778 BTU/ft-lb
constant
lengeh, m (it)
Mach number
number; rotor speed, rpm
constant
overall pressure ratio
total pressure, ~ / r n ~ (lb/f t 2,
thrust, N (lb); total temperature, OK (OR)
tip diameter, m (it)
velocity, m/sec (it /sec)
mass, kg (weight, lb)
air£ low, kglsec (lb/sec) kg 0~112 lb 0 ~ 1 1 2
mass flow function, N sec ( ) (see eq. (10)) lb sec
ratio of specific heats
see equation (A10)
iii
X d i f f u s e r ha l f ang le , deg
A turbine s tage loading parameter ( see eq. (Al ) )
ri i sentropic e f f i c i e n c y
Subscripts
C compressor, core cowl
CX core exhaust
E engine
F fan f a c e , fan
FX fan exhaust flowpath
H hub
HP T high-pressure turbine
I i n l e t
LPT low-pressure turbine
SLS sea l e v e l s t a t i c standard day
STG s tages
T t o t a l ; turbine
TIP rotor t i p
Superscript
- average value
ANALYSIS OF TURBOFAN PROPULSION SYSL'EM WEIGHT AND DIMENSIONS
Mark H. Waters and Edward T. Scha i re r
Ames Research Center
SUMMARY
A An important f a c t o r o f t e n overlooked i n a i r c r a f t pre l iminary des ign s t u d i e s i s t h e t r a d e o f f t h a t e x i s t s i n propuls ion system weight and s i z e . This i s an eve!; more important f a c t o r when t h e r e i s an emphasis on reduced a i r c r a f t noise . Engines designed s p e c i f i c a l l y f o r low n o i s e may have s i g n i f - i c a n t p e n a l t i e s i n propuls ion system weight and s i z e . The o b j e c t i v e of t h i s paper i s t o summarize weight and dimensional r e l a t i o n s h i p s t h a t can be used i n a i r c r a f t pre l iminary des ign s t u d i e s . These r e l a t i o n s h i p s must be r e l a - t i v e l y simple t o prove u s e f u l t o t h e p re l iminary des igner , but they must be s u f f i c i e n t l y d e t a i l e d t o provide meaniqgful des ign t r a d e o f f s . A l l weight and dimensional r e l a t i o n s h i p s w i l l be developed from d a t a bases of e x i s t i n g and conceptual turbofan engines. The t o t a l propuls ion system w i l l be considered inc lud ing both engine and n a c e l l e , and a l l e s t i m a t i n g r e l a t i o n s w i l l stem from phys ica l p r i n c i p l e s , not s t a t i s t i c a l c o r r e l a t i o n s .
INTRODUCTION
There a r e s e v e r a l a s p e c t s of the c y c l e s e l e c t i o n , des ign, and i n s t a l l a - t i o n of t h e turbofan propuls ion system t h a t should be considered i n t h e pre- l iminary des ign of t r a n s p o r t a i r c r a f t . This i s p a r t i c u l a r l y t r u e i f t h e a n a l y s t has t h e freedom t o s e l e c t the engine c y c l e and s i z e r a t h e r than deal - ing wi th e x i s t i n g turbofans t h a t a r e i n product ion o r under development.
Cruise f u e l consumption and engine weight a r e t h e most obvious parameters of i n t e r e s t . However, t h e geometry of t h e engine - f r o n t a l a r e a and l e n g t h - is a l s o important s i n c e i t has a d i r e c t bea r ing on t h e n a c e l l e design. Nacel le weight ( inc lud ing t h e t h r u s t r e v e r s e r ) and drag ( s p i l l a g e , cowl, f r i c t i o n , and b o a t t a i l ) a r e d i r e c t l y inf luenced by t h e engine geometry; u n f o r t u n a t e l y , t h e o v e r a l l propuls ion system weight t r a d e o f f s wi th engine f u e l economy and n a c e l l e drag a r e sometimes overlooked.
Recent s t u d i e s of commercial a i r t r a n s p o r t s have f r e q u e n t l y eva lua ted a i r c r a f t no i se . J e t no i se is b e s t c o n t r o l l e d by t h e c y c l e s e l e c t i o n t h a t reduces t h e core j e t v e l o c i t y . The a l t e r n a t i v e i s a mechanical suppressor . Eiore important f o r the h igher bypass r a t i o tu rbofan engines is f a n n o i s e , which is c o n t r o l l e d p r imar i ly by reducing t h e f a n t i p speed and/or i n s t a l l i n g a c o u s t i c t rea tment m a t e r i a l i n t h e i n l e t and fan exhaust duc t ing . A l l of these c y c l e and des ign a l t e r n a t i v e s a f f e c t both engine weight and geometry, and, i n the case of a c o u s t i c s p l i t t e r s i n t h e fan i n l e t and exhaus t , engine performance. The va r ious t r a d e o f f s between n o i s e , ~ n g i n e weight , and s i z e a r e d i f f i c u l t t o e v a l u a t e i n pre l iminary des ign and a r e f requen t ly overlooked.
F i n a l l y , t h e r e i s a n emers ing unde r s t and ing of t h e des ign t r a d e o f f s between e n g i n e performance and weight and eng ine maintenance r equ i r emen t s . r o r educe eng ine weight and s i z e f o r a g iven t h r u s t , t h e r e has been a long- s t a n d i n g e f f o r t t o i n c r e a s e t h e t u r b i n e i n l e t t empera tu re i n t u r b i n e e n g i n e s , which i n c r e a s e s s p e c i f i c t h r u s t ( t h r u s t l a i r f l o w ) . Shor t d u r a t i o n t a k e o f f r a t i n g s and f l a t r a t i n g s t o m a i n t a i n t a k e o f f t h r u s t on h o t days a r e common- p l a c e i n a l l modern t u r b o f a n eng ines . However, eng ine d e t e r i o r a t i o n can be t r a c e d t o t h e number of h igh t empera tu re c y c l e s expe r i enced by t h e e n g i n e ; from t h e s t a n d p o i n t of l i f e t i m e a i r l i n e c o s t s , i t may be d e s i r a b l e t o s a c r i - f i c e eng ine peak t h r u s t performance t o r educe t h e number of eng ine removals f o r maintenance. Any r e d u c t i o n i n s p e c i f i c t h r u s t , of c o u r s e , i n c r e a s e s I
weight and s i z e . T h i s problem is d i s c u s s e d i n r e f e r e n c e 1, bu t i n t h e o p i n i o n of t h e a u t h o r s , c o n s i d e r a b l y more e v a l u a t i o n of modern t u r b o f a n eng ine l i f e - t ime maintenance r equ i r emen t s i s needed b e f o r e t h i s t r a d e o f f can be unders tood and a p p l i e d i n p r e l i m i n a r y d e s i g n s t u d i e s .
The o b j e c t i v e of t h i s paper i s t o summarize t u r b o f a n p r o p u l s i o n sys tem weight and d imens iona l r e l a t i o n s h i p s t h a t can be used i n p r e l i m i n a r y t r a n s p o r t a i r c r a f t d e s i g n s t u d i e s . The t v p e of p r o p u l s i o n sys tem cons ide red i s t h e modern t u r b o f a n eng ine having a x i a l f low compressors and i n s t a l l e d i n pod-type n a c e l l e s . The t o t a l p r o p u l s i o n sys tem is c o n s i d e r e d , i n c l u d i n g bo th eng ine and n a c e l l e . A l l e s t i m a t i n g r e l a t i o n s s tem from p h y s i c a l p r i n c i p l e s , n o t s t a t i s t i c a l c o r r e l a t i o n s .
DATA SOURCES
The p r o d u c t i o n e n g i n e s used i n t h i s s t u d y a r e l i s t e d i n t a b l e 1, ? l o n g w i t h t h e p e r t i n e n t weight and geome t r i c d a t a used f o r each. A s t u d y of t h i s n a t u r e depends g r e a t l y on a c c u r a t e d a t a s o u r c e s ; t h u s , r e f e r e n c e s 2-19 a r e no t ed throughout t h i s t a b l e t o i d e n t i f y o u r s o u r c e s . Two c o n c e p t u a l e n g i n e s , t h e P r a t t and Whitney STF 477 and an undes igna t ed Genera l E l e c t r i c e n g i n e , bo th r e c e n t l y completed i n s t u d i e s f o r NASA ( r e f s . 20 and 21 ) , have been inc luded i n t h i s s t u d y . The d a t a f o r t h e s e e n g i n e s a r e a l s o noted i n t a b l e 1.
P ropu l s ion sys tem d a t a from t h e NASA-sponsored Qu ie t Clean Shor t Haul Exper imenta l Engine Study Program (QCSEE) have been used e x t e n s i v e l y i n t h i s s t u d y . I n t a s k 1 of t h e QCSEE s t u d y program, one of t h e c o n t r a c t o r s , D e t r o i t D i e s e l A l l i s o n , D iv i s ion of Gene ra l Motors , developed a broad d a t a base of c a n d i d a t e eng ine performance, w e i g h t s , d imens ions , n o i s e , and c o s t s . Tab le 2 lists t h e c o n c e p t u a l eng ines and t h e d i s t i n g u i s h i n g c y c l e pa rame te r s f o r each . The A l l i s o n d e s i g n ~ c i o n s f o r t h e e n g i n e s have been p re se rved . More d e t a i l s on t h e s e s t u d y e n g i r z s can be found i n r e f e r e n c e 22.
ENGINE DRY WEIGHT
S p e c i f i c Weight
The most common measure of b a s i c engine weight ( f r e q u e n t l y r e f e r r e d t o a s dry weight) f-s the engine s p e c i f i c weight. This parameter i s de f ined a s the engine dry weight d ivided by t h e maximum r a t e d t h r u s t a t sea - l eve l s t a t i c cond i t ions . The obvious advantage of t h i s parameter is i ts s i m p l i c i t y - weight is t i e d d i r e c t l y t o t h e r a t e d t h r u s t of t h e engine. However, t h e
b engine flow passages a r e s i z e d by dir flow and t h e thermodynamics i n t h e engine; thus , s p e c i f i c weight is c l e a r l y coupled t o t h e engine c y c l e through the s p e c i f i c t h r u s t parameter.
I n a d d i t i o n t o t h i s connection wi th t h e c y c l e , s p e c i f i c weight tends t o s c a l e wi th engine s i z e . An elementary a n a l y s i s h i g h l i g h t s t h i s p o i n t : assume engine weight W s c a l e s w i t h t h e cube of t h e c h a r a c t e r i s t i c dimension D, and t h r u s t T s c a l e s w i t h t h e square of D.
Thus, s p e c i f i c weight s c a l e s wi th the square r o o t of r a t e d t h r u s t . This is commonly r e f e r r e d t o a s t h e " t h r e e halves" law.
A more genera l a n a l y s i s fcl.lows:
Obviously i f n = 2 , s p e c i f i c weight is independent of t h r u s t .
The d a t a i n f i g u r e 1 a r e s p e c i f i c weight va lues f o r commercial t u r b o j e t engines (bypass r a t i o = 0) . Also shown a r e a family of s t r a i g h t l i n e s which, on t h e log-log p l o t , correspond t o va lues of t h e exponent n from 2.0 t o 3.0. (Note t h a t the l i n e f o r n = 3 .0 p l o t s very n i c e l y through these d a t a . ) Thus, t u r b o j e t s appear t o fo l low t h e three-halves law c l o s e l y , a t l e a s t over the t h r u s t range shown.
More recen t commercial turbofan engines a r e p l o t t e d log-log on f i g u r e 2 . These d a t a a r e grouped i n t o low bypass r a t i o tu rbofans (bypass r a t i o l e s s than 2 . 0 ) and high bypass r a t i o (bypass r a t l o g r e a t e r than 4.0 but l e s s than 8 .0 ) . Althoupl~ t h e r e is s c a t t e r i n t h e d a t a , a r educ t ion i n t h e s c a l i n g s e n s i t i v i t y wi th h igher bypass r a t i o turbofans is obvious. From t h i s i t appears t h a t the s p e c i f i c weight f o r t h e h igh bypass r a t i o tu rbofans i s inde- pendent of t h r u s t s i z e , aga in over t h e t h r u s t range shown.
b
Figure 3 i s a c r o s s p l o t of f i g u r e 2, which approximates t h e s c a l i n g exponent n wi th the engine bypass r a t i o a t sea - l eve l s t a t i c cond i t ions . The band is shown r a t h e r than a s i n g l e curve , but t h e lower curve seems t h e a
most r e a l i s t i c . No e x t r a p o l a t i o n below n = 2 is shown, but it is conceiv- a b l e t h a t very high bypass r a t i o turbofans could s c a l e t o reduced s p e c i f i c weight i n the range of t h r u s t r a t i n g s given.
Also shown on f i g u r e 3 a r e va lues of minimum t h r u s t f o r s c a l i n g purposes. I t i s obvious t h a t t h i s simple s c a l e e f f e c t w i l l not cont inue t o be v a l i d f o r engines of smal le r and smal le r sea - l eve l t h r u s t l e v e l s . There a r e many pos- s i b l e reasons f o r t h i s t r end . For example: t h e des ign of p a r t s f o r smal l engines is a f f e c t e d by minimum gauge requirements of t h e m a t e r i a l s ; if b lade weights no longer s c a l e i n p ropor t ion , then hub-tip r a t i o i n c r e a s e s t o keep hub s t r e s s e s t o l e r a b l e , thus dec reas ing t h e weight flow per u n i t a r e z and i n c r e a s i n g s p e c i f i c weight; accessory systems may not s c a l e a s t h e gas gener- a t o r and thus become a l a r g e r percentage of t h e weight f o r smal l engines ; compressor and turbi.ie e f f i c i e n c i e s dec rease due t o t h e boundary l a y e r prob- lems i n smal l passages. For t h e s e and f o r o t h e r reasons , one can p r e d i c t t h a t t h e r e w i l l be some engine s c a l e s i z e where t h e engine s p e c i f i c weight i s a minimum.
Thus, t h e two curves on f i g u r e s 2 and 3 form the b a s i s of a simple model t o p r e d i c t and s c a l e turbofan engine s p e c i f i c t h r u s t . Assuming t h a t t h e tech- nology l e v e l of t h e engine is approximately t h e same, and t h e c y c l e does noi go t o bypass r a t i o s g r e a t e r than those shown, t h i s model may be adequate f o r p re l iminary des ign purposes. However, f o r advanced conceDt s t u d i e s , a more d e t a i l e d eva lua t ion of t h e engine dry weight i s d e s i r a b l e .
Reference 23, an e x c e l l e n t e v a l u a t i o n of turbofan engine weight , demon- s t r a t e s how t h e a i r f l o w can be used t o e s t i m a t e engine weight. This i s q u i t e l o g i c a l s i n c e component s i z e s a r e keyed t o t h e des ign po in t a i r f l o w . The r a t i o of t h r u s t t o weight i s then computed us ing s t andard cyc le a n a l y s i s t o compute s p e c i f i c t h r u s t :
Sca l ing of Engine k ? i g h t With Airflow
The parametr ic engine d a t a developed by Al l i son i n t h e QCSEE s tudy a f f o r d a unique oppor tun i ty t o e v a l u a t e turbofan engine weight t r e n d s with a i r f l o w because t h e r e i s a r e l a t i v e l y l a r g e d a t a sample f o r d i f f e r e n t c y c l e s and t h r u s t l e v e l s . Also, the sample is c o n s i s t e n t i n terms of the des ign
assumptions and weight accounting. To keep t h e weight e s t ima t ing model versa- t i l e and y e t simple enough t o be usab le , t h e turbofan dry weight i s broken i n t o t h r e e components:
Fun weight - which inc ludes both t h e fan s e c t i o n and t h e low-pressure o r f an t u r b i n e s e c t i c n . The fan s e c t i o n inc ludes t h e fan r o t o r , f an housing, and forward suppor t . The low-pressure t u r b i n e s e c t i o n inc ludes t h e t r a n s i t i o n duc t ing between t h e high- and low-pressure t u r b i n e s , low- p r e s s u r e t u r b i n e r o t o r , low-pressure t u r b i n e c a s e and vanes, and r e a r bear ing support . Although none of t h e s tudy engines had boos te r s t a g e s on t h e low r o t o r spoo l , these would b e included w i t h t h e f a n system weight.
Core weight - which inc ludes t h e compressor r o t o r , c a s e and vanes, combustor-diffuser and burner , h igh-pressure t u r b i n e r o t o r , c a s e and vanes.
Remaining weight - which inc ludes t h e f u e l system, l u b r i c a t i o n system, i g n i t i o n system, b leed a i r system, c o n t r o l s , accessory gearbox, f an gear ing assembly and mounting system, and engine b u i l d i n g i tems such a s t h e cons tan t speed d r i v e l i m i t , h y d r a u l i c pump, s t a r t e r , g e n e r a t o r , e t c .
Each of t h e s e t h r e e c a t e g o r i e s i s c o r r e l a t e d wi th a i r f l o w with the i n t e n t i o n of developing a l o g i c a l way of b u i l d i n g up an e s t i m a t e of engine weight based on engineer ing p r i n c i p l e s .
Figure 4 c o r r e l a t e s f an weight wi th t h e t o t a l a i r f l o w pumped through t h e fan . The symbols on t h e f i g u r e r e l a t e t o t h e s tudy engines t a b u l a t e d i n t a b l e 2. Three t h r u s t r a t i n g s a r e included f o r each engine cyc le : 35,584, 88,960, 133,440 N (8,000, 20,000, and 30,000 l b ) sea - l eve l s t a t i c s t andard day t h r u s t . The c o r r e l a t i o n s inc lude f a n t i p speed and f i r s t s t a g e hub-tip r a t i o , which a r e both f a c t o r s i n t h e s t r e s s i n t h e f a n b lade r o o t c r e a t e d by c e n t r i f u g a l fo rce . Also included i n t h e c o r r e l a t i o n a r e t h e number of s t a g e s i n the f a n p l u s t h e low-pressure t u r b i n e . (Low-pressure booster s t a g e s would a l s o be counted he re . ) Figure 4 i s a log-log p l o t and c a r e has been taken t o p l o t a s t r a i g h t l i n e through each s e t of t h r e e po in t s . The equat ion f o r t h e s e l i n e s is noted on t h e f i g u r e .
The c o r r e l a t i o n appears reasonably good over a range of a i r f l o w from 90.7 t o 771 kg l sec (200 t o 1700 l b l s e c ) . Obviously, t o be u s e f u l , reasonable e s t i m a t e s f o r t h e v a r i o u s f a c t o r s i n t h e c o r r e l a t i o n s must be made. An appen- d i x i n t h i s r e p o r t summarizes engine des ign d a t a (such a s t i p speeds and hub- t i p r a t i o s ) which a r e necessary t o apply t h e c o r r e l a t i o n s developed he re . Also qiven i n t h e appendix a r e t h e procedures f o r c a l c u l a t i n g t h e number of compressor and t u r b i n e s t a g e s .
The c o r r e l a t i o n f o r t h e core engine weight is shown i n f i g u r e 5. Here, t h e c o r r e l a t i o n inc ludes only t h e number of s t a g e s i n t h e compressor and high- p ressure tu rb ines . The independent parameter i s c o r e a i r f l o w , which is c a l - c u l a t e d d i r e c t l y from t h e t o t a l a i r flow and t h e bypass r a t i o B:
The c o r r e l a t i o n is not a s good as t h a t f o r t h e fan weight , and obviously more geometric f a c t o r s could be included t o improve i t . However, t h e cor re - l a t i o n appears t o be accep tab le f o r t h e purpose of pre l iminary des ign e s t i - mating. Again, a log-log p l o t has been used, and t h e equat ion is noted on t h e f i g u r e .
The t h i r d element i n t h e weight bui ldup of t h e t o t a l engine is t h e .. remaining weight, which is shown a s a r a t i o wi th t h e core engine weight i n t
f i g u r e 6. The c o r r e l a t i o n , aga in wi th c o r e a i r f l o w , is good and g r a p h i c a l l y demonstrates t h e s i g n i f i c a n c e of t h e subsystem weights t o t h e t o t a l weight. These components tend t o s c a l e very l i t t l e wi th s i z e and, f o r smal le r engines , become a g r e a t e r weight f a c t o r than t h e core engine. The d i f f e r e n c e i n t h e two curves , one f o r geared f a n engines and t h e o t h e r f o r nongeared engines , r e f l e c t s t h e gearbox weight. The weight t r adeof f between geared tu rbofan engines and nongeared turbofans , which of course r e q u i r e more low-pressure t u r b i n e s t a g e s , can be addressed wi th t h e s e c o r r e l a t i o n s and is presented i n t h e Resu l t s s e c t i o n .
ENGINE FRONTAL DIMENSIONS
Engine Face Diameter
A c r i t i c a l dimension i n the engine is t h e diameter a t t h e fan face . Even with low bypass turbofan engines t h i s i s t h e maximvq diameter i n t h e engine , and d i c t a t e s much of the n a c e l l e geometry.
Axia l Mach numbers do not vary widely a t the engine f a c e ; 0 .45 t o 0.60 is t h e approximate spread. Thus, t h e diameter can be c a l c u l a t e d e a s i l y , knowing t h e a i r f l o w a t sea - l eve l s t a t i c cond i t ions and t h e hub-tip r a t i o a t the f a c e of t h e f i r s t s t a g e of t h e f a n ( see appendix).
Assuming
( i n Engl ish u n i t s t h e cons tan t 0.0698 becomes 0.919)
a t sea - l eve l s t a t i c (P - 101,309 ~ / r n ~ (2116 l b / f t 2 ) ; TF = 288 'K (519 O R ) ) ,
F
e .g . , a t h / t = 0.38 and MF = 0.5, WaSLS/AF = 152.3 kg/sec/m2 (31.2 l b l s e c l f t 2 ) . Thus, t h e engine f a c e f r o n t a l a r e a and t h e f a n diameter a r e computed d i r e c t l y
6
from t h e a i r f l o w . This r e l a t i o n is shown i n f i g u r e 7 wi th l i n e s of cons tan t a x i a l Mach number and a c t u a l engine da ta .
Once aga in t h e hub-tip r a t i o becomes an important f a c t o r , and one can observe an obvious weight t r a d e o f f i n i ts s e l e c t i o n . The f a n weight i n c r e a s e s wi th reduced h / t because of b lade roo t s t r e s s e s ( f i g u r e 4 ) . However, a t smal le r hub-tip r a t i o s , t h e diameter of t h e engine is reduced, a l lowing a smal le r n a c e l l e s i z e and both weight and drag reduc t ions . This t r adeof f is discussed i n t h e Resu l t s s e c t i o n .
Engine Core Diameter
The diameter a t t h e face of t h e engine core is no t a s c r i t i c a l a dcmen- s i o n a s t h a t of the f a n face , but i t is used l a t e r t o p r e d i c t t h e l eng th of t h e t r a n s i t i o n duc t ing from t h e fan e x i t t o t h e f a c e of t h e core . Unfor- t u n a t e l y , t h i s d iameter is not a s p r e d i c t a b l e because of t h e wider v a r i a t i o n i n t h e hub-tip r a t i o a t t h e core face and t h e p r e s s u r e and temperature r i s e between t h e fan and t h e core when booster s t a g e s a r e added i n t k e t r a n s i t i o n duct ing. However, i f one assumes t h a t the a x i a l Mach number a t t h e face of t h e core is equa l t o t h a t a t the fan face , t h e fo l lowing express ion can be developed:
The p ressure and temperature r i s e a t the c o r e f a c e is due t o both t h e fan and boos te r s t a g e s , i f t h e r e a r e any. Assuming an i s e n t r o p i c e f f i c i e n c y of 862, the fo l lowing r e l a t i o n s h i p can be e a s i l y computed.
To use t h i s equa t ion , e s t i m a t e t h e hub-tip r a t i o s and p ressure r a t i o between t h e engine f a c e and t h e core face from t h e appendix.
Turbine Ex i t Diameter
The t u r b i n e e x 4 t d iameter is used i n t h e c a l c u l a t i o n of t h e number of t u r b i n e s t a g e s . The appendix d e t a i l s t h e c a l c u l a t i o n procedure, which is s i m i l a r t o t h a t shown above except t h a t p r e s s u r e and temperature r a t i o s must be computed a c r o s s the whole engine.
For t u r b o j e t s , t h i s dimension is t h e maximum engine diameter and t h u s is a f a c t o r i n determining n a c e l l e dimensions. For ' o r b y p a s s r a t i o tu rbofans having mixed flow nozz les , t h e f a n exhaust duc t ing over t h e t u r b i n e may r e q u i r e a diameter l a r g e r than t h e fan f a c e t o avoid high subsonic i n t e r n a l Mach numbers. Again, n a c e l l e d i m e n s i ; . ! ~ ~ w i l l be a f f e c t e d . Also, t h e t u r b i n e exhaust d iameter is used t o e s t i m a t e t l a c o r e nozzle l eng th t h a t is d i scussed i n t h e n a c e l l e geometry s e c t i o n .
ENGINE LENGTH
The c a l c u l a t i o n of eng ine l e n g t h becomes somewhat more compl i c r t ed i n h igh bypass r a t i o t u r b o f a n eng ines because of t h e need t o r e d . . , r t h e d i a m e t e r s i? t h e f low p a t h going from t h e f a n e x i t t o t h e c o r e . T h i s r e g i o n is o f t e n c a l l e d t h e "gocseneck." To account f o r t h e l e n g t h c o n s i s t e n t l y , t h i s paper d i v i d e s t h e engine i n t o f o u r r e g i o n s : f a n s t a g e s (from t h e eng ine f r o n t f l a n g e t o t h e f a c e of t h e l a s t f a n s t a g e ) ; t h e gooseneck (from t h e f a c e of t h e ' l a s t f a n s t a g e t o t h e f a c e of t h e eng ine c o r e ) ; t h e c o r e (from t h e f a c e O L t h e eng ine c o r e t o t h e r e a r f l a n g e of t h e h igh -p re s su re t u r b i n e ) ; and t h e low- c'.
p r e s s u r e t u r b i n e s t a g e s (from t h e h igh -p re s su re t u r b i n e e x i t t o t h e r e a r 1
f l a n g e of t h e eng ine , which i n c l u d e s any t r a n s i t i o n d u c t i n g between t u r b i n e s ) . These r e g i o n s a r e d e p i c t e d i n s k e t c h ( a ) . Note t h a t by t h i s d e f i n i t i o n t h e f a n l e n g t h goes t o z e r o f o r a s i n g l e s t a g e f an . I n c l u d i n g t h e f i n a l f a n s t a g e w i t h t h e gooseneck avo ids d e f i n i t i o n problems i n t r y i n g t o account f o r d i f - f e r e n t geome t r i c d e s i g n v a r i a t i o n s i n s p l i t t i n g t h e f low i n t o t h e c o r e .
Ske tch ( a )
Both t h e i n l e t and n o z z l e a r e cons jde red i n a l a t e r d i s c u s s i o n of n a c e l l e l e n g t h .
Fan Length
I n e s t i m a t i n g t h e l e n g t h of a m u l t i s t a g e f a n , t h e l e n g t h p e r s t a g e i s cocs ide red . The f i r s t s t a g e s of s e v e r a l t u r b o j e t eng ine compressors a r e a l s o i n c l u d e d i n t h e d a t a shown on f i g u r e 8. The c o r r e l a t i o n o f f a n (compressor ) l e n g t h p e r s t a g e is p l o t t e d a g a i n s t an e f f e c t i v e engine f a c e f low pa th d iameter t o i n d i c a t e s c a l e s i z e . The sp read of t h e d a t a i n f i g u r e 8 is s i g n i f i c a n t w i t h t h e r a t i o of l e n g t h t o e f f e c t i v e d i ame te r p e r u n i t s t a g e v a r y i n g between 0.18 and 0.32. A mean va lue of 0.25 would g i v e a r e a s o n a b l e e s t i m a t e , b u t t h e i n c r e a s e d s p a c i n g between f a n s t a g e s may b e a d e s i g n requi rement f o r reduced f a n n o i s e .
Gooseneck Length
The l e n g t h of t h e gooseneck i s d i c t a t e d p r i m a r i l y by t h e r e d u c t i o n i n f low pa th d i ame te r from t h e f a c e t o t h e c o r e . The neces sa ry t u r n i n g o f t h e
flow path must be c a r e f u l l y contoured t o avoid l o c a l separat4.0n and r e s u l t i n g t o t a l p ressure l o s s e s . Frequent ly , engines a r e designed wi th fan e x i t guide vanes, o r co re boos te r s t a g e s loca ted i n t h e gooseneck. These a d d i t i o n s may add t o t h e l eng th un less t h e flow is turned through t h e s e s t a g e s t o r t j u c c t h e flow path diameter.
F igure 9 is a c o r r e l a t i o n t h a t a l lows t h e gooseneck l eng th t o be calcu- l a t e d once t h e fan face diameter and t h e c o r e f a c e diameter a r e known. A s
8 t h e core diameter i n c r e a s e s r e l a t i v e t o t h e fan d .meter ( a s wi th low bypass
r a t i o tu rbofans ) , t h e gooseneck l eng th - a s def ined he re - approaches t h e l eng th of a s i n g l e f an s t a g e , a s noted on t h e f i g u r e . However, f o r t h e high
* bypass r a t i o tu rbofans , where t h e core diameter may be as low a s one- th i rd t h e fan diameter , the gooseneck l eng th becomes s i g n i f i c a n t .
Two engines f a l l consi2erably o u t s i d e t h e c o r r e l a t i o n band shown on f i g u r e 9. The s tudy engine shown has an e x i t guide vane wi th a l l f low pzth t u r n i n g a f t e r t h i s vane; t h e production engine has s e v e r a l core boos te r s t a g e s i n t h e flow path be fo re diameter-reduction t akes p lace . Thus, t h e approach taken by t h e engine des igner can add s i g n i f i c a n t l y t o t h e gooseneck l eng th . However, t h e r e a r e d e f i n i t e minimum leng th l i m i t s represented by t h e lower s i d e of the c o r r e l a t i o n band.
Core Engine Length
The l eng th of t h e c o r e engine c o r r e l a t e s w e l l wi th t h e log of t h e c o r e compressor p r e s s u r e r a t i o , a s shown i n f i g u r e 10. The square roo t of the core a i r f l o w i n d i c a t e s t h e s c a l e s i z e af the core . The log of t h e compressor p r e s s u r e r a t i o r e l a t e s t h e compression r a t i o t o t h e number of compression s t a g e s , which has a d i r e c t bea r ing on t h e l eng th of t h e core .
Note t h a t t h e three-spool RB211 turbofan engine f a l l s below the t r end l i n e i n t h e f i g u r e . Adding a t h i r d spoo l improves compressor s t a g e work i n t h e high-pressure end of t h e compressor and thus reduces the number of com- p ressor s t a g e s requ i red . Such is t h e case f o r t h e RB211 engine; t h e e f f e c t on core engine l eng th i s r e f l e c t e d i n f i g u r e 10.
Low-Pressure Turbine Length
The format f o r c o r r e l a t i n g the low-pressure t u r b i n e l eng th is i d e n t i c a l t o t h a t f o r t h e fan: l eng th per s t a g e i s p l o t t e d a g a i n s t t h e e f f e c t i v e diameter a t t h e t u r b i n e e x i t . The c o r r e l a t i o n is shown on f i g u r e 11; t h e spread of d a t a , which is s i g n i f i c a n t , is s i m i l a r t o t h a t f o r t h e fan l eng th per s t a g e ( f i g . 8) . Again, a r a t i o of low-pressure t u r b i n e l eng th p e r s t a g e t o t h e e f f e c t i v e e x i t diameter of 0.25 is an accep tab le meav va lue es t ima te .
There is some r a t i o n a l explanat ion f o r t h e d a t a spread. Thk low-pressure t u r b i n e l eng th inc ludes any t r a n s i t i o n duct ing between t h e e x i c of t h e h igh p ressure t u r b i n e and the low-pressure t u r b i n e en t rance . However, t h e a d d i t i o n of the t r a n s i t i o n p i e c e is a t the op t ion of the des igner , a s demonstrated by ske tch (b ) .
n TURBOFAN I
----
Sketch ( b j
Turbofan I has no i n t e r t u r b i n e t r a n s i t i o n p iece , and t h e low-pressure t u r b i n e is designed wi th a constant hub diameter and cons tan t ly inc reas ing t i p diam- c t e r s . Turbofan I1 has a t r a n s i t i o n p i e c e and t h e t i p diameter of t h e low- pressure t u r b i n e is constant wi th success ive ly reduced hub diameters. These a r e two extreme examples. A s h o r t e r t r a n s i t i o n p iece combining inc reas ing t i p diameter and reduced hub diameter is a common design approach.
The t radeof f involved is minimum engine l eng th ve rsus higher t u r b i n e work per s t age . I n t e r e s t i n g l y , t h e JT9D is an example of a Turbofan I, and t h e CF6 is an example of a Turbofan 11. The t u r b i n e l eng ths p l o t t e d f o r t h e s e two engines i n f i g u r e 11 r e f l e c t t h e design approach taken.
NACELLE DIMENSIONS
Calcu la t ion of t h e n a c e l l e geometry is a necessary s t e p t o compute n a c e l l e weights. I n a d d i t i o n , n a c e l l e geometry must be c a l c u l a t e d t o compute t h e e x t e r n a l drag of t h e n a c e l l e . Sketch ( c ) i d e n t i f i e s t h e key dimensions used i n t h i s paper i n de f in ing the n a c e l l e geometry f o r both s h o r t f an duct n a c e l l e s and long fan duct nace l l es . To be c o n s i s t e n t wi th t h e engine gccr ict ry already presented, t h e c r i t i c a l dimensions t h a t lead t o t h e geometry of t h e n a c e l l e a r e t h e f a n and t u r b i n e diameters and t h e engine l eng th ( fan f a c e of t u r b i n e r e a r f l ange) .
It is d i f f i c u l t t o ob ta in c o n s i s t e n t d a t a t o compare n a c e l l e geometry and weights on e x i s t i n g production a i r c r a f t . For tuna te ly , a r e c e n t l y com- p l e t e d NASA study ( r e f . 19) has done j u s t t h a t ; wi th very few except ions , a l l da ta shown f o r n a c e l l e dimensions and weights a r e taken from t h a t r e p o r t . These d a t a a r e t abu la ted i n t a b l e 3.
** LE L1 *LFX+Lc*
SHORT FAN DUCT NACELLE
LONG FAN DUCT NACELLE
Sketch ( c )
I n l e t Length
It is f a i r l y s t ra igh t fo rward t o compute t h e l eng th of t h e i n l e t based on t h e requirement t o d i f f u s e t h e engine a i r f l o w from t h e v e l o c i t y at t h e i n l e t t h r o a t t o t h e face of t h e engine. However, t h i s i s a c r i t i c a l region f o r t h e a d d i t i o n of a c o u s t i c t rea tment , and i n t h e f u t u r e , n a c e l l e i n l e t s may be
1 d i c t a t e d more by t h i s requirement than by a i r f l o w d i f f u s i o n .
The ske tch (d) d e f i n e s t h e f a c t o r s needed t o compute t h e i n l e t length .
A t t a k e o f f , t h e i n l e t t h r o a t i s t y p i c a l l y s i z e d f o r a l o L a l Mach number of 0.65 t o 0.70. A t t h e face of t h e engine, t h e Mach number is nominally 0.50 (see f i g . 7 ) . However, f o r proper d i f f u s i o n t o minimize t h e combined e f f e c t s of flow s e p a r a t i o n and w a l l f r i c t i o n , t h e e f f e c t i v e d i f f u s i o n cone angle should be approximately 7'. Thus, once t h e engine f a c e hub-tip r a t i o is
specified (tvpically 0.36 to 0.40 - see appendix) then the ratio of length to diameter can be computed directly.
Sketch (d)
wavT - = Wff , as defined previously with y = 1.4 XP
From continuitv,
From geometry
Thus,
1 - ( f / ~ 1 - (hit)']) 1 1 2 =I - i D~ 2 tan h
For a typical case, assume
3 = 0.475, WffF = 0.029 kg ' K ~ ~ ~ / N sec (0.38 lb '~'/~/lb sec) F
Then,
The d a t a i n f i g u r e 12 represen t cur ren t n a c e l l e designs. Since t h e parameters i n t h e equat ion above can be juggled t o change t h e i n l e t l eng th t o diameter r a t i o , and t h e spread of d a t a is not s u r p r i s i n g . A s mentioned pre- v ious ly , longer i n l e t s a r e u s u a l l y designed t o reduce fan noise .
Fan Cowl Length
The fan cowl, by d e f i n i t i o n h e r e , inc ludes t h e p o r t i o n of t h e n a c e l l e t h a t wraps around t h e f a n and fan exhaust duct ing ahead of t h e f a n nozzle and t h r u s t r everse r . This d e f i n i t i o n i s l i m i t e d t o t h e s h o r t f an duct n a c e l l e only. For the long duct n a c e l l e , t h e l eng th of t h e engine def ines t h i s por- t i o n of t h e n a c e l l e ( see sketch ( c ) ) .
The d a t a i n f i g u r e 1 3 r e l a t e t h e fan cowl l eng th t o t h e engine length . Note t h a t t h e d a t a group near a length-to-length r a t i o of 0.4. It is apparent t h a t t h e des ign o b j e c t i v e f o r t h e s h o r t duc t n a c e l l e is t o enclose t h e f a n flow t o a po in t ahead of t h e maximum core diameter a t t h e tu rb ine . The t h r u s t reverser-nozzle system then must extend f a r enough back t o a l low t h e f a n flow t o expand over t h e b o a t t a i l of t h e core nozzle.
Fan Thrust Reverser and Exhaust Nozzle Length
Whereas t h e f a n n a c e l l e l e n ~ t h is r e l a t i v e l y p r e d i c t a b l e , t h e l eng th of t h e f a n t h r u s t r e v e r s e r is n o t , because t h e des igner has some op t ion i n con- t o u r i n g t h e core b o a t t a i l , a s shown by t h e two examples i n ske tch (e ) .
Thus, t h e l eng th LFX is no t very p r e d i c t a b l e , as demonstrated i n f i g u r e 14. For t h e s h o r t f an duct n a c e l l e , t h e r a t i o of t h e f a n exhaust l eng th t o the f a n diameter v a r i e s from <0.4 t o X . 7 . However, p icking a nominal value such a s 0.55 docs not n e c e s s a r i l y l ead t o unacceptable weight es t imates , s i n c e t h e c o r e cowl length LC is computed as
Thus, a s h o r t e s t imate f o r t h e l eng th LFX gives a long es t imate f o r t h e l eng th LC and a f f e c t s t h e n a c e l l e weight es t imate accordingly ( see next s e c t i o n ) .
NACELLE WEIGHTS
To account f o r t o t a l n a c e l l e weight, t h e n a c e l l e is subdivided i n t o t h e cowling, f a n exhaust system, and core exhaust system. Thus, t h e cowling inc ludes t h e i n l e t cowl su r face , t h e cowl s u r f a c e covering t h e f a n and fan exhaust duc t ing ahead of t h e fan nozzle and t h r u s t r e v e r s e r , and the exposed core cowl s u r f a c e from t h e plane of t h e f a n nozz le e x i t t o t h e a t t a c h p o i n t f o r t h e core nozzle and t h r u s t r e v e r s e r .
Surface a r e a is used t o c o r r e l a t e weight, and, r a t h e r than in t roduce t h e complexity of t h e a c t u a l nonaxisymmetric shape of t h e n a c e l l e , simple c y l i n d e r s a r e assumed, using t h e f a n diameter , t h e t u r b i n e diameter , and t h e l e n g t h s from t h e previous sec t ion .
Cowl weight d a t a a r e given i n f i g u r e 16; a u n i t weight of 17.1 t o 19.6 kg/m2 (3.5 t o 4.0 l b / f t 2 ) appears reasonable .
S i m i l a r l y , t h e exhaust system weights can be es t imated from s u r f a c e a r e a c a l c u l a t i o n s . The l eng ths from t h e previous s e c t i o n a r e used, a long wi th t h e fan and t u r b i n e diameters f o r t h e f a n and core exhaust systems, r e s p e c t i v e l y . The d a t a a r e given i n f i g u r e 17, a u n i t weight of 73.2 kg/m2 (15 l b / f t 2 ) i s a r e p r e s e n t a t i v e val-ue f o r es t imat ion purposes.
RESULTS
This s e c t i o n demonstrates how t h e c o r r e l a t i o n s t h a t have been developed can be used i n a model t o es t imate weights and t h e r e l a t i v e e f f e c t on n a c e l l e drag. The model has d e f i c i e n c i e s , and they a r e h igh l igh ted i n t h e fol lowing d i scuss ion .
Since a i r f l o w is t h e primary f a c t o r a f f e c t i n g engine weights wi th t h e c o r r e l a t i o n s developed, a necessary s t e p is the c a l c u l a t i o n of s p e c i f i c t h r u s t ( t h r u s t l a i r f low) a t sea- level s t a t i c s tandard day cond i t ions . The f o l - lowing l is t of cyc les used i n t h i s s e c t i o n a l s o lists t h e s p e c i f i c t h r u s t f o r each.
Cycle
1 2 3 4 5 6 7
Bypass r a t i o
5.9 7.5 9 .0 2.8 3.2 4.3 5.0
Turbine i n l e t , temp, O K (OR)
1563 (2813) 1563 (2813) 1563 (2813) 1563 (2813) 1563 (2813) 1563 (2813) 1563 (2813)
Fan pressure r a t i o
1.56 1.32 1.20 a. 0 2.5 2.0 1.75
i
S p e c i f i c t h r u s t
30.6 26.5 22.8 38.2 39.1 35.2 33.4
Overal l p r e s s u r e r a t i o - - 24.2 24.2 24.2 24.2 24.2 24.2 24.2
Cycle 1 i n t h e above t a b l e is t h e c y c l e f o r t h e General E l e c t r i c CF6-6 turboran engine. This c y c l e , a long w i t h t h e flow-path geometry and Mach num- b e r s and f a n and t u r b i n e t ipspeeds of the CF6, was used a s a c a l i b r a t i o n po in t f o r a computer program t h a t c a l c u l a t e s engine and n a c e l l e weights , geometry, and n a c e l l e drag.
Figure 18 shows how t h i s program e s t i m a t e s s p e c i f i c weight over a range of r a t e d t h r u s t from 44,480 t o 222,400 N (10,000 t o 5C,000 l b ) . The CF6 c a l i b r a t i o n po in t is noted and each of t h e t h r e e cyc les1 shown s c a l e s i m i - l a r l y , i n c r e a s i n g s l i g h t l y wi th increased t h r u s t r a t i n g . Based on t h e exper- i ence wi th e x i s t i n g turbofan engines , i t was a n t i c i p a t e d t h a t t h e low bypass r a t i o c y c l e engine would s c a l e more s t r o n g l y wi th s i z e (see t h e f i r s t p a r t o f t h e s e c t i o n on Engine Dry Weight). Since t h i s d i d n o t r e s u l t , one must suspec t t h a t t h e model i s d e f i c i e n t i n p r e d i c t i n g t h e weight of low bypass r a t i o turbofans . The probable source of e r r o r is t h e s c a l i n g of c o r e weight wi th a i r f l o w ( f i g . 5) .
Also n o t e i n f i g u r e 1 8 t h a t two curves a r e shown f o r t h e h igh bypass r a t i o (9) engine. One curve used a t u r b i n e work f a c t o r l i m i t of 2.0 ( see appendix) and an unusually l a r g e number of low p r e s s u r e s t a g e s was computed (15). The second curve is t h e r e s u l t of i n c r e a s i n g t h e t u r b i n e work f a c t o r t o 3.5.2 This reduced t h e number of low-pressure t u r b i n e s t a g e s t o n ine , a more reasonab le , bu t s t i l l h igh , number.
An a l t e r n a t i v e t o many low-pressure t u r b i n e s t a g e s is a gearbox between t h e f s n t h e low-pressure t u r b i n e . A s a r u l e of thumb, tu rbofan engines w i t h a a c y c l e having a bypass r a t i o of 9 o r g r e a t e r should be geared t o avoid weight p e n a l t i e s due t o an excess ive number of t u r b i n e s t a g e s . Since t h e remaining engine weight ( f i g . 6) r e f l e c t s engines wi th and wi thout a gearbox, t h e engine weight model was used t o t e s t t h i s r u l e of thumb. The r e s u l t s , shown on f i g u r e 19, i n d i c a t e t h a t t h e weight t r a d e o f f becomes favorab le f o r a geared fan a t a bypass r a t i o much lower than 9.0. Note t h a t s e v e r a l d i f f e r e n t con- d i t i o n s f o r t h e bypass 9.0 nongeared engine a r e shown: (1) f a n t ipspeed held cons tan t a t t h e CF6 va lue (437 mlsec (1436 f p s ) ) , (2) t ipspeed according t o f i g u r e 22 (271 mlsec (890 i p s ) ) , (3) low-pressure t u r b i n e work f a c t o r inc reased from 2.0 t o 3.5 (maximum). The c y c l e s a r e noted by number on t h e f i g u r e . The number of f an s t a g e s and low-pressure t u r b i n e s t a g e s f o r each a r e a s fo l lows:
l ~ h e CF6 fan and t u r b i n e hub-tip r a t i o s and Mach numbers and t u r b i n e work f a c t o r were held cons tan t f o r each cyc le . However, f an t ipspeed v a r i e d wi th f a n p r e s s u r e r a t i o , a s shown on f i g u r e 22 i n t h e appendix.
2 ~ h i s h igh wgrk f a c t o r would l ead t o very high swirl i n t h e flow a t t h e t u r b i n e discharge . A f i x e d s t a g e of s t a t o r s ("half s tage") would probably be requ i red t o s t r a i g h t e n ou t t h e flow. Th i s has no t been considered i n t h e weight comparisons.
The CF6 c a l i b r a t i o n p o i n t is once aga in noted and t h e d i s t u r b i n g r e s u l t i s t h a t t h e geared engine weight appears t o be es t ima ted too low. It :hould be po in ted ou t t h a t t h e engine d a t a base from which t h e s e c o r r e l a t i o n s were developed have no geared engines wi th a bypass r a t i o l e s s than 9.0. With high bypass r a t i o engines , t h e gearbox i s housed i n t h e c a v i t y i n s i d e t h e fan hub and gooseneck, wi th no apparent p e n a l t i e s t o frame o r bea r ing des ign. A s t h e bypass r a t i o is reduced, t h i s may no t be t h e case , and weight p e n a l t i e s may r e s u l t due t o s p e c i a l i n s t a l l a t i o n s f o r t h e gearbox. I n any even t , t h e geared turbofan engine weights p r e d i c t e d by t h e computer model a r e suspec t .
1 2 3 4 5 6 7 3' 3"
The n a c e l l e weight and c r u i s e drag a r e shown on f i g u r e s 20 and 21, r e s p e c t i v e l y . These d a t a a r e c o n s i s t e n t wi th t h e nongeared engines of f i g u r e 19 (cyc le 3 f o r bypass r a t i o 9.0). A well-defined break occurs i n f i g u r e 19 where t h e number of f a n s t a g e s changes. These b reakpo in t s show up i n t h e n a c e l l e weight and drag curves a s w e l l , a l though smoother cu rves have been f a i r e d . I n g e n e r a l , a s t h e bypass r a t i o i n c r e a s e s , cowl weight i n c r e a s e s s l i g h t l y and t h e r e is a marked t r a d e o f f between t h e weight of t h e f a n exhaust system and the core exhaust system. The n e t r e s u l t is a modest i n c r e a s e i n n a c e l l e weight wi th h igher bypass r a t i o and a r a t h e r s i g n i f i c a n t i n c r e a s e i n n a c e l l e c r u i s e drag.
CCNCLUDING REMARKS
Fan s t a g e s (geared and nongeared)
1 1 1 3 2 2 2 1 1
The o b j e c t i v e of t h i s s tudy was t o p r e s e n t a methodology t h a t is reason- a b l y d e t a i l e d i n t h e p r e d i c t i o n of propuls ion system weights and dimensions. The type of propuls ion addressed is l i m i t e d t o subsonic tu rbofan engines i n s t a l l e d i n pod-type n a c e l l e s . Although d e t a i l e d , t h e methodology i s genera l enough t o provide t h e proper weight and dimension t r ends wi th d i f f e r e n t turbo- fan c y c l e s wi thout r e s o r t i n g t o d e t a i l e d l ayou t s of t h e engine flowpath.
LPT s t a g e s
- *ongeared Geared - 5 3 6 2 6 2 3 4 3 3 4 3 4 3
15 9
I n t h e opinion of t h e a u t h o r s , t h i s o b j e c t i v e can b e met, but t h e methodology presented h e r e f a l l s s h o r t of doing s o . One obvious shortcoming i n r e t r o s p e c t is t h e over - re l i ance on f a c t o r s t h a t a f f e c t b lade s t r e s s when c o r r e l a t i n g fan and low-pressure t u r b i n e weight. This approach addresses
the blades, disks, and hub of the rotating machinery, but ignores the fact that the component weight also includes stators, frames, shafts, and bearings. Thus, a more accurate model may have to distinguish between rotating and non- rotating elements of each component to estimate weight properly. In addition, it may be necessary to distinguish between the fan and the low-pressure tur- bine and not treat the two as one component when correlating weight. With the correlation presented in this paper, a turbine stage is equated equally with a fan stage. This certainly is not true for high bypass ratio engines, which have very large fans with respect to any other component in the engine.
APPENDIX
This appendix summarizes turbofan engine des ign d a t a needed t o compute t h e c o r r e l a t i o n f a c t o r s f o r f an weight, c o r e weight , and s e v e r a l dimensional parameters. Data f o r both product ion engines and s e v e r a l of t h e QCSEE s tudy engines a r e used, a s noted i n t h e t a b l e s .
Table 4 p r e s e n t s d iameters , hub-t ip r a t i o s , and t ipspeeds f o r t h e f a n f a c e , high-pressure t u r b i n e face , and low-pressure t u r b i n e e x i t . These d a t a were gleaned from a v a r i e t y of sources and i n some c a s e s were approximated from a v a i l a b l e engine l ayou t s . The f a c e hub-tip r a t i o i s used d i r e c t l y i n t h e c o r r e l a t i o n f o r f a n weight, and a v a l u e of 0.- :> 0.45 is a represen ta - t i v e value . Fan t ipspeed is a l s o used d i r e c t l y i n the f a n weight c o r r e l a t i o n , and f i g u r e 22 shows how t h i s parameter tends t o vary w i t h f a n s t a g e loading (p ressure r a t i o ) . A c l e a r d i s t i n c t i o n bewteen s i n g l e s t a g e and m u l t i s t a g e f a n s is evident . The ~ n l y except ion i n t h e t r e n d s is t h e r e l a t i v e l y h igh t ipspeed shown f o r t h e s i n g l e s t a g e f a n JT15D, which is a much smal le r engine than a l l o t h e r s l i s t e d i n t h i s appendix.
Table 5 shows p r e s s u r e r a t i o s and numbers of s t a g e s f o r both f a n s and compressors. This i n t u r n l e a d s t o a c a l c u l a t i o n of p r e s s u r e r i s e p e r s t a g e , which is a l s o shown. I n t h e compressor, a p r e s s u r e rise per s t a g e of 1 . 2 is r e p r e s e n t a t i v e of c u r r e n t tu rbofan engine compressor des ign.
Tables 6 and 7 show t h e parameters used i n computing average s t a g e load- ing f o r t h e h igh-pressure and low-pressure t u r b i n e s , r e s p e c t i v e l y , and t a b l e 8 shows parameters used t o compute t h e low-pressure t u r b i n e e x i t Mach number. An e s t i m a t e of average s t a g e loading is a necessary s t e p i n computing t h e number of t u r b i n e s t a g e s .
Ca lcu la t ion of t h e Number of HPT Stages
The c a l c u l a t i o n of t h e number c f high-pressure t u r b i n e s t a g e s is s t r a i g h t f o r w a r d , assuming t h e t ipspeeds and hub-tip r a t i o s a r e known. From t a b l e 6 , va lues of 427 m/sec (1400 f t l s e c ) and 0.86, r e s p e c t i v e l y , a r e repre- s e n t a t i v e . A r u l e of thumb t h a t a p p l i e s w e l l f o r high-pressure t u r b i n e s is t h a t t h e average s tage- loading parameter should no t exceed a va lue of 2. T h u ~ , ~
AH^^ 1 NSTG 7 -
HPT 'HPT
With A = 2.0, round t h i s c a l c u l a t i o n t o t h e next h i g h e s t i n t e g e r , e .g . , NSTG = 1 .2 + NSTG = 2. Note t h a t --
31n Engl ish u n i t s mul t ip ly AHHpT(BTU/lb) by gJ = 32.2 x 778.
Calcu la t ion of Number of LPT S tages
The c a l c u l a t i o n of t h e low-pressure t u r b i n e s t a g e s is somewhat more complicated f o r nongeared turbofan engines because t h e t ipspeed of t h e f a n governs t h e t ipspeeds i n t h e LPT. Again, t h e average s t a g e load ing parameter A is assumed and f r e q u e n t l y is g r e a t e r than 2.0 ( t a b l e 7 ) .
With
1 + h/ tLpT 5 = v
LPT TIPLpT 2
It is t h e c a l c u l a t i o n of VTIp t h a t adds t o t h e complexity
LPT
f o r
Mp = 0.5, Wffp = 0.030 kg 'IC112/~ s e c (0.39 l b ' ~ ~ / ~ / l b sac ) (y = 1 . 4 )
f o r
MtpT ' 0.4, Wf LPT = 0.026 kg O K ~ / ~ / N sac (0.335 l b ' ~ ' / ~ / l b Set) (Y = 1 - 3 5 )
For c l a r i f i c a t i o n use t h e fo l lowing s t a t i o n s : C
1 fan f a c e
4 high-pressure t u r b i n e i n l e t
6 low-pressure t u r b i n e e x i t
Turbine i n l e t temperature T4 and the o v e r a l l p r e s s u r e r a t i o OPR a r e s p e c i f i e d f o r t h e engine. Thus, T4/T1 = T4(OK)/288; P4/P1 = 0.96 OPR (accounts f o r combus t o r p r e s s u r e drop).
' 4 - = 0.96 OPR (accounts f o r combustor 1 pressure drop)
t y p i c a l value of E IEp = 113011005 = 1.12 (0.27/0.24 = 1.12) P~ C
This parameter cannot be c a l c u l a t e d i n a simple c losed form without i n t r o - ducing a s i g n i f i c a n t e r r o r due t o t h e changing gas p r o p e r t i e s dur ing t h e t u r b i n e expansion. The s e r i e s of curves i n f i g u r e 23 represen t t h i s func t ion f o r qT = 0.90. The hub-tip r a t i o s can be es t ima ted from t a b l e s 6 and 7.
($) = 0.55 - 0.60 LPT
Although t h i s c a l c u l a t i o n appears formidable, i t l ends i t s e l f t o a s imple computer program. Obviously, t h e problems a r i s e from s e l e c t i n g t h e proper geometric inpu t s .
For geared fane , t h e c a l c u l a t i o n s of LPT s t a g e s r e v e r t t o t h e s imple r form f o r t h e HPT s i n c e LP t u r b i n e t ipspeed can be s e t independently from t h e f a n t ipspeed . Note t h e h igher LPT t ipspeeds f o r t h e geared fan engines ! i s t ed i n t a b l e 6.
REFERENCES
Swan, W. C. ; Bouwer, D. W. ; ?;I,! Tolle, I?. F. : Life Cycle Cost Impact on 1. Design Considerations for Civil Transport Aircraft Propulsion Systems.
Third International Symposium on Air Breathing Engines, Munich, Germany, March 7-12, 1976.
2, Prctt and Whitney Turbine Engine Guide. Pratt and Whitney Aircraft, November, 1975.
3. Taylor, J. W. R., ed.: Jane's All the World's Aircraft, 1975-76. Franklin Watts Inc. (New York), 1975.
4. Hague, D. S.; bandenberg, J. D.; and Woodbury, N. W.: Mu1ti::ariate Analysis, Retrieval, and Storage System (MARS). Vol. I V , Turbojet and Turbofan Data Base (By Engine). NASA CR-137674, 19?4.
5. Froduct Information. Aircraft Engine Group, General Electric Company, August 1975.
6. Taylor, J. W. R. , ed. : Jane's All the World's Aircraft, 1966-67. Franklin Watts Inc. (New York), 1966.
7. Taylor, J. W. R., ed. : Jane's All the World's Aircraft, 1967-68. Franklin Watts Inc. (New York) , 1967.
8. ALF 502 Turbofan Engine. Avco Lycouing Division, Stratford, Connecticct, 1974.
9. CF6 High Bypass Turbofans. General Electric Aircraft Engine Croup, CincinnatiiLynn, 1975.
10. JT15D Commercial Turbofan: Status and Information Report. United Air- craft of Canada Limlted, Longview, Quebec, 1970.
11. Thomas, W. H.: Unique Features of the TF41 Turbofan. Paper 690687, presented at SAE National Aeronautic and Space Engineering and Manufac- turing Meeting, Los Angeles, Calif., Oct. 1969.
12. Fink, D. E.: M49 Larzac Testing Under Way. Aviation Week and Space Technology, Nov. 23, 1970, pp. 40-44.
13. TF34-GE-130 Installation Data. Ceneral Electric Aircraft Engine Croup, Lynn, Mass., 1973.
14. Yaffee, M. L.: Lycoming Enters Turbofan Arena. Aviation Week and Space Technology, Aug. 21, 1972, pp. 44-47.
15. Yaffee, M. L.: Pratt and Whitney Builds on JT9D Concept. Aviation Week and Space Technology, June 24, 1974, pp. 49-54.
16. Keen, J. M. S. : Development of the Rolls-Royce RB.211 Turbofan for Air- line Operation. Paper 700292, presented at SAE National Air Transpor- tation Meeting, New York, N. Y., April 1970.
17. RB.211 Development Program on Schedule. American Aviation, Sept. 30, 1968, p. 32.
18. RB.401 Turbofan Engine. Rolls-Royce (1971) Limited Bristol Engine Divi- sion, Bristol, 1975.
19. Parametric Study of Transport Aircraft Systems Weight and Cost. PRC Systems Sciences Company, Report No. R-1816, 1974.
20. pray, D. E.: Study of Unconver.,ional Aircraft Engine Designed for Low Energy Consumption. NASA CR-135065, 1976.
21. Nietzel, R. E.; Hirsckron, R.; and Johnston, R. P.: Study of Turbofan Engines Designed for Low Energy Consumption. NASA CR-135053, 1976.
22. Helms, H. E.: Quiet Clean STOL Experimental Engine Study Program, Task I - Parametric Propulsion System Studies. NAS3-16727, 1972.
23. Gerend, Robert P.; and Roundhill, John P.: Correlation of Gas Turbine Engine Weights and Dimensions. AIAA Paper 70-669, AIAA Sixth Pro- pulsion Joint Specialist Conference, San Diego, Calif., June, 1970.
TABL
E 1.-
EN
GIN
E DA
TA
Low
B
yp
ass
Tu
rbo
fan
s
TF4
1-A
-1
64
,49
6
(14
,50
0)
11
7
(25
8)
r
Th
rust
, S
LS
, N
(lb
)
Air
flo
w,
SLS ,
kg
/se
c (
lb/s
ec
)
Sy
pa
ss r
ati
o,
SLS
Dry
w
eig
ht,
kg
(l
b)
Fan
ti
p d
iam
, m
(i
n.)
Tu
rbin
e
tip
dia
m,
las
t st
ag
e,
m
(in.
En
gin
e le
ng
th,
fac
e t
o f
lan
ge
, m
(i
n.)
Num
ber
sta
ge
s
Fan
LPC/HPC
HPT
/LPT
References
b
M49
-01
0
10
,23
0
(2,3
00
)
2 6
(57
)
CF
700-
20-2
1
20
,01
6
(4,5
00
)
5 8
(12
8)
1.9
7
334
(73
7)
1 (a
ft)
8
2 5
CJ8
05-2
3B
71
,61
3
(16
,10
0)
19
1
(42
2)
1.4
6
1,7
08
(3
,76
6)
3.3
3
(13
1)
l(a
ft)
1
7
3/ 1
6
JT3D
-3B
Oa
80
,06
4
(18
,00
0)
21
1
(46
5)
1,9
68
(4
,34
0)
1.2
7
(50
)
0.9
39
(3
7)
3.4
0
(13
4)
2
ti/
7
1/ 3
2,4
1,4
40
0
-76
I I
;:
JT8D
-7
&
62
,27
2
(14
,00
0)
14
3
(31
5)
1.1
1,4
53
(3
,20
5)
1.0
4
(41
)
0.8
13
(3
2)
2.8
4
(11
2)
2 4
/ 7
1/3
2,3
(3,1
75
)
0.9
65
(3
8)
0.7
11
(2
8)
2.6
2
(10
3) 3
2/1
1
2/ 2
3,1
1
(57
3)
0,4
32
(1
7)
1.1
4
(45
)
2
4
1/ 1
12
,3
JT8D
-15
La
68
,94
4
(15
,50
0)
14
7
(32
4)
0.9
9
1,5
01
(3
,30
9)
1.0
4
(41
)
0.8
13
(3
2)
2.8
4
(11
2)
2
4/7
1/3
2,3
JT15
D-1
0
9,7
86
(2
,20
0)
34
(7
5)
3.2
8
229
(50
6)
0.5
33
(2
1)
1.2
9
(51
)
1
1 c
en
t
1/2
3,l
O
TF
E73
1-2
&3
15
,56
8
(3,5
00
)
5 1
(11
3)
2.6
6
322
(71
0)
0.7
11
(2
8)
1.2
7
(50
)
1
4 a
x,
1 c
en
t 1
/3
3
TABLE 1.-
ENGINE DATA -
CONTINUED
Low Bypass Turbofans -
Concluded
Thrust, SLS, N (lb)
Airflow, SLS,
kg/sec (lb/sec)
Bypass ratio, SLS
Dry weight, kg (lb)
Fan tip diam, m (in.)
Turbine tip diam, last
stage, m (in.)
Engine length, face-to-
flange, m (in.)
Number stages
Fan
LPCIHPC
HPT/LPT
References
- -
SPEY MK512
53,198
(11,960)
1,167
(2 , 574) 5
12
2/2 3
SNECMA M53
54,933
(12,350)
8 4
(185)
1,419
(3,130)
4.83
(190)
3
5
2 3
TF30-P-408
&
59,603
(13,400)
1,180
(2,602)
2.66
(105) 2
717
113
2,3
A
TAB
LE 1.-
EN
GIN
E DA
TA -
CO
NTI
NU
ED
Hig
h B
yp
ass
Tu
rbo
fan
s
Th
rust
, S
LS
, N
(It)
)
Air
flo
w,
SL
S,
kg
/se
c
(Ib
lse
c)
By
pas
s ra
tio
, SL
S
Dry
we
igh
t,
kg
(l
b)
Fan
ti
p d
iam
, m
(in
.)
Tu
rbin
e t
ip d
iam
, la
st
sta
ge
, m
(i
n.
En
gin
e le
ng
th,
fac
e
to f
lan
ge
, m
(i
n.
j
Num
ber
sta
ge
s
Fan
L
PC/H
PC
HP
T/L
YT
Re
fere
nc
es
CF
6-50
m
2
17
,95
2
(49
,00
0)
652
(1,4
39
)
4.4
3,7
89
(8
,35
5)
2.1
8
(86
)
1.2
2
(48
)
4.2
7
(16
8)
1
31
14
214
5,9
CFM
56
a 9
7,8
56
(2
2,0
00
)
354
(78
0)
5.9
1,8R2
(4,1
50
)!
1.8
0
(71
)
1
319
11
4
3,s
. .
AL
F502
H
4 2
8,9
12
(6
,50
0)
10
9
(24
0)
6.1
56
5
(1,2
45
)
1.0
4
(41
)
1.4
2
(56
) 1
7 a
xia
l 1 c
en-
trif
ug
al
212
3,8
,14
TF
34-1
00
* 4
0,3
21
(9
,06
5)
15
1
(33
3)
5.2
647
(1,4
27
)
1.0
9
(43
)
0.6
8
(27
)
1.9
8
(78
)
1
14
214
1
CF6
-6D
a
17
7,9
20
(4
0,0
00
)
59
3
(1,3
07
)
5.9
3,5
21
(7
,76
5)
2.1
8
(86
)
1.2
2
(48
)
4.4
4
(17
5)
1
1/1
6
21
5
JT9D
-7
+ 2
13
,05
9
(4 7
,90
0)
65
3
(1,4
40
)
5.1
5
3,9
82
(8
,78
0)
2.3
3
(92
)
1.3
5
(53
)
3.2
5
(12
8)
1
3/ 1
1
214
2,3
,4
JT9D
-59
A
23
5,7
44
(5
3,0
00
)
74
3
(1,6
39
)
4.9
4,1
45
(9
,14
0)
2.3
9
(94
)
1.3
7
(54
)
3.3
5
(13
2)
1
4/1
1
2/4
2,3
,15
5
5
,9
RB
.211
-22
18
6,8
16
(4
2,0
00
)
62
6
(1,3
80
)
5.0
2,8
81
(6
,35
3)
2.1
8
(86
)
1.2
7
(50
)
3.0
2
(11
9)
1
716
11
1/3
3,1
6,1
7
RB
.401
-07
h
24
,01
9
(5,4
00
)
4.5
4 76
(1
,05
0) 1
9
1 / 2
3,1
8
TAB
LE 1.-
EN
GIN
E DA
TA -
CO
NTI
NU
ED
Hig
h B
yp
ass
Tu
rbo
fan
s -
Co
ncl
ud
ed
Th
rust
, S
LS
, N
(l
b)
Air
flo
w,
SL
S,
kg
/se
c
(lb
/se
c)
By
pas
s ra
tio
, SL
S
Dry
we
igh
t,
kg
(lb)
Fan
ti
p d
im,
m
(in
.)
Tu
rbin
e t
ip d
iam
, la
st
sta
ge
, m
(i
n.)
En
gin
e le
ng
th,
face
-to
- fl
an
ge
, m
(i
n. )
Num
ber
sta
ge
s F
an
LPC
/NPC
H
PT
/LP
T
Re
fere
nc
es
JTlO
D
a
10
8,9
76
(2
4,5
00
)
5.3
5
2,1
77
(4
,80
0) 1
12
21
4 3
ST
F47
7 Ir
l~d,
09
4
(26
,55
0)
1,7
87
(3
,94
0)
1.9
1
(75)
1.0
2
(40
)
2.87
(1
13
)
1
3/1
0
2/5
19
G.E
. A
dv.
Eng
.
14
7,6
74
(3
3,2
00
)
7.5
2.0
0
(79
)
1.0
4
(41
)
2.5
9
(10
2) 1
3/9
1
14
TAB
LE 1.-
ENG
INE
DA
TA -
CON
CLU
DED
Tu
rbo
j ets
Th
rust
, S
LS
, N
(l
b)
Alr
fla
w,
SL
S,
kg
lse
c
(lb
lse
c)
Dry
w
eig
ht,
k
g
(lb
)
ist
sta
ge
dia
m,m
(in
.)
Tu
rbin
e
tip
dia
m,
last
sta
ge
, m
(i
n.)
En
gin
e le
ng
th,
fac
e-
to-f
lan
ge
, m
(i
n.)
Num
ber
sta
ge
s
LPC
/HE'
C
HP
TIL
PT
R
efe
ren
ce
s I
CJ8
05-3
0
49
,81
7
(11
,20
0)
76
(16
8
1,2
70
(2
,80
0)
17
3
6
CJ6
10
-1
E
12
,67
7
(2,8
50
)
20
(44
)
18
1
(39
9)
8 2 5
CJ6
10-8
d
13
,78
9
(3,1
00
)
20
(44
)
18
5
(40
7)
8
2
5
JT4A
-3
8
70
,27
8
(15
,80
0)
2,2
77
(5
,02
0)
3.6
6
(14
4)
8/ 7
1
12
7
J52-
P-6
A
' A
37
,60
8
(8,5
00
)
93
2
(2,0
56
)
3.0
22
(1
19
)
12
3
JT12
A-8
A
14
,67
8
(3,3
00
)
212
(46
8)
0.4
83
(1
9)
1.5
7
(62
)
9 2
2,3
JT3C
-6
A
60
,04
8
(13
,50
0)
9 1
(20
0)
1,9
20
(4
,23
4)
4.2
4
(16
7)
9/
7 1
12
7
TAB
LE
2.-
ST
UD
Y
EN
GIN
ES
PD
287
Gea
red
(S
LS
thru
st
=
88
,96
0
N
(20
,00
0
lb))
r
Air
flo
w,
SL
S,
kg
/se
c
(lb
lse
c)
By
pas
s ra
tio
, SL
S
Dry
we
igh
t,
kg
(l
b)
Fan
we
igh
t,
fan
+ L
PT
, k
g
(lb
)
Co
re w
eig
ht,
co
mp
+ co
mb
+ H
PT,
kg
(l
b)
Rem
ain
ing
we
igh
t,
kg
(l
b)
Fan
ti
p d
iam
, u
(in
.)
Tu
rbin
e t
ip d
im,
in (i
n.)
En
gin
e le
ng
th,
fac
e-
to-f
lan
ge
, m
(i
n.)
Num
ber
sta
ge
s
Fan
L
PCIH
PC
HPT
/LPT
-2 1
0
50
6
(11
17
)
21
.5
12
25
(2
70
1)
744
(16
41
)
18
9
(41
7)
29
2
(64
3)
2.1
1
(83
)
0.7
1
(28
)
2.6
9
(10
6)
1
11
21
3
-2 3
0
342
(75
5)
9.0
14
04
(3
09
5)
815
(17
96
)
26
0
(5 7
3)
329
(72
6)
1.7
0
(67
)
0.9
1
(36
)
2.6
9
(10
6)
1
10
21
2
-22
42
3
(93
3)
14
.8
12
63
(2
78
5)
736
(16
24
)
21
9
(48
2)
30
8
(67
9)
1.9
3
(76
)
0.7
9
(31
)
2.7
2
(10
7)
1
11
21
3
-24 A
50
6
(11
15
)
18
.2
12
66
(2
79
1)
749
(16
51
)
224
(49
3)
29
4
(64
9)
2.1
1
(83
)
0.7
6
(30
)
2.9
5
(11
6)
1
11
21
3
-25
0
50
7
(11
18
)
26
.0
11
90
(2
62
3)
74 1
(1
63
4)
15
8
(34
9)
29
0
(64
0)
2.1
1
(83
)
0.6
8
(27
)
2.7
4
(10
8)
1
11
21
3
-26
b
50
7
(11
19
)
29
.4
-27 0
50
6
(11
17
)
22
.0
-28
0
50
6
(l1
16
)
11
73
1
24
3
(25
88
) 1 (2
74
1)
- 30 0
49
8
(10
99
)
12
24
'O"
I
(27
00
)
740
(16
32
)
14
3
(31
5)
29
1
(64
1)
2.1
1
(83
)
0.6
6
(26
)
2.6
9
(10
6)
1
11
21
3
20.7
74
5
(16
38
)
19
2
74
9
(16
52
)
19
7
(43
5)
29
7
(65
4)
2.1
1
(83
)
0.7
4
(29
)
2.8
4
(11
2)
1
9 21
3
724
(15
96
)
19
3
(42
3)
29
0
1 (4
26
) 2
92
(63
9)
2.1
1
(83
)
0.6
8
(27
)
2.7
2
(10
7)
1
I1
21
3
(54
3)
2.1
1
(83
)
0.7
1
(28
)
2.6
4
(10
4)
1
11
21
3
TAB
LE
2.-
STU
DY
E
NG
INE
S P
D28
7 -
CON
CLU
DED
No
ng
eare
d
(SL
S th
rus
t =
8
8,9
60
N
(20
,00
0
lb))
-48
D
2 39
(5
28
)
3.3
15
01
(3
31
0)
99
7
(21
99
)l
29
7
(65
5)
20
7
(45
6)
1.4
2
(56
)
1.0
7
(42
)
3.3
8
(13
3)
3 9 21
3
-47 0
2 37
(5
24
)
3.6
15
72
(3
46
8)
10
30
(2
27
2)
32
6
:720
)
21
6
(47
6)
1.4
0
(55
)
1.0
9
(43
)
3.5
1
(13
8)
3 7 2/3
-
Air
flo
w,
SL
S,
kg
/se
c
(lb
/se
c)
By
pas
s ra
tio
, SL
S
Dry
we
igh
t,
kg
(lb
)
Fan
we
igh
t,
fan
+ L
PT
, kg (I
ll)
Cc
rew
eig
ht,
co
mp
+co
mb
+H
PT
, kg
(lb
)
Rem
ain
ing
we
igh
t, kg
(lb
)
Fan
tip
dia
m,
m
(in
.)
Tu
rbin
e t
ip d
im,
m
(in
.)
En
gin
e le
ng
th,
fac
e-t
o-f
lan
ge
, m
(i
n.)
Num
ber
sta
ge
s
Fan
L
PCIH
PC
HPT
/LPT
-42
0
2 3
8
(52
5)
3.5
15
22
(3
35
6)
10
09
(2
22
5)
30
3
(66
8)
21
0
(46
3)
1.4
0
(55
)
1.0
7
(42
)
3.3
3
(13
1)
3 8
2/3
-44
A
24
3
(53
5)
2.9
15
71
(3
46
4)
10
09
(2
22
6)
35
1
(77
3)
21
1
(46
5)
1.4
2
(56
)
1.0
9
(43
)
3.3
8
(13
3)
3 8
21
2
-4 3
D 22
2
(49
0)
2.8
16
45
(3
62
7)
11
17
(2
46
4)
30
5
(67
2)
22
2
(49
0)
1.3
7
(54
)
1.1
2
(44
)
3.4
3
(13
5)
3
8
21
3
-45
0
2 34
(5
16
)
4.2
14
73
(3
24
7)
10
04
(2
21
4)
25
7
(56
6)
21
2
(46
7)
1.4
0
(55
)
1.0
7
(42
)
3.3
5
(13
2)
3 8
21
4
TABL
E 3.
- N
AC
ELLE
DA
TA
JT9D
-7
(B74
7)
1.2
2
(48
)
1.1
2
(44
)
0.8
6
(34
)
1.2
7
(50
)
1.0
9
(43
)
20
RB
.211
-22
(L1
01
1)
1.2
4
(49
)
1.2
7
(50
)
1.6
5
(65
)
0.1
0
(4)
344
(76
0)
682
(15
03
)
22
1
(48
7)
2 0
CF
6-50
I
1.5
2
(63
)
1.2
4
(49
)
1.1
4
(45
)
2 1
CF6
-6D
(D
C-1
0-10
)
1.5
0
(59
)
1.5
5
(61
)
1.1
7
(46
)
1.7
3
(68
)
0.6
3
(25
)
50
9
(11
24
)
60
7
(13
38
)
207
(45
6)
2 0
JT3D
-3B
(D
C-8
-62)
tar
0.9
1
(36
)
3.27
(1
29
)
0
(0)
1.4
7
(58
)
330
(72
7)
507
(11
17
)
2 0
JT9D
-15
(DC
-10-
40)
+ 1
.80
(7
1)
1.3
7
(54
)
1.5
7
(62
)
0.3
0
(13)
1.4
5
(57
)
50 5
(1
11
3)
797
(17
58
)
390
(b5
9)
2 0
JT3D
-3B
(D
C-8
-55)
Q
1.1
7
(46
)
1.2
4
(49
)
0.8
9
(35
1.1
4
(45
)
1.3
7
(54
)
2 1
8
(48
1)
248
(54
7)
3 30
(7
28
)
20
Dim
ensi
on
s In
let
len
gth
, m
(i
n.)
Fan
sh
rou
d l
en
gth
, m
(i
n.)
Fan
th
urs
t re
ve
rse
r le
ng
th,
m
(in
.)
Co
re c
ow
l le
ng
th,
m
(in
.)
Pri
mar
y
thru
st
rev
ers
er
len
gth
, m
(i
n.)
Wei
gh
ts
Cow
l,
kg
(lb
)
Fan
th
rus
t re
ve
rse
r,
kg
(l
b)
Pri
mar
y
thru
st
rev
ers
er,
k
g
(Ib
)
Re
fere
nc
es
JT8D
-9
(DC
-9-3
0)
0.7
11
(2
8)
3.0
5
(12
0)
0
(0)
1.3
5
(53
)
15
5
(34
1)
17
8
(39
2)
2 0
TAB
LE
4.-
F.W
A
ND
TU
RB
INE
DATA
En
gin
e
CF6
-6
CF
6-50
TF
34-1
00
JT9D
- 7
JT8D
-15
AL
F502
H
JT15
D-1
TF
41-A
-1
PD
287-
21
PD
287-
22
PD
287-
23
PD
287-
62
PD
287-
43
a~
ntr
an
ce
b
~x
it
o
f L
PT
Dia
mete
rs,
m
(in
.)
an^
H
PT
~
LP
T~
2.1
9
0.8
6
1.2
3
(86
.4)
(33
.9)
(48
.4)
2.1
9
0.8
6
1.2
4
(86
.4)
(33
.9)
(48
.7)
1.0
8
0.4
8
0.6
9
(42
.5)
(18
.8)
(27
.0)
2.3
4
0.9
4
1.3
4
(92
.3)
(37
.0)
(52
.6)
1.0
4
0.6
8
0.8
1
(41
.0)
(26
.6)
(32
.0)
1.0
4
(41
.0)
0.5
3
0.3
1
0.3
8
(21
.0)
(12
.1)
(15
.0)
Hub
- to
- ti
p
an^
H
PT a
LP
T~
0.3
8
0.8
6
0.5
8
0.3
8
0.8
6
0.6
0
0.4
2
0.8
9
0.7
1
0.3
7
0.8
7
0.5
7
0.3
7
0.7
5
0.4
9
0.4
2
0.4
0
0.7
7
0.6
0
Tip
spe
ed
s,
m/s
ec
(f t
/se
c)
an
^ H
PT
~
4 38
44
7
-2
24 5
(1
43
6)
(14
68
) (8
04
)
457
4 74
(1
50
1)
(15
55
) (8
46
)
416
44
7
258
I 2
58
1
(13
65
) (1
46
8)
409
354
()
I (1
34
1)
(11
63
)
452
(14
84
)
(77
2)
I (1
15
8)
38
8
(12
74
) i i
447
52 7
31
9 (1
46
6)
(17
29
) (1
G47
)
419
374
324
(14
72)
(12
26
) (1
06
8)
279
(91
5)
316
34 5
(1
04
0)
(11
34
)
399
(13
10
)
4 3
6 (1
43
0)
46
6
(15
30
)
I 0
.96
0
.56
0
.69
(3
7.7
) (2
2)
(2 7
)
2.1
0
(82
.8)
1.9
2
0.7
9
(75
.6)
(31
.0)
0.3
7
0.8
6
0.5
3
0.4
5
0.4
5
0.5
0
1.7
0
1 0.4
5
(67
.0)
1.4
1
(55
.4)
1.3
6
(53
.6)
0.4
5
0.4
5
of
fan
an
d H
PT
TABLE 5.- FAN AND COMPRESSOR STAGE PRESSURE RATIOS
a ~ u m b e r of s t a g e s inc ludes boost s t a g e s from fan r o t o r a s w e l l a s compressor and high p ressure compressor s t a g e s .
Engine
TF34-100 CF6-6D CF6-50A CF34 JT9D- 7 JT15D-1 ALF502H TF41-A-1 TF41-A-2 JT12A-8 JT8D-15
-
1
Compressor
CPR
13.2 15.51 16.6 12.5 13.9
8 .4 8.59 6.8
17.64
Fan
FPR
1.5 1.56 1.69 1.40 1.6 1.5 1.45 2.37 2.49
2.04
Stagesa
14 17 17 14 14
1 3 1 3 9 11
PRISTC:
1.202 1.174 1.179 1.197 1.206
1.177 1.180 1.237 1.298
Stages
1 1 1 1 1 1 1 2 2
2
PRISTC,
1 .5 1.56 1.69 1.40 1.6 1 . 5 1.45 1.54 1.58
1.428
m Q
QI m 9 3 2 4 u s OD m m m
b . . 4 4 N 4 d N
m u ( Y h l m m
n n n n
e 2 n w + 4 Y
1
c.c 3
L1 % E
n m * . * 4 .,oi us m s
. Q\h m h m o r l N
I N N a m ogo url m m m m rncu @ w r'JE 4% TIz NE NS NZ NZ
w
n n .j n n e 4 n
n i x % og* m G W G O D h U U *ID m e Y)m o m
m-t U O m N N O 4 0 - 7 4 m 0 m d m s m d m z
Y w
n n n h n n h n r . h i : m r O d U N \ 3 O C
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h S J. 4 00 P. h m m
n 2 z :
U U OD
d e 3
'0 U) u OD W k 2 g 2 ( ~ r l r ( m
Q In h h U h
d 0 . 0 . h
m a o m m m m U) E h " h * h C l h h h rln *8< 24 yh q o m a h e m m OD h N P I
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n
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n n n N n O n d h a O N mrn U n
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d h N n U h m n r n h N n
n n n m n n Q I O O N N N
U O 0000 0000 OQI 4QI UQ) QI9 rzrn N m u9 m u aom "0e 00e 005 QIZ "05 00e
r r n n n n n n U v l r l 9 N N
m e * N r n * U V , Q I * 9u 0 9 QI .* Y;Z d\d o m uu .N N A
$ 25 w
n r L ) u t n a a 4 a a w
g O J ~ Y
005 Ns dc 95 hs W i s
n n h N n o n
N
m o 4 m m u - 0 0 h 00m m m rnm m u d m u r n mcu r n 3 42 *r( C(N m t ; N Z V u
8 d N d u
d - V)
' A d w o o U d W
w
4
TUR
BO
JETS
A
XIA
L FL
OW
C
OM
PRES
SOR
S
MAX
R
ATE
D S
EA
LEV
EL
STA
TIC
TH
RU
ST,
N
Fig
ure
1.-
T
urb
ojet
en
gin
e s
pe
cif
ic w
eig
hts
.
TUR
BOFA
NS
BPR
< 2.0
* kk T
UR
BO
FAN
S
*---
----
--
3
- BP
R >
4.0
< 8.0
I b
L I
I
I
I
I I
II
I
I I
1
1000
0 20000
50000 1
0000
0 300000
MA
X R
ATE
D S
EA
LE
VE
L S
TATI
C T
HR
UST
, N
Figure
2.-
Turbofan engine specific weights.
SP
EC
IFIC
W
EIG
HT
SC
ALI
NG
E
XP
ON
EN
T, n
lu
* .P
0
UI
0
SLS
TH
RU
ST-
LO
WER
LIM
IT
FOR
SC
ALI
NG
, N
10 50 100 CORE AIRFLOW (SLS), kglsec
Figure 5.- Core engine weights.
0 25 50 75 100 CORE AIRFLOW (SLS), kg/sec
Figure 6 . - Engine remaining weights.
GO
OSE
NEC
K
.3
.4
.5
.6
CORE
DIA
MET
ER /
FAN
DIA
ME
TER
Fig
ure
9.
- Gooseneck l
eng
ths.
2 4 6 8 10
(CORE AIRFLOW, SLSF', J q , h/sec)'''
Figure 10.- Core engine lengths.
in.
0 .50 I .O EFFECTIVE TURBINE EXIT DIAMETER,
Figure 11.- Lorpre88ure turbine length per stage.
9 S a m
SUR
FAC
E A
RE
A,
FAN
TH
RU
ST
RE
VE
RS
ER
S,
A~~
~ 9 ,
Fig
ure
17.-
Fan exhaust
system weights.
CORE ENGINE EXHAUST SYSTEM WEIGHT, WCX , kg
1 I
I I
I 0
50,0
00
100,
000
20
0,0
00
SE
A
LEV
EL
STA
TIC
TH
RU
ST,
N
Ftgure 18.-
Predicted engine specific weights.
Fig
ure
19.-
Pre
dic
ted
en
gin
e w
eig
hts
: g
eare
d v
s n
onge
ared
.
V~
lpL
Pr =
l I34 ft
/s
(346 m
/s)
GEA
RED
ONL
Y -
gJ
AH
A~~
~ /F
AN
TI
PSPEED
, ft
/s
(m/s
) 2.
0/14
36 (4
36) N
ON
-GEA
RED
-
,A
2.0
/890 (2
71)
t 8 8
'-4 3.5
/890
(271
) /*
2.
0/14
36 (
436) G
EARED
- @
a @a 0
G3 a
(9'
0''
I I
I I
I
I5,O
OO
0
2 4
6
8
10 B
YPA
SS R
ATIO
60
00
0'
Y 4
00
0
kg
I
S2
w s
W + 2
000
a
Z
W
0 -
lop
00
- a
- 5
,000
- -
I I
I I
I
2
4
6
8
10 B
YP
AS
S R
ATI
O
Fig
ure
2
1.-
P
red
icte
d n
ac
ell
e d
rag.
3000
\ /
500 I000 \NLET T E M P E R ~ T U R E * ~ ~ Figure 23 .- Turbine re^^^^^ ratio *
